null User manual

null  User manual
A1-F18EA-NFM-000
NATOPS FLIGHT MANUAL
NAVY MODEL
F/A-18E/F
165533 AND UP
AIRCRAFT
THIS PUBLICATION SUPERSEDES A1-F18EA-NFM-000
DATED 1 DECEMBER 2004 CHANGED 1 AUGUST 2006
THIS PUBLICATION IS INCOMPLETE WITHOUT
A1-F18EA-NFM-200
DISTRIBUTION STATEMENT C. Distribution authorized to U.S. Government
agencies only and their contractors to protect publications required for official
use or for administrative or operational purposes only, determined on
15 September 2008. Other requests for this document shall be referred to
Commander, Naval Air Systems Command (PMA-265), RADM William A.
Moffett Bldg, 47123 Buse Rd, Bldg 2272, Patuxent River, MD 20670-1547.
DESTRUCTION NOTICE - For unclassified, limited documents, destroy by any
method that will prevent disclosure of contents or reconstruction of the
document.
ISSUED BY AUTHORITY OF THE CHIEF OF NAVAL OPERATIONS AND
UNDER THE DIRECTION OF THE COMMANDER
NAVAL AIR SYSTEMS COMMAND.
0801LP1088214
1 (Reverse Blank)
15 SEPTEMBER 2008
NAVAIR A1--F18EA--NFM--000
DEPARTMENT OF THE NAVY
NAVAL AIR SYSTEMS COMMAND
RADM WILLIAM A. MOFFETT BUILDING
47123 BUSE ROAD, BLDG 2272
PATUXENT RIVER, MD 20670-1547
15 September 2008
LETTER OF PROMULGATION
1. The Naval Air Training and Operating Procedures Standardization (NATOPS) Program is a
positive approach toward improving combat readiness and achieving a substantial reduction in the
aircraft mishap rate. Standardization, based on professional knowledge and experience, provides the
basis for development of an efficient and sound operational procedure. The standardization program
is not planned to stifle individual initiative, but rather to aid the Commanding Officer in increasing
the unit’s combat potential without reducing command prestige or responsibility.
2. This manual standardizes ground and flight procedures but does not include tactical doctrine.
Compliance with the stipulated manual requirements and procedures is mandatory except as
authorized herein. In order to remain effective, NATOPS must be dynamic and stimulate rather than
suppress individual thinking. Since aviation is a continuing, progressive profession, it is both
desirable and necessary that new ideas and new techniques be expeditiously evaluated and
incorporated if proven to be sound. To this end, Commanding Officers of aviation units are
authorized to modify procedures contained herein, in accordance with the waiver provisions
established by OPNAV Instruction 3710.7, for the purpose of assessing new ideas prior to initiating
recommendations for permanent changes. This manual is prepared and kept current by the users in
order to achieve maximum readiness and safety in the most efficient and economical manner. Should
conflict exist between the training and operating procedures found in this manual and those found
in other publications, this manual will govern.
3. Checklists and other pertinent extracts from this publication necessary to normal operations and
training should be made and carried for use in naval aircraft.
S. R. EASTBURG
Rear Admiral, United States Navy
By direction of
Commander, Naval Air Systems Command
3/(4 blank)
ORIGINAL
A1-F18EA-NFM-000
NATOPS Flight Manual
CONTENTS
Page
No.
PART I
THE AIRCRAFT
CHAPTER 1
The Aircraft
1.1
1.1.1
1.1.2
1.1.3
1.1.4
AIRCRAFT DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Meet The Super Hornet. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Aircraft Gross Weight.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
F/A-18F. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Radar Cross Section (RCS) Reduction. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
1.2
BLOCK NUMBERS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-1-4
CHAPTER 2
Systems
2.1
2.1.1
2.1.2
POWER PLANT SYSTEMS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-2-1
Engines. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-2-1
ATC - Automatic Throttle Control. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-2-10
2.2
2.2.1
2.2.2
2.2.3
2.2.4
2.2.5
2.2.6
2.2.7
2.2.8
2.2.9
FUEL SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engine Feed System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fuel Transfer System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fuel Tank Pressurization and Vent. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Thermal Management System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Refueling System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fuel Dump System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fuel Quantity Indicating System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fuel Low Level Indicating System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fuel System Related Cautions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-11
I-2-12
I-2-13
I-2-16
I-2-17
I-2-18
I-2-18
I-2-19
I-2-23
I-2-24
2.3
2.3.1
2.3.2
2.3.3
2.3.4
FPAS
FPAS
FPAS
FPAS
FPAS
- FLIGHT PERFORMANCE ADVISORY SYSTEM . . . . . . . . .
Display. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
CLIMB Option.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HOME Waypoint Selection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HOME FUEL Caution. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-24
I-2-24
I-2-25
I-2-25
I-2-25
2.4
2.4.1
2.4.2
2.4.3
2.4.4
2.4.5
2.4.6
SECONDARY POWER SYSTEM. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AMAD - Airframe Mounted Accessory Drive. . . . . . . . . . . . . . . . . . . . . . . . .
APU - Auxiliary Power Unit. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
APU Switch. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ATS Air Sources. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AUG PULL.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AMAD Related Cautions.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-27
I-2-27
I-2-28
I-2-28
I-2-28
I-2-29
I-2-29
5
I-1-1
I-1-1
I-1-2
I-1-2
I-1-3
ORIGINAL
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Page
No.
2.5
2.5.1
2.5.2
2.5.3
2.5.4
2.5.5
2.5.6
ELECTRICAL POWER SUPPLY SYSTEM . . . . . . . . . . . . . . . . . . . . . . .
Electrical RESET Button. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AC Electrical Power. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
DC Electrical Power. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
External Electrical Power. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Circuit Breakers. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Electrical System Cautions and Caution Lights. . . . . . . . . . . . . . . . . . . . . .
I-2-30
I-2-30
I-2-30
I-2-32
I-2-34
I-2-36
I-2-36
2.6
2.6.1
2.6.2
2.6.3
LIGHTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Exterior Lighting. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Interior Lighting. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Interior Lighting (F/A-18F). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-36
I-2-36
I-2-40
I-2-42
2.7
2.7.1
2.7.2
2.7.3
HYDRAULIC POWER SUPPLY SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . .
Hydraulic System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Hydraulic Accumulators. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Hydraulic System Related Cautions and Caution Light.. . . . . . . . . . . . . .
I-2-42
I-2-42
I-2-45
I-2-46
2.8
2.8.1
2.8.2
2.8.3
2.8.4
2.8.5
UTILITY HYDRAULIC FUNCTIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Landing Gear System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Nosewheel Steering System (NWS). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Wheel Brake System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Launch Bar System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Arresting Hook System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-46
I-2-46
I-2-49
I-2-50
I-2-56
I-2-58
2.9
2.9.1
2.9.2
2.9.3
2.9.4
WING FOLD SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Wingfold Mechanism. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Wingfold Operation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
WINGFOLD Switch. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Wingfold Overheat Cutout Protection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-59
I-2-59
I-2-59
I-2-60
I-2-60
2.10
2.10.1
2.10.2
2.10.3
2.10.4
2.10.5
2.10.6
2.10.7
2.10.8
2.10.9
2.10.10
2.10.11
2.10.12
2.10.13
2.10.14
FCS - FLIGHT CONTROL SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Flight Control Surfaces. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCCs - Flight Control Computers. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Redundancy and Survivability. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
CAS Operating Modes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Control Augmentation System (CAS). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Speedbrake Function. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
G-Limiter Considerations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Air Data Function. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Flight Controls. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Yaw Rate Warning Tone.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AOA Warning Tone.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Spin Recovery System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Stabilator Failure Control Law Reconfiguration. . . . . . . . . . . . . . . . . . . . . .
GAIN ORIDE. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-60
I-2-61
I-2-61
I-2-63
I-2-64
I-2-65
I-2-67
I-2-69
I-2-72
I-2-73
I-2-76
I-2-76
I-2-76
I-2-78
I-2-78
6
ORIGINAL
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Page
No.
2.10.15
2.10.16
FCS Failures.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-2-79
FCS Status Display. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-2-82
2.11
2.11.1
2.11.2
2.11.3
2.11.4
2.11.5
2.11.6
2.11.7
2.11.8
AFCS - AUTOMATIC FLIGHT CONTROL SYSTEM . . . . . . . . . . . . .
AFCS Mode Selection.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Basic Autopilot. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AFCS Mode Deselection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Pitch-Axis Pilot Relief Modes.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Roll-Axis Pilot Relief Modes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
CPL - Coupled Steering Modes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Coupled Data Link Modes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AFCS Related Caution and Advisories. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-84
I-2-84
I-2-86
I-2-86
I-2-86
I-2-86
I-2-87
I-2-87
I-2-87
2.12
2.12.1
2.12.2
2.12.3
2.12.4
2.12.5
2.12.6
2.12.7
2.12.8
WEAPON SYSTEMS CONTROLS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Stick Grip Switches/Controls (Front Cockpit). . . . . . . . . . . . . . . . . . . . . . . .
Stick Grip Switches/Controls (Trainer Configured F/A-18F).. . . . . . . . .
Throttle Grip Switches/Controls (Front Cockpit).. . . . . . . . . . . . . . . . . . . .
Throttle Grip Switches/Controls (Trainer Configured Rear Cockpit).
Hand Controllers (Missionized Rear Cockpit Lots 21 thru 25). . . . . . . .
Hand Controllers (Rear Cockpit LOT 26 AND UP). . . . . . . . . . . . . . . . . .
ALE-47 DISP Switch. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Grab Handle Chaff/Flare Dispense Switches (Rear Cockpit LOTs 21
thru 25).. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Grab Handle Chaff/Flare Dispense Switches (Rear Cockpit LOT 26
and up). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-87
I-2-88
I-2-90
I-2-90
I-2-91
I-2-91
I-2-95
I-2-95
I-2-95
2.13
2.13.1
2.13.2
2.13.3
2.13.4
2.13.5
2.13.6
2.13.7
2.13.8
2.13.9
2.13.10
2.13.11
2.13.12
2.13.13
2.13.14
ECS - ENVIRONMENTAL CONTROL SYSTEM . . . . . . . . . . . . . . . . . .
Bleed Air Shutoff Valves.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Bleed Air Subsystem.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Primary Heat Exchanger.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Warm Air Subsystems. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Air Conditioning System (ACS) Pack. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Secondary Heat Exchanger. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Avionics Cooling Fans. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ECS Operating Modes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Cabin Pressurization. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Windshield Anti-ice and Rain Removal. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Anti-g System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ECS RESET and AV COOL Switch.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
LCS - Liquid Cooling System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ECS Related Warnings, Cautions, and Advisories. . . . . . . . . . . . . . . . . . . .
I-2-95
I-2-95
I-2-96
I-2-97
I-2-97
I-2-97
I-2-98
I-2-98
I-2-99
I-2-103
I-2-104
I-2-104
I-2-105
I-2-105
I-2-106
2.14
2.14.1
2.14.2
OXYGEN SYSTEMS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-2-107
On Board Oxygen Generating System (OBOGS).. . . . . . . . . . . . . . . . . . . . . I-2-107
Emergency Oxygen. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-2-110
2.12.9
7
I-2-95
ORIGINAL
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Page
No.
2.15
2.15.1
2.15.2
2.15.3
2.15.4
2.15.5
2.15.6
2.15.7
2.15.8
2.15.9
FIRE DETECTION, FIRE EXTINGUISHING, AND BLEED AIR
LEAK DETECTION SYSTEMS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FIRE Lights. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
APU FIRE Light. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FIRE Warning Voice Alerts. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FIRE EXTGH READY/DISCH Light. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
APU Fire Extinguishing System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engine/AMAD Fire Extinguishing System. . . . . . . . . . . . . . . . . . . . . . . . . . .
FIRE Detection System Test.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Bleed Air Leak Detection (BALD) System. . . . . . . . . . . . . . . . . . . . . . . . . . .
DBFS - Dry Bay Fire Suppression System. . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-111
I-2-111
I-2-112
I-2-112
I-2-112
I-2-113
I-2-113
I-2-113
I-2-114
I-2-115
2.16
2.16.1
2.16.2
2.16.3
2.16.4
ENTRANCE/EGRESS SYSTEMS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Canopy System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Boarding Ladder.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Ejection Seat. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Ejection Seat System (F/A-18F). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-115
I-2-115
I-2-118
I-2-119
I-2-126
2.17
2.17.1
2.17.2
2.17.3
2.17.4
I-2-127
I-2-127
I-2-129
I-2-130
2.17.5
EMERGENCY EQUIPMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Jettison Systems. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Warnings/Cautions/Advisories. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Voice Alert System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Terrain Awareness Warning System (TAWS) (MC OFP 18E+ and
H2E AND UP). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
GPWS - Ground Proximity Warning System. . . . . . . . . . . . . . . . . . . . . . . . .
I-2-131
I-2-135
2.18
2.18.1
2.18.2
2.18.3
2.18.4
2.18.5
2.18.6
INSTRUMENTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Standby Attitude Reference Indicator.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Standby Airspeed Indicator. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Standby Altimeter. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Standby Rate of Climb Indicator.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Standby Magnetic Compass. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Angle Of Attack Indexer. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-140
I-2-140
I-2-140
I-2-140
I-2-141
I-2-141
I-2-141
2.19
2.19.1
2.19.2
2.19.3
2.19.4
2.19.5
2.19.6
2.19.7
2.19.8
AVIONICS SUBSYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Mission Computer System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Master Modes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Cockpit Controls and Displays. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Multipurpose Display Group. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Up Front Control Display (UFCD). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
SDC - Signal Data Computer. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
CVRS - Cockpit Video Recording System. . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fast Tactical Imaging Set (FTI-II). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-142
I-2-142
I-2-146
I-2-146
I-2-146
I-2-166
I-2-171
I-2-171
I-2-178
2.20
TACTICAL AIRCRAFT MOVING MAP CAPABILITY
(TAMMAC) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-2-178
8
ORIGINAL
A1-F18EA-NFM-000
Page
No.
2.20.1
2.20.2
2.20.3
2.20.4
2.20.5
2.20.6
2.20.7
TAMMAC Status Monitoring. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AMU Maintenance Format Options and Display Information. . . . . . . .
Map Theater Data Loading. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Map Loading Format Options.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Map Loading Format Status Information. . . . . . . . . . . . . . . . . . . . . . . . . . . .
Map Loading Interruptions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AMU/PC Cards Cautions and Advisories.. . . . . . . . . . . . . . . . . . . . . . . . . . . .
2.21
2.21.1
2.21.2
COUNTERMEASURES DISPENSING SYSTEM . . . . . . . . . . . . . . . . . . I-2-185
ALE-47 Countermeasures Dispensing Set. . . . . . . . . . . . . . . . . . . . . . . . . . . . I-2-185
ALE-50 Decoy Dispensing Set. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-2-185
2.22
2.22.1
2.22.2
2.22.3
2.22.4
2.22.5
2.22.6
BIT-STATUS MONITORING SUBSYSTEM . . . . . . . . . . . . . . . . . . . . . .
FIRAMS - Flight Incident Recorder and Aircraft Monitoring Set. . . .
DFIRS - Deployable Flight Incident Recorder Set.. . . . . . . . . . . . . . . . . . .
Avionics BIT. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Non-Avionic BIT. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Status Monitoring Backup. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Non-BIT Status. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-186
I-2-186
I-2-186
I-2-186
I-2-201
I-2-201
I-2-202
2.23
2.23.1
JOINT HELMET MOUNTED CUEING SYSTEM (JHMCS) . . . . . .
Helmet Mounted Display (HMD)/Aft Helmet Mounted Display
(AHMD).. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Electronics Unit (EU).. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Cockpit Unit (CU). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Magnetic Transmitter Unit (MTU). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Boresight Reference Unit (BRU). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Seat Position Sensor (SPS). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HMD/AHMD OFF/BRT Knobs. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HUD Video Record Panel.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Cautions/Advisories. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Configuration Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Built-In Test (BIT). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
JHMCS Alignment. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HMD/AHMD Symbology. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Navigation Master Mode.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Mission Computer Failure. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Electronic Unit Failure. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Helmet Tracker Failure. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Helmet Not Installed. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-205
2.23.2
2.23.3
2.23.4
2.23.5
2.23.6
2.23.7
2.23.8
2.23.9
2.23.10
2.23.11
2.23.12
2.23.13
2.23.14
2.23.15
2.23.16
2.23.17
2.23.18
9
I-2-179
I-2-181
I-2-181
I-2-181
I-2-182
I-2-183
I-2-184
I-2-205
I-2-208
I-2-208
I-2-208
I-2-208
I-2-208
I-2-208
I-2-208
I-2-208
I-2-211
I-2-211
I-2-213
I-2-214
I-2-215
I-2-218
I-2-218
I-2-221
I-2-221
ORIGINAL
A1-F18EA-NFM-000
Page
No.
CHAPTER 3
Servicing and Handling
3.1
SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-3-1
CHAPTER 4
Operating Limitations
4.1
4.1.1
4.1.2
4.1.3
4.1.4
4.1.5
4.1.6
4.1.7
4.1.8
4.1.9
LIMITATIONS OF THE BASIC AIRCRAFT . . . . . . . . . . . . . . . . . . . . . .
Engine Operation Limitations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
CG Limitations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Airspeed Limitations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Gross Weight and Lateral Weight Asymmetry Limitations. . . . . . . . . . .
AOA Limitations - Flaps AUTO. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Acceleration Limitations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Limitations with Flaps HALF or FULL. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Refueling Limitation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Prohibited Maneuvers. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.2
4.2.1
4.2.2
EXTERNAL STORES LIMITATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-4-13
ARS Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-4-17
ATFLIR Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-4-19
4.3
4.3.1
OPERATING LIMITATIONS (LOT 21) . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-4-19
Prohibited Maneuvers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-4-19
PART II
INDOCTRINATION
CHAPTER 5
Indoctrination
5.1
5.1.1
5.1.2
INITIAL QUALIFICATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . II-5-1
Minimum Ground Training Requirements. . . . . . . . . . . . . . . . . . . . . . . . . . . II-5-1
Minimum Flight Training Requirements. . . . . . . . . . . . . . . . . . . . . . . . . . . . . II-5-1
5.2
FOLLOW-ON TRAINING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . II-5-1
5.3
5.3.1
CURRENCY REQUIREMENTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . II-5-2
Regaining Currency. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . II-5-2
5.4
5.4.1
5.4.2
5.4.3
REQUIREMENTS FOR VARIOUS FLIGHT PHASES . . . . . . . . . . . . .
Instrument Evaluation Flights. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Instrument Qualification. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Ceiling/Visibility Requirements. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5.5
WAIVERS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . II-5-4
5.6
PERSONAL FLYING EQUIPMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . II-5-4
10
I-4-1
I-4-1
I-4-1
I-4-1
I-4-3
I-4-5
I-4-5
I-4-9
I-4-9
I-4-9
II-5-2
II-5-2
II-5-2
II-5-3
ORIGINAL
A1-F18EA-NFM-000
Page
No.
PART III
NORMAL PROCEDURES
CHAPTER 6
Flight Preparation
6.1
6.1.1
6.1.2
MISSION PLANNING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-6-1
General. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-6-1
Flight Codes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-6-1
6.2
6.2.1
6.2.2
BRIEFING/DEBRIEFING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-6-1
Briefing.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-6-1
Debriefing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-6-4
CHAPTER 7
Shore-Based Procedures
7.1
7.1.1
7.1.2
7.1.3
7.1.4
7.1.5
7.1.6
PREFLIGHT CHECKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
In Maintenance Control.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Inspection of RCS Reduction Features. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Exterior Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Before Entering Cockpit. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Interior Checks - Pilot. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Interior Checks - WSO. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.2
7.2.1
7.2.2
ENGINE START . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-7-18
Intercockpit Communications (F/A-18F). . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-7-19
Engine Start Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-7-19
7.3
BEFORE TAXI CHECKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-7-23
7.4
TAXI CHECKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-7-30
7.5
7.5.1
7.5.2
7.5.3
7.5.4
TAKEOFF . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Before Takeoff Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Normal Takeoff.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Crosswind Takeoff. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
After Takeoff Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
III-7-30
III-7-30
III-7-32
III-7-34
III-7-34
7.6
7.6.1
7.6.2
7.6.3
AIRBORNE CHECKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Climb.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
10,000 Foot Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Cruise. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
III-7-34
III-7-34
III-7-34
III-7-35
7.7
7.7.1
7.7.2
7.7.3
7.7.4
7.7.5
LANDING CHECKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Descent/Penetration. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
VFR Landing Pattern Entry. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
VFR Landing Pattern and Approach. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Pattern Adjustments.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Final Approach. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
III-7-36
III-7-36
III-7-39
III-7-39
III-7-40
III-7-40
11
III-7-1
III-7-1
III-7-1
III-7-2
III-7-8
III-7-11
III-7-16
ORIGINAL
A1-F18EA-NFM-000
Page
No.
7.7.6
7.7.7
7.7.8
7.7.9
7.7.10
7.7.11
7.7.12
7.7.13
ATC Approaches.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FPAH/ROLL - ATC Approaches.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Full Stop Landings. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Braking Technique. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Heavy Gross Weight Landings. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Crosswind Landings. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Wet Runway Landings. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Asymmetric Stores Landings. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
III-7-40
III-7-40
III-7-41
III-7-42
III-7-43
III-7-44
III-7-45
III-7-45
7.8
7.8.1
7.8.2
7.8.3
7.8.4
POST-FLIGHT CHECKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
After Landing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Hot Refueling.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Before Engine Shutdown Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engine Shutdown Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
III-7-45
III-7-45
III-7-47
III-7-47
III-7-49
CHAPTER 8
Carrier-Based Procedures
8.1
GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-8-1
8.2
8.2.1
8.2.2
8.2.3
8.2.4
8.2.5
8.2.6
8.2.7
8.2.8
DAY OPERATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Preflight Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Hangar Deck Operation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engine Start. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Catapult Trim. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Taxi. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Takeoff Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Catapult Launch. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Landing Pattern. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8.3
ACL MODE 1 AND 1A APPROACHES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-8-15
8.4
ACL MODE 2 APPROACH . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-8-18
8.5
ARRESTED LANDING AND EXIT FROM THE LANDING
AREA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-8-20
8.6
SECTION CCA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-8-21
8.7
8.7.1
8.7.2
8.7.3
8.7.4
8.7.5
8.7.6
8.7.7
8.7.8
NIGHT OPERATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
General. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Preflight. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Before Taxi. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Taxi. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Catapult Hook-Up. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Catapult Launch. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Catapult Suspend. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Night Landings. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12
III-8-1
III-8-1
III-8-1
III-8-2
III-8-3
III-8-9
III-8-9
III-8-10
III-8-12
III-8-21
III-8-21
III-8-21
III-8-21
III-8-21
III-8-21
III-8-21
III-8-23
III-8-23
ORIGINAL
A1-F18EA-NFM-000
Page
No.
8.7.9
Arrestment and Exit From the Landing Area. . . . . . . . . . . . . . . . . . . . . . . . III-8-23
CHAPTER 9
Special Procedures
9.1
9.1.1
9.1.2
9.1.3
9.1.4
9.1.5
FORMATION FLIGHT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Formation Taxi/Takeoff. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Aborted Takeoff. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Parade.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Balanced Cruise Formation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Section Approaches/Landing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
9.2
9.2.1
9.2.2
AIR REFUELING (RECEIVER) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-9-5
Air Refueling Checklist. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-9-5
Refueling Technique. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-9-6
9.3
9.3.1
9.3.2
9.3.3
9.3.4
AIR REFUELING (TANKER) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Air Refueling Store (ARS). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ARS (Tanker) Procedures. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ARS Jettison. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ARS Limitations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
III-9-9
III-9-9
III-9-12
III-9-18
III-9-18
9.4
9.4.1
9.4.2
9.4.3
9.4.4
9.4.5
NIGHT VISION DEVICE (NVD) OPERATIONS . . . . . . . . . . . . . . . . . .
Effects on Vision. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Effects of Light. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Weather Conditions.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Object/Target Detection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Flight Preparation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
III-9-18
III-9-18
III-9-18
III-9-19
III-9-19
III-9-19
9.5
9.5.1
9.5.2
9.5.3
9.5.4
9.5.5
SHORT AIRFIELD FOR TACTICAL SUPPORT (SATS)
PROCEDURES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Landing Pattern. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Approach. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Waveoff. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Arrested Landing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Bolter. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
III-9-19
III-9-19
III-9-19
III-9-19
III-9-19
III-9-20
9.6
HOT SEAT PROCEDURE. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-9-20
9.7
9.7.1
9.7.2
ALERT SCRAMBLE LAUNCH PROCEDURES. . . . . . . . . . . . . . . . . . . . III-9-20
Setting the Alert. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-9-20
Alert Five Launch. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-9-21
9.8
9.8.1
9.8.2
9.8.3
AIRBORNE HMD ACCURACY CHECKS . . . . . . . . . . . . . . . . . . . . . . . . . .
HMD Alignment. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Airborne HMD Accuracy Check with Radar. . . . . . . . . . . . . . . . . . . . . . . . .
Airborne HMD Accuracy Check with CATM/AIM-9X. . . . . . . . . . . . . . .
13
III-9-1
III-9-1
III-9-1
III-9-1
III-9-4
III-9-4
III-9-22
III-9-22
III-9-23
III-9-23
ORIGINAL
A1-F18EA-NFM-000
Page
No.
CHAPTER 10
Functional Checkflight Procedures
10.1
10.1.1
10.1.2
10.1.3
GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engine Functional Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Flight Control System Functional Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . .
Landing Gear Functional Checks.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
10.2
FCF REQUIREMENTS. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-10-2
10.3
FCF QUALIFICATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-10-2
10.4
FCF PROCEDURES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-10-2
10.5
10.5.1
10.5.2
10.5.3
10.5.4
10.5.5
10.5.6
10.5.7
10.5.8
10.5.9
10.5.10
10.5.11
10.5.12
10.5.13
10.5.14
10.5.15
10.5.16
10.5.17
10.5.18
FCF CHECKLIST - PROFILE A . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Plane Captain Brief.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Preflight Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Pre-Start Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engine Start Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Post-Start Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Before Taxi Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Taxi Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Shipboard Taxi/Takeoff Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Shorebased Takeoff Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
After Takeoff Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Medium Altitude Checks (above 10,000 feet). . . . . . . . . . . . . . . . . . . . . . . . .
10,000 Feet Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
High Altitude (above 30,000 feet). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
10,000 Feet to Landing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Landing Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
After Landing Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Before Engine Shutdown Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engine Shutdown Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
10.6
10.6.1
FCF CHECKLIST - PROFILE C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-10-31
10,000 Feet Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-10-31
10.7
10.7.1
10.7.2
10.7.3
10.7.4
FCF CHECKLIST - PROFILE D (REAR COCKPIT) . . . . . . . . . . . . . . .
Preflight Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Before Taxi Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Taxi Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Medium Altitude Checks (above 10,000 feet). . . . . . . . . . . . . . . . . . . . . . . .
10.8
10.8.1
FCF CHECKLIST - PROFILE E . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-10-34
10,000 Feet Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-10-34
14
III-10-1
III-10-1
III-10-1
III-10-2
III-10-4
III-10-4
III-10-4
III-10-5
III-10-6
III-10-7
III-10-15
III-10-19
III-10-20
III-10-21
III-10-21
III-10-22
III-10-24
III-10-27
III-10-28
III-10-29
III-10-29
III-10-29
III-10-30
III-10-32
III-10-32
III-10-32
III-10-32
III-10-33
ORIGINAL
A1-F18EA-NFM-000
Page
No.
PART IV
FLIGHT CHARACTERISTICS
CHAPTER 11
Flight Characteristics
11.1
11.1.1
11.1.2
11.1.3
11.1.4
11.1.5
11.1.6
HANDLING QUALITIES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Flight Control Mode Effects on Handling Qualities. . . . . . . . . . . . . . . . . .
Handling Qualities with Flaps HALF or FULL.. . . . . . . . . . . . . . . . . . . . . .
Flaps AUTO Handling Qualities. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FLIR Carriage Handling Qualities. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-Wet (4-EFT and ARS) Loading Handling Qualities. . . . . . . . . . . . . . . .
Dual Midboard with Outboard Stores Handling Qualities. . . . . . . . . . . .
IV-11-1
IV-11-1
IV-11-1
IV-11-2
IV-11-5
IV-11-6
IV-11-7
11.2
11.2.1
11.2.2
11.2.3
AIR COMBAT MANEUVERING. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Air-to-Air Gun Tracking. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Over-the-Top Maneuvering. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Slow Speed Maneuvering. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IV-11-7
IV-11-7
IV-11-7
IV-11-7
11.3
11.3.1
11.3.2
11.3.3
OUT-OF-CONTROL FLIGHT (OCF) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Departure Resistance. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Departure Characteristics. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Spin Characteristics.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IV-11-8
IV-11-8
IV-11-8
IV-11-9
11.4
11.4.1
11.4.2
11.4.3
11.4.4
11.4.5
11.4.6
DEGRADED MODE HANDLING QUALITIES. . . . . . . . . . . . . . . . . . . . .
Single Engine Operation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Leading Edge Flap Asymmetry. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Trailing Edge Flap Failure. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Stabilator Failure. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
GAIN ORIDE. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AHRS Failure Flying Qualities.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IV-11-10
IV-11-10
IV-11-11
IV-11-13
IV-11-14
IV-11-15
IV-11-16
PART V
EMERGENCY PROCEDURES
EMERGENCY INDEX
CHAPTER 12
General Emergencies
12.1
12.1.1
12.1.2
GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-12-1
Immediate Action Items. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-12-1
Warnings, Cautions, and Advisories. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-12-1
CHAPTER 13
Ground Emergencies
13.1
LOSS OF DC ESSENTIAL BUS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-13-1
13.2
ENGINE FAILS TO START/HUNG START . . . . . . . . . . . . . . . . . . . . . . V-13-1
13.3
HOT START . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-13-1
15
ORIGINAL
A1-F18EA-NFM-000
Page
No.
13.4
13.4.1
GROUND FIRE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-13-2
HOT BRAKES/BRAKE FIRE. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-13-2
13.5
EMERGENCY EGRESS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-13-3
13.6
BRAKE FAILURE/EMERGENCY BRAKES . . . . . . . . . . . . . . . . . . . . . . . V-13-4
CHAPTER 14
Takeoff Emergencies
14.1
EMERGENCY CATAPULT FLYAWAY
14.2
ABORT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-14-2
14.3
GO AROUND . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-14-3
14.4
LOSS OF DIRECTIONAL CONTROL DURING TAKEOFF OR
LANDING (BLOWN TIRE, NWS FAILURE) . . . . . . . . . . . . . . . . . . . . . . V-14-4
14.5
LANDING GEAR FAILS TO RETRACT . . . . . . . . . . . . . . . . . . . . . . . . . . V-14-5
CHAPTER 15
Inflight Emergencies
15.1
AFTERBURNER FAILURE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-1
15.2
RESTART . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-1
15.3
FUSELAGE FUEL LEAK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-4
15.4
HYDRAULIC FAILURES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-5
15.5
DOUBLE TRANSFORMER-RECTIFIER FAILURE . . . . . . . . . . . . . . . V-15-10
15.6
COCKPIT TEMPERATURE HIGH . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-15
15.7
COCKPIT SMOKE, FUMES, OR FIRE. . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-16
15.8
HYPOXIA/LOW MASK FLOW/NO MASK FLOW . . . . . . . . . . . . . . . . . V-15-17
15.9
LOSS OF CABIN PRESSURIZATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-18
15.10
DISPLAY MALFUNCTION - NON-AMCD AIRCRAFT . . . . . . . . . . . . V-15-18
15.11
DISPLAY MALFUNCTION - AMCD AIRCRAFT . . . . . . . . . . . . . . . . . . V-15-19
15.12
DUAL MISSION COMPUTER (MC) FAILURE - AMCD
AIRCRAFT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-19
15.13
OUT-OF-CONTROL FLIGHT (OCF) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-19
16
. . . . . . . . . . . . . . . . . . . . . . . . . V-14-1
ORIGINAL
A1-F18EA-NFM-000
Page
No.
15.13.1
15.13.2
15.13.3
15.13.4
Departure from Controlled Flight. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Spin. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
OCF Recovery Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Post Departure Dive Recovery. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
15.14
CONTROLLABILITY CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-22
15.15
EXTERNAL STORES JETTISON . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-23
15.16
15.16.1
15.16.2
15.16.3
15.16.4
ARS MALFUNCTIONS. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ARS Hose Fails to Retract. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ARS Refueling Hose Jettison. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ARS Hydraulic Pressure Light. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
No RDY Light. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
15.17
EMERGENCY TANKER DISENGAGEMENT . . . . . . . . . . . . . . . . . . . . . V-15-26
15.18
FCS FAILURE INDICATIONS AND EFFECTS . . . . . . . . . . . . . . . . . . . . V-15-27
15.19
15.19.1
15.19.2
15.19.3
15.19.4
ANGLE OF ATTACK (AOA) FAILURE . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AOA PROBE DAMAGE OR BINDING.. . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AOA PROBE SELECTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
SINGLE AOA FAILURE ON TAKEOFF . . . . . . . . . . . . . . . . . . . . . . . . . . .
DUAL AOA FAILURE ON TAKEOFF. . . . . . . . . . . . . . . . . . . . . . . . . . . . .
15.20
15.20.1
AILERON HINGE FAILURE - SUSPECTED, INBOARD. . . . . . . . . . V-15-47
Suspected Inboard Aileron Hinge Failure Corrective Action. . . . . . . . . . V-15-48
CHAPTER 16
Landing Emergencies
16.1
SINGLE ENGINE FAILURE IN LANDING CONFIGURATION . . V-16-1
16.2
SINGLE ENGINE APPROACH AND LANDING. . . . . . . . . . . . . . . . . . . V-16-1
16.3
SINGLE ENGINE WAVEOFF/BOLTER . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-3
16.4
FORCED LANDING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-4
16.5
LANDING GEAR UNSAFE/FAILS TO EXTEND . . . . . . . . . . . . . . . . . . V-16-4
16.6
LANDING GEAR EMERGENCY EXTENSION . . . . . . . . . . . . . . . . . . . V-16-5
16.7
PLANING LINK FAILURE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-12
16.8
16.8.1
16.8.2
ARRESTMENT - FIELD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-13
Arresting Gear Types. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-13
Arrestment Decision. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-14
17
V-15-19
V-15-20
V-15-20
V-15-21
V-15-25
V-15-25
V-15-25
V-15-25
V-15-26
V-15-44
V-15-45
V-15-46
V-15-46
V-15-47
ORIGINAL
A1-F18EA-NFM-000
Page
No.
16.8.3
16.8.4
Arrestment - Short Field.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-14
Arrestment - Long Field. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-14
16.9
BARRICADE ARRESTMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-14
16.10
CV RECOVERY MATRIX . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-15
CHAPTER 17
Ejection
17.1
17.1.1
17.1.2
17.1.3
17.1.4
EJECTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Ejection Seat Restrictions.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Low Altitude Ejection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
High Altitude Ejection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Ejection Procedures.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
17.2
DITCHING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-17-4
17.3
SEAWATER ENTRY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-17-4
CHAPTER 18
Immediate Action
18.1
GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-1
18.2
APU FIRE LIGHT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-1
18.3
DUAL L BLEED and R BLEED WARNING LIGHTS//L/R ATS
CAUTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-1
18.4
SINGLE L BLEED or R BLEED WARNING LIGHT . . . . . . . . . . . . . . V-18-1
18.5
FIRE LIGHT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-2
18.6
ENGINE CAUTIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-2
18.7
L/R FUEL INLT CAUTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-2
18.8
HYD1 (2) HOT CAUTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-2
18.9
OBOGS DEGD CAUTION//HYPOXIA/LOW MASK FLOW/NO
MASK FLOW//LOSS OF CABIN PRESSURIZATION/CABIN
CAUTION LIGHT BELOW 47,000 FEET. . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-2
18.10
HOT START . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-3
18.11
BRAKE FAILURE/EMERGENCY BRAKES . . . . . . . . . . . . . . . . . . . . . . V-18-3
18.12
EMERGENCY CATAPULT FLYAWAY . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-3
18
V-17-1
V-17-1
V-17-3
V-17-4
V-17-4
ORIGINAL
A1-F18EA-NFM-000
Page
No.
18.13
ABORT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-3
18.14
LOSS OF DIRECTIONAL CONTROL DURING TAKEOFF OR
LANDING/ PLANING LINK FAILURE. . . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-3
18.15
COCKPIT SMOKE, FUMES, OR FIRE . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-4
18.16
OCF RECOVERY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-18-4
18.17
SINGLE ENGINE FAILURE IN LANDING CONFIGURATION . . V-18-4
PART VI
ALL WEATHER PROCEDURES
CHAPTER 19
Instrument Flight
19.1
19.1.1
19.1.2
19.1.3
INSTRUMENT FLIGHT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Before Takeoff.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Inflight. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Approaches. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
19.2
DEGRADED SYSTEMS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VI-19-2
CHAPTER 20
Extreme Weather Procedures
20.1
20.1.1
20.1.2
20.1.3
ICE AND RAIN. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Ground Operation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
In Flight. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Landing in Heavy Rain. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
20.2
TURBULENT AIR AND THUNDERSTORM OPERATION . . . . . . VI-20-4
CHAPTER 21
Hot Weather Procedures
21.1
GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VI-21-1
21.2
GROUND OPERATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VI-21-1
21.3
IN FLIGHT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VI-21-1
21.4
DESCENT/RECOVERY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VI-21-2
21.5
AFTER LANDING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VI-21-2
CHAPTER 22
Cold Weather Procedures
22.1
EXTERIOR INSPECTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VI-22-1
19
VI-19-1
VI-19-1
VI-19-1
VI-19-2
VI-20-1
VI-20-1
VI-20-1
VI-20-3
ORIGINAL
A1-F18EA-NFM-000
Page
No.
22.2
BEFORE ENTERING COCKPIT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VI-22-1
22.3
INTERIOR CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VI-22-1
22.4
ENGINE START . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VI-22-1
22.5
BEFORE TAXI . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VI-22-1
22.6
TAKEOFF . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VI-22-2
PART VII
COMM-NAV EQUIPMENT
AND PROCEDURES
CHAPTER 23
Communication-Identification Equipment
23.1
MULTIFUNCTION INFORMATION DISTRIBUTION SYSTEM
(MIDS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-23-1
23.2
23.2.1
ICS - INTERCOM SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-23-1
ICS Function Selector Switch (F/A-18F). . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-23-2
23.3
23.3.1
23.3.2
23.3.3
VHF/UHF COMMUNICATION SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . .
VHF/UHF Controls and Indicators. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
COMM 1 and 2 Operation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Anti-Jam Operation.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
VII-23-2
VII-23-3
VII-23-8
VII-23-13
23.4
23.4.1
VII-23-14
23.4.3
23.4.4
HAVE QUICK SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Have Quick Menu From AJ Menu Fixed Frequency COMM
Sublevel. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
SINCGARS Menu From AJ Menu Fixed Frequency COMM
Sublevel. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
COMM Sublevel - Presets 1 Thru 20 (Anti-Jam). . . . . . . . . . . . . . . . . . . . .
AJ MENU From Have Quick COMM Sublevel.. . . . . . . . . . . . . . . . . . . . . .
23.5
23.5.1
23.5.2
SINCGARS SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-23-20
AJ MENU From SINCGARS COMM Sublevel. . . . . . . . . . . . . . . . . . . . . . VII-23-21
Guard Transmit on Top Level CNI Format. . . . . . . . . . . . . . . . . . . . . . . . . . VII-23-23
23.6
23.6.1
23.6.2
23.6.3
KY-58
KY-58
KY-58
KY-58
- SECURE SPEECH SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Control Panel Assembly. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Controls and Indicators.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Operation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
VII-23-23
VII-23-23
VII-23-25
VII-23-25
23.7
23.7.1
23.7.2
23.7.3
DIGITAL COMMUNICATION SYSTEM (DCS). . . . . . . . . . . . . . . . . . . .
VHF/UHF Communication System - DCS. . . . . . . . . . . . . . . . . . . . . . . . . . .
DCS COMSEC (Communications Security). . . . . . . . . . . . . . . . . . . . . . . . . .
UFCD - DCS. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
VII-23-27
VII-23-27
VII-23-28
VII-23-29
23.4.2
20
VII-23-14
VII-23-17
VII-23-18
VII-23-20
ORIGINAL
A1-F18EA-NFM-000
Page
No.
23.7.4
23.7.5
Relay Mode Of Operation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-23-29
DCS - Avionics Subsystem. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-23-31
23.8
IDENTIFICATION FRIEND OR FOE (IFF)/COMBINED
INTERROGATOR TRANSPONDER (CIT) . . . . . . . . . . . . . . . . . . . . . . . .
IFF Transponder. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IFF Interrogator. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IFF BIT. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IFF Programming (IFF PROG) Page.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IFF Controls. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IFF Related Cautions and Advisories. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
23.8.1
23.8.2
23.8.3
23.8.4
23.8.5
23.8.6
VII-23-32
VII-23-32
VII-23-43
VII-23-43
VII-23-43
VII-23-46
VII-23-48
23.9
COMMUNICATION-NAVIGATION-IDENTIFICATION
INTERFACE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-23-48
CHAPTER 24
Navigation Equipment
24.1
24.1.1
24.1.2
24.1.3
24.1.4
24.1.5
24.1.6
24.1.7
NAVIGATION CONTROLS AND INDICATORS . . . . . . . . . . . . . . . . . . .
UFCD Navigation Controls. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Moving Map - Digital Map Set (DMS). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HSI Format. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HSI Format on a DDI. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HUD - Navigation Information. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
CRS Select Switch. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IFF Control Panel. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
VII-24-1
VII-24-1
VII-24-1
VII-24-2
VII-24-9
VII-24-9
VII-24-10
VII-24-10
24.2
24.2.1
24.2.2
24.2.3
24.2.4
24.2.5
24.2.6
24.2.7
24.2.8
24.2.9
24.2.10
24.2.11
24.2.12
24.2.13
24.2.14
24.2.15
24.2.16
INERTIAL NAVIGATION SYSTEM (INS)/GLOBAL
POSITIONING SYSTEM (GPS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Inertial Navigation System (INS). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Global Positioning System (GPS). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Inertial Alignment Modes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24.2.4 Navigation System BIT. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
NAVCK Display. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
GPS Page. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
INS Knob. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
INS/GPS Related Cautions and Advisories. . . . . . . . . . . . . . . . . . . . . . . . . .
Waypoints, Offset Aimpoints (OAP), and Offsets.. . . . . . . . . . . . . . . . . . . .
Aircraft (A/C) Programming. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Position Keeping. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Position Updating. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
NAV/TAC Bank Limit Options. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Steering. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Designation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
INS Updates (not available in AINS). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
VII-24-10
VII-24-11
VII-24-11
VII-24-11
VII-24-23
VII-24-23
VII-24-24
VII-24-25
VII-24-25
VII-24-26
VII-24-35
VII-24-37
VII-24-38
VII-24-39
VII-24-39
VII-24-47
VII-24-50
24.3
ADF (AUTOMATIC DIRECTION FINDER) . . . . . . . . . . . . . . . . . . . . . . . VII-24-51
21
ORIGINAL
A1-F18EA-NFM-000
Page
No.
24.4
24.4.1
24.4.2
24.4.3
24.4.4
24.4.5
TACAN
TACAN
TACAN
TACAN
TACAN
TACAN
(TACTICAL AIR NAVIGATION). . . . . . . . . . . . . . . . . . . . . . . . . .
BIT.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Mode Selection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Programming. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Position Keeping. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Position Updating. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
VII-24-52
VII-24-53
VII-24-53
VII-24-53
VII-24-54
VII-24-54
24.5
24.5.1
24.5.2
24.5.3
24.5.4
24.5.5
INSTRUMENT CARRIER LANDING SYSTEM (ICLS) . . . . . . . . . . .
ICLS Receiver. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ICLS Decoder. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ICLS BIT.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ICLS Initialization. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ICLS Steering. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
VII-24-54
VII-24-54
VII-24-58
VII-24-58
VII-24-58
VII-24-59
24.6
24.6.1
DATA LINK SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-24-59
Automatic Carrier Landing Mode. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-24-59
24.7
24.7.1
24.7.2
24.7.3
NAVIGATION DATA ENTRY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Standard Data Entry. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fast Data Entry. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Data Entry Using the Shifted Keypad. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
CHAPTER 25
Backup/Degraded Operations
25.1
25.1.1
25.1.2
MISSION COMPUTER FAILURE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-25-1
Mission Computer Failure (Non-AMCD Aircraft). . . . . . . . . . . . . . . . . . . VII-25-1
Mission Computer Failure (AMCD Aircraft) . . . . . . . . . . . . . . . . . . . . . . . . . VII-25-1
25.2
25.2.1
25.2.2
25.2.3
25.2.4
BACKUP ATTITUDE AND NAVIGATION SYSTEM . . . . . . . . . . . . .
Standby Attitude Reference Indicator.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Static Power Inverter. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Magnetic Azimuth Detector. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Backup Attitude and Navigation System Controls and Indicators. . . .
25.3
25.3.1
25.3.2
NAVIGATION BACKUP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-25-4
Navigation Controls and Indicators. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-25-5
Backup Heading Mode Control. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-25-5
25.4
BACKUP FREQUENCY CONTROL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-25-6
CHAPTER 26
Visual Communications
CHAPTER 27
Deck/Ground Handling Signals
PART VIII
WEAPONS SYSTEMS
PART IX
FLIGHT CREW COORDINATION
22
VII-24-70
VII-24-78
VII-24-78
VII-24-78
VII-25-2
VII-25-2
VII-25-3
VII-25-3
VII-25-3
ORIGINAL
A1-F18EA-NFM-000
Page
No.
CHAPTER 28
Aircrew Coordination
28.1
DEFINITION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-28-1
28.2
28.2.1
28.2.2
28.2.3
28.2.4
28.2.5
28.2.6
28.2.7
28.2.8
CRITICAL SKILLS OF AIRCREW COORDINATION . . . . . . . . . . . . .
Situation Awareness. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Assertiveness. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Decision-Making. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Communication. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Leadership.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Adaptability/Flexibility. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Mission Analysis. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Factors That Degrade Aircrew Coordination. . . . . . . . . . . . . . . . . . . . . . . . .
IX-28-1
IX-28-1
IX-28-1
IX-28-1
IX-28-1
IX-28-1
IX-28-1
IX-28-2
IX-28-2
28.3
28.3.1
28.3.2
28.3.3
28.3.4
FLIGHT MEMBER POSITIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Mission Commander. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Pilot In Command. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Formation Leader.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Weapon Systems Operator (WSO). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IX-28-2
IX-28-2
IX-28-2
IX-28-2
IX-28-3
28.4
28.4.1
28.4.2
28.4.3
28.4.4
28.4.5
28.4.6
AIRCREW RESPONSIBILITIES BY FLIGHT PHASE. . . . . . . . . . . . .
Mission Planning and Briefing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Pretakeoff. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Takeoff/Departure. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Enroute.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Recovery.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Mission Critique. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IX-28-3
IX-28-3
IX-28-3
IX-28-3
IX-28-3
IX-28-4
IX-28-4
28.5
28.5.1
28.5.2
SPECIAL CONSIDERATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-28-4
Functional Checkflights. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-28-4
Formation Flights.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-28-4
28.6
EMERGENCIES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-28-4
CHAPTER 29
Crew Coordination Standards
29.1
PHILOSOPHY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-1
29.2
MISSION PLANNING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-1
29.3
CREW BRIEFING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-1
29.4
29.4.1
29.4.2
COMMUNICATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-1
Intra-Cockpit Communications (ICS). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-1
Guidelines for Effective Communications. . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-2
29.5
PRE-FLIGHT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-2
23
ORIGINAL
A1-F18EA-NFM-000
Page
No.
29.6
29.6.1
29.6.2
29.6.3
29.6.4
29.6.5
29.6.6
29.6.7
29.6.8
FLIGHT PHASES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Start/Taxi/Takeoff. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Departure. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Rendezvous and Formation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Cruise. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Inflight Refueling. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Approach. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Landing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Post Flight. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
29.7
DEBRIEFING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-7
29.8
29.8.1
29.8.2
SPECIAL CONSIDERATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-7
Functional Check Flights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-7
Emergencies. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-7
29.9
STANDARD ICS TERMS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-8
PART X
NATOPS EVALUATION
CHAPTER 30
NATOPS Evaluation
30.1
30.1.1
30.1.2
CONCEPT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . X-30-1
Implementation.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . X-30-1
Definitions.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . X-30-1
30.2
30.2.1
GROUND EVALUATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . X-30-2
General. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . X-30-2
30.3
30.3.1
30.3.2
30.3.3
30.3.4
30.3.5
30.3.6
30.3.7
30.3.8
30.3.9
30.3.10
30.3.11
30.3.12
30.3.13
FLIGHT EVALUATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Mission Planning/Briefing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Preflight/Line Operations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Taxi. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Takeoff (*). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Climb/Cruise. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Approach/Landing (*). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Communications. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Emergency/Malfunction Procedures (*). . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Post Flight Procedures.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Mission Evaluation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Applicable Publications. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Flight Evaluation Grading Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Flight Evaluation Grade Determination. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30.4
NATOPS EVALUATION QUESTION BANK . . . . . . . . . . . . . . . . . . . . . . X-30-6
24
IX-29-2
IX-29-2
IX-29-3
IX-29-3
IX-29-4
IX-29-4
IX-29-5
IX-29-6
IX-29-7
X-30-3
X-30-3
X-30-3
X-30-4
X-30-4
X-30-4
X-30-4
X-30-4
X-30-4
X-30-4
X-30-4
X-30-4
X-30-5
X-30-5
ORIGINAL
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Page
No.
PART XI
PERFORMANCE DATA
APPENDIX A
APPENDIX B
Aircraft Differences(LOTs 21 - 22 Aircraft)
FOLDOUTS
25
ORIGINAL
A1-F18EA-NFM-000
LIST OF ILLUSTRATIONS
Page
No.
PART I
THE AIRCRAFT
CHAPTER 1
The Aircraft
Figure
Figure
Figure
Figure
General Arrangement. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Approximate Dimensions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Radar Cross Section (RCS) Reduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
LOT NUMBER/BUNO . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
1-1
1-2
1-3
1-4
CHAPTER 2
Systems
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Throttle Grips (Front Cockpit) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engine Fuel Display (EFD) - Engine Parameters . . . . . . . . . . . . . . . . . . . .
Engine Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Tank 1 and 4 Fuel CG Control and FUEL XFER Caution Schedule .
Fuel Quantity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engine Fuel Display (EFD) - Fuel Parameters . . . . . . . . . . . . . . . . . . . . . . .
FUEL Display. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FPAS Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Secondary Power Supply . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Simplified Electrical Schematic. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCC Electrical Redundancy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Ground Power Panel and Placard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Circuit Breaker Panels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Exterior Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ID Strobe Patterns . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Hydraulic Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Wheel Brake and Anti-skid System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Emergency/Parking Brake Handle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Flight Control System Functional Diagram . . . . . . . . . . . . . . . . . . . . . . . . . .
Flap Schedules . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
G-Limiter G-Bucket Reductions in Maximum Commandable
G-Level . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Stick and Pedal Travel Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Stick Grip FCS Controls. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
SPIN Recovery Display. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
GAIN ORIDE Flap Positions and Gain Schedules . . . . . . . . . . . . . . . . . .
FCS Related Cautions and Cockpit Indications . . . . . . . . . . . . . . . . . . . . . .
FCS Status Display. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AFCS Controls and Indicators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Stick Grip Switches/Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Throttle Grip Switches/Controls (Front Cockpit) . . . . . . . . . . . . . . . . . . . .
Hand Controllers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
2-1
2-2
2-3
2-4
2-5
2-6
2-7
2-8
2-9
2-10
2-11
2-12
2-13
2-14
2-15
2-16
2-17
2-18
2-20
2-21
2-22
2-23
2-24
2-25
2-26
2-27
2-28
2-29
2-30
2-31
2-32
26
I-1-1
I-1-2
I-1-3
I-1-5
I-2-7
I-2-8
I-2-10
I-2-15
I-2-20
I-2-22
I-2-23
I-2-26
I-2-27
I-2-31
I-2-34
I-2-35
I-2-36
I-2-37
I-2-39
I-2-43
I-2-52
I-2-55
I-2-62
I-2-67
I-2-71
I-2-73
I-2-75
I-2-77
I-2-79
I-2-81
I-2-82
I-2-85
I-2-88
I-2-90
I-2-93
ORIGINAL
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No.
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
2-33
2-34
2-35
2-36
2-37
2-38
2-39
2-40
2-41
2-42
2-43
2-44
2-45
2-46
2-47
2-48
2-49
2-50
2-51
2-52
2-53
2-54
2-55
2-56
2-57
2-58
Figure
Figure
Figure
Figure
Figure
Figure
2-59
2-60
2-61
2-62
2-63
2-64
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
2-65
2-66
2-67
2-68
2-69
2-70
2-71
2-72
2-73
2-74
2-75
2-76
OBOGS Monitor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Canopy Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Boarding Ladder . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
SJU-17 and SJU-17A Ejection Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Leg Restraint System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Survival Kit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HSI-DATA-A/C Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
TAWS HUD Visual Recovery Warning - Pull Up (VRT). . . . . . . . . . . . .
TAWS HUD Visual Recovery Warning - Pull Up (ORT). . . . . . . . . . . . .
TAWS HUD Visual Recovery Warning (VRT) . . . . . . . . . . . . . . . . . . . . . . .
GPWS HUD Roll Warning Cues. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
GPWS Aural Cues. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Angle of Attack Indexer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
MUMI Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
MENU Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Electronic Attitude Display Indicator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8 x 10 Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
MPCD Controls and HSI Symbology . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HUD Controls. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HUD Symbology . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HUD Symbology Degrades . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Alphanumeric Entry Format. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Up Front Control Display (UFCD) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Forward CVRS Control Panel (LOTs 21-22) . . . . . . . . . . . . . . . . . . . . . . . . .
Forward CVRS Control Panel (LOTs 23-25 Before AFC 445) . . . . . . . .
Aft CVRS Control Panel (LOTs 21-22) and LOTs 23-25 Before AFC
445. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Forward CVRS Control Panel (LOTs 23-24 After AFC 445). . . . . . . . . .
Forward CVRS Control Panel (LOT 25 After AFC 445 . . . . . . . . . . . . . .
Aft CVRS Control Panel (LOT 23-25 After AFC 445). . . . . . . . . . . . . . . .
Fwd CVRS Control Panel (LOT 26 AND UP) . . . . . . . . . . . . . . . . . . . . . . .
Aft CVRS Control Panel (LOTs 26-29 Before AFC 445) . . . . . . . . . . . . .
Aft CVRS Control Panel (LOTs 26-29 After AFC 445 AND LOT 30
AND UP) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Video Display Routing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AMU Maintenance Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Map Loading Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Flight Aids Reversion Mechanization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Equipment Status Messages . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Caution/Advisory Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
BIT Control Display. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
MPCD and UFCD Test Patterns . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
EFD Test Pattern . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
UFCD Test Pattern . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
CONFIG Display (Sample) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
INS Postflight Data Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27
I-2-108
I-2-117
I-2-120
I-2-121
I-2-123
I-2-124
I-2-132
I-2-134
I-2-135
I-2-136
I-2-138
I-2-139
I-2-141
I-2-144
I-2-149
I-2-150
I-2-152
I-2-154
I-2-156
I-2-159
I-2-164
I-2-168
I-2-169
I-2-173
I-2-174
I-2-174
I-2-175
I-2-175
I-2-176
I-2-176
I-2-177
I-2-177
I-2-180
I-2-181
I-2-182
I-2-187
I-2-189
I-2-190
I-2-191
I-2-198
I-2-199
I-2-200
I-2-202
I-2-204
ORIGINAL
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Figure
Figure
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Figure
2-77
2-78
2-79
2-80
2-81
2-82
2-83
2-84
JHMCS Upper HVI Routing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HMD Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Boresight Reference Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Displays BIT Sublevel. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HMD/AHMD Test Patterns . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Coarse Alignment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fine Alignment. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Alignment Verification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I-2-207
I-2-209
I-2-210
I-2-211
I-2-212
I-2-216
I-2-219
I-2-221
CHAPTER 4
Operating Limitations
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure 4-12
Figure 4-13
Figure 4-14
Engine Operation Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Basic Aircraft Airspeed Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Subsystem Airspeed Limitations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Gross Weight and Lateral Weight Asymmetry Limitations . . . . . . . . . . .
Asymmetric Stores Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AOA Limitations - Flaps AUTO . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Acceleration Limitations - Basic Aircraft (with or without empty
pylons) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Limitations with Flaps HALF or FULL . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Station 6 480-Gal External Fuel Tank Carriage Limits E . . . . . . . . . . . .
Station 4/8 480-Gal External Fuel Tank Carriage Limits (Rev B
Pylons) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Station 4/8 480-Gal External Fuel Tank Carriage Limits (Rev A
Pylons) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Air Refueling Store Carriage Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ARS Operating Limitations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ATFLIR Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
PART II
INDOCTRINATION
CHAPTER 5
Indoctrination
Figure 5-1
Figure 5-2
Requirements for Various Flight Phases During Initial Training . . . . . II-5-2
Pilot Ceiling and Visibility Restrictions Prior to Instrument
Qualification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . II-5-3
PART III
NORMAL PROCEDURES
CHAPTER 7
Shore-Based Procedures
Figure 7-1
Figure 7-2
Figure 7-3
Exterior Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-7-2
Checklist Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-7-31
Typical Field Landing Pattern. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-7-38
4-1
4-2
4-3
4-4
4-5
4-6
4-7
Figure 4-8
Figure 4-9
Figure 4-10
Figure 4-11
28
I-4-1
I-4-2
I-4-2
I-4-3
I-4-4
I-4-5
I-4-6
I-4-9
I-4-13
I-4-14
I-4-15
I-4-16
I-4-17
I-4-19
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CHAPTER 8
Carrier-Based Procedures
Figure
Figure
Figure
Figure
Figure
Launch Trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Carrier Landing Pattern . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Carrier Controlled Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ACL Mode 1 and 1A Approaches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ACL Mode 2 Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8-1
8-2
8-3
8-4
8-5
III-8-4
III-8-13
III-8-17
III-8-19
III-8-22
CHAPTER 9
Special Procedures
Figure 9-1
Figure 9-2
Figure 9-3
Formation Takeoff Runway Alignments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-9-2
Formations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-9-3
ARS Control Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-9-9
CHAPTER 10
Functional Checkflight Procedures
Figure 10-1
Functional Checkflight Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III-10-3
PART IV
FLIGHT CHARACTERISTICS
CHAPTER 11
Flight Characteristics
Figure 11-1
AHRS Channel Failure Indication and Effects . . . . . . . . . . . . . . . . . . . . . . . IV-11-17
PART V
EMERGENCY PROCEDURES
CHAPTER 12
General Emergencies
Figure 12-1
Warning/Caution/Advisory Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-12-3
CHAPTER 14
Takeoff Emergencies
Figure 14-1
Maximum Weight for 100 fpm Single Engine Rate of Climb . . . . . . . . . V-14-4
CHAPTER 15
Inflight Emergencies
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Spooldown Restart Envelope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Windmill Restart Envelope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Crossbleed Restart Envelope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
APU Restart Envelope
.........................................
Hydraulic Flow Diagram (Aircraft on Deck) . . . . . . . . . . . . . . . . . . . . . . . . .
Hydraulic Subsystems Malfunction Guide . . . . . . . . . . . . . . . . . . . . . . . . . . .
Flight Control Effects Due To Hydraulic Failure . . . . . . . . . . . . . . . . . . . .
Emergency Power Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
External Stores Jettison Chart. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects - Channel 2. . . . . . . . . . . . . . . . . . . . .
15-1
15-2
15-3
15-4
15-5
15-6
15-7
15-8
15-9
15-10
15-11
29
V-15-2
V-15-2
V-15-3
V-15-3
V-15-6
V-15-7
V-15-8
V-15-11
V-15-24
V-15-27
V-15-28
ORIGINAL
A1-F18EA-NFM-000
Page
No.
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
15-12
15-13
15-14
15-15
15-16
15-17
15-18
15-19
15-20
15-21
15-22
15-23
15-24
15-25
15-26
15-27
15-28
15-29
15-30
15-31
Figure 15-32
Figure 15-33
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
15-34
15-35
15-36
15-37
15-38
15-39
15-40
15-41
FCS Failure Indications and Effects - Channel 3. . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects - Channel 4. . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects - Channels 1 and 2 . . . . . . . . . . . . .
FCS Failure Indications and Effects - Channels 1 and 3 . . . . . . . . . . . . .
FCS Failure Indications and Effects - Channels 1 and 4 . . . . . . . . . . . . .
FCS Failure Indications and Effects - Channels 2 and 3 . . . . . . . . . . . . .
FCS Failure Indications and Effects - Channels 2 and 4 . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects - Aileron Channels 1 and 2 or 3
and 4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects - Aileron Channels 1 and 4 or 2
and 3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects - LEF Channels 1 and 4 or 2
and 3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects - AOA Channel 4 . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
CHAPTER 16
Landing Emergencies
Figure
Figure
Figure
Figure
Figure
Figure
Maximum Single Engine Recovery Weight - Military Thrust. . . . . . . . .
Maximum Single Engine Recovery Weight - Maximum Thrust. . . . . . .
Landing Gear Emergency Flow Chart. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Landing Gear Malfunction – Landing Guide . . . . . . . . . . . . . . . . . . . . . . . . .
Field Arresting Gear Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
CV Recovery Matrix. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
16-1
16-2
16-3
16-4
16-5
16-6
V-15-28
V-15-29
V-15-29
V-15-30
V-15-30
V-15-31
V-15-32
V-15-33
V-15-33
V-15-34
V-15-34
V-15-35
V-15-35
V-15-36
V-15-36
V-15-37
V-15-37
V-15-38
V-15-38
V-15-39
V-15-39
V-15-40
V-15-40
V-15-41
V-15-41
V-15-42
V-15-42
V-15-43
V-15-43
V-15-44
V-16-7
V-16-8
V-16-9
V-16-10
V-16-16
V-16-18
CHAPTER 17
Ejection
Figure 17-1
Sink Rate Effects on Minimum Ejection Altitude . . . . . . . . . . . . . . . . . . . . V-17-5
30
ORIGINAL
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Figure
Figure
Figure
Figure
17-2
17-3
17-4
17-5
Airspeed and Bank Angle Effects on Minimum Ejection Altitude . . . .
Airspeed and Dive Angle Effects on Minimum Ejection Altitude . . . . .
Ejection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Ditching Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
V-17-7
V-17-9
V-17-11
V-17-24
PART VI
ALL WEATHER PROCEDURES
CHAPTER 20
Extreme Weather Procedures
Figure 20-1
Icing Danger Zone . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VI-20-2
PART VII
COMM-NAV EQUIPMENT
AND PROCEDURES
CHAPTER 23
Communication-Identification Equipment
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
23-1
23-2
23-3
23-4
23-5
23-6
23-7
23-8
23-9
23-10
23-11
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
23-12
23-13
23-14
23-15
23-16
23-17
23-18
23-19
23-20
23-21
23-22
23-23
23-24
23-25
23-26
23-27
23-28
Aft Cockpit Rudder Pedal Switches - Lot 26 and up . . . . . . . . . . . . . . . . .
Top Level CNI Format on Up Front Control Display (UFCD) . . . . . . .
COMM Sublevel - Ship Maritime Preset . . . . . . . . . . . . . . . . . . . . . . . . . . . .
COMM Sublevel - Guard and Cue Presets . . . . . . . . . . . . . . . . . . . . . . . . . . .
COMM Sublevel - Manual Preset. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
COMM Sublevel - Presets 1 Thru 20 (Fixed Frequency) . . . . . . . . . . . . .
AJ MENU From Fixed Frequency COMM Sublevel . . . . . . . . . . . . . . . . .
Have Quick Fixed Frequency Sublevel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Have Quick WOD Loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Have Quick TNET Loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Single Channel Ground and Airborne Radio System (SINCGARS)
MENU Fixed Frequency Sublevel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
COMM Sublevel - Presets 1 to 20 (Anti-Jam). . . . . . . . . . . . . . . . . . . . . . . .
AJ MENU From Have Quick COMM Sublevel . . . . . . . . . . . . . . . . . . . . . .
AJ Have Quick Menu Sublevel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AJ MENU From SINCGARS COMM Sublevel . . . . . . . . . . . . . . . . . . . . . .
AJ SINCGARS ERF Sublevel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Guard Transmit on Top Level CNI Format. . . . . . . . . . . . . . . . . . . . . . . . . .
Guard Transmit on COMM Sublevel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
KY-58 Control Panel Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
DCS Frequencies Conversions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
DCS KEY Display. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Relay Bandwidth Limitations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
MUMI Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
BIT Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IFF Transponder UFCD Sublevel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Mode 3 and Mode C Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Electronic Boresight Constant (EBC) Display with W on W . . . . . . . . .
XPOND Format Selection to Enable Mode S Enhanced and
Squitter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
31
VII-23-3
VII-23-5
VII-23-9
VII-23-10
VII-23-11
VII-23-12
VII-23-13
VII-23-15
VII-23-16
VII-23-17
VII-23-18
VII-23-19
VII-23-20
VII-23-21
VII-23-22
VII-23-22
VII-23-24
VII-23-24
VII-23-25
VII-23-27
VII-23-30
VII-23-31
VII-23-33
VII-23-34
VII-23-36
VII-23-37
VII-23-39
VII-23-40
ORIGINAL
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Figure
Figure
Figure
Figure
Figure
23-29
23-30
23-31
23-32
23-33
Return to XPOND Format with Squitter Disabled . . . . . . . . . . . . . . . . . .
Mode S UFCD Sublevel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Mode S Address Not Available on XPOND Sublevel . . . . . . . . . . . . . . . .
Mode S A/C Call Sign Sublevel. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
IFF PROG EDIT Sublevel - POS Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
CHAPTER 24
Navigation Equipment
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Navigation Controls and Indicators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
TAMMAC Mode Selections . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
GPS Page . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
INS CV Align . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
INS Ground Align . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
INS In-Flight Align . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
NAVCK Display. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
GPS Waypoint Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
INS Programming . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Position Keeping . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Position Updating Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Waypoint/OAP Direct Great Circle Steering . . . . . . . . . . . . . . . . . . . . . . . . .
Waypoint/OAP Course Line Steering . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AUTO Sequential Steering . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Navigation Designation (WYPT DSG) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Overfly Designation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
TACAN Mode Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
TACAN Programming. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
TACAN Direct Great Circle Steering . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
TACAN Course Line Steering . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ICLS Initialization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ICLS Steering . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
DDI SA ACL Display. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
HUD ACL Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Traffic Control Couple Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ACL Mode 1 Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ACL Mode 2 Steering Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
T/C Guidance to Marshal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
ACL Control - Marshal to Touchdown. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-1
24-2
24-3
24-4
24-5
24-6
24-7
24-8
24-9
24-10
24-11
24-12
24-13
24-14
24-15
24-16
24-17
24-18
24-19
24-20
24-21
24-22
24-23
24-24
24-25
24-26
24-27
24-28
24-29
VII-23-41
VII-23-42
VII-23-43
VII-23-45
VII-23-46
VII-24-3
VII-24-5
VII-24-12
VII-24-15
VII-24-16
VII-24-21
VII-24-24
VII-24-27
VII-24-31
VII-24-37
VII-24-40
VII-24-42
VII-24-43
VII-24-46
VII-24-48
VII-24-49
VII-24-53
VII-24-55
VII-24-56
VII-24-57
VII-24-58
VII-24-60
VII-24-62
VII-24-65
VII-24-67
VII-24-68
VII-24-70
VII-24-71
VII-24-74
CHAPTER 25
Backup/Degraded Operations
Figure 25-1
Figure 25-2
SDC Backup HUD Display (AMCD Aircraft) . . . . . . . . . . . . . . . . . . . . . . . . VII-25-2
Standby Attitude Reference Indicator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-25-4
CHAPTER 26
Visual Communications
Figure 26-1
Visual Communications. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-26-1
32
ORIGINAL
A1-F18EA-NFM-000
Page
No.
CHAPTER 27
Deck/Ground Handling Signals
Figure 27-1
Deck Ground Handling Signals . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII-27-2
PART VIII
WEAPONS SYSTEMS
PART IX
FLIGHT CREW COORDINATION
CHAPTER 29
Crew Coordination Standards
Figure 29-1
Figure 29-2
Approach Briefs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-6
Standard ICS Terms . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX-29-8
PART X
NATOPS EVALUATION
PART XI
Figure B2−41
Figure B2-60
Figure B15-37
Figure B15-38
Figure B15-39
Figure B15-40
Figure B15-43
Figure B15-44
Figure B-FO-1
Figure B-FO-2
PERFORMANCE DATA
CFIT Conditions, Voice Warning, and Repetition Rates . . . . . . . . . . . . .
HMD Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FCS Failure Indications and Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Cockpit F/A-18E/F . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Rear Cockpit F/A-18F . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
B-3
B-4
B-7
B-7
B-8
B-8
B-9
B-9
B-10
B-12
Cockpit F/A-18E/F (Lots 21 thru 25) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Rear Cockpit F/A-18F (Lots 21 thru 25) . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Cockpit F/A-18E/F (Lot 26 AND UP) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Rear Cockpit F/A-18F (Lot 26 AND UP) . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Electrical Bus Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Ejection Seat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Environmental Control System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FO-3
FO-7
FO-11
FO-15
FO-19
FO-21
FO-25
FO-29
FOLDOUTS
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
FO-1
FO-2
FO-3
FO-4
FO-5
FO-6
FO-7
FO-8
33 (Reverse Blank)
ORIGINAL
A1-F18EA-NFM-000
RECORD OF CHANGES
Change No. and
Date of Change
Date of Entry
35 (Reverse Blank)
Page Count Verified by
(Signature)
ORIGINAL
A1-F18EA-NFM-000
INTERIM CHANGE SUMMARY
The following Interim Changes have been canceled or previously incorporated in this manual:
INTERIM
CHANGE
NUMBER
1 thru 27
REMARKS/PURPOSE
Previously Incorporated or Canceled
The following Interim Changes have been incorporated in this Change/Revision:
INTERIM
CHANGE
NUMBER
REMARKS/PURPOSE
28
Cockpit Smoke and Fumes/Single GEN FAIL
29
AOA FCS Procedures, Cabin Smoke and Fumes Boldface
30
Aerobraking Technique
31
FCC S/W Update/CPWS FCF Fix
32
Multi-subject IC
Interim Changes Outstanding - To be maintained by the custodian of this manual:
INTERIM
CHANGE
NUMBER
33
34
ORIGINATOR/DATE
(or DATE/TIME GROUP)
082001Z JUN 09
262001Z AUG 09
PAGES
AFFECTED
I-4-20, I-4-21,
III-8-23, VI19-2
REMARKS/PURPOSE
AESA Radar Limitations and
Considerations
I-2-83 I-4-10/ FCC Memory Inspect Warning
11, III-7-39
37 (Reverse Blank)
ORIGINAL W/IC 34
P 262001Z AUG 09
FM COMNAVAIRSYSCOM PATUXENT RIVER MD//4.0P//
TO ALL HORNET AIRCRAFT ACTIVITIES
INFO COMNAVAIRSYSCOM PATUXENT RIVER MD//5.0F/4.1//
COMNAVAIRFOR SAN DIEGO CA//N455/N421B//
COMNAVSAFECEN NORFOLK VA//11//
PEOTACAIR PATUXENT RIVER MD//PMA265//
FLTREADCEN SOUTHWEST SAN DIEGO CA//6.2//
FLTREADCEN SOUTHEAST JACKSONVILLE FL//3.3.3//
STRKFITRON ONE TWO TWO
DCMA BOEING ST. LOUIS//RDOAA/RDDF/RDDP//
SECINFO/U/-//
MSGID/GENADMIN,USMTF,2008/COMNAVAIRSYSCOM AIR-4.0P//
SUBJ/FA-18EF AIRCRAFT NATOPS PUBLICATIONS INTERIM CHANGE//
REF/A/DESC:EML/COMNAVAIRFOR/20AUG2009/-/NOTAL//
REF/B/DESC:DOC/COMNAVAIRSYSCOM/20AUG2009/-NOTAL//
REF/C/DESC:A1-F18EA-NFM-000/COMNAVAIRSYSCOM/15SEP2008//
REF/D/DESC:A1-F18EA-NFM-500/COMNAVAIRSYSCOM/15SEP2008//
NARR/REF A IS COG CONCURRENCE.
REF B IS AIRS 2009-184.
REF C IS NATOPS FLIGHT MANUAL NAVY MODEL F/A-18E/F 165533 AND UP
AIRCRAFT, DTD 15 SEP 2008.
REF D IS NATOPS POCKET CHECKLIST F/A-18E/F AIRCRAFT, DTD 15 SEP 2008.//
POC/THOMAS ELLIS/FC FACILITATOR/UNIT:PMA-265/NAME:PATUXENT RIVER MD
/TEL:301-342-3149/TEL:DSN 342-3149/EMAIL:[email protected]//
GENTEXT/REMARKS/1. THIS MESSAGE IS ISSUED IN RESPONSE TO REFS A AND
B. THIS MESSAGE ISSUES INTERIM CHANGE (IC) NUMBER 34 TO REF C AND
IC NUMBER 27 TO REF D.
2. SUMMARY.
A. THESE CHANGES TO REFS C AND D ADD WARNINGS FOR IN-FLIGHT
MEMORY INSPECT (MI) OF FLIGHT CONTROL COMPUTER (FCC).
B. REPLACEMENT PAGES CONTAINING THESE CHANGES FOR DOWNLOADING
AND INSERTION INTO REFS C AND D WILL BE ATTACHED TO THIS
INTERIM CHANGE MESSAGE WHEN IT IS POSTED ON THE NATEC AND
AIRWORTHINESS WEBSITES (SEE LAST PARA BELOW).
3. THE REPLACEMENT PAGES IMPACT THE FOLLOWING NATOPS FLIGHT
MANUAL AND ASSOCIATED CHECKLIST. THE REPLACEMENT PAGE PACKAGE
INCLUDES THE FOLLOWING PAGES:
A. REF C (F/A-18E/F NFM), PAGES: 37(REVERSE BLANK), I-2-83,
I-2-84, I-4-9 THRU I-4-12, III-7-39 AND III-7-40.
B. REF D (F/A-18E/F PCL), PAGES: B/(C BLANK), 35 AND 36.
C. TO ENSURE THE PDF PAGES PRINT TO SCALE: SELECT PRINT
AND VIEWING PRINT SETUP WINDOW, ENSURE 'NONE' IS
SELECTED IN THE PAGE SCALING DROPDOWN.
4. POINTS OF CONTACT:
A. LT DAMON LOVELESS, VFA-122,NATOPS PROGRAM MANAGER,TEL DSN
949-1960 OR COMM (559)998-1960,EMAIL:[email protected]
B. NAVAIR POCS:
(1) MARTY SCANLON, NATOPS IC COORDINATOR, TEL DSN 757-6045
OR COMM (301) 757-6045, EMAIL: [email protected]
(2) ED HOVANESIAN, F/A-18E/F CLASS DESK, TEL DSN 757-7573 OR
COMM (301)757-7573, EMAIL:EDWIN.HOVANESIAN(AT)NAVY.MIL.
(3) KRISTIN SWIFT, AIR-4.0P NATOPS CHIEF ENGINEER,TEL DSN
995-4193 OR (301)995-4193,EMAIL:KRISTIN.SWIFT(AT)NAVY.MIL.
(4) LCDR BEN KELSEY, 4.0P NATOPS OFFICER, TEL DSN 995-2502
OR COMM (301) 995-2502, EMAIL:[email protected]
_______________________________________________________________________
NAVAIR 262001Z AUG 09
Page 1 of 2
A1-F18EA-NFM-000 IC 34
A1-F18EA-NFM-500 IC 27
(5) AIRWORTHINESS GLOBAL CUSTOMER SUPPORT TEAM,
TEL: 301-757-0187, E-MAIL:AIRWORTHINESS(AT)NAVY.MIL.
5. THIS MESSAGE WILL BE POSTED ON THE NATEC WEBSITE,
WWW.MYNATEC.NAVY.MIL WITHIN 48 HOURS OF RELEASE. NEW NATOPS IC
MESSAGES MAY BE FOUND IN TWO PLACES ON THIS WEBSITE:
A. IN THE NATOPS IC DATABASE FOUND UNDER THE TMAPS OPTION.
B. IN THE AFFECTED PUBLICATIONS(S) JUST AFTER THE IC SUMMARY
PAGE. IF THE IC MESSAGE INCLUDES REPLACEMENT PAGES, THEY
WILL BE ADDITIONALLY PLACED WITHIN THE MANUAL AND REPLACED
PAGES DELETED. MESSAGES ARE NORMALLY POSTED IN THE DATABASE
BEFORE APPEARING IN THE PUBLICATION. THIS MESSAGE WILL ALSO
BE POSTED ON THE AIRWORTHINESS WEBSITE,
AIRWORTHINESS.NAVAIR.NAVY.MIL. IF UNABLE TO VIEW THIS MESSAGE ON
EITHER THE NATEC OR AIRWORTHINESS WEBSITES, INFORM THE NATOPS
GLOBAL CUSTOMER SUPPORT TEAM AT (301) 342-0870, DSN 342-0870, OR
BY EMAIL AT NATOPS(AT)NAVY.MIL.
C. INFORMATION REGARDING THE AIRWORTHINESS PROCESS, INCLUDING
A LISTING OF ALL CURRENT INTERIM FLIGHT CLEARANCES, NATOPS
AND NATIP PRODUCTS ISSUED BY NAVAIR 4.0P, CAN BE FOUND AT
OUR WEBSITE: AIRWORTHINESS.NAVAIR.NAVY.MIL.
D. E-POWER FOLDER NUMBER 874764. AIRWORTHINESS TRACKING NUMBER
35317.//
BT
#0001
NNNN
KRISTIN SWIFT, NATOPS CHIEF ENGINEER, 4.0P, 08/26/2009
_______________________________________________________________________
NAVAIR 262001Z AUG 09
Page 2 of 2
A1-F18EA-NFM-000 IC 34
A1-F18EA-NFM-500 IC 27
P 082001Z JUN 09
FROM COMNAVAIRSYSCOM PATUXENT RIVER MD//4.0P//
TO ALL HORNET AIRCRAFT ACTIVITIES
INFO COMNAVAIRSYSCOM PATUXENT RIVER MD//5.0F/4.1//
COMNAVAIRFOR SAN DIEGO CA//N455/N421B//
COMNAVSAFECEN NORFOLK VA//11//
PEOTACAIR PATUXENT RIVER MD//PMA265//
FLTREADCEN SOUTHWEST SAN DIEGO CA//6.2//
FLTREADCENSOUTHEAST JACKSONVILLE FL//3.3.3//
STRKFITRON ONE TWO TWO
DCMA BOEING ST. LOUIS OPS//RDOAA/RDDF/RDDP//
SECINFO/U/-//
MSGID/GENADMIN,USMTF,2008/COMNAVAIRSYSCOM AIR-4.0P//
SUBJ/FA-18EF AIRCRAFT NATOPS PUBLICATIONS INTERIM CHANGE//
REF/A/DESC:EML/COMNAVAIRFOR/01JUN2009//
REF/B/DESC:DOC/COMNAVAIRSYSCOM/25MAY2009//
REF/C/DESC:A1-F18EA-NFM-000/COMNAVAIRSYSCOM/15SEP2008//
REF/D/DESC:A1-F18EA-NFM-500/COMNAVAIRSYSCOM/15SEP2008//
NARR/REF A IS COG CONCURRENCE.
REF B IS AIRS 2009-113.
REF C IS NATOPS FLIGHT MANUAL, A1-F18EA-NFM-000, DTD 15 SEP 2008.
REF D IS NATOPS POCKET CHECKLIST (PCL), A1-F18EA-NFM-500, DTD
15 SEP 2008.//
POC/THOMAS ELLIS/FC FACILITATOR/UNIT:PMA-265/NAME:PATUXENT RIVER MD
/TEL:301-342-3149/TEL:DSN 342-3149/EMAIL:[email protected]//
GENTEXT/REMARKS/1. THIS MESSAGE IS ISSUED IN RESPONSE TO REFS A AND
B. THIS MESSAGE ISSUES INTERIM CHANGE (IC) NUMBER 33 TO REF C, AND
IC NUMBER 26 TO REF D.
2. SUMMARY.
A. THESE CHANGES TO REFS C AND D INCORPORATE AESA RADAR
LIMITATIONS.
B. REPLACEMENT PAGES CONTAINING THESE CHANGES FOR DOWNLOADING
AND INSERTION INTO REFS C AND D WILL BE ATTACHED TO THIS
INTERIM CHANGE MESSAGE WHEN IT IS POSTED ON THE NATEC AND
AIRWORTHINESS WEBSITES (SEE LAST PARA BELOW).
3. THE REPLACEMENT PAGES IMPACT THE FOLLOWING NATOPS FLIGHT
MANUAL AND ASSOCIATED CHECKLIST. THE REPLACEMENT PAGE PACKAGE
INCLUDES THE FOLLOWING PAGES:
A. REF C (F/A-18E/F NFM), PAGES: 37(REVERSE BLANK), I-4-19,
I-4-20, I-4-21(REVERSE BLANK) III-8-23(REVERSE BLANK),
VI-19-1 AND V1-19-2.
B. REF D (F/A-18E/F PCL), PAGES: B/(C BLANK), 44A AND 44B.
C. TO ENSURE THE PDF PAGES PRINT TO SCALE: SELECT PRINT
AND VIEWING PRINT SETUP WINDOW, ENSURE "NONE" IS
SELECTED IN THE PAGE SCALING DROPDOWN.
4. POINTS OF CONTACT:
A. LT DAMON LOVELESS, VFA-122, NATOPS PROGRAM MANAGER, TEL
DSN 949-1960 OR COMM (559) 998-1960, EMAIL:
[email protected]
B. NAVAIR POCS:
(1) MARTY SCANLON, NATOPS IC COORDINATOR, TEL DSN 757-6045
OR COMM (301) 757-6045, EMAIL: [email protected]
(2) ED HOVANESIAN, F/A-18E/F CLASS DESK, TEL DSN
757-7573 OR COMM (301) 757-7573, EMAIL:
[email protected]
(3) KRISTIN SWIFT, AIR-4.0P NATOPS CHIEF ENGINEER, TEL
DSN 995-4193 OR COMM (301) 995-4193,
_______________________________________________________________________
NAVAIR 082001Z JUN 09
Page 1 of 2
A1-F18EA-NFM-000 IC 33
A1-F18EA-NFM-000 IC 26
EMAIL: [email protected]
(5) LCDR BEN KELSEY, 4.0P NATOPS OFFICER,
DSN 995-2502, COM 301-995-2505, EMAIL:
[email protected]
(4) AIRWORTHINESS GLOBAL CUSTOMER SUPPORT TEAM,
COMM (301) 757-0187, E-MAIL: [email protected]
5. THIS MESSAGE WILL BE POSTED ON THE NATEC WEBSITE,
WWW.MYNATEC.NAVY.MIL WITHIN 48 HOURS OF RELEASE. NEW NATOPS IC
MESSAGES MAY BE FOUND IN TWO PLACES ON THIS WEBSITE:
A. IN THE NATOPS IC DATABASE FOUND UNDER THE TMAPS OPTION.
B. IN THE AFFECTED PUBLICATIONS(S) JUST AFTER THE IC SUMMARY
PAGE. IF THE IC MESSAGE INCLUDES REPLACEMENT PAGES, THEY
WILL BE ADDITIONALLY PLACED WITHIN THE MANUAL AND REPLACED
PAGES DELETED. MESSAGES ARE NORMALLY POSTED IN THE DATABASE
BEFORE APPEARING IN THE PUBLICATION. THIS MESSAGE WILL ALSO
BE POSTED ON THE AIRWORTHINESS WEBSITE,AIRWORTHINESS.NAVAIR.
NAVY.MIL. IF UNABLE TO VIEW THIS MESSAGE ON EITHER THE NATEC
OR AIRWORTHINESS WEBSITES, INFORM THE NATOPS GLOBAL CUSTOMER
SUPPORT TEAM AT (301) 342-0870, DSN 342-0870, OR BY EMAIL AT
[email protected]
C. INFORMATION REGARDING THE AIRWORTHINESS PROCESS, INCLUDING
A LISTING OF ALL CURRENT INTERIM FLIGHT CLEARANCES, NATOPS
AND NATIP PRODUCTS ISSUED BY NAVAIR 4.0P, CAN BE FOUND AT OUR
WEBSITE: AIRWORTHINESS.NAVAIR.NAVY.MIL.
D. E-POWER FOLDER NUMBER 860893. AIRWORTHINESS TRACKING NUMBER
34551.//
KRISTIN SWIFT, NATOPS CHIEF ENGINEER, 4.0P 06/08/2009
_______________________________________________________________________
NAVAIR 082001Z JUN 09
Page 2 of 2
A1-F18EA-NFM-000 IC 33
A1-F18EA-NFM-000 IC 26
A1-F18EA-NFM-000
SUMMARY OF APPLICABLE TECHNICAL
DIRECTIVES
Information relating to the following technical directives has been incorporated in this manual
Change
Number
ECP
Number
Visual Identification
Effectivity
Multifunctional Information
Distribution System (MIDS)
updates for F/A-18E/F aircraft
Comm switch change on throttle
(R) LOT 22 LOT 23
(P) LOT 24 &
UP
6061
Automatic Direction Finder
(ADF) replacement
None
6022
Incorporation of Advanced
Mission Computers and
Displays
DDIs with Contrast Rocker
Switch.
(P) LOT 25 &
UP
6038R1
Lot 26 Structural and System
Provisions for Block 2
Front cockpit Video Record
Panel, Rear cockpit Hand Controllers.
(P) LOT 16 &
UP
6038R2
Lot 26 Avionics Installation
Provisions for Block 2
TAMMAC Digital Video Map
Computer (DVMC), Rear cockpit 8 x 10 Display.
(R) LOT 26 LOT 27
(P) LOT 28 &
UP
6176R2
Omnibus Software Update
DDI Configuration Display:
MC1/MC2 H2E+
(R) LOT 25 &
UP
(R) LOT 26 LOT 28
(P) LOT 29 &
UP
0577R3
Description
(P) F/A-18E
165867 & UP
(P) F/A-18F
165883 & UP
AFC 443
6212
Solid State Recorder
Secure Erase button on right
hand forward vertical console.
AFC 365
AFC 366
6165
Canopy Switch Guard, Boarding
Ladder Switch
Switches inside Door 9
(R)LOT 21 LOT 25
6217
Cabin Pressurization Warning
System
Caution Lights Panel
(P) F/A-18E
166784 & UP
(P) F/A-18F
166804 & UP
(R) Retrofit
(P) Production
39
CHANGE 3
A1-F18EA-NFM-000
Change
Number
ECP
Number
Visual Identification
Effectivity
6201
ANAV
None
(P) LOT 30 &
UP
6272
Omnibus Software Update
DDI Configuration Display:
MC1/MC2 H4E
(P) LOT 30 &
UP
Aft Seat JHMCS
BRU mounted on aft UFCD
(P) LOT 30 &
UP
(R) LOT 23 29
AN/AVX-4 Fast Tactical Imaging Set
Digital Imaging Processor on
RH console
(R) LOT 24
570R4
AFC 395
Description
960
Information relating to the following recent technical directives will be incorporated in the future
change.
Change
Number
ECP
Number
Description
Visual Identification
40
Effectivity
CHANGE 5
A1-F18EA-NFM-000
LIST OF ABBREVIATIONS/ACRONYMS
A/A
AACQ
AB
AB LIM
A/C
ac
ACCUM
ACI
ACLS
ACM
ACNTR
ADB
ADF
ADV
AFCS
A/G
AGI
AGL
AHRS
AIL
AIM
AINS
ALDDI
ALR-67
ALT
AMAD
AMCD
AMPCD
AMPD
AMU
AN/ALE-47
AN/APN-194
AN/ASN-139
AOA
AOB
A/P
APU
AQ
ARDDI
ARI
ARS
ASL
ASRM
A
air-to-air
automatic acquisition mode
afterburner
afterburner limiting
aircraft
alternating current
accumulator
amplifier control intercommunication
automatic carrier landing system
air combat maneuvering
aft center (8 x 10) display
aircraft discrepancy book
automatic direction finding
advisory
automatic flight control system
air-to-ground
armament gas ingestion
above ground level
attitude heading reference set
aileron
air intercept missile
aided INS
aft left digital display indicator
radar warning receiver
altitude
airframe mounted accessory drive
advanced mission computers and displays
aft multipurpose color display
advanced multipurpose display
advanced memory unit
countermeasures dispensing set
radar altimeter set
inertial navigation system
angle of attack
angle of bank
autopilot
auxiliary power unit
align quality
aft right digital display indicator
attitude reference indicator
air refueling store
azimuth steering line
automatic spin recovery mode
41
ORIGINAL
A1-F18EA-NFM-000
ATARS
ATS
ATSCV
ATTH
AUFCD
AUG
AUR
AUTO
AVMUX
BALT
BCN
BINGO
BIT
BLD
BLIM
BLIN
BNK
BRG
BRK
BRT
BST
°C
CAS
CAUT
CB
CCW
CD
CDP
C/F
CFIT
CG
CH or CHAN
CHKLST
CIT
CK
CKPT
CLR
CNI
COMM
CONT PVU
CPL
CPLD
CPWS
CRS
CSC
CSEL
CV
advanced tactical air reconnaissance system
air turbine starter
air turbine starter control valve
attitude hold
aft upfront control display
augment
aural
automatic
avionics multiplex
B
barometric altimeter
beacon
minimum fuel required to return to base
built in test
bleed
bank limit
BIT logic inspection
bank
bearing
brake
bright
boresight acquisition mode
C
degrees Celsius
control augmentation system
caution
circuit breaker
counterclockwise
countdown
compressor discharge pressure
chaff/flare
controlled flight into terrain
center of gravity
channel
checklist
combined interrogator /transponder
check
cockpit
clear
communication, navigation, and identification
communication radio
continuous precision velocity update
couple
coupled
cabin pressurization warning system
course
communication system control
course select
carrier
42
ORIGINAL
A1-F18EA-NFM-000
CVRS
DA
DBS
DBFS
DC
dc
DCS
DDI
DECD
DEGD
Δ P (Delta P)
DF
DISCH
D/L
DME
DMS
DN
DSU
DTED
DT2
DVMC
EADI
EBB
ECS
EFD
EFT
EGT
EMCON
EMIS
ENG
ENT
EPR
ERF
EST
ET
EXT
EXTD
°F
FADEC
FCC
FCCA
FCCB
FCES
FCF
FCLP
cockpit video recording system
D
density altitude
doppler beam sharpening
dry bay fire suppression
designator controller
direct current
decompression sickness
digital display indicator
digital expandable color display
degraded
hydraulic filter indicator
direction finding
discharge
data link
distance measuring equipment
digital map set
down
data storage unit
digital terrain elevation data
second designated target
digital video map computer
E
electronic attitude display indicator
essential bus backup
environmental control system
engine fuel display
external fuel tank
exhaust gas temperature
emission control
electro magnetic interference shield
engine
enter
engine pressure ratio
electronic fill remote
estimated
elapsed time
external
extend
F
degrees Fahrenheit
full authority digital engine control
flight control computer
flight control computer A
fight control computer B
flight control electronic system
functional checkflight
field carrier landing practice
43
ORIGINAL
A1-F18EA-NFM-000
FCNS
FCS
FE
FF
FIP
FIRAMS
FLBIT
FLIR
FO
FOD
FOV
FPAH
FPAS
fpm
FPT
F-QTY
FRS
ft
FUS
G or g
GACQ
GB
GCU
GEN
GEN TIE
G-LIM
G-LOC
GND
GPS
GPWS
GRCV
GW
GXMT
HDG
HDG/SLV
HI
HMD
HOBS
HOTAS
HQ
HRC
HSEL
HSI
HSIB
HSVN
HUD
fiber channel network switch
flight control system
fighter escort configuration
fuel flow
form-in-place
fight incident recording and monitoring system
fuel low BIT
forward looking infrared
foldout
foreign object damage
field of view
flight path attitude hold
flight performance advisory system
feet per minute
first pilot time
fuel quantity
fleet replacement squadron
foot, feet
fuselage
G
gravity
gun acquisition mode
gyro bias
generator converter unit
generator
generator tie
g-limiter
g-induced loss of consciousness
ground
global positioning system
ground proximity warning system
guard receive
gross weight
guard transmit
H
heading
heading slaved
high
helmet mounted display
high off-boresight
hands on throttle and stick
have quick
helmet release connector
heading select
horizontal situation indicator
high speed interface bus
high speed video network
head-up display
44
ORIGINAL
A1-F18EA-NFM-000
HYD
HYD1
HYD2
IAF
IBIT
ICLS
ICS
ID
IDECM
IFA
IFF
IFR
ILS
IMC
IMN
IMU
INOP
INS
INST
INV
IP
I/P (IDENT)
IR
IRC
ISOL
ITB
IWSO
JETT
KCAS
KGS
KIAS
kt
KTAS
L
lb(s)
L&S
L ACC
LAMPD
LATLN
LBA
L BAR
LCS
LDC
hydraulic, hydraulic system
hydraulic system 1
hydraulic system 2
I
initial approach fix
initiated built in test
instrument carrier landing system
intercockpit communication system
identification
integrated defensive electronic countermeasures
inflight alignment
identification friend or foe
instrument flight rules
instrument landing system
instrument meteorological conditions
indicated mach number
inertial measurement unit
inoperative
inertial navigation system
instrument
invalid
instructor pilot
identification of position
infrared
in-line release connector
isolate
image transfer bus
instructor WSO
J
jettison
K
knots calibrated air speed
knots ground speed
knots indicated air speed
knots
knots true airspeed
L
left
pound(s)
launch & steering target
lateral accelerometer
left advanced multipurpose display
latitude longitude
limit basic aircraft
launch bar
liquid cooling system
left designator controller
45
ORIGINAL
A1-F18EA-NFM-000
LDDI
LDG
LED
LEF
LEU
LEX
LG
LI
LM
LO
LO
LON
LPU
LSO
LT
LTOD
MAC
MAD
MAX
MC
MC1
MC2
MER
MFS
MIDS
MIL
min
MMP
MNTCD
MPCD
MSL
MSNCD
MSRM
MTRS or m
MU
MUMI
MUX
MVAR
N1
N2
N ACC
NACES
NATOPS
NAV
ND
nm
left digital display indicator
landing
leading edge down
leading edge flaps
leading edge up
leading edge extension
landing gear
left inboard
left midboard
left outboard
low
limit of NATOPS
life preserver unit
landing signal officer
light
local time of day
M
mean aerodynamic chord
magnetic azimuth detector
maximum afterburner thrust
mission computer
mission computer 1
mission computer 2
multiple ejector rack
multifunction switch
multifunctional information distribution system
military thrust
minimum, minutes
maintenance monitor panel
maintenance card
multipurpose color display
mean sea level
mission card
manual spin recovery mode
meters
memory unit
memory unit mission initialization
multiplex bus
magnetic variation
N
fan rpm
compressor rpm
normal accelerometer
navy aircrew common ejection seat
naval air training and operations procedures standardization
navigation
nose down
nautical miles
46
ORIGINAL
A1-F18EA-NFM-000
NORM
NOTAMS
NOZ
NU
NVD
NVIS
NWS
Nz REF
OAP
OAT
OBOGS
OCF
OFP
ORIDE
OVFLY
OVRSPD
OXY
PA
PBIT
PC
P CAS
PCL
PIO
PLF
PMG
PNL
POS
pph
ppm
PR
PROC
psi
PTS
PTS
QDC
QTY
R
RALT
RAM
RAMPD
RAT
RATS
R CAS
RCDR
RCS
normal
notice(s) to airmen
nozzle
nose up
night vision devices
night vision imaging system
nosewheel steering
reference load factor
O
offset aim point
outside air temperature
onboard oxygen generating system
out of control flight
operational flight program
override guide
overfly
overspeed
oxygen
P
powered approach
periodic BIT
plane captain
pitch control augmentation system
pocket checklist
pilot induced oscillation
parachute landing fall
permanent magnet generator
panel
position
pounds per hour
pounds per minute
pressure
processor
pounds per square inch
power transmission shaft
pressure transmitter set
Q
quick disconnect connector
quantity
R
right
radar altimeter
radar absorbing material
right advanced multipurpose display
ram air turbine
reduced authority thrust system
roll control augmentation system
recorder
radar cross section
47
ORIGINAL
A1-F18EA-NFM-000
RCVY
RDC
RDDI
RDR
REC
REC
RECCE
REJ
RI
RLG
RLS
R-LIM
RM
RMM
RNG
RO
ROE
ROMA
RP
rpm
RSET
RSRI
R/T
RUD
RVSM
RWR
RWSO
SCT
SDC
SDCR
SEAWARS
SEC
SEQ
SMS
SOP
SPD
SPD BRK
SPN
SRM
SSR
STAB
STBY
STD HDG
STT
SUPT
S/W
SW
recovery
right designator controller
right digital display indicator
radar
radar elevation control
record or receive
reconnaissance
reject
right inboard
ring laser gyro
reservoir level sensing
roll rate limiter
right midboard
removable memory module
range
right outboard
rules of engagement
removable optics module assembly
replacement pilot
revolutions per minute
reset
rolling-surface-to- rudder interconnect
receive/transmit
rudder
reduced vertical separation minimum
radar warning receiver
replacement WSO
S
special crew time
signal data computer
signal data computer replacement
seawater parachute release mechanism
source error correction
sequence
stores management set
standard operating procedures
speed
speedbrake
spin
spin recovery mode
solid state recorder
stabilator
standby
stored heading
single target track
support
software
switch
48
ORIGINAL
A1-F18EA-NFM-000
T1
TAC
TAMMAC
TAS
TAWS
TCN or TACAN
TCV
TDC
TDP
TED
TEF
TEU
TEMP
T&G
THA
TK PRESS
T/O
TOT
TR
TTG
UA
UFCD
UHF
UNLK
UPDT
UTM
vac
VACQ
vdc
VEL
VER
VFR
VHF
VIB
VMC
VOL
VTR
VVSLV
W
WACQ
W DIR
W SPD
WARN
WDSHLD
T
engine inlet temperature
tactical
tactical aircraft moving map capability
true air speed
terrain awareness warning system
tactical air navigation
thermal control valve
throttle designator controller
turbine discharge pressure
trailing edge down
trailing edge flaps
trailing edge up
temperature
touch and go
throttle handle angle
fuel tank pressure
takeoff
time on target
transformer rectifier
time to go
U
up-AUTO
upfront control display
ultra high frequency
unlock
update
universal transverse mercator
V
volts alternating current
vertical acquisition mode
volts direct current
velocity
vertical ejector rack
vsual flight rules
very high frequency
vibration
visual meteorological conditions
volume
video tape recorder
velocity vector slave
W
waterline symbol
wide acquisition mode
wind direction
wind speed
warning
windshield
49
ORIGINAL
A1-F18EA-NFM-000
WonW
WoffW
WSO
WYPT
weight on wheels
weight off wheels
Weapon Systems Officer
waypoint
XFER
transfer
X
Y CAS
yd
ZTOD
Y
yaw control augmentation system
yards
Z
zulu time of day
50
ORIGINAL
A1-F18EA-NFM-000
PREFACE
SCOPE
The NATOPS Flight Manual is issued by the authority of the Chief of Naval Operations and under
the direction of Commander, Naval Air Systems Command in conjunction with the Naval Air Training
and Operating Procedures Standardization (NATOPS) Program. This manual contains information on
all aircraft systems, performance data, and operating procedures required for safe and effective
operations, however, it is not a substitute for sound judgement. Compound emergencies, available
facilities, adverse weather or terrain, or considerations affecting the lives and property of others may
require modification of the procedures contained herein. Read this manual from cover to cover as it is
each aircrew’s responsibility to have a complete knowledge of its contents.
APPLICABLE PUBLICATIONS
The following applicable publications complement this manual:
•
•
•
•
•
•
•
•
•
A1-F18EA-NFM-200 (Performance Data)
A1-F18EA-NFM-500 (Pocket Checklist)
A1-F18EA-NFM-600 (Servicing Checklist)
A1-F18EA-NFM-700 (Functional Checkflight Checklist)
A1-F18EA-TAC-000 (Volume I Tactical Manual)
A1-F18EA-TAC-010 (Volume II Tactical Manual)
A1-F18EA-TAC-100 (Volume III Tactical Manual)
A1-F18EA-TAC-020 (Volume IV Tactical Manual)
A1-F18EA-TAC-300 (Tactical Manual Pocket Guide)
HOW TO GET COPIES
Additional copies of this manual and changes thereto may be procured through the local supply
system from NAVICP Philadelphia via DAAS in accordance with NAVSUP P−409 (MILSTRIP/
MILSTRAP), or a requisition can be submitted to Naval Supply Systems Command via the Naval
Logistics Library (NLL) website, https://nll.ahf.nmci.navy.mil. This publication is also available to
view and download from the Naval Air Technical Data and Engineering Service Command (NATEC)
website, https://www.natec.navy.mil or at the NAVAIR Airworthiness website https://
airworthiness.navair.navy.mil.
AUTOMATIC DISTRIBUTION (WITH UPDATES).
This publication and changes to it are automatically sent to activities that are established on the
Automatic Distribution Requirements List (ADRL) maintained by Naval Air Technical Data and
Engineering Service Command (NATEC), in San Diego, CA. If there is continuing need for this
publication, each activity’s Central Technical Publication Librarian must coordinate with the
NATOPS Model Manager of this publication to ensure appropriate numbers of this and associated
derivative manuals are included in the automatic mailing of updates to this publication.
NOTE
Activities not coordinating with the NATOPS Model Manager unit for
more than 12 months may be dropped from distribution.
51
ORIGINAL
A1-F18EA-NFM-000
UPDATING THE MANUAL
To ensure that the manual contains the latest procedures and information, NATOPS review
conferences are held in accordance with OPNAVINST 3710.7 series.
CHANGE RECOMMENDATIONS
Recommended changes to this manual or other NATOPS publications may be submitted by anyone
in accordance with OPNAVINST 3710.7 series. Change recommendations of any nature, (URGENT/
PRIORITY/ROUTINE) should be submitted directly to the Model Manager via the Airworthiness
website (https://airworthiness.navair.navy.mil) into the AIRS (Airworthiness Issue Resolution System) database. The AIRS is an application that allows the Model Manager and the NATOPS Office,
Naval Air Systems Command (NAVAIR) AIR−4.0P to track all change recommendations with regards
to NATOPS products. The Model Manager will be automatically notified via email when a new
recommendation is submitted. A classified side of the website is available to submit classified
information via the SIPRNET. Routine change recommendations can also be submitted directly to the
Model Manager (via your unit NATOPS Officer if applicable) on OPNAV Form 3710/6 shown on the
next page.
The address of the Model Manager of this aircraft/publication is:
Commanding Officer
VFA-122
U. S. Naval Air Station
Lemoore, CA 93245-5122
Attn: F/A-18E/F Model Manager
DSN: 949-2341
Commercial: (559) 998-2341
Email: [email protected]
52
ORIGINAL
A1-F18EA-NFM-000
NATOPS CHANGE RECOMMENDATION
OPNAV/FORM 3710/6(4-90) S/N 0107-LF-009-7900
DATE
TO BE FILLED IN BY ORIGINATOR AND FORWARDED TO MODEL MANAGER
FROM (originator)
Unit
TO (Model Manager)
Unit
Complete Name of Manual/Checklist
Revision Date
Change Date
Section/Chapter
Page
Paragraph
Recommendation (be specific)
r
CHECK IF CONTINUED ON BACK
Justification
Signature
Rank
Title
Address of Unit of Command
TO BE FILLED IN BY MODEL MANAGER (Return to Originator)
FROM
Date
TO
Reference
(a) Your change Recommendation Dated
r
Your change recommendation dated
is acknowledged. It will be held for action of the
review conference planned for
r
to be held at
Your change recommendation is reclassified URGENT and forwarded for approval to
by my DTG
/S/
AIRCRAFT
MODEL MANAGER
53
ORIGINAL
A1-F18EA-NFM-000
YOUR RESPONSIBILITY
NATOPS Flight Manuals are kept current through an active manual change program. Any
corrections, additions, or constructive suggestions for improvement of content of the manual should be
submitted by routine or urgent change recommendation, as appropriate, at once.
NATOPS FLIGHT MANUAL INTERIM CHANGES
Interim changes are changes or corrections to NATOPS manuals promulgated by CNO or
COMNAVAIRSYSCOM. Interim changes are issued either as printed pages, or as a Naval message.
The Interim Change Summary page is provided as a record of all interim changes. Upon receipt of a
change or revision, the custodian of the manual should check the updated Interim Change Summary
to ascertain that all outstanding interim changes have been either incorporated or canceled; those not
incorporated shall be recorded as outstanding in the section provided.
CHANGE SYMBOLS
Revised text is indicated by a black vertical line in either margin of the page, adjacent to the affected
text, like the one printed next to this paragraph. The change symbol identifies the addition of either
new information, a changed procedure, the correction of an error, or a rephrasing of the previous
material.
WARNING, CAUTIONS, AND NOTES
The following definitions apply to ‘‘WARNINGS’’, ‘‘CAUTIONS’’, and ‘‘NOTES’’ found throughout
the manual.
An operating procedure, practice, or condition, etc., which may result in
injury or death if not carefully observed or followed.
An operating procedure, practice, or condition, etc., which may result in
damage to equipment if not carefully observed or followed.
NOTE
An operating procedure, practice, or condition, etc., which is essential
to emphasize.
WORDING
The concept of word usage and intended meaning which has been adhered to in preparing this
54
ORIGINAL
A1-F18EA-NFM-000
manual is as follows:
″Land as soon as possible″ means to land at the first site which a safe landing can be made.
″Land as soon as practical″ means extended flight is not recommended. The landing site and
duration of flight is at the discretion of the pilot in command.
‘‘Shall’’ has been used only when application of a procedure is mandatory.
‘‘Should’’ has been used only when application of a procedure is recommended.
‘‘May’’ and ‘‘need not’’ have been used only when application of a procedure is optional.
‘‘Will’’ has been used only to indicate futurity, never to indicate any degree of requirement for
application of a procedure.
AIRSPEED
All airspeeds in this manual are in knots calibrated airspeed (KCAS) unless stated in other terms.
TERMINOLOGY
To standardize terminology throughout this publication, the following guidelines should be followed:
a. When specifying a switch, handle, or knob to be actuated in an emergency procedure, the name
of the switch, handle or knob should be written as it is labeled in the cockpit (i.e. LDG GEAR
handle).
b. When referencing a position of a switch, handle, or knob, the label as shown in the cockpit
should be used (i.e. ECS MODE switch − OFF/RAM).
c. LOT numbers should be used vice BUNO numbers when the entire LOT is affected. For
multiple LOTs use LOTs XX−XX or LOT XX and up as appropriate.
d. For MC OFP’s use terminology such as ″Prior to MC OFP 18E″ or ″MC OFP 18E and up″ to
avoid requiring a NATOPS change with each subsequent OFP release.
e. Procedures which are nested in other procedures such as the Emergency Oxygen Procedure
should contain only immediate action items.
f. When emergency procedures are referenced in the PCL, page numbers should be included to
facilitate quick location of the referenced procedure.
MANUAL DEVELOPMENT
This NATOPS Flight Manual was prepared using a concept that provides the aircrew with
information for operation of the aircraft, but detailed operation and interaction is not provided. This
concept was selected for a number of reasons: reader interest increases as the size of a technical
publication decreases, comprehension increases as the technical complexity decreases, and accidents
decrease as reader interest and comprehension increase. To implement this streamlined concept,
observance of the following rules was attempted:
a. Aircrew shall be considered to have above-average intelligence and normal (average) common
sense.
55
ORIGINAL
A1-F18EA-NFM-000
b.
No values (pressure, temperature, quantity, etc.) which cannot be read in the cockpit are
stated, except where such use provides the pilot with a value judgement. Only the information
required to fly the airplane is provided.
c.
Notes, Cautions, and Warnings are held to an absolute minimum, since almost everything in
the manual could be considered a subject for a Note, Caution, or Warning.
d.
No procedural data are contained in the Descriptive Section, and no abnormal procedures
(Hot Starts, etc.) are contained in the Normal Procedures Section.
e.
Notes, Cautions and Warnings are not used to emphasize new data.
f.
Multiple failures (emergencies) are not covered.
g.
Simple words in preference to more complex or quasi-technical words are used and
unnecessary and/or confusing word modifiers are avoided.
A careful study of the NATOPS Flight Manual will probably disclose a violation of each rule stated.
In some cases this is the result of a conscious decision to make an exception to the rule. In many cases,
it only demonstrates the constant attention and skill level that must be maintained to prevent slipping
back into the old way of doing things.
In other words, the ‘‘Streamlined’’ look is not an accident, it takes constant attention for the
NATOPS Flight Manual to keep the lean and simple concept and to provide the aircrew with the
information required.
56
ORIGINAL
A1-F18EA-NFM-000
58 (Obverse Blank)
ORIGINAL
A1-F18EA-NFM-000
PART I
THE AIRCRAFT
Chapter 1 - Aircraft
Chapter 2 - Systems
Chapter 3 - Servicing and Handling
Chapter 4 - Operating Limitations
59 (Reverse Blank)
ORIGINAL
A1-F18EA-NFM-000
CHAPTER 1
The Aircraft
1.1 AIRCRAFT DESCRIPTION
1.1.1 Meet The Super Hornet. The F/A-18E/F Super Hornet is a carrier based strike fighter aircraft
built by McDonnell Douglas Corporation. The general arrangement, approximate dimensions, and
cockpit layout are shown in Figure 1-1, Figure 1-2, and the Cockpit Foldout section, respectively. The
multi-mission aircraft has an internal 20 mm gun and can carry AIM-7, AIM-9, and AIM-120 air-to-air
missiles; and numerous air-to-ground weapons. The aircraft fuel load may be increased with the
addition of up to five external fuel tanks. The aircraft can be configured as an airborne tanker by
carrying a centerline mounted air refueling store.
Figure 1-1. General Arrangement
The aircraft is powered by two General Electric F414-GE-400 turbofan engines utilizing Full
Authority Digital Engine Control (FADEC). The aircraft features a variable camber mid-wing with
I-1-1
ORIGINAL
A1-F18EA-NFM-000
Figure 1-2. Approximate Dimensions
leading edge extensions (LEX) mounted on each side of the fuselage. Twin vertical tails are angled
outboard 20 degrees from the vertical.
The aircraft is designed with relaxed static stability to increase maneuverability and to reduce
approach and landing speed. The aircraft is controlled by a digital fly-by-wire Flight Control System
through hydraulically actuated flight control surfaces: ailerons, twin rudders, leading edge flaps,
trailing edge flaps, LEX spoilers, and differential stabilators. The leading edge of the wing incorporates
a ‘‘snag,’’ which increases outboard wing area and increases roll authority in the approach and landing
configuration. A speed brake function is provided by differential deflection of the primary flight
control surfaces.
The pressurized cockpit is enclosed by an electrically operated clamshell canopy. An aircraft
mounted auxiliary power unit (APU) provides self-contained start capability for the engines.
1.1.2 Aircraft Gross Weight. Basic weight is approximately 31,500 pounds for the F/A-18E and
32,000 pounds for the F/A-18F. Refer to applicable DD 365-3 for accurate aircraft weight.
1.1.3 F/A-18F. The F/A-18F is the two seat model of the Super Hornet and is configured with
tandem cockpits. The rear cockpit can be configured with a stick, throttles, and rudder pedals (trainer
configuration); or with two hand controllers, a UFCD adapter, and foot-operated communication
switches (missionized configuration). The rear cockpit controls and displays operate independently of
those in the front cockpit.
I-1-2
ORIGINAL
A1-F18EA-NFM-000
Figure 1-3. Radar Cross Section (RCS) Reduction
1.1.4 Radar Cross Section (RCS) Reduction. RCS reduction is a significant feature of the F/A18E/F. While the maintenance community is tasked with maintaining the RCS features of the aircraft,
it is in the best interests of the aircrew community to take an active role to ensure the survivability
characteristics of the aircraft are retained.
RCS reduction is accomplished through numerous airframe design features. See figure 1-3. The
baseline feature is planform alignment of as many surface edges as feasible. The outer moldline of the
aircraft is treated to make it a smooth, conductive surface in order to reduce radar scattering.
Treatment entails metalizing the navigation lights, canopy, and windshield. Permanent joints and
gaps around infrequently opened panels are filled with a form-in-place (FIP) sealant, which is blended
flush and conductively painted. Gaps around frequently opened panels are filled with a conductive FIP
(CFIP) sealant, which allows for easier repair. Conductive tape is applied to a few gaps where there is
no substructure to support FIP material, such as along LEX edges. Conductive tape can also be used
to quickly repair damaged FIP joints.
Since CFIP in the gaps around frequently opened panels will experience the most wear and tear, a
corrosion-proof radar absorbing material (RAM) is applied in front of many of these gaps. RAM is also
applied (1) on the inlet lip and duct, (2) as diamond-shaped patches around drain holes, and (3) in
various locations that tend to highly scatter radar energy such as around pitot tubes, vertical tail
openings, vents and screens, flap hinges and fairings, and portions of the pylons and external tanks. A
multi-layer RAM is used in a few locations, such as around AOA probes and on the top, front surface
of the pylons.
I-1-3
ORIGINAL
A1-F18EA-NFM-000
Gaps around landing gear doors are treated in two ways. Nose landing gear doors use flexible
conductive blade seals on leading and trailing edges; main landing gear door edges are wrapped with
RAM. Scattering from trailing edges (i.e., trailing edge flaps and rudders) is controlled by a radar
absorbing boot which is bonded to the surface. Scattering from the back edge of the windshield is
controlled by a gray, laminated material called the aft arch termination strip.
The engine inlet ducts incorporate a device to minimize engine front face scattering. The edge of the
canopy incorporates a conductive bulb seal to block radar reflections from that joint. Conductive bulb
seals are also used where there is significant structural flexure, such as at the wing-to-LEX interface.
Eleven electro magnetic interference shields (EMIS) III radar shields are permanently installed on
the radar antenna hardware. To allow the aircraft to achieve its full RCS reduction potential, a
missionized kit consisting of twelve more EMIS III radar bulkhead shields, are installed for combat
missions only. Additionally, SUU-79 pylons can be fitted with a set of low observable (LO) hardware.
1.2 BLOCK NUMBERS
See figure 1-4 for block numbers and bureau numbers for each lot of aircraft. The baseline for this
NATOPS manual is aircraft 165779 (Lot 23) and up. Differences from the baseline aircraft for aircraft
165533 thru 165679 (Lot 21 and Lot 22) are presented in Appendix B of this manual.
I-1-4
ORIGINAL
A1-F18EA-NFM-000
LOT
F/A-18E
F/A-18F
LOT 21
165533 - 165540
165541 - 165544
LOT 22
165660 - 165667
165668 - 165679
LOT 23
165779 - 165792
165793 - 165808
LOT 24
165860 - 165874
165875 - 165895
LOT 25
165896 - 165909
165910 - 165934
LOT 26
166420 - 166448
166449 - 166467
LOT 27
166598 - 166609
166610 - 166642
LOT 28
166643 - 166657
166658 - 166684
LOT 29
166775 - 166789
166790 - 166816
LOT 30
166817 - 166841
166842 - 166854
Figure 1-4. LOT NUMBER/BUNO
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A1-F18EA-NFM-000
CHAPTER 2
Systems
2.1 POWER PLANT SYSTEMS
2.1.1 Engines. The aircraft is powered by two General Electric F414-GE-400 engines. The engines
are low bypass, axial-flow, twin-spool turbofans with afterburner. The three stage fan (low pressure
compressor) and the seven stage high pressure compressor are each driven by a single stage turbine.
The basic functions are supported by the engine driven accessory gearbox which drives the engine fuel
pump, variable exhaust nozzle (VEN/start) fuel pump, lubrication and oil scavenge pump, engine fuel
control, and alternator. Fuel flow from the VEN/start pump is used to drive the VEN actuator and to
provide initial fuel pressure for main engine start.
The uninstalled military thrust (MIL) of each F414-GE-400 engine is approximately 13,900 pounds
with maximum afterburner thrust (MAX) in the 20,700 pound class.
An inlet device is installed in each engine intake to reduce the aircraft radar signature and to
improve survivability.
2.1.1.1 FADEC - Full Authority Digital Engine Control. Engine operation is controlled by a full
authority digital engine control (FADEC), mounted on the engine casing. Each FADEC computer has
two central processor units, channel A (CH A) and channel B (CH B), and is integrated with the
Mission Computers (MCs), flight control computers (FCCs), and throttles. Normally, both FADEC
channels monitor engine and control system operation with one channel in control and the other in
standby. The channel currently in control is boxed on the SUPT MENU/ENG display.
In the event of a control system failure, the FADEC automatically selects the channel with better
capability. Manual FADEC channel transfer can be commanded by selecting the CH A or CH B
pushbutton on the ENG display. When the throttle is at or above IDLE, the FADEC transfers control
to the other channel only if the requested channel’s health is no worse than the channel in control.
FADEC software implements the engine control schedules by modulating fuel flow and engine
geometry for the current flight conditions and the ″requested″ throttle setting. With the throttles
matched, engine parameters may vary significantly between the engines as control schedules are
adjusted for optimum performance. Therefore, a mismatch between engine parameters is not a sign of
degraded performance as long as ENG STATUS is NORM. FADEC cooling is provided by the fuel
system.
2.1.1.1.1 FADEC Power Sources. Prior to first engine start, the battery is used to power CH A of
both FADECs. When N2 reaches 10% rpm during start, the engine driven alternator comes online and
powers both channels of its corresponding FADEC. When an aircraft generator comes online (N2
greater than 60%), the aircraft’s electrical system provides power to both channels of the other FADEC
as well. With both engines operating, each engine driven alternator is the primary source of FADEC
power with the aircraft’s 28 vdc essential bus as backup. When both channels of a FADEC are powered
after initial start, the FADEC automatically switches operation to the channel which was not in control
during the last flight/engine run.
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Ten seconds after reaching idle power, the FADEC attempts to switch to the 28 vdc essential bus to
verify that backup power is available. If backup power is inadequate/inoperative, the FADEC declares
a channel degrade (dual channel lineout) and sets a 6A8 or 6C8 MSP code (L or R FADEC/aircraft 28V
fail). These degrade indications appear if the first engine is started with the corresponding GEN switch
OFF, and requires a FADEC reset to restore normal engine indications.
2.1.1.1.2 Engine Status. Engine status is reported by the FADEC and appears on the ENG
STATUS line of the ENG display. The levels of engine performance capability, listed in descending
order, are:
NORM
Engine performance is normal
PERF90
10% or less thrust loss and/or slower engine transients. Afterburner is not
inhibited
AB FAIL
No afterburner capability
THRUST
Engine thrust is limited to between 40% and 90% and significantly slower
transients
IDLE
Engine is limited to idle power only
SHUTDOWN
Engine automatically shut down
2.1.1.1.3 FADEC/Engine Degrades. FADEC/engine degrades fall into two categories: minor failures
which do not affect engine operability and significant failures that do affect engine operability.
Due to a high level of FADEC redundancy, most minor control system failures do not cause any
degradation in engine performance (ENG STATUS remains NORM). Cockpit indications for these
types of failures are slightly different depending on whether the aircraft is inflight or on the ground.
Inflight a. FADEC and BIT advisories.
b. ENG STATUS - NORM.
c. OP GO indication for the affected engine channel on the BIT/HYDRO MECH display.
On the ground or below 80 KCAS after landing a. FADEC and BIT advisories.
b. ENG STATUS - NORM.
c. DEGD replaces OP GO for the affected engine channel on the BIT/HYDRO MECH display.
d. Both CH A and CH B lined out on the ENG display.
A FADEC OP GO or DEGD with an ENG STATUS of NORM is an indication of a loss of control
system redundancy and not of a loss of engine operability. Therefore, a FADEC OP GO inflight should
be considered informative and should not mandate a mission abort. However, on the ground, the
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FADEC DEGD and dual channel lineout indications are intended to prevent takeoff with a known loss
of redundancy and maintenance action is required. Therefore, takeoff with a FADEC DEGD indication
(dual channel line out) is prohibited.
Significant failures which do cause degradation in engine performance have the following indications
both inflight and on the ground:
a. L or R ENG caution and voice alert.
b. FADEC and BIT advisories.
c. ENG STATUS change on the ENG display.
d. DEGD indication for the affected engine channel on the BIT/HYDRO MECH display.
e. Both CH A and CH B lined out on the ENG display.
2.1.1.1.4 FADEC Reset. When a throttle is OFF, the corresponding FADEC uses a channel change
request as a FADEC software reset. FADEC reset capability is provided to clear a DEGD indication
which was caused by a momentary FADEC fault (e.g., startup power transient). A FADEC reset should
only be performed for DEGD indications which occur on the ground and which do not result in a
change of ENG STATUS. For a FADEC OP GO inflight, engine shutdown and FADEC reset is not
recommended as the engine is functioning normally. In all other circumstances, a FADEC reset should
not be attempted, particularly airborne, as any degrade in ENG STATUS is most likely indicative of
the failure of a mechanical component. Under these conditions, the engine may fail to restart following
a shutdown and FADEC reset attempt.
2.1.1.1.5 Ignition System. The FADEC provides simultaneous control of the main and afterburner
igniters via the engine driven alternator and ignition exciter. Ignition (both main engine and
afterburner) is commanded whenever:
a. N2 rpm is between 10% and 45% during engine start.
b. Flameout occurs.
c. Throttle is advanced into afterburner, remaining on until afterburner light off is sensed.
d. The gun is fired, or any wing pylon mounted A/A or forward firing A/G weapon is launched.
Ignition remains on for 5 seconds.
2.1.1.1.6 Oil Pressure Sensing System. The oil pressure sensing system utilizes an oil pressure
transducer and a separate oil pressure switch. The transducer provides an oil pressure value for display
in the cockpit. The oil pressure switch provides an additional source to confirm the presence of oil
pressure if the oil pressure transducer fails.
If a L or R OIL PR caution is set with a valid cockpit readout, actual engine oil pressure is below
scheduled limits. If the cockpit readout is zero with no caution set, the oil pressure transducer has
failed and the pressure switch is inhibiting the caution.
2.1.1.1.7 Engine Idle Schedules. Each FADEC modifies engine idle schedules based on weight on
wheels (WonW) status, aircraft flight condition, engine bleed demand, and environmental control
system (ECS) mode of operation. The purpose of idle scheduling is to ensure that engine bleed output
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A1-F18EA-NFM-000
is always sufficient to run the ECS, particularly the aircrew onboard oxygen generating system
(OBOGS).
Baseline idle schedules are used during normal engine operation and moderate environmental
conditions, and are set as a function of pressure altitude and WonW status. With WonW and airspeed
below 80 knots, the FADEC commands ground idle by reducing the engine compressor discharge
pressure (CDP) (typically a slight decrease in N2 rpm) from the inflight idle setting. The throttle
handle angle (THA) for ground and inflight idle are identical.
In LOT 26 and up, with WoffW, spin recovery mode not engaged, and inflight refueling probe and
landing gear (and hook in LOT 24 and below) retracted, baseline flight idle thrust LOT 26 and up is
higher than in LOT 25 and below. Ground idle thrust is not affected.
Alternate idle schedules are used when engine bleed demands are high (e.g., hot ambient
temperatures, high ECS output, RECCE configuration, engine or windshield anti-ice operation, or
ECS OFF/RAM mode). When an alternate idle schedule is requested, the FADEC increases the
minimum CDP that is commanded at idle power (typically a slight increase in N2 rpm), which may also
result in a noticeable decrease in throttle response at the lower end of the throttle range. With the
throttles near flight idle, a small engine transient may be noticed when an alternate idle schedule is
activated or when a transition between alternate idle schedules occurs.
In LOT 26 and up, when an alternate idle schedule is activated flight idle thrust can be double that
for LOT 25 and below aircraft.
Alternate idle schedules are deactivated with WonW, when spin recovery mode is engaged, or when
the inflight refueling probe or landing gear (or hook in LOT 24 and below) are extended. When
alternate idle schedules are deactivated in LOT 26 and up, a small noticeable engine transient may
occur with the throttle near flight idle, and the resulting flight idle thrust is identical to LOT 25 and
below baseline flight idle thrust.
2.1.1.1.8 Fan Speed Lockup. The fan speed lockup system prevents inlet instability (buzz) at high
Mach by holding engine speed and airflow at military power levels when the throttle(s) are retarded
below MIL. Speed lockup is activated when the aircraft accelerates above Mach 1.23 and deactivated
when the aircraft decelerates below Mach 1.18.
2.1.1.1.9 SETLIM - Supersonic Engine Thrust Limiting. SETLIM minimizes the potential for an
aircraft departure due to asymmetric thrust following an engine stall or flameout at certain supersonic
(Mach greater than 1.8) or high-q conditions equivalent to approximately 700 KCAS at sea level or 750
KCAS at 25,000 feet. If the FADEC detects a stall or flameout condition, this function terminates
afterburner operation in both engines. Normal afterburner operation is restored 12 seconds after
engine recovery or immediately when airspeed drops below Mach 1.7 and q-conditions are equivalent
to approximately 650 KCAS at sea level or 710 KCAS at 25,000 feet.
2.1.1.1.10 RATS - Reduced Authority Thrust System. The reduced authority thrust system (RATS)
reduces the wind-over-deck required for carrier landings by rapidly reducing thrust at the beginning
of a successful arrestment, reducing the energy absorbed by the arresting gear. RATS logic, only
resident in MC1, declares a successful arrestment if the landing gear and arresting hook are down and
longitudinal deceleration is more than 1.0g (a typical arrestment is approximately 3g). MC1 sends a
″set RATS on″ signal to the FADECs, which reduce thrust to approximately 70% of MIL power. RATS
logic also senses WonW, wheel speed (less than 20 knots), and THA to prevent the engines from
spooling back to MIL power at the end of cable pullout. RATS operation is canceled when the throttles
are reduced to IDLE (THA less than 10°). RATS operation is inhibited during single engine operation.
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A1-F18EA-NFM-000
RATS operation can be overridden by advancing the throttles to full afterburner (THA within 2° of
the MAX stop). With RATS enabled, afterburner operation is inhibited if the throttles are
subsequently advanced to afterburner (below the MAX stop). If the throttles are in afterburner (below
the MAX stop) during an arrested landing, RATS functions normally, rolling back both the main
engine and the afterburner.
2.1.1.1.11 AGI (Armament Gas Ingestion) Protection. AGI protection provides preemptive engine
ignition in case armament gas ingestion causes an engine flameout. As mechanized, AGI protection is
a backup to the FADEC’s inherent flameout detection and relight logic. While flight test data indicates
that the system may not be needed, AGI protection remains functional.
The AGI signal is sent by the Stores Management Set (SMS) to the FADECs and is used to initiate
engine ignition (both engines) for 5 seconds. The signal is sent when the gun is fired, or or any wing
pylon mounted A/A or forward firing A/G weapon is launched. AGI is functional in the SIM mode as
well as the tactical mode.
2.1.1.1.12 IBU (Increased Bleed Usage) Signal In LOT 26 and up, during high bleed flow rates, the
mission computers may send the FADECs an IBU signal to prevent engine turbine overheating. When
IBU scheduling is active, hot day MIL and MAX thrust may be reduced by up to 1.5% compared to
LOT 25 and below aircraft.
2.1.1.1.13 ABLIM - Afterburner Limiting Function. The ABLIM function limits engine power to
half afterburner with the throttles at MAX to prevent engine stalls due to exhaust gas ingestion. The
system is only to be used during carrier-based operations. The function is pilot selectable with WonW.
The system defaults to disabled (unboxed) after engine start. The ABLIM function is activated by
selecting (boxing) the ABLIM option on the CHKLIST format with the FLAP switch in HALF or
FULL. The ABLIM advisory is set to confirm that the function has been activated on both engines.
With the function activated, only half afterburner power is available with the throttles at MAX.
Indicated fuel flows are reduced from 35,000 to 45,000 pph to about 25,000 pph. The function is
automatically deactivated with acceleration due to a catapult launch, at 80 KCAS, or with WoffW. The
ABLIM function is disabled with a FCC CH 1, FCC CH 2, FCC CH 4, MC1, or FADEC failure.
2.1.1.2 Engine Related Cautions and Advisories. The following engine related cautions and advisories are described in the Warning/Caution/Advisory Displays in Part V:
D
D
D
D
D
D
L or R EGT HIGH
L or R ENG
L or R ENG VIB
FADEC HOT (ground only)
L or R FLAMEOUT
NO RATS
D
D
D
D
D
D
L or R OIL HOT (engine or AMAD)
L or R OIL PR
L or R OVRSPD
L or R STALL
ABLIM advisory
FADEC advisory
2.1.1.3 Engine Anti Ice. Each engine supplies its own bleed air for engine and inlet device anti-ice.
The engine anti-ice system is normally controlled by the ENG ANTI ICE switch. However, after engine
start (initial FADEC power-up or following a FADEC reset), the engine anti-ice system automatically
turns on 45 seconds after the engine reaches idle power and remains on for 30 seconds, provided the
throttle remains at IDLE. The appropriate LHEAT or RHEAT advisory is displayed during this
anti-ice functional test.
With the ENG ANTI ICE switch ON, anti-ice air flows as long as INLET TEMP is between -40
and +15°C. Outside of these limits, anti-ice airflow is terminated immediately if airborne, or after 60
seconds with WonW. When anti-ice air is flowing, N2 rpm increases approximately 2%, and EGT
increases approximately 5°C.
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A1-F18EA-NFM-000
The inlet device has an anti-ice leak detection system. The system detects hot air leaks from the
device air manifold or in the cavity between the device and the airframe and sets the L or R DEVC
BLD caution. A hot air leak into the cavity reduces device anti-ice capability and may structurally
damage the device and surrounding structure.
2.1.1.3.1 ENG ANTI ICE Switch. The ENG ANTI ICE switch is located on the ECS panel on the right
console.
ON
Activates the engine anti-ice system
OFF
Deactivates the engine anti-ice system
TEST
Checks ice detector operation and displays the INLET ICE caution (indicating proper
operation).
2.1.1.3.2 Engine Anti-Ice Advisories. The L HEAT and R HEAT advisories are displayed when the
engine anti-ice system is activated (ENG ANTI ICE switch ON or anti-ice functional test) and no
system failures are detected. If an engine anti-ice failure occurs with the system turned on, the HEAT
FAIL caution is displayed and the corresponding L HEAT or R HEAT advisory is removed.
If an engine anti-ice failure occurs with the ENG ANTI ICE switch OFF, the HEAT advisory is
displayed. This advisory indicates that anti-ice operation is degraded or not available if selected.
2.1.1.3.3 Engine Anti-Ice Related Cautions and Advisories. The following engine anti-ice related
cautions and advisories are described in the Warning/Caution/Advisory Displays in Part V:
D L or R ANTI ICE caution
D L or R DEVC BLD caution
D HEAT FAIL caution
D INLET ICE caution
D L HEAT or R HEAT advisory
D HEAT advisory
2.1.1.4 Engine Controls and Displays.
2.1.1.4.1 ENG CRANK Switch. The ENG CRANK switch is described in the Secondary Power
System section.
2.1.1.4.2 Throttles. Two throttles, one for each engine, are located on the left console. Throttle
movement is transmitted electrically to the corresponding FADEC for thrust modulation and to the
FCCs for autothrottle operation. There is no mechanical linkage between the throttles and the engines.
During engine start, advancing the throttles from OFF to IDLE opens the engine fuel control shutoff
valves and, when commanded by the FADEC, provides fuel flow to the engines.
Afterburner operation is initiated by advancing the throttles through the MIL detent into the
afterburner range. During catapult launch or carrier touchdown (WonW and launch bar or arresting
hook extended), an afterburner lockout mechanism extends to preclude inadvertent afterburner
selection. In such cases, the throttles can be moved to the afterburner range by raising the finger lifts
on the front of each throttle or by applying a force of approximately 30 pounds.
During engine shutdown, the finger lifts must be raised to move the throttles to OFF, closing the
engine fuel control shutoff valves. The throttle grips (figure 2-1) contain switches that allow control of
various systems without moving the hand from the throttles.
2.1.1.4.3 Throttles (Trainer Configured F/A-18F). The rear cockpit of the trainer configured
F/A-18F contains an additional set of flight controls: control stick, throttles, and rudder pedals. The
rear cockpit throttles, located on the left console, are mechanically connected to those in the front
cockpit and provide thrust modulation from IDLE to MAX. The rear throttles do not contain finger
lifts, so the engines cannot be secured from the rear cockpit. The rear throttle grips are slightly
different than those in the front cockpit. The ATC engage/disengage switch is not functional; the
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-1. Throttle Grips (Front Cockpit)
chaff/flare/ALE-50 switch is not installed; and the speed brake switch is momentary action only. In
general, systems controlled by throttle switches respond to the last crewmember action taken from
either cockpit.
2.1.1.4.4 EFD - Engine Fuel Display, Engine Parameters. The EFD, located on the main instrument
panel below the left digital display indicator (LDDI), is a night vision imaging system (NVIS)
compatible, monochromatic, liquid-crystal, grey/black display powered by the Signal Data Computer
(SDC).
The EFD normally displays critical engine parameters in the bottom half of the display and fuel
quantities in the top half. The fuel portion of the EFD is described in the Fuel System section. Invalid
parameters are usually displayed as 999 or 9999 in inverse video; out of limit parameters are always
highlighted by inverse video. Nozzle position is displayed both graphically and digitally in percent
open.
On battery power prior to APU start, the EFD either displays only RPM and TEMP or the entire
top level format (figure 2-2). If the entire top level format is not displayed on battery power, it will be
displayed when the APU switch is selected ON. When the APU switch is selected ON, only RPM,
TEMP, and OIL pressure are valid. When the first engine alternator comes online at 10% N2 rpm, the
FF parameter also becomes valid. When the first generator comes online at 60% N2 rpm, all
parameters for both engines become valid. If the EFD locks up or blanks completely during engine
start power transients, the display can be reset by selecting the SDC RESET option from the SUPT
MENU/FUEL display.
The EFD displays the following engine parameters within the listed display tolerances:
RPM
Compressor rpm (N2 ) (0 to 127%) - Displays RPM in inverse video format above
102%
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ORIGINAL
A1-F18EA-NFM-000
TEMP
Compensated turbine exhaust gas temperature (EGT) (186 to 1,088°C) - Displays 9999
in inverse video above 1,100°C
FF
Total commanded fuel flow including afterburner (0 or 400 to 65,000 pph in 100 pph
increments).
NOTE
Engine fuel flow is calculated from commanded engine fuel metering
valve position. In failure modes, fuel flow can be indicated on the EFD
even though no actual fuel is flowing.
OIL
Oil pressure (0 to 200 psi)
NOZ
VEN position (0 to 101% open)
Figure 2-2. Engine Fuel Display (EFD) - Engine Parameters
During first engine battery start, the EFD RPM indication typically jumps from 0 to either 5% or
10%, and lightoff is indicated by TEMP rising from a minimum reported value of approximately
190°C. Each engine has three sources of N2 rpm: two engine alternator sensors and one accessory
gearbox sensor. The accessory gearbox sensor can provide rpm readings down to only 5% and is the
initial source of engine RPM. Readings from the alternator sensors are not available until the
alternator comes online above 10% N2 rpm. Input for the TEMP parameter is provided by a
compensated EGT algorithm in the FADEC. When actual EGT is below accuracy tolerances (e.g.,
engine shutdown), the FADEC limits the minimum reported TEMP (approximately 190°C). The FF
parameter is a calculated number based on metering valve position, not an actual measurement of fuel
flow. Consequently, an indication of fuel flow may be present when there is no flow, such as when the
throttle is above IDLE with the engine off.
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ORIGINAL
A1-F18EA-NFM-000
2.1.1.4.5 ENG Display. The ENG display (figure 2-3) is selected by pushing the ENG option from the
SUPT MENU. The ENG display shows the following engine and thermal management system
parameters:
ENG STATUS
The current level of engine performance provided by the control system
INLET TEMP
Engine inlet temperature (°C)
N1 RPM
Fan speed (% rpm)
N2 RPM
Compressor speed (% rpm)
EGT
Exhaust gas temperature (°C)
FF
Total commanded fuel flow (pph)
NOZ POS
Nozzle position (% open)
OIL PRESS
Engine oil pressure (psi)
THRUST
Takeoff thrust (%), referenced to hot day MIL power (blanked inflight)
FAN VIB
Fan vibration (inches/second)
CORE VIB
Core vibration (inches/second)
EPR
Engine pressure ratio (exhaust pressure to engine inlet pressure).
CDP
Compressor discharge pressure (psia)
CPR
Compressor pressure ratio
THA
Throttle handle angle (deg)
AMAD OIL TEMP
AMAD oil temperature (°C)
ENG OIL TEMP
Engine oil temperature (°C)
FUEL INLET
TEMP
Engine inlet fuel temperature (°C)
FUEL NOZ TEMP
Engine nozzle fuel temperature (°C)
FEED TANK
TEMP
Feed tank fuel temperature (°C)
When an engine or thermal management system related caution appears, the MENU option at the
bottom of each DDI is replaced with the ENG option, providing one pushbutton access to the ENG
display. The value of the out of limit parameter which triggered the caution is displayed in red and
highlighted by carets on either side. The CH A and CH B options at the top of the display are used to
command a manual FADEC channel transfer, and the active FADEC channel for each engine is boxed.
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-3. Engine Display
The selected fuel grade is displayed top center and also on the takeoff side of the CHKLIST display.
The FUEL option enables the JP-5, 8, and JP-4 options, each of which must be selected twice to change
the fuel grade. The fuel grade selected should reflect the majority of fuel in the aircraft. JP-5, -8 should
be selected when using JET-A, A+, or A1. The engines can be started and operated at ground idle with
either fuel grade selected; however, for operations above IDLE, the correct fuel grade must be selected
to ensure proper engine operation and to avoid engine combustor rumble. The fuel grade selected is
also used by the FADEC to control the thermal control valve (TCV) setting and determines the
maximum fuel temperature to be sent to the engines. Incorrect fuel grade selection can adversely
impact the fuel thermal management system and result in a premature FUEL HOT caution. The
RECORD option, boxed when selected, saves a 30 second record of display and engine data (15 seconds
pre- and 15 seconds post-event) to the memory unit (MU). The DFIRS DWNLD option downloads
DFIRS data to the MU.
2.1.2 ATC - Automatic Throttle Control. The ATC system has two operating modes: approach and
cruise. The system automatically modulates engine thrust between flight IDLE and MIL power in
order to maintain on-speed angle of attack (AOA) in the approach mode or calibrated airspeed
(existing at the time of engagement) in the cruise mode.
During ATC operation, engine commands are sent to the FADEC directly from the FCCs instead of
the throttles. FCC generated engine commands are limited to a range slightly above idle to slightly
below MIL. The throttles are continuously positioned by an FCC commanded backdrive unit to match
the throttles with the current engine command and to provide feedback to the pilot.
2.1.2.1 ATC Engagement. Pressing and releasing the ATC button on the left throttle engages the
approach mode with the FLAP switch in HALF or FULL and the cruise mode with the FLAP switch
in AUTO. When either mode is engaged, an ATC advisory is displayed on the HUD. Because ATC
mode engagement and ATC HUD advisories are not commanded until release of the ATC button, the
pilot may need to deliberately pause after press and release to avoid inadvertant ATC disengagement/
re-engagement. Automatic transition between the two modes or engagement during single engine
operation is not possible. Engaging ATC with the friction lever in the full aft position and with the
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throttles at mid-range power provides optimum pilot feedback with the smallest engagement power
transients.
2.1.2.2 ATC Disengagement. If either mode does not engage when selected, or automatically
disengages after engagement, the ATC advisory flashes for 10 seconds and is removed from the HUD.
Disengagement for any reason requires reengagement to restore ATC operation. Normal disengagement is accomplished by re-actuation of the ATC button or by applying a force of approximately 12
pounds (friction off) to either throttle for greater than 0.20 seconds. This force is sufficient to permit
the pilot’s hand to follow throttle movement without causing disengagement. Holding the throttles
against the MIL or IDLE stop during ATC disengagement commands a rapid acceleration or
deceleration to the commanded power setting instead of a smooth transition.
2.1.2.3 ATC Automatic Disengagement. The ATC system automatically disengages for the following
reasons:
Either mode •
•
•
•
•
•
•
Any ATC system internal failure
ATC button failure
FADEC failure
FCC CH 2 or CH 4 failure
Backdrive failure
THA split greater than 3° for more than 1 second
FLAP switch position change between AUTO and HALF or FULL
Approach mode only •
•
•
•
•
AOA, pitch rate, or Nz sensor failure
Bank angle in excess of 70°
Flap blowup at 250 KCAS
Gain ORIDE selection
Weight on wheels
Cruise mode only • FCC calibrated airspeed failure
2.1.2.4 ATC Related Cautions. The ATC FAIL caution is described in the Warning/Caution/
Advisory Displays in Part V.
2.2 FUEL SYSTEM
The aircraft is fitted with four internal fuselage tanks (Tanks 1 through 4), two internal wing tanks
(left and right), two fuselage vent tanks, and two vertical vent tanks. Tanks 2 and 3 are engine feed
tanks while Tanks 1, 4, and the wing tanks are transfer tanks. Total fuel can be increased by the
carriage of up to four 480 gallon external fuel tanks on the centerline, inboard, and midboard pylons.
The aircraft can also be configured as an airborne tanker with the carriage of a centerline mounted air
refueling store (ARS). All tanks, internal and external, may be refueled on the ground through a
single-point refueling receptacle or inflight through the inflight refueling probe.
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The aircraft’s fuel system is composed of the following subsystems: engine feed, motive flow, fuel
transfer, tank pressurization and vent, thermal management, refueling, fuel dump, fuel quantity
indicating, and fuel low level indicating. Refer to Fuel System, Foldout Section, for simplified
schematics.
2.2.1 Engine Feed System. Each engine feed system contains an airframe mounted accessory drive
(AMAD) driven motive flow/boost pump, a feed tank with an internal motive flow powered turbo
pump, and an engine feed shutoff valve. For survivability, the left and right feed systems are normally
separated but can be interconnected by a normally closed crossfeed valve and a normally closed feed
tank interconnect valve.
2.2.1.1 Motive Flow/Boost Pumps. Each AMAD drives a two-stage motive flow/boost pump. The
first stage supplies low pressure fuel to its respective engine mounted fuel pump, while the second stage
supplies high pressure fuel to the motive flow system. Fuel from the motive flow system is used to cool
accessories, power the feed tank turbo pumps and certain transfer/scavenge pumps, and control certain
transfer valves.
2.2.1.2 Feed Tanks. During normal operation, each engine receives fuel from separate fuel feed lines.
Tank 2 supplies fuel to the left engine; Tank 3 to the right. A motive flow powered turbo pump in each
feed tank supplies fuel to its respective motive flow/boost pump.
Each feed tank has a horizontal baffle which traps fuel, providing a minimum of 10 seconds of
negative g flight at MAX power. No sustained zero g capability is provided, and prolonged transitions
through zero g (greater than 2 seconds) may produce a L and/or R BOOST LO caution.
If a feed tank turbo pump fails, fuel is suction fed to the motive flow/boost pump. In this case, flight
at high altitude with high feed tank fuel temperatures may not supply enough fuel for high power
settings.
2.2.1.3 Feed Shutoff Valves. In the event of a fire or fuselage fuel leak, engine feed shutoff valves
provide the capability to isolate a fuel feed system immediately downstream of the feed tank. Pressing
the L or R FIRE warning light electrically closes the corresponding engine feed shutoff valve, isolating
that fuel feed system.
2.2.1.4 Crossfeed Valve. The crossfeed valve, normally closed, allows a single motive flow/boost
pump to feed both engines when boost pressure is lost on one side (e.g., single engine shutdown, a leak,
motive flow/boost pump failure, or feed tank depletion). A loss of boost pressure downstream of the
motive flow/boost pump sets the L or R BOOST LO caution and opens the crossfeed valve. An open
crossfeed valve allows the output from the good motive flow/boost pump to supply fuel to the opposite
engine at rates sufficient for at least MIL power.
Pressing the L or R FIRE warning light electrically closes (inhibits opening) the crossfeed valve,
isolating the two fuel feed systems.
2.2.1.5 Interconnect Valve. A feed tank interconnect valve, installed between Tanks 2 and 3, is used
to control gravity transfer/balancing between the two feed tanks. During normal operation, the dual
flapper-type valve is held closed by motive flow pressure on either side (left motive flow on the Tank
2 side and right motive flow on the Tank 3 side), and no fuel gravity transfers.
If motive flow is lost on one side (e.g., single engine shutdown), the valve opens to make sure that
feed tank fuel is available to the opposite engine. For instance, if motive flow is lost on the right side,
the Tank 3 side of the valve opens, allowing fuel to gravity transfer to Tank 2 anytime the Tank 3 fuel
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A1-F18EA-NFM-000
level is higher. If Tank 3 has a fuel leak (e.g., battle damage), motive flow pressure on the Tank 2 side
of the valve prevents Tank 2 fuel from gravity transferring into the leak.
2.2.1.6 Feed Tank Balancing. The SDC incorporates feed tank balancing logic, designed to keep
Tanks 2 and 3 within 100 lb of each other. With a normally operating fuel system, balancing begins
after Tank 4 is effectively empty (less than about 300 lb) and the feed tanks begin to deplete below full.
If a feed tank imbalance reaches 100 lb, the SDC shuts off the corresponding Tank 4 scavenge pump
until the imbalance is 50 lb in the opposite direction. With WoffW, feed tank balancing continues until
either feed tank reaches FUEL LO level (approximately 1,125 lb). Feed tank balancing stops at FUEL
LO to make sure tank 4 fuel is transferred to both feed tanks in case one feed tank is damaged and is
leaking. After transitioning to WonW, balancing is reinitiated and continues until either feed tank is
below 300 lb.
In the event of a fuel transfer failure (e.g., a feed tank begins to deplete with fuel in Tank 4), feed
tank balancing begins when either feed tank drops below approximately 2,100 lb for 1 minute. This
mechanization attempts to minimize the effect of the fuel transfer failure by reducing the resulting
feed tank split.
2.2.1.7 Feed Tank Imbalance with One Engine at Idle. If one engine is intentionally reduced to
idle/low power or is commanded to IDLE by the FADEC, a higher rate of fuel depletion can be
expected from the ″good″ engine’s feed tank. At internal fuel weights below approximately 4,900 lb
(transfer fuel depleted), a fuel split can be expected to develop between the feed tanks (interconnect
valve is closed). If fuel burn continues to approximately 2,450 lb, the good engine feed tank depletes
and runs dry. The motive flow/boost pump output pressure on the good side drops, sets the L or R
BOOST LO caution, and opens the crossfeed valve. The good engine feeds from the opposite feed tank
through the crossfeed valve.
When driven by an idling engine, a motive flow/boost pump can support fuel flow up to 28,000 pph
through the crossfeed valve (MIL power fuel flow is approximately 12,000 pph at sea level, standard
day). If the fuel flow demand on the usable engine exceeds 28,000 pph (midrange afterburner), motive
flow/boost pump output pressure drops, setting the other BOOST LO caution, closing the crossfeed
valve, and starving the good engine. MAX power, single engine fuel flow is approximately 38,500 pph
at sea level, 0.2M, standard day (approach conditions).
Selecting afterburner on the good engine with its feed tank reading empty
results in engine flameout if fuel flow exceeds 28,000 pph.
The only way to balance a growing feed tank split is to shutdown the idling engine. This opens both
the interconnect and crossfeed valves. The risk of balancing is a loss of hydraulic and electrical
redundancy provided from the engine if left at idle.
2.2.2 Fuel Transfer System. The fuel transfer system, controlled by the SDC, is designed to keep the
feed tanks full or near full during normal engine operation. Fuel is routed from Tanks 1 and 4, the
internal wing tanks, and external fuel tanks, if installed, through three independent sets of transfer
lines. Additionally, the SDC schedules Tank 1 and 4 transfer to control fuel center of gravity (CG).
2.2.2.1 Fuel Transfer - Tanks 1 and 4. Fuel is transferred from Tanks 1 and 4 to the feed tanks by
two dual-speed electric transfer pumps, one in each tank. The low speed setting is used for normal
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transfer. The high speed setting is used during high fuel flow conditions such as afterburner operation,
ARS replenishment, or fuel dump. The one exception to this is that the Tank 1 transfer pump remains
in low speed setting during afterburner operation. During normal operation, each pump pressurizes the
Tank 1 and 4 transfer line as long as its tank has transfer fuel available. The SDC shuts down the
electric transfer pumps when the respective tanks are dry (Tank 1 empty, Tank 4 approximately 300
lb).
Jet level sensors (JLS) in the feed tanks control the flow of transfer fuel from the Tank 1 and 4
transfer line. For instance, Tank 2 does not accept fuel until its fuel quantity drops to approximately
2,100 lb, uncovering the JLS and opening the transfer valve. Tank 2 accepts fuel until its fuel quantity
reaches approximately 2,450 lb, covering the JLS and closing the transfer valve. Therefore, during
normal operation, Tank 2 fuel level cycles between 2,100 and 2,450 lb as long as transfer fuel is
available (JLS cycling).
Flapper valves in Tanks 1 and 4 provide a backup gravity transfer capability in certain circumstances. The flapper valve in Tank 4 is free flowing, gravity transferring to Tank 3 any time the Tank
4 fuel level is higher. Therefore, Tank 4 tends to keep Tank 3 full (near 2,600 lb) until the Tank 4 fuel
level drops below that of Tank 3 (wing tank fuel depleted). The flapper valve in Tank 1 is controlled
by left motive flow. The valve can be opened by the SDC following a Tank 1 transfer pump failure or
by loss of motive flow (left engine shutdown).
Since the Tank 4 transfer pump is not located on the bottom level of the tank, two motive flow
powered scavenge pumps, one routed to Tank 2 and the other to Tank 3, are installed to transfer the
last 300 lb of Tank 4. With empty transfer tanks, an excessive feed tank fuel split following symmetric
engine operation may indicate a Tank 4 scavenge pump failure. There is no SDC monitoring of the
Tank 4 scavenge pumps.
The Tank 1 and 4 transfer pumps are also used to dump fuel through the dump valve and to transfer
fuel to the ARS through the ARS replenishment valve.
2.2.2.1.1 Fuel Transfer Schedule/CG Control. The SDC implements a fuel transfer schedule (figure
2-4) designed to keep aircraft CG at an optimum location. The system periodically shuts off the Tank
1 transfer pump to keep Tank 1 and Tank 4 properly balanced. Fuel transfer scheduling operates until
Tank 4 drops below 300 lb or the FUEL LO caution comes on. When Tank 4 reaches 300 lb, Tank 1
should indicate 250 lb or below.
The FUEL XFER caution is set when Tank 1 and 4 fuel is not scheduling properly or wing tank
imbalance exceeds 350 lb. The caution is inhibited when the inflight refueling probe is extended.
2.2.2.2 Fuel Transfer - Internal Wing Tanks. Fuel is transferred from the wing tanks to Tank 4 by
two motive flow powered ejector pumps, one in each tank. When Tank 4 is less than full, the SDC opens
both wing motive flow control valves, which direct motive flow to the ejector pumps and transfer fuel
from the wing tanks to Tank 4. When Tank 4 is full, the motive flow control valves are closed and
normal wing transfer is inhibited.
If motive flow is lost on one side (single engine shutdown), the cross-motive shutoff valve opens so
that one motive flow system can power the ejector pumps in both wing tanks. If both motive flow
systems are lost, the wing tanks gravity transfer to Tank 4. Bank angle changes or a steady sideslip may
be required to gravity transfer all available wing fuel.
2.2.2.2.1 Wing Tank Balancing. The SDC incorporates wing tank balancing logic designed to keep
wing tank asymmetry below 200 lb. If wing tank asymmetry exceeds 200 lb, the SDC shuts off fuel
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-4. Tank 1 and 4 Fuel CG Control and FUEL XFER Caution Schedule
transfer from the lower tank by closing its wing motive flow control valve. If wing tank asymmetry
exceeds 350 lb for 15 seconds, the FUEL XFER caution appears.
Wing tank balancing also occurs during refueling, where the SDC alternately opens/closes the wing
refuel valves attempting to keep the wing tanks within 200 lb. The FUEL XFER caution is not set
during refueling. If SDC balancing logic cannot keep the wings from refueling asymmetrically (greater
than 350 lb), the FUEL XFER caution is set when the inflight refueling probe is retracted.
2.2.2.2.2 INTR WING Control Switch. The INTR WING control switch, located on the EXT LT
panel on the left console, is used to isolate the wing tanks (e.g., following battle damage).
INHIBIT
Prevents normal transfer and refueling of the wing tanks (closes both wing motive
control valves, both wing refuel valves, and switches both diverter valves, located
in Tank 3, from the wing tanks to the feed tanks).
NORM
Permits normal transfer and refueling of the internal wing tanks.
2.2.2.3 External Fuel Transfer. External fuel is transferred by regulated engine bleed air pressure
applied to all installed external tanks with WoffW. External tank pressurization is terminated for
inflight refueling (PROBE switch in EXTEND) and for arrested landing (both HOOK and LDG
GEAR handles down). With pressurization applied, external fuel transfer is controlled by the three
EXT TANKS transfer switches.
During external transfer, fuel is routed through the aircraft’s refuel/defuel line. Refuel valves in each
tank open only if commanded by the SDC and space is available. At MIL power and below, the SDC
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A1-F18EA-NFM-000
only allows external fuel to transfer to Tank 1 and the wing tanks. In afterburner, the SDC allows
external fuel to transfer to any internal tank that can accept it.
NOTE
In Lots 21 - 25, SDC CG control logic (CG Restart) inhibits external
fuel transfer in flight and on deck with ORIDE selected until Tank 1
fuel quantity depletes to approximately 900 lbs. After external fuel
transfer begins, erroneous empty (<200 lbs) external fuel quantity
readings, which may occur during climbs/dives, can reinitiate CG
Restart logic inhibiting external fuel transfer until Tank 1 fuel
quantity depletes again to approximately 900 lbs. To initiate
immediate external fuel transfer in flight or to check external fuel
transfer on deck (i.e., override CG Restart logic), the IFR probe must
be extended for at least 25 seconds prior to retraction with at least one
EXT TANKS switch in ORIDE.
2.2.2.3.1 EXT TANKS Transfer Switches. The three EXT TANKS transfer switches are located on
the FUEL panel on the left console and are labeled LM/RM, LI/RI, and CTR (left and right midboard,
left and right inboard, and centerline tanks, respectively). With the external tanks pressurized, fuel
transfers when the FUEL LO caution is displayed regardless of the position of the EXT TANKS
transfer switches.
ORIDE Applies pressurization and transfers fuel from all external tanks whose switches are not
in STOP. May be used to transfer external fuel during extended ground operations
(EXT TANK caution). Overrides any SDC stop transfer command.
NORM
Permits normal transfer and refueling of controlled external tank(s).
STOP
Prevents transfer and refueling of controlled external tank(s) except with a FUEL LO
caution.
NOTE
• For ARS configured aircraft: If fueling of the ARS is not desired during
aerial refueling, as the receiver, CTR ORIDE must be selected since
CTR STOP will not prevent fuel from entering the ARS. Selecting
CTR ORIDE will pressurize all external fuel tanks and significantly
reduce refueling rate.
• If the ARS control panel is installed, centerline external tank transfer
is inhibited unless the STORE switch is in FROM. ORIDE on the CTR
EXT TANKS switch will have no effect.
2.2.3 Fuel Tank Pressurization and Vent. In Lots 21 thru 23, the internal fuel tank pressurization
system provides regulated engine bleed air pressure to all internal tanks to prevent fuel boil-off at
altitude. Bleed air pressure is applied to all internal tanks with WoffW. Pressurization is terminated
for inflight refueling (PROBE switch in EXTEND) and for arrested landing (both HOOK and LDG
GEAR handles down). In Lots 24 and up, internal tank pressurization is supplied by ram air only from
the vertical tail vents and no longer augmented by engine bleed.
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ORIGINAL
A1-F18EA-NFM-000
The vent system provides over-pressure and over-fill relief for the internal tanks. Pressurization is
applied to the vent lines in the two fuselage vent tanks. The vent lines connect all internal tanks and
are ported through the fuselage and vertical tail vent tanks to outlets located on the side of the vertical
tail. Normally, the vent lines contain only pressurized air; however, if a refuel valve failure overfills an
internal tank, fuel flows through the vent lines to the fuselage vent tanks. Two motive flow powered
vent tank scavenge pumps return this fuel from the fuselage vent tanks to the feed tanks.
Additionally, the vent system provides pressure relief of the internal fuel tanks during climbs and
vacuum relief during descents, if the pressurization system fails.
2.2.4 Thermal Management System. The thermal management system uses fuel from the high
pressure fuel stage of the motive flow/boost pump to cool the FADECs, liquid coolant, and AMAD and
hydraulic oils. The high pressure motive flow output of the pump has four branches.
The first branch is used to run motive flow powered pumps and valves in the fuel system. The second
branch directs cooling flow to the FADEC and exits into the fuel recirculation return line.
The third branch runs through the liquid coolant/fuel heat exchanger and the combined AMAD oil
and hydraulic oil/fuel heat exchanger in order to cool those fluids. A hot fuel diverter valve in the third
branch either directs fuel away from the engine and into the fuel recirculation return line or directs fuel
to the engines where it is combined with fuel feed from the motive flow/boost pump and burned.
Recirculated fuel first passes through a fuel/air heat exchanger bypass valve, which either directs the
fuel through or around the fuel/air heat exchanger. When sufficient recirculation fuel flow is present,
the SDC may open the heat exchanger bypass valve (Mach greater than approximately 0.35) to direct
hot fuel through the fuel/air heat exchanger. Then, the fuel passes through a diverter valve, located in
Tank 3, which either directs the fuel to the wing tanks or the feed tanks. Fuel directed to the wing tanks
then flows to Tank 4 for additional cooling. When Tank 4 is below 300 lb, recirculation fuel is returned
to the feed tanks by the Tank 4 scavenge pumps. When recirculation fuel is diverted to the wing tanks,
modulation of wing fuel quantities on cockpit fuel displays is noticeable at lower wing fuel levels.
The engine thermal control valve (TCV), located in the engine fuel control unit, maintains engine
combustor nozzle, engine lube oil, and aircraft accessories within their temperature limits. When the
system determines that more cooling is required (typically due to hot weather or low fuel levels), the
TCV opens, directing feed fuel into the recirculation return line. With the TCV open, greater cooling
flow is induced through the engine lube oil and aircraft accessory heat exchangers, and the FADEC,
ultimately reducing system temperatures.
The fourth branch runs to the cross cooling valve which opens following a motive flow system failure,
allowing one motive flow system to cool both FADECs and both engines’ accessories.
During ground operations, when temperatures exceed 40°C (30°C in LOT 26 and up), the liquid
coolant pump and Liquid Cooling System (LCS) ground cooling fan may be commanded on (if not
already on) to provide LCS cooling of the fuel system. In LOT 25 and below, the bypass valve, which
allows liquid coolant to the liquid coolant/fuel heat exchangers, is solely controlled by the liquid
coolant temperature, so the amount of fuel cooling provided (if any) depends on the temperature of
each fluid. In LOT 26 and up, the bypass valve is controlled by the ECS controller and will only direct
liquid coolant to the liquid coolant/fuel heat exchanger if the RADAR knob is in OFF.
NOTE
In LOT 26 and up, the RADAR knob must be in OFF in order to
provide any postflight LCS fuel cooling.
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ORIGINAL
A1-F18EA-NFM-000
In either case, placing the RADAR knob to OFF removes the radar as a heat source and should
extend ground operating time.
2.2.4.1 Fuel/Air Heat Exchanger. A fuel/air heat exchanger is located above each engine inlet, near
the leading edge. When the heat exchanger bypass valve is open, the heat exchanger uses inlet air to
provide additional fuel cooling. Air is drawn from the inlet through several banks of small pin holes
(bleed plates) and is exhausted through the spoiler opening on the upper surface of the LEX/fuselage.
2.2.4.2 Fuel/Air Heat Exchanger Leak Detection. Since a leak in the fuel/air heat exchanger can
result in engine fuel ingestion through the bleed plates, a leak detection system is incorporated. If a fuel
leak is detected during fuel/air heat exchanger operation, the SDC closes the heat exchanger bypass
valve and isolates the heat exchanger.
Leak detection logic is only capable of detecting a leak greater than approximately 400 pph in the
fuel air heat exchanger. Fuel flow through the heat exchangers is inhibited below approximately Mach
0.35 and anytime the ECS auxiliary scoops are deployed to guard against potential for engine inlet fuel
ingestion.
A leak of less than 400 pph can, however, be discovered during the post-flight switching valve checks.
Following engine shutdown, the SDC opens the cross cooling valve and the heat exchanger bypass valve
on the non-operating side for 20 seconds. If ground crew observe fuel exiting from the fuel/air heat
exchanger drainage ports (bottom, inboard edge of the inlet), a leak exists.
2.2.5 Refueling System. The aircraft can be refueled on deck through a single point refueling
receptacle or inflight through a hydraulically actuated inflight refueling probe. The refueling
receptacle is located behind door 8R on the forward right fuselage. The refueling probe is located on
the upper right side of the fuselage forward of the windscreen. A fuel pressure regulator/surge
suppressor is installed downstream of the refueling probe in order to control pressure spikes associated
with inflight refueling. Fuel from the single point receptacle or the refueling probe enters the
refuel/defuel line and is routed to all internal and external tanks. During refueling, the SDC opens all
refuel valves, allowing fuel to transfer into all internal tanks. External tank pressurization is
terminated when the probe is extended, allowing the refuel/defuel line to fill all installed external tanks
(EXT TANKS switch(es) not in STOP).
2.2.5.1 PROBE Switch. The guarded PROBE switch, located on the FUEL panel on the left console,
is used to extend and retract the inflight refueling probe.
EXTEND Extends the inflight refueling probe using HYD 2A pressure, energizes the probe
light (external lights master switch in NORM), and depressurizes all internal and
external tanks.
RETRACT Retracts the inflight refueling probe using HYD 2A pressure, deenergizes the probe
light, and repressurizes the internal and external tanks. The probe cannot be
retracted if HYD 2A pressure is not available.
EMERG
EXTD
Emergency extends the inflight refueling probe using either HYD 2B or APU accumulator pressure, energizes the probe light (external lights master switch in NORM),
and depressurizes all internal and external tanks.
2.2.6 Fuel Dump System. The fuel dump system allows all fuel except feed tank fuel to be dumped
overboard. The dump valve, controlled by the DUMP switch, is located in the Tank 1 and 4 transfer
line. With the dump valve open, the Tank 1 and 4 transfer pumps (high-speed setting) force fuel out
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ORIGINAL
A1-F18EA-NFM-000
the dump outlet, located on the trailing edge of each vertical tail. Wing tank fuel is dumped by
transferring to Tank 4 with the INTR WING switch in NORM. External fuel is dumped by
transferring to Tanks 1 and 4 only with the EXT TANKS switch(es) in NORM or ORIDE.
NOTE
Anytime four external fuel tanks are loaded on wing stations (3, 4, 8,
and 9), selecting ORIDE on LI/RI external transfer switch will improve
dump performance and external transfer rate by commanding
simultaneous transfer of all external tanks vs. normal transfer sequence
(tanks on Stations 3/9 must be empty prior to tanks on Stations 4/8
transferring). Performing this function imposes airspeed limitations
defined in Figure 4-12.
Dump rate is typically in excess of two engine MAX power fuel flow, approximately 1,300 lb per
minute (78,000 pph). Fuel dumping continues until:
a. The DUMP switch is placed to OFF.
b. The BINGO caution comes on.
c. Tanks 1 and 4 are empty and all available fuel from internal wing and external tanks has been
depleted.
d. The FUEL LO caution comes on.
Simultaneous selection of fuel dump and afterburner during high AOA
maneuvering may ignite fuel and cause fuselage damage.
2.2.6.1 Fuel DUMP Switch. The lever locked fuel DUMP switch, located on the FUEL panel on the
left console, is spring loaded to the OFF position and electrically held in the ON position.
ON
Opens the dump valve, allowing transfer tank and external tank fuel to be dumped.
The switch reverts to OFF with a BINGO or FUEL LO caution. With either caution,
holding the switch in the ON position with Tank 1 and/or 4 fuel available reinitiates
fuel dump.
OFF
Dump valve closed
2.2.7 Fuel Quantity Indicating System. The fuel quantity indicating system measures the individual
fuel quantities in all internal and external fuel tanks and provides cockpit readouts for individual
tanks, total internal fuel, and total fuel onboard. Quantities are displayed on the EFD and the FUEL
display, rounded to the nearest 10 pounds (figure 2-5). Actual fuel tank probe readings are displayed
on the FUEL QTY display selected from the SUPT MENU/BIT/STATUS MONITOR display.
While the volume of fuel with full tanks does not change, the fuel quantities listed in pounds vary
with changes in temperature and fuel density. Full internal fuel quantity can vary from 13,250 to 15,960
lb for the F/A-18E or from 12,410 to 14,940 lb for the F/A-18F at fuel temperatures of 100° and -40°F,
respectively. Figure 2-5 lists standard day fuel quantities for JP-5, JP-8, and JP-4.
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A1-F18EA-NFM-000
Figure 2-5. Fuel Quantity
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2.2.7.1 EFD - Engine Fuel Display, Fuel Parameters. The EFD displays fuel, maintenance code, and
consumables information in green digits on a black background (figure 2-6). The top level EFD format
graphically displays total internal fuel, bingo level, and fuel quantities for up to five external fuel tanks.
The shaded regions of each graphic display the ratio of fuel available to fuel capacity. Digital readouts
of total fuel (large numbers), total INT fuel, external fuel tank quantities, and current BINGO setting
are also provided but are truncated to 100 lb increments.
The internal fuel format graphically displays only internal fuel tank quantities. Digital readouts of
feed tank fuel quantities are also provided, truncated to 100 lb increments.
During battery start of the engines, the EFD displays either three or five external fuel tanks if one
or more external fuel tanks are installed. For instance, if a full centerline tank is installed, the EFD
displays a full centerline graphic and 3.2 lb fuel. The inboard/midboard tanks appear empty reading
0.0 lb fuel. When the SMS completes startup BIT and initial inventory, the EFD displays only the
installed tank(s).
The aircraft fuel load can be checked on battery power by placing the ENG CRANK switch to L or
R, by starting the APU, or by resetting the MSP codes in the nose wheelwell. In the F/A-18F, the total
fuel load and the tank one fuel ratio will be in error, since the SDC defaults to F/A-18E settings while
on battery power. The total fuel load will appear 200 to 400 lb high, and the Tank 1 graphic (as
displayed on the internal format) will indicate approximately two-thirds when Tank 1 is, in fact, full.
When the MCs become operative with ac power applied, the total fuel load will be correct, and the EFD
will show the correct Tank 1 fuel ratio.
Bingo level is adjusted in hundred pound increments by rotation of the BINGO knob and thousand
pound increments by pull and rotation (clockwise is increasing). The MODE button toggles the EFD
between the top level and internal fuel formats. Pressing the MODE button for greater than 1.5
seconds selects the MSP format. With the MSP format displayed, the BINGO knob cycles between
MSP code pages if more than one is present. Pressing the MODE button again for greater than 1.5
seconds runs and displays the results of a consumables test. Subsequent actuation for less than 1.5
seconds returns to the last displayed top level or internal fuel format. Clockwise rotation of the BRT
knob increases the display’s brightness.
In the F/A-18F, the formats can be displayed on the aft EFD independently from the front EFD.
2.2.7.2 FUEL Display. The FUEL display (figure 2-7) is selected by the FUEL option on the SUPT
MENU. The FUEL display lists TOTAL fuel (internal and external), total INTERNAL fuel, the
available fuel in each tank, and the current BINGO setting. A moving caret is shown on the right side
of each tank, indicating the ratio of fuel available to tank capacity.
The SDC continuously monitors the validity of the fuel probes installed in each tank. If all probes
in a tank (except the feed tanks) are declared invalid, the SDC displays 0 lb fuel and the INV (invalid)
cue next to the tank. If one or more probes in a multi-probe tank are declared invalid, the SDC displays
the total of the valid probes only and the EST (estimated) cue next to the tank. If a feed tank fuel probe
is invalid, the SDC displays 1,125 lb (0 lb if the FUEL LO caution is set) and the EST cue. The TOTAL
and INTERNAL fuel values indicate the sum of all valid and estimated tank quantities with EST cues
(INV cues if any tank is INV).
The FLBIT option on the FUEL display is used to initiate a BIT of the fuel low level indicating
system. The FLBIT option remains boxed during BIT. A satisfactory test results in a FUEL LO
caution and voice alert within 13 seconds of BIT initiation. FLBIT cannot be initiated with a FUEL
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Figure 2-6. Engine Fuel Display (EFD) - Fuel Parameters
LO caution set or an SDC failure. An SDC RESET option is provided to command an SDC software
reset.
2.2.7.3 F-QTY Advisory. The F-QTY advisory indicates an SDC or fuel quantity indicating system
failure which affects the accurate display of fuel quantity or CG information. The advisory is activated
when:
a. The MC loses communication with the SDC.
b. The SDC reports an internal or gauging system failure.
c. Any tank quantity is INV.
d. The SDC reports its output discretes are not working.
If a F-QTY advisory results from loss of MC communication with the SDC or the SDC reporting an
internal or gauging system failure, the MC is unable to report actual fuel quantities and the following
are displayed on the FUEL display (GLIM 7.5G caution):
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Figure 2-7. FUEL Display
a. All fuel quantities (except TOTAL) are held at their last displayed value.
b. TOTAL fuel is estimated by the MC based on the last valid fuel quantity and engine fuel flow.
c. A flashing ESTIMATED cue is displayed along with a minutes and seconds (XX:XX) timer
which indicates the duration since the displayed fuel quantities were last updated.
2.2.8 Fuel Low Level Indicating System. The fuel low level indicating system is completely
independent of the fuel quantity indicating system. When the fuel level in either feed tank drops to
approximately 1,125 lb, a FUEL LO caution, caution light, and voice alert are activated, and the
affected fuel tank quantity is displayed in inverse video on the EFD. If a low level indication was set
by a transient condition, such as prolonged negative g flight, the cautions remain on for 1 minute after
the low level indication is removed.
If the FUEL LO cautions are set, assume that at least one feed tank is
below approximately 1,125 lb regardless of displayed fuel quantity
indications.
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2.2.9 Fuel System Related Cautions. The following fuel system related cautions are described in the
Warning/Caution/Advisory Displays in Part V:
D
D
D
D
D
D
D
D
L or R BOOST LO
L or R FUEL HOT
L or R FUEL INLT
FUEL XFER
EXT XFER
EXT TANK
TK PRES LO
TK PRES HI
D
D
D
D
D
D
D
L or R THERMAL
FADEC HOT (WonW, IDLE or above)
PROBE UNLK
REFUEL DR
DUMP OPEN
FUEL LO caution, caution light, and voice alert
BINGO caution and voice alert
2.3 FPAS - FLIGHT PERFORMANCE ADVISORY SYSTEM
NOTE
FPAS is functional in MC OFP H1E AND UP.
The flight performance advisory system (FPAS) is provided to aid the pilot in making time, fuel, and
distance calculations. Readouts for maximum range and maximum endurance are provided for three
flight conditions: current Mach and altitude, optimum Mach at the current altitude, and optimum
Mach and altitude. These three readouts can be used to adjust the aircraft flight profile to meet
mission requirements. Additionally, FPAS provides fuel remaining at arrival and recommended
distance to begin descent from the selected waypoint or TACAN station. Range, time, altitude, Mach,
and fuel are calculated by the FPAS algorithm and appear on the FPAS display.
2.3.1 FPAS Display. The FPAS display (figure 2-8) appears when the FPAS option is selected from
the SUPT MENU. The display is divided into five areas: the current range and endurance area, the
waypoint/TACAN steering area, the fuel flow area, the optimum range and endurance area, and the
default area. With engines running and WonW, only the optimum and default areas are valid. With
WoffW, all five areas are valid. Waypoint/TACAN steering information is provided with WoffW and
waypoint or TACAN steering selected (boxed) on the HSI display.
2.3.1.1 FPAS CURRENT RANGE and ENDURANCE. The CURRENT RANGE area displays three
calculations: the range in nautical miles TO 2000 LB fuel remaining at the current altitude and Mach,
the BEST MACH to fly at the current altitude to maximize range, and the range TO 2000 LB fuel
remaining if the BEST MACH is flown. If parameters used to calculate current range become invalid,
X’s replace the current range value, and an FPAS advisory replaces the current endurance value.
The CURRENT ENDURANCE area displays three calculations: the time in hours and minutes TO
2000 LB fuel remaining at the current altitude and Mach, the BEST MACH to fly at the current
altitude to maximize endurance, and the endurance TO 2000 LB fuel remaining if the BEST MACH
is flown. If parameters used to calculate current endurance become invalid, X’s replace the current
endurance value.
If IMN exceeds Mach 0.9, a MACH advisory replaces the current range value, and a LIM advisory
replaces the current endurance value. When total fuel drops below 2,500 lbs, the FPAS calculations
shift to 0 lbs remaining, and the TO 2000 LB legend changes to TO 0 LB.
2.3.1.2 FPAS Waypoint/Tacan Steering and the HSI Display. If waypoint or TACAN steering is
selected (boxed) on the HSI display, the selected waypoint or TACAN station is displayed under the
NAV TO legend, and the arrival time and fuel remaining at arrival (at the current flight conditions)
are displayed under the TIME and FUEL REMAIN legends, respectively.
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Fuel remaining at arrival and recommended distance to begin descent from the selected steering
source are displayed on the HSI display. If the recommended distance to begin descent is greater than
99 miles, 99 is displayed.
If FPAS waypoint/TACAN steering parameters become invalid, X’s replace the fuel remaining and
descent distance values. If IMN exceeds Mach 0.9, fuel remaining values are blanked. If fuel remaining
at the selected steering source is less than the TO 2000 LB or TO 0 LB legend, the WYPT number, the
TO 2000 (0) LB legend, and the fuel remaining value flash on the FPAS display and the fuel remaining
value flashes on the HSI display. If fuel remaining is calculated to be less than 0 lbs, 0 is displayed.
2.3.1.3 FPAS Fuel Flow. The total fuel flow rate (both engines) is displayed in pounds per nautical
mile under the LB/NM legend whenever the engines are running.
2.3.1.4 FPAS OPTIMUM RANGE and ENDURANCE. The OPTIMUM RANGE area displays three
calculations: the optimum ALTITUDE and MACH to fly to achieve the displayed maximum range TO
2000 LB. If parameters used to calculate optimum range become invalid, X’s replace the altitude,
Mach, and range values.
The OPTIMUM ENDURANCE area displays three calculations: the optimum ALTITUDE and
MACH to fly to achieve the displayed maximum endurance time TO 2000 LB. If parameters used to
calculate optimum range or optimum endurance become invalid, X’s replace the altitude, Mach, range,
and time values.
When total fuel onboard drops below 2,500 lbs, the FPAS calculations shift to 0 lbs remaining, and
the TO 2000 LB legend changes to TO 0 LB.
2.3.1.5 FPAS Default Area. If outside air temperature, stores drag, or fuel flow become invalid, the
TEMP, DRAG, or FF advisories are displayed next to the DEFAULT legend. These parameters, if
invalid, do not have a fatal impact on FPAS calculations.
2.3.2 FPAS CLIMB Option. The CLIMB option is available for selection in the NAV master mode.
Pressing the CLIMB option on the FPAS display enables the climb airspeed prompt, displayed above
the airspeed box in the HUD (HUD reject switch in the NORM position). When selected, CLIMB is
boxed and the climb airspeed prompt indicates the desired calibrated airspeed for an optimum climb
profile.
2.3.3 FPAS HOME Waypoint Selection. The HOME option arrows are used to increment/decrement
the home waypoint for use in FPAS fuel-on-deck (HOME FUEL caution) calculations. The selected
home waypoint is displayed above the HOME legend on the FPAS display. The home waypoint
defaults to 0 at power up and must be changed if another waypoint is desired (0 to 59). Decrementing
the home waypoint from 0 selects 59. If invalid parameters prevent FPAS from calculating the HOME
FUEL caution, the FPAS DDI advisory is displayed, the home waypoint is X’d, and the option arrows
are removed.
2.3.4 FPAS HOME FUEL Caution. When FPAS calculated fuel remaining at the selected home
waypoint reaches 2,000 lbs, the HOME FUEL caution is displayed. HOME FUEL caution logic is
disabled with WonW, the refueling probe extended, the landing gear cycled down then up, or within
5 seconds after a home waypoint change.
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Figure 2-8. FPAS Displays
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Figure 2-9. Secondary Power Supply
2.4 SECONDARY POWER SYSTEM
The aircraft secondary power system contains two airframe mounted accessory drives (AMAD) and
a single auxiliary power unit (APU). Figure 2-9 shows the major components of the secondary power
system.
2.4.1 AMAD - Airframe Mounted Accessory Drive. During normal operation, each AMAD is
mechanically driven by its corresponding engine through a power transmission shaft and is used to
drive a fuel boost/motive flow pump, an ac/dc electrical generator, and a 3000/5000 psi hydraulic pump.
Pneumatic pressure is used to rotate an air turbine starter (ATS) on each AMAD for engine crank/start
capability.
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For ground maintenance use, either AMAD can be decoupled from its engine, allowing pneumatic
pressure to drive the AMAD and its accessories.
NOTE
Failure of the power transmission shaft (PTS) results in the display of
the associated GEN, BOOST LO, and both HYD circuit cautions.
2.4.2 APU - Auxiliary Power Unit. The APU is a small gas turbine engine used to generate a source
of air to power the ATS for normal engine start or to provide an alternate air source for the
environmental control system (ECS). The APU is located between the engines, with intake and exhaust
facing downwards.
The aircraft battery provides electrical power for APU ignition and start. A hydraulic motor powered
by the APU accumulator is used to start the APU. The APU receives fuel from the left engine feed line
upstream of the left engine feed shutoff valve. During normal operation, the APU shaft turns a
separate compressor which supplies air for main engine start or alternate ECS operation.
If an APU fire or overheat condition is detected on the ground, the APU fire extinguishing system
automatically shuts the APU down and, after 10 seconds, discharges the extinguisher bottle.
2.4.3 APU Switch. The APU switch, located on the left console, is spring loaded to the OFF position
and is electrically held in the ON position.
ON
Automatic start and normal APU operation. The switch returns to OFF 1 minute after
the second aircraft generator comes online (BLEED AIR knob not in AUG PULL).
OFF
Manual APU shutdown
To prevent an APU running engagement and to prevent APU exhaust
torching, a minimum of 2 minutes must elapse between APU shutdown
and another APU start.
2.4.3.1 APU READY Light. The APU READY light, located on the left console adjacent to the APU
switch, comes on when the APU has completed the start cycle and is capable of supporting engine
crank.
2.4.4 ATS Air Sources. Pneumatic pressure from one of three sources can be used to power the ATS
for engine crank/start: APU compressor air, opposite engine bleed air (crossbleed), or external air.
The APU compressor is the primary engine crank air source. With the APU online, the ECS air
isolation valve is closed and the ENG CRANK switch opens the desired air turbine starter control valve
(ATSCV), allowing APU compressor output to turn the ATS, AMAD, and engine core.
A crossbleed start can be utilized when one engine is operating and the APU is shutdown. The
operating engine should be set to a minimum of 80% N2 rpm to make sure bleed air output is sufficient
to crank the opposite engine. For a crossbleed start, the ENG CRANK switch opens the ECS air
isolation valve and the ATSCV, allowing compressor bleed air pressure to turn the ATS, AMAD, and
engine core. If one engine fails inflight and the engine core is rotating freely, crossbleed air may be used
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A1-F18EA-NFM-000
to rotate the AMAD and retain some fuel, electrical, and hydraulic system output.
ATS exhaust may blister paint and cause possible door damage on the aft
underside of the fuselage during extended crossbleed operation of a failed
engine.
An external air source may also be used to start one or both engines. External air is applied to the
aircraft through a connection in the right main wheelwell. Both engine bleed air valves must be closed
(BLEED AIR knob OFF) to make sure that external air is the sole air source for engine start. For an
external air start, the ENG CRANK switch opens the ATSCV, allowing external air pressure to turn
the ATS, AMAD, and engine core.
2.4.4.1 ATS Protection. Each ATS has two sources of overspeed cutout protection. The primary
source is the corresponding generator, and the backup source is the frequency sensing relay (FSR). The
FSR monitors ATS speed and provides the signal which electrically holds the ENG CRANK switch.
When the generator comes on the line at 60% N2 rpm, it removes power from the ATSCV and the FSR,
which releases the ENG CRANK switch. If the primary cutout does not function (GEN switch OFF or
major generator malfunction), the FSR releases the ENG CRANK switch when it senses 63% N2 rpm.
Regardless of the engine start air source utilized, the corresponding GEN
switch should be ON, as the generator provides primary overspeed cutout
protection for the ATS.
2.4.4.2 ENG CRANK Switch. The ENG CRANK switch, located on the left console, is spring loaded
to the OFF position and is electrically held in the L or R position.
L
Opens the left ATSCV and/or the ECS air isolation valve to direct pneumatic pressure
to the ATS for left engine crank.
OFF
Closes both ATSCVs and the ECS air isolation valve. When the left or right generator
comes online following engine start, the switch automatically returns from L or R to the
OFF position.
R
Opens the right ATSCV and/or the ECS air isolation valve to direct pneumatic pressure
to the ATS for right engine crank.
2.4.5 AUG PULL. During extended ground operations, APU compressor air may be used instead of
engine bleed air to run the ECS and cool the avionics (BLEED AIR knob in AUG PULL). AUG PULL
operation is discussed in the ECS section.
2.4.6 AMAD Related Cautions. The following AMAD related cautions are described in the Warning/
Caution/Advisory Displays in Part V:
• L or R OIL HOT
• L or R AMAD PR
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• L or R ATS
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A1-F18EA-NFM-000
2.5 ELECTRICAL POWER SUPPLY SYSTEM
The electrical power supply system consists of two generators, two transformer-rectifiers (TR), one
battery with dedicated battery charger, and a power distribution (bus) system (figure 2-10). Each
generator provides a primary ac source and three isolated dc sources from a permanent magnet
generator (PMG). During normal operation, the left generator powers only the left buses while the
right generator powers only the right buses. If one generator fails, the other generator is capable of
carrying the entire electrical load of the aircraft. Battery power is provided for normal engine start.
External electrical power can be applied to power the entire system on the ground. The bus system
consists of the left and right 115 vac buses, right 26 vac bus, left and right 28 vdc buses, 28 vdc essential
bus and a 28 vdc maintenance bus. See figure 2-10 for a simplified schematic and Electrical System,
foldout section, for the specific systems powered by each bus.
2.5.1 Electrical RESET Button. The electrical system RESET button is located on the electrical
power panel on the right console. This button provides master reset capability for any failed generator
or electrical system relay without interrupting operational circuits.
2.5.2 AC Electrical Power. The two generators are the primary source of ac electrical power. With the
GEN switch in NORM, each generator comes online at approximately 60%N2 rpm as long as voltage
and frequency are within limits. Each generator supplies ac power to an independent 115 vac bus. In
addition, the right 115 vac bus powers a 26 vac bus through a dedicated transformer.
2.5.2.1 GEN Switches. Two generator control switches, labeled L GEN and R GEN, are located on
the electrical power panel on the right console.
NORM
Provides normal generator operation.
OFF
Removes the generator ac source from the bus system.
2.5.2.2 Electrical Fault Protection Circuitry. The electrical system provides fault protection with
generator isolation, bus tie, generator automatic reset, and ac bus isolation circuitry.
If a generator fault occurs, generator isolation circuitry removes the affected generator from its buses
(L or R GEN caution and caution light). All generator faults except N2 underspeed require manual
generator reset (GEN switch cycled to OFF then NORM or RESET button pressed). Generator reset
is successful only if the out-of-tolerance or fault condition has cleared. If the generator fault remains,
bus tie circuitry allows the remaining generator to power all electrical buses. During an N2 underspeed
condition, the affected generator is automatically restored when rpm returns to normal range.
If a short or overload condition (bus or equipment fault) occurs on a bus (the R 115 vac bus, for
instance), the following sequence is initiated. The right generator attempts to power through the short
and, if unsuccessful, trips offline. The left generator is then connected to the right buses by the bus tie
circuitry, attempts to power through the short, and, if unsuccessful, also trips offline. Approximately
1 second after the dual generator outage, the generator automatic reset logic resets both generators. If
the bus or equipment fault has cleared, both generators remain online to power their respective buses.
If the bus or equipment fault remains, the right generator trips offline again, but bus isolation circuitry
now prevents the left generator from picking up the right buses (GEN TIE caution light). The left
generator remains online to power the left buses and the R 28 vdc bus. The R 115 vac bus, R 26 vac bus,
and battery charger are unpowered, and the battery runs the maintenance bus. This entire process may
take as long as 16 seconds.
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Figure 2-10. Simplified Electrical Schematic
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For the ac bus isolation and generator automatic reset circuits to operate, the GEN TIE switch must
be in NORM, the BATT switch must be ON and the PARK BRAKE handle must not be set. With the
PARK BRAKE handle set, the generators do not reset following a dual outage.
2.5.2.3 GEN TIE Caution Light. During initial engine start (battery or external power) GEN TIE
circuitry requires a set PARK BRAKE handle to properly function. If the PARK BRAKE handle is not
set during right (first) engine start, a GEN TIE caution light comes on when the right generator comes
online. For a battery start, setting the PARK BRAKE handle and cycling the R GEN switch reties the
left and right buses and clears any avionics faults that would otherwise occur. For an external power
start, setting the PARK BRAKE handle, disconnecting external power, and cycling the GEN TIE
switch reties the left and right buses.
2.5.2.4 GEN TIE Switch. The red-guarded GEN TIE switch is located on the left console outboard of
the exterior lights panel.
RESET
Resets the bus tie circuitry. Reset is accomplished by cycling the switch to RESET
then NORM.
NORM
With the BATT switch ON, enables the bus tie, ac bus isolation and generator
automatic reset circuits.
If the left and right buses are isolated because of a detected fault (e.g., R
GEN caution and GEN TIE caution light), cycling the GEN TIE switch
reenergizes the faulty bus/equipment and may cause further damage or
loss of the remaining generator.
2.5.3 DC Electrical Power. DC electrical power is provided by two TRs, three dc outputs from each
PMG, the battery, and the battery charger.
2.5.3.1 TR - Transformer-rectifiers. While each TR is powered by its respective 115 vac bus, the
output of each TR is connected in parallel, powering both the left and right 28 vdc buses and providing
primary power for the essential bus. If one TR fails, the other powers the entire dc system. There is no
cockpit warning of a single TR failure (no caution and no MSP code). TR operation is checked on
maintenance phase inspections.
2.5.3.2 PMG - Permanent Magnet Generator. The PMG in each generator provides three dc sources,
two for FCC channels and one for essential bus backup. The left PMG provides the primary power
source for FCC A (channels 1 and 2), while the right PMG provides the primary power source for FCC
B (channels 3 and 4). Regardless of generator control switch position, the PMGs come online when the
engine reaches approximately 50% N2 rpm on spool up and remain on until 20% N2 rpm on spool
down.
2.5.3.3 Battery. The primary operational use of the battery is engine start. The battery powers the
maintenance bus directly, allowing operation of the canopy, ladder, and maintenance monitor in the
absence of ac electrical power. With the BATT switch ON (first engine start), the battery also powers
the essential bus. In the unlikely event of a total ac/dc failure inflight, the battery provides the last
level of essential bus backup capability, providing about 5 to 10 minutes of power for the FCCs, after
which aircraft control is lost.
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A1-F18EA-NFM-000
Regardless of BATT switch position, the battery charger supplies charging power to the battery
anytime the right 115 vac bus is energized.
2.5.3.4 FCC Electrical Redundancy. FCC electrical redundancy is provided by several sources of dc
power (figure 2-11). The primary dc source for each FCC channel is its respective PMG output. If a
PMG output should fail, that FCC channel is powered by the essential bus, which also has several
sources of redundancy (the TRs, the PMGs, and the battery). The BATT switch must be ON for either
a PMG or the battery to power the essential bus. Additionally, ‘‘keep alive’’ circuits connected directly
to the maintenance bus provide each FCC channel with an uninterrupted backup power source during
normal bus power transients.
2.5.3.5 BATT Switch. The BATT switch is located on the electrical power panel on the right console.
ON
Allows the battery or either PMG to power the essential bus when TR power is not
available.
OFF
Prevents the battery or either PMG from powering the essential bus when TR
power is not available.
2.5.3.6 Automatic Battery Cutoff. The automatic battery cutoff circuit is provided to conserve
battery power. On the ground with the BATT switch ON, the circuit disconnects the battery from the
essential bus and returns the BATT switch to OFF 2 minutes after ac power is removed from the
aircraft. When battery cutoff is activated, the battery can be reconnected to the essential bus by
reselecting the BATT switch ON. The automatic battery cutoff circuit is disabled when the APU comes
online.
2.5.3.7 Battery Gauge. A battery gauge is installed on the electrical power panel on the forward right
console in the front cockpit only. Depending on generator status, the battery gauge provides an
indication of either essential bus voltage (both GENs offline) or maintenance bus voltage (either GEN
online).
With both GENs offline and the BATT switch ON (e.g., prior to first engine start), the battery gauge
is connected to the essential bus and indicates battery voltage. Nominal voltage for a ″good″ battery
should be 23 to 24 vdc. Minimum battery voltage is that which provides a successful engine start (e.g.,
APU remains online and the EFD remains powered to provide indications of RPM and TEMP). EFD
blanking and/or uncommanded APU shutdown should be anticipated with battery voltage at or below
18 vdc.
With at least one GEN online, the battery gauge is connected to the maintenance bus and indicates
28 vdc output of the battery charger. If the battery gauge fails to jump to approximately 28 vdc with
one GEN online, a battery charger malfunction has occurred which requires maintenance action prior
to flight.
If a dual GEN failure occurs, the battery gauge is reconnected to the essential bus and must be
referenced to determine the essential bus power source. If the battery gauge remains at approximately
28 vdc, an EBB PMG is powering the essential bus, and FCC operating time is not limited. However,
if the battery gauge indicates 24 vdc or below, the battery is powering the essential bus, and FCC
operating time is limited to about 5 to 10 minutes. The FCCs should continue to operate down to a
battery gauge voltage of approximately 18 vdc.
2.5.3.8 BATT SW Caution and Caution Light. The BATT SW caution and caution light are only set
to alert the aircrew of an improperly placed BATT switch. The battery gauge must be referenced to
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A1-F18EA-NFM-000
determine whether an EBB PMG or the battery is powering the essential bus following a dual GEN
failure. These cautions are set in only two circumstances.
1. The BATT switch is ON on the ground in the absence of ac power (e.g., first engine start). The
battery is depleting and the switch should be placed to OFF unless APU start is about to be made.
2. The BATT switch is OFF inflight and should be placed to ON to provide essential bus backup
capability from the PMGs and the battery.
2.5.4 External Electrical Power. External electrical power may be connected to the aircraft bus
system through an external power receptacle located on the left forward fuselage. If external power is
not of the proper quality, the external power monitor prevents application of power to the aircraft.
Actuation of 1 to 4 ground power switches is required to energize certain aircraft systems following
application of external power.
The aircraft buses are energized by external power in the same manner as if a generator were
operating provided the BATT switch is OFF or the PARKING BRAKE is set.
2.5.4.1 External Power Switch. The external power switch, located on the ground power panel on the
left console, is spring loaded to the NORM position (figure 2-12).
Figure 2-11. FCC Electrical Redundancy
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A1-F18EA-NFM-000
Figure 2-12. Ground Power Panel and Placard
RESET
Momentary actuation allows external power to be applied.
NORM
Aircraft buses are energized by external power, provided the switch was first positioned to RESET. The switch returns to OFF when external power is disconnected.
OFF
Disconnects external power from the aircraft.
2.5.4.2 Ground Power Switches. The four ground power switches are located on the ground power
panel on the left console (figure 2-12). Each switch controls a group of systems and/or instruments, as
listed on a placard above the panel.
A ON
Only systems/instruments listed for the A position are energized by external
power.
AUTO
All controlled systems/instruments are deenergized with external power on the aircraft. When a generator comes online, the switch(es) automatically revert to
AUTO.
B ON
All controlled systems/instruments (both A and B) are energized by external
power.
The first ground power switch placed to ON must be held for 3 seconds to complete an avionics
overheat BIT. If an avionics overheat condition is present, the switch(es) revert to AUTO and cannot
be returned to ON until the condition is corrected.
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Figure 2-13. Circuit Breaker Panels
2.5.5 Circuit Breakers. The circuit breaker panels (figure 2-13), located under each side of the
canopy sill outboard of the left and right consoles, contain the following circuit breakers:
Left Side
Right Side
LAUNCH BAR
FCS CHAN 2
FCS CHAN 1
FCS CHAN 3
FCS CHAN 4
HOOK
LG
2.5.6 Electrical System Cautions and Caution Lights. The following electrical system cautions and
caution lights are described in the Warning/Caution/Advisory Displays in Part V:
• L or R GEN caution and caution light
• GEN TIE caution light
• BATT SW caution and caution light
• L or R DC FAIL caution
2.6 LIGHTING
2.6.1 Exterior Lighting. Exterior lighting is utilized to highlight aircraft position and aspect to other
aircraft, to provide AOA indications to a landing signal officer (LSO), to light the aircraft path for
in-flight refueling, landing or taxi, and to distinguish the F/A-18E/F from other F/A-18 models. The
following exterior lights are provided: strobe lights, position lights, formation lights, approach lights,
refueling probe light, and landing/taxi light (figure 2-14).
Strobe lights and formation lights have two operating modes, normal and NVIS.
2.6.1.1 Exterior Lights Master Switch. The exterior lights master switch, located on the outboard
side of the left throttle grip, provides master control of all exterior lighting except the approach and
landing/taxi lights.
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Figure 2-14. Exterior Lights
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NORM
(Forward)
Power is available for controlled lighting (strobe and formation lights in normal
mode)
NVIS
(Center)
Power is available only to the strobe and formation lights in NVIS mode
OFF
(Aft)
Power is removed from all controlled lighting. Required position for Day ID light
strobe power ON.
2.6.1.2 Pattern Strobe Lights. Two red anti-collision strobe lights, one on each outboard vertical tail,
are provided to highlight aircraft position during both day and night operations. Each strobe light
contains two bulbs, one normal and one infrared (IR). The external lights master switch determines
which pair of bulbs are powered by the STROBE switch.
2.6.1.2.1 STROBE Switch. The STROBE switch, located on the EXT LT panel on the left console,
is used to apply power and control the brightness of the strobe lights.
BRT
Strobe lights on at full intensity (normal or NVIS mode)
OFF
Strobe lights off
DIM
Strobe lights on at reduced intensity (normal or NVIS mode)
2.6.1.2.2 IDENT Knob. Pattern selection is controlled by the IDENT knob on the exterior lights
panel. The IDENT knob can be set to select strobe patterns of NORM, or A thru F. For night carrier
landings the IDENT knob should be in the NORM position. Refer to figure 2-15 for possible strobe
patterns. Possible strobe patterns are as follows:
NORM
Strobe light flashes three times, pauses for 2.44 seconds, then repeats pattern
A
Strobe light flashes two times, pauses 1.92 seconds, then repeats pattern
B
Strobe light flashes once, pauses 0.64 seconds, flashes two times, pauses 2.56 seconds,
then repeats pattern
C
Strobe light flashes two times, pauses 0.64 seconds, flashes once, pauses 2.56 seconds,
then repeats pattern
D
Strobe light flashes three times, pauses 2.88 seconds, flashes once, pauses 0.64 seconds,
then repeats pattern
E
Strobe light flashes three times, pauses 0.64 seconds, flashes two times, pauses 3.2 seconds, then repeats pattern
F
Strobe light flashes two times, pauses 0.64 seconds, flashes two times, pauses 2.88 seconds, then repeats pattern
2.6.1.3 Position Lights. The position lights are provided to highlight the aircraft aspect during night
or reduced visibility operation. There are seven position lights, three red, three green, and one white.
The white position light is on the tail. Three red position lights are installed on the left side. One on
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the wingtip, one on the LEX just forward of the wing root and one under the wing at the aileron hinge.
Green position lights are installed at the same locations on the right side.
2.6.1.3.1 POSITION Lights Knob. The POSITION lights knob, located on the EXT LT panel on the
left console, is used to apply power and control the brightness of the position lights (external lights
master switch NORM). The knob provides variable lighting intensity between the OFF and BRT
positions.
2.6.1.4 Formation Lights. Ten formation ″strip″ lights, five on each side of the aircraft, are provided
to highlight the aircraft aspect during night or low visibility formation flight. Strip lights are located
on the forward fuselage forward of the LEX, on the wingtip above the missile rail, on the wingtip below
the missile rail, on the aft fuselage below the vertical tail, and on the vertical tail. Each formation light
contains two lighting strips, one normal and one IR. The external lights master switch determines
which set of strips are powered by the FORMATION lights knob.
2.6.1.4.1 FORMATION Lights Knob. The FORMATION lights knob, located on the EXT LT panel
on the left console, is used to apply power and control the brightness of the formation strip lights. The
knob provides variable lighting intensity between the OFF and BRT positions in either the normal or
NVIS mode.
2.6.1.5 Approach Lights. The approach lights, located on the nose gear strut, provide AOA
indications to an LSO during carrier landings. Three approach lights are provided to indicate a fast
(red), on-speed (amber), or slow (green) AOA condition. The approach lights are powered with WoffW
and all landing gear down and locked. Therefore, the approach lights are an external indication that the
Figure 2-15. ID Strobe Patterns
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A1-F18EA-NFM-000
landing gear are down and locked. The approach lights flash when the HOOK BYPASS switch is in the
CARRIER position and the arresting hook is not down, indicating to an LSO that the hook must be
lowered for a carrier arrestment. The approach lights are dimmed by the WARN/CAUT lights knob.
2.6.1.6 HOOK BYPASS Switch. The HOOK BYPASS switch, located on the lower left main
instrument panel, is spring loaded to the CARRIER position and is electrically held in the FIELD
position.
FIELD
Approach lights and AOA indexers do not flash regardless of hook position. The
switch reverts to the CARRIER position if the hook is lowered.
CARRIER
Approach lights and AOA indexers flash if the hook is not down
2.6.1.7 Landing/Taxi Light. A landing/taxi light, located on the nose gear strut, is used to light the
flightpath during landing or a taxiway/runway during ground operations.
2.6.1.7.1 LDG/TAXI Light Switch. The LDG/TAXI light switch is located on the lower left main
instrument panel.
ON
Landing/taxi light on with the LDG GEAR handle DN and the nose gear down and
locked
OFF
Landing/taxi light off
2.6.1.8 Day ID Light. A high intensity white strobe light is mounted below the approach light on the
nose gear strut. The strobe light operates with the landing gear down and locked, WoffW, and the
exterior lights master switch in the OFF (aft) position.
2.6.1.8.1 Day ID Test Switch. A day ID test switch, located in the nose wheel on frame 233, is
provided to test the day ID strobe light. The switch operates only with ac power applied to the aircraft.
TEST
The day ID strobe light comes on, ID LT is displayed on the LDDI, and the master
caution light and aural tone come on.
OFF
The switch is spring loaded to the OFF position.
2.6.2 Interior Lighting. Interior lighting is utilized to provide adjustable cockpit lighting for the main
instrument panel and consoles during night or low light operations. All controls for interior lighting are
located on the INTR LT panel on the right console, except for the utility flood light, the AOA indexers,
and the five cockpit displays.
2.6.2.1 MODE Switch. The MODE switch, located on the INTR LT panel, is used to select one of
three cockpit lighting modes; allowing the pilot to optimize interior lighting for current ambient light
conditions.
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ORIGINAL
A1-F18EA-NFM-000
NVG
Reduces the brightness range for the warning, caution, and advisory lights, the UFCD,
MPCD, EFD, (LOT 25 AND UP) AMPD, and (LOT 26 AND UP) 8 x 10 display. This
disables the integral console lights and the white floodlights; and enables six NVG compatible floodlights to illuminate the consoles.
NITE
Reduces the brightness range for the warning, caution, and advisory lights, the UFCD,
MPCD, EFD, (LOT 25 AND UP) AMPD, and (LOT 26 AND UP) 8 x 10 display.
DAY
Provides the maximum brightness range for all interior lighting.
The UFCDs, MPCD, EFD, (LOT 25 AND UP) AMPD, and (LOT 26 AND UP) 8 x 10 display reset
to DAY mode brightness after aircraft shutdown. Following electrical power interruption, these
displays reset to DAY mode brightness with the mode switch in DAY or NITE positions, and to NITE
mode brightness with the mode switch in NVG position.
2.6.2.2 CONSOLES Lighting Knob. The CONSOLES lighting knob, located on the INTR LT panel,
is used to control the brightness of the integral lighting for the left and right consoles, the hydraulic
pressure gauge, and both circuit breaker panels. Clockwise rotation of the knob increases console
lighting intensity from the OFF to BRT positions. The CONSOLES knob and integral console lighting
are disabled in the NVG mode.
2.6.2.3 INST PNL Lighting Knob. The INST PNL lighting knob, located on the INTR LT panel, is
used to control the brightness of the integral lighting for the main instrument panel and the standby
magnetic compass. Clockwise rotation of the knob increases main instrument panel lighting intensity
from the OFF to BRT positions. The strobe function of the SHOOT light is disabled when the
instrument lights are on.
2.6.2.4 FLOOD Lights Knob. The FLOOD knob, located on the INTR LT panel, is used to control the
brightness of the white cockpit floodlights. Eight floodlights are provided for secondary lighting; three
above each console and one on either side of the main instrument panel. Clockwise rotation of the knob
increases floodlight intensity from the OFF to BRT positions. The FLOOD knob and all white
floodlights are disabled in the NVG mode. There is no brightness control for the six NVG floodlights.
2.6.2.5 CHART Light Knob. The CHART light knob, located on the INTR LT panel, is used to
control the brightness of the NVG compatible chart light. The chart light is located on the canopy bow
at the 10:30 position and rotates in two axes. Clockwise rotation of the knob increases chart light
intensity from the OFF to BRT positions.
2.6.2.6 Utility Floodlight. The utility floodlight, normally stowed above the right console, provides a
portable source of secondary lighting. An attached alligator clip allows the light to be fastened at
various locations in the cockpit. The light contains a knob which provides variable lighting intensity
from off to bright and a button which, when pressed, illuminates the light at full intensity. The light
also contains a rotary selector for white or NVG compatible green lighting.
2.6.2.7 Emergency Instrument Lights. The emergency instrument lights, located on the left and
right sides of the main instrument panel, illuminate the EFD and standby flight instruments in the
absence of ac electrical power. The lights come on anytime the PMGs or the battery are powering the
essential bus. There is no separate cockpit control for the emergency instrument lights.
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A1-F18EA-NFM-000
2.6.2.8 Engine Instrument Light. The engine instrument light, located on the left side of the main
instrument panel, provides lighting for the EFD during battery start of the first engine. The light
comes on when the APU switch is placed to ON.
2.6.2.9 WARN/CAUT Lights Knob. The WARN/CAUT knob, located on the INTR LT panel, is used
to control the brightness of the warning, caution, and advisory lights in the reduced brightness range.
Clockwise rotation of the knob increases warning, caution, and advisory light intensity from the OFF
to BRT positions. The brightness is maximum in the DAY mode and in the reduced brightness range
in the NITE and NVG modes.
Following a power interruption in either the DAY or NITE mode, the warning, caution, and advisory
lights default to the maximum brightness range. Following a power interruption in the NVG mode, the
warning, caution, and advisory lights remain in the reduced brightness range.
2.6.2.10 LT TEST Switch. The LT TEST switch, located on the INTR LT panel, is spring loaded to
the OFF position. The switch is used to test important cockpit lighting to verify bulb integrity prior to
flight. The switch requires ac electrical power to operate.
TEST
Powers all operating warning, caution, and advisory lights, the AOA indexer lights, the
integral background lighting on the EFD (BINGO, MODE, and BRT), changes MENU
to ENG on the DDIs, provides a CHECK SEAT caution in the F/A-18E, and annunciates the landing gear warning tone.
OFF
Lights test off
2.6.3 Interior Lighting (F/A-18F). All controls for the interior lights of the rear cockpit are located
on the INTR LT panel on the right console. The controls operate in the same manner as those in the
front cockpit with two exceptions. There is no MODE switch on the rear cockpit INTR LT panel, and
the rear cockpit LT TEST switch does not illuminate the AOA indexer lights or annunciate the landing
gear warning tone.
2.7 HYDRAULIC POWER SUPPLY SYSTEM
The hydraulic power supply system is a dual pressure system (3,000 and 5,000 psi). The aircraft uses
hydraulic power to actuate primary flight control surfaces and to run the following utility hydraulic
functions: landing gear, wheel brakes and anti-skid, hook, launch bar, refueling probe, nosewheel
steering (NWS), gun, and parking brake. Two hydraulic accumulators provide emergency hydraulic
power for critical utility functions.
2.7.1 Hydraulic System. The hydraulic power supply system incorporates two independent hydraulic systems, HYD 1 and HYD 2 (figure 2-16). Each system is divided into two branches providing four
independent hydraulic circuits identified as 1A and 1B for the left system and 2A and 2B for the right
system. HYD 1 circuits are dedicated solely to flight controls. HYD 2A powers both flight controls and
most utility hydraulic functions. HYD 2B powers the flight controls and arresting hook and pressurizes
both the APU and emergency brake accumulators.
All flight control surface actuators are powered by one HYD 1 circuit and one HYD 2 circuit, either
simultaneously or through hydraulic switching valves.
The utility system operates at 3,000 psi only. Two pressure reducers, one on HYD 2A and one on
HYD 2B reduce utility circuit pressure to 3,000 psi when pump output is 5,000 psi.
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A1-F18EA-NFM-000
Figure 2-16. Hydraulic Flow
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A1-F18EA-NFM-000
2.7.1.1 Hydraulic Pumps. Each system is pressurized by a single, dual pressure (3,000 and 5,000 psi),
variable displacement pump mounted on an AMAD. Pump output pressure is commanded by the
FCCs based on aircraft flight condition; with 5,000 psi utilized during high-speed flight when air loads
are high. One pump is capable of powering the entire flight control system in the event of a single
system failure. A hydraulic pressure transducer relays system pressure to a hydraulic pressure gauge
in the cockpit. Hydraulic cautions (HYD 1A, HYD 1B, HYD 2A, HYD 2B) are set when individual
hydraulic pressure switches detect circuit pressure below 1,400 psi.
2.7.1.1.1 Hydraulic Pressure Gauge. The hydraulic pressure gauge is located on the lower right main
instrument panel. The gauge has individual needles for indicating HYD 1 and HYD 2 system pressure.
Tick marks for 3,000 and 5,000 psi are provided. Two white bands indicate the range of acceptable
operation pressure (2,600 to 3,300 psi and 4,500 to 5,400 psi). Since the gauge is ac powered, actual
hydraulic readings are not provided until the first generator is online following engine start. During
shutdown, pressure readings freeze when the last generator drops offline.
2.7.1.2 Hydraulic Reservoirs. Hydraulic fluid is supplied to each system by a separate hydraulic
reservoir. The HYD 2 reservoir is larger than the HYD 1 reservoir in order to accommodate the utility
system.
2.7.1.2.1 Reservoir Level Sensing (RLS) System. Each reservoir incorporates an RLS system,
designed to isolate a leak in either system circuit. When reservoir fluid level drops to approximately
50%, RLS shuts off circuit A (HYD 1A or HYD 2A caution). If fluid level continues to deplete to
approximately 30%, RLS restores circuit A and shuts off circuit B (HYD 1B or HYD 2B caution). If
alternate circuit shutdown fails to isolate the leak, RLS restores circuit B (no cautions) at approximately 15%, providing hydraulic pressure to both systems until fluid depletion (both cautions).
2.7.1.3 Switching Valves. Hydraulic switching valves are utilized to provide backup hydraulic power
to actuators that are not powered simultaneously by both systems. Two hydraulic circuits, a primary
and a backup, provide power to each switching valve.
2.7.1.3.1 Switching Valve Operation. Following a drop in primary circuit pressure (less than 900
(±100 psi)), the switching valve automatically shuts off the primary circuit and tests downstream
pressure to make sure its actuator(s) was not the leakage source which caused the primary circuit loss.
Concurrently, FCC monitoring detects the pressure loss and inhibits FCC actuator failure detection
logic for 10 seconds to allow the switching valve time to function.
If the actuator(s) passes this leak detection test, the switching valve allows the backup circuit to
provide hydraulic power. If the actuator(s) fails the test, the switching valve isolates both circuits to
prevent additional loss of the backup circuit. At the expiration of the 10 second timer, the FCCs no
longer inhibit actuator failure detection logic. With both circuits isolated, this logic Xs LEF actuators
immediately and rudder or aileron actuators only when the actuator is subsequently commanded to
move. It is for this reason that the FLAP switch is cycled during the post-flight switching valve check.
Switching valve operation is completely hydro-mechanical, separate from electrical inputs or FCS
reset commands. As mechanized, there is no hazard associated with multiple reset attempts to regain
an Xd surface following a hydraulic circuit failure.
Preference is given to the primary circuit at all times. The switching valve transfers to the primary
circuit any time primary circuit pressure recovers above 2,000 psi, regardless of the valve position or the
backup circuit pressure level.
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A1-F18EA-NFM-000
NOTE
If the leak detection test failed due to cavitation in the actuator, the
switching valve resets, runs another test, and completes the switch to
the backup circuit. This self-resetting feature may require several
minutes to complete depending on surface loading, hydraulic system
pressure, and system temperature. In this case, an FCS reset restores
an X’d surface if and when the switching valve successfully transfers to
the backup circuit.
2.7.1.4 Hydraulic Isolation Valves. HYD 2 utility functions that are required only during takeoff,
landing, and ground operations are downstream of isolation valves. The forward isolation valve is
closed when the LDG GEAR handle is UP and all three landing gear are up and locked; isolating the
nose gear, NWS, launch bar, wheel brakes, and anti-skid. The aft isolation valve and the arming valve
are open with WonW and are normally closed inflight; isolating arresting hook retraction, parking
brakes, and emergency brakes. The aft isolation valve is manually opened inflight by hook retraction
or by holding the HYD ISOL switch in ORIDE. The arming valve is manually opened inflight by
emergency gear or emergency probe extension.
2.7.2 Hydraulic Accumulators. Two hydraulic accumulators are provided in the HYD 2B circuit; the
auxiliary power unit (APU) accumulator and the brake accumulator.
The APU accumulator provides hydraulic pressure to start the APU. With a HYD 2 failure, pressure
from the APU accumulator can be used to:
a. Emergency extend the landing gear or refueling probe inflight.
b. Provide emergency nosewheel steering on the ground.
c. Aid the brake accumulator with emergency braking.
On the ground with engines shutdown, the brake accumulator provides hydraulic pressure to set the
parking brake. With a HYD 2 failure, pressure from the brake accumulator can be used to provide
emergency braking. A fully charged brake accumulator provides a minimum of ten full brake
applications.
The APU and brake accumulator charges are maintained against normal leakage and temperature
fluctuations by a trickle-charge restrictor connected to HYD 2A. Additionally, both accumulators may
be manually recharged inflight using HYD 2B pressure by placing the HYD ISOL switch to ORIDE.
This procedure recharges the brake accumulator if and only if the arming valve is open (emergency gear
or emergency probe extension previously selected). Both accumulators can be charged on the ground
by a hand pump located in the right main landing gear wheelwell.
2.7.2.1 Brake Accumulator Pressure Gauge. The brake accumulator pressure gauge is located on
the lower left main instrument panel and is redlined to indicate pressure below 2,000 psi. The BRK
ACCUM caution is displayed when brake accumulator pressure drops below 2,000 psi. The caution and
redlined pressure indication are warnings that approximately five full brake applications remain before
the brake accumulator is empty. When ac power is not applied, power to the gauge is controlled by the
BRK PRESS switch.
2.7.2.1.1 BRK PRESS Switch. The BRK PRESS switch, located on the forward left console, is spring
loaded to the aft position.
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A1-F18EA-NFM-000
Forward
(unmarked)
Applies maintenance bus power to the brake accumulator pressure gauge when ac
power is not applied.
Aft
(unmarked)
Brake accumulator pressure gauge unpowered when ac power is not applied.
2.7.2.2 HYD ISOL Switch. The HYD ISOL switch, located on the aft left console, is spring loaded to
the NORM position.
ORIDE
Opens the aft isolation valve in flight allowing HYD 2B pressure to recharge the
brake and/or APU accumulators. Following emergency landing gear extension, the
switch may need to be held for up to 20 seconds to remove the APU ACCUM caution and provide a full charge (up to 40 seconds following an in-flight APU start).
NORM
Allows normal aft isolation valve functioning.
If an APU ACCUM caution appears in flight and is not related to
emergency gear/probe extension or APU start, it may indicate a possible
leak in the isolated HYD 2B system. A BRK ACCUM caution in flight is
not normal and may indicate a possible leak in the isolated HYD 2B
system.
2.7.3 Hydraulic System Related Cautions and Caution Light. The following hydraulic system
related cautions and caution light are described in the Warning/Caution/Advisory Displays in Part V:
D HYD 1A, HYD 1B, HYD 2A, HYD 2B
D HYD 5000
D HYD 1 HOT, HYD 2 HOT
D BRK ACCUM
D APU ACCUM caution and caution light
2.8 UTILITY HYDRAULIC FUNCTIONS
The utility hydraulic functions are powered by HYD 2 and include landing gear extension and
retraction, nosewheel steering, wheel braking and anti-skid, launch bar extension, arresting hook
retraction, and in-flight refueling probe extension and retraction. Operation of the in-flight refueling
probe is described in the Fuel System section.
2.8.1 Landing Gear System. The landing gear is a tricycle design and includes a nose landing gear
with steerable nosewheel and two fixed main landing gear. The nose landing gear retracts forward,
while the main landing gear retract aft and inwards. When the landing gear is extended, all landing gear
doors remain open.
2.8.1.1 Planing Links. Each main landing gear assembly incorporates a planing link, which is
designed to properly align the main wheels after landing gear extension. The joint which connects the
wheel to the main landing gear lever is designed to rotate off-axis, so that the wheel fits properly into
the main landing gear wheelwell. The planing link rotates the main wheel from its stowed orientation,
aligns it with the longitudinal axis of the aircraft, and locks over-center. A planing link proximity
switch is used to verify proper planing link position and thereby proper wheel alignment. A flashing
main landing gear position light is used to provide an indication of a planing link failure.
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A1-F18EA-NFM-000
2.8.1.2 Normal Landing Gear Extension and Retraction. Normal landing gear extension and
retraction is electrically controlled by the LDG GEAR handle and uses hydraulic pressure from HYD
2A. With weight off the nose gear and the launch bar retracted, moving the LDG GEAR handle to the
UP position sends an electrical signal to the landing gear selector valves to initiate normal landing gear
retraction. Likewise, moving the LDG GEAR handle to the DN position sends an electrical signal to the
landing gear selector valves to initiate normal landing gear extension
If the launch bar does not return to the up and locked position after catapult launch or the nose gear
indicates WonW, the nose landing gear cannot be retracted. In either case, placing the LDG GEAR
handle UP will raise the main landing gear and leave the nose landing gear extended.
2.8.1.3 Emergency Landing Gear Extension. Emergency landing gear extension is mechanically
controlled by the LDG GEAR handle (front cockpit) or the EMERG LDG GEAR handle (rear cockpit
Lots 21 thru 25) and uses hydraulic pressure provided by the APU accumulator. Both handles are
mechanically connected to the landing gear emergency selector valves by a series of levers and cables.
Emergency extension is mechanically activated by rotating the LDG GEAR handle 90° clockwise and
pulling to detent (approximately 1.5 inches) or by pulling the EMERG LDG GEAR handle to the
detent.
Emergency landing gear extension opens the hydraulic arming valve and directs APU accumulator
pressure to the emergency selector valves. APU accumulator pressure is used to unlock the doors,
release the landing gear uplocks, and is applied to the drag brace locking actuator and sidebrace
downlock actuator. The nose landing gear extends by freefall aided by airloads and the drag brace
locking actuator. The main landing gear extends by freefall aided by the sidebrace downlock actuator.
Emergency extension can be performed with the LDG GEAR handle either UP or DN (DN is
recommended).
2.8.1.4 LDG GEAR Handle. The wheel-shaped LDG GEAR handle, located on the lower left main
instrument panel in the front cockpit, is used to control landing gear extension and retraction. A
downlock solenoid in the LDG GEAR handle assembly prevents gear retraction with WonW by
preventing movement of the handle from the DN position.
UP
With WoffW and the launch bar retracted, electrically initiates normal
landing gear retraction.
DN
Electrically initiates normal landing gear extension.
Emergency
(Rotate handle 90°
clockwise and pull
to the detent)
Mechanically initiates emergency landing gear extension.
2.8.1.5 DOWNLOCK ORIDE Button. The DOWNLOCK ORIDE button, located on the lower left
main instrument panel outboard of the LDG GEAR handle, is used to override the downlock solenoid.
If the downlock solenoid does not retract with WoffW (LDG GEAR handle cannot be moved from the
DN position), a failure has occurred in the downlock circuitry (Landing Gear Control Unit). If the
landing gear indicate three down and locked, cycling the landing gear handle is not recommended, as
proper landing gear functioning is questionable. However, if dictated by an emergency situation,
pressing and holding the DOWNLOCK ORIDE button will retract the mechanical stop and allow the
LDG GEAR handle to be moved to the UP position.
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ORIGINAL
A1-F18EA-NFM-000
The LDG GEAR handle must be in the full down position for the mechanical stop to properly engage
after landing (WonW).
If the DOWNLOCK ORIDE button is pressed or the mechanical stop is
not fully engaged, the LDG GEAR handle can be raised on the ground,
and the main landing gear will retract. The nose gear will not retract with
weight on the nose gear.
2.8.1.6 EMERG LDG GEAR Handle (F/A-18F). The EMERG LDG GEAR handle is located on the
lower left main instrument panel in the rear cockpit. Pulling the handle to the detent mechanically
initiates emergency landing gear extension.
2.8.1.7 Landing Gear Control Unit (LGCU). The LGCU monitors the position of the landing gear and
launch bar systems, provides cockpit indications of gear/launch bar position, and provides outputs to
various aircraft systems which are dependent on gear position (e.g., FCC A and B, the SMS, and the
SDC). The LGCU does not control landing gear extension and retraction.
The LGCU receives inputs from the LDG GEAR handle, the LAUNCH BAR switch, and the
following proximity switches: launch bar, landing gear uplocks, landing gear downlocks, planing links,
and WonW. The LGCU controls the red and green L BAR warning/advisory lights, the landing gear
position lights, the light in the gear handle, the landing gear warning tone, the downlock solenoid, and
all inputs to the FCCs and the SMS. The LGCU also performs a self-BIT and a functional check of all
proximity switches, providing MSP code input to the SDC.
2.8.1.8 Landing Gear Warning Light and Warning Tone. The landing gear warning light is a red light
located inside the LDG GEAR handle. The landing gear warning tone is a beeping tone heard in the
headset. The landing gear warning light and warning tone serve three purposes: to indicate a mismatch
between LDG GEAR handle position and actual gear position, to warn of a planing link failure, and to
provide a ″wheels warning.″
A steady warning light comes on whenever the landing gear is in transit and remains on until all
three gear are down and locked (LDG GEAR handle DN) or all gear doors are closed and locked (LDG
GEAR handle UP). If the landing gear is unsafe, the landing gear warning light remains on. The
warning tone is inhibited for 15 seconds to allow for normal landing gear extension and retraction. If
the warning light remains on for 15 seconds, the warning tone is annunciated to provide an aural
indication of unsafe landing gear position.
If a left or right planing link failure occurs with the landing gear down and locked (planing link
proximity switch not properly activated), the landing gear warning light will come on immediately
accompanied by the warning tone.
Lastly, when the LDG GEAR handle is UP, a flashing warning light accompanied by the warning
tone will be activated when airspeed is below 175 KCAS, altitude is less than 7,500 feet, and rate of
descent is greater than 250 fpm. This ″wheels warning″ is provided as a cue to check the position of the
landing gear at flight conditions where the LDG GEAR handle should normally be DN. The wheels
warning is also activated if calibrated airspeed and/or barometric altitude data are lost. In this case, the
standby airspeed and/or altitude indicators should be referenced prior to silencing the warning tone.
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A1-F18EA-NFM-000
2.8.1.8.1 WARN TONE SIL Button. The WARN TONE SIL button, located to the left of the LDG
GEAR handle, is used to silence the landing gear warning tone.
2.8.1.8.2 Landing Gear UNSAFE Light (F/A-18F). The landing gear UNSAFE light is a red light
located on the upper left main instrument panel in the rear cockpit. The light indicates a mismatch
between LDG GEAR handle position and actual gear position (e.g., gear in transit). The light does not
illuminate for a planing link failure, wheels warning, or loss of air data.
2.8.1.9 Landing Gear Position Lights. Three green landing gear position lights, located on the lower
left main instrument panel, are labeled NOSE, LEFT, and RIGHT. When the LDG GEAR handle is
DN, steady lights indicate that the corresponding landing gear is down and locked. The LEFT and
RIGHT landing gear position lights are also used to indicate a planing link failure. If the main landing
gear are down and locked but a planing link proximity switch is not properly activated, the
corresponding position light will flash.
A landing gear position of three down and locked is indicated by three steady green position lights
with the landing gear warning light out. Additionally, when illuminated inflight, the approach lights
provide an external indication that the landing gear is down and locked.
If a landing gear position light is out with the LDG GEAR handle DN and the landing gear warning
light out, a LT TEST should be performed to test the integrity of the position light bulb. If the bulb
tests bad, it is safe to assume that the gear is down and locked. During day operations, if all three
position lights appear to be out/dim, make sure the interior lights MODE switch is in the DAY position.
A landing gear position of three up and locked is indicated by the landing gear warning light out with
all three position lights out.
If one or more landing gear indicates unsafe, a visual inspection can only
confirm general position and obvious damage. There is no external
indication of a locked landing gear.
2.8.1.9.1 Landing Gear Position Lights (F/A-18F). Three green landing gear position lights, labeled
NOSE, LEFT, and RIGHT, are located on the upper left main instrument panel in the rear cockpit.
These lights have the same functionality as those in the front cockpit.
2.8.2 Nosewheel Steering System (NWS). The NWS system is used to provide directional control
and shimmy damping during ground operations. The NWS hydraulic power unit, attached to the nose
landing gear strut, is electrically controlled by commands from the FCCs and is hydraulically actuated
by pressure from HYD 2A (primary) or HYD 2B/ APU accumulator (backup). In the event of a HYD
2A failure, a pressure-biased shuttle valve routes HYD 2B pressure (if available), or APU accumulator
pressure to the NWS unit for backup operation. The FCCs accept input from the rudder pedals to
provide NWS commands.
The NWS system has two modes, NWS (low) and NWS HI. In the low mode (NWS cue in the HUD),
full rudder pedal deflection commands approximately 22.5° of nosewheel deflection. In the high mode
(NWS HI cue in the HUD), full rudder pedal deflection commands approximately 75° of nosewheel
deflection. The NWS system (low gain) incorporates a yaw rate feedback input from the FCCs, which
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ORIGINAL
A1-F18EA-NFM-000
is designed to suppress directional PIO tendencies by increasing directional damping during takeoff
and landing roll.
With loss of yaw rate information to the FCCs, directional PIO may occur
during aggressive ground tracking.
If the NWS system fails, the NWS caution is displayed and the NWS or NWS HI cue is removed
from the HUD. When failed, the NWS system reverts to a 360° free-swiveling mode.
2.8.2.1 NWS Engagement/Disengagement. With WonW, manual NWS engagement is provided by
actuation of the NWS/undesignate button. The method required to engage each of the two NWS
modes (low and high) is dependent on wing lock/unlock status.
With the wings spread and locked and NWS disengaged, the first momentary press and release of the
NWS button engages full-time NWS (low). NWS HI is engaged by subsequent press and hold of the
NWS button. With NWS disengaged, press and hold for greater than 1 second also engages NWS HI.
If the NWS button is released, the system reverts to NWS (low).
With the wings unlocked and NWS disengaged, the first momentary press and release of the NWS
button still engages full-time NWS (low). However, subsequent press and release engages full-time
NWS HI, providing hands-free NWS HI capability for operations in the carrier environment. If the
wings are subsequently spread and locked, NWS reverts to the low mode.
During landing, full-time NWS (low) is automatically engaged when the nose landing gear and at
least one main landing gear transition to WonW. If NWS is engaged with both HYD 2A and 2B failures,
the NWS or NWS HI cue will flash in the HUD as an indication that APU accumulator pressure is
depleting.
NWS is manually disengaged by pressing the paddle switch. NWS is automatically disengaged for
catapult launch, when the launch bar is extended. With the launch bar extended, NWS (low) can be
momentarily engaged to position the launch bar by press and hold of the NWS button. Additionally,
NWS is automatically disengaged when the nose landing gear transitions to WoffW during takeoff or
when power is removed from the FCCs.
2.8.2.2 Emergency High Gain NWS. With a FCS CH 2 or FCS CH 4 failure, normal nosewheel
steering is lost. Emergency high gain NWS can be regained by pulling the failed channel circuit
breaker, unlocking the wings, and momentarily pressing the nosewheel steering button.
When emergency high gain NWS mode is entered, NWS indications may
not be displayed on the HUD. As a result, inadvertent nosewheel steering
actuation may injure ground personnel.
2.8.3 Wheel Brake System. The aircraft’s wheel brake system provides normal braking, anti-skid,
emergency braking, a parking brake, and main wheel anti-spin. Normal braking utilizes HYD 2A
pressure and is capable of functioning with a separate anti-skid system. The anti-skid system, when
enabled, provides maximum braking effectiveness on wet runways or during heavy braking by
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A1-F18EA-NFM-000
preventing wheel skid. When selected, emergency braking utilizes HYD 2B pressure, if available, or
brake and APU accumulator pressure to provide backup braking capability following a HYD 2A
failure. The anti-spin function stops main landing gear wheel rotation prior to landing gear retraction.
2.8.3.1 Wheel Brake Assembly. Each main landing gear wheel is fitted with hydraulically actuated
multiple disk brakes. There are two independent sets of brake lines running to each wheel brake
assembly: the normal brake line pressurized by HYD 2A and the emergency brake line pressurized by
HYD 2B or the brake and APU accumulators. See figure 2-17. Only one set of brake lines can be
pressurized at any given time. A shuttle valve on each wheel brake assembly switches from normal to
emergency brake pressure, depending on which is applied.
Each wheel brake assembly has a brake wear indicator pin, located on the inboard side of the wheel.
When the brakes are applied and the indicator pin is flush or below flush with the brake housing, the
brake pads require changing.
Each wheel assembly incorporates a fuse plug which is designed to melt and deflate the tire at
temperatures below those which would result in a catastrophic tire blowout.
2.8.3.2 Wheel Brake Operation. Each main wheel brake is controlled by a separate brake pedal,
integrated into the rudder/brake pedal mechanism. Pilot applied force to the top of each brake pedal
is transmitted by a series of cables and pulleys directly to the brake control hydraulic servovalves,
located in the nose wheelwell. The amount of hydraulic pressure applied to the wheel brakes by the
servovalves is directly proportional to brake pedal force. Dual brake pedal action provides symmetric
braking, while individual brake pedal action provides differential braking. In the F/A-18F (trainer
configuration), a second set of cables are routed to the servovalves from the rear cockpit brake pedals.
The servovalves are controlled by the pilot applying the most brake pedal force.
2.8.3.3 Normal Braking. Normal braking is enabled when HYD 2A is operable and the EMERG
BRK handle(s) are in the stowed position. The emergency brake valve is closed and the emergency
brake lines are unpressurized. During normal braking, HYD 2A pressure is applied through the left and
right servovalves proportional to the amount of pilot applied brake pedal force and is routed to the
main wheel brakes. When the ANTI SKID switch is ON, the anti-skid system modulates pilot applied
brake pressure in order to prevent wheel skid. When the ANTI SKID switch is OFF, the pilot must
regulate brake pedal force to prevent wheel skid.
2.8.3.4 Anti-skid System. The anti-skid system performs 4 basic functions which are designed to
maximize braking effectiveness during landing rollout: touchdown protection, wheel spin-up override,
skid control, and locked wheel protection. The anti-skid system is enabled when the ANTI SKID
switch is ON and the LDG GEAR handle is DN.
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Figure 2-17. Wheel Brake and Anti-skid System
Failure of either wheel speed sensor can lead to an anti−skid failure,
resulting in a complete loss of brakes. Placing the ANTI SKID switch to
the OFF position or pulling the EMERG BRK handle will bypass the
faulty system and restore braking ability. Judicious braking must be
used, as the anti−skid system is not available. Refer to the BRAKE
FAILURE/EMERGENCY BRAKES procedure.
The system contains two wheel speed sensors, an anti-skid control unit, and an anti-skid control
valve. The anti-skid control unit senses wheel speed and operates by electronically limiting the amount
of HYD 2A pressure that is applied to the wheel brakes through the anti-skid control valve and the
normal brake lines. Anti-skid is not available when emergency brakes are selected.
Touchdown protection delays initial brake application on landing by completely dumping brake
pressure until (1) weight is on the right main landing gear and wheel speed is over 50 knots or (2), if
a wet runway delays wheel spin-up, for 3 seconds after landing. This function prevents landing with
locked main wheels (tire blowouts) even if full brake pedal force is applied at touchdown.
Wheel spin-up override is activated at 50 knots wheel speed to allow normal braking if the right
WonW switch fails.
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Skid control is enabled when sensed wheel speed differs from what the anti-skid control unit
determines it should be (e.g., hydroplaning is detected). If the system detects wheel skid, anti-skid
limits the amount of HYD 2A pressure applied to both brakes as required to prevent skidding. If no
skid exists, full pilot-applied brake pressure is routed to the brakes.
If the speed of one wheel drops 40% below the other wheel, locked wheel protection dumps brake
pressure to both wheels until the speed of the slower wheel returns above 40% of the other. Locked
wheel protection is removed below 35 knots, so that full braking performance (including locking a tire)
is available for taxi and turning operations. Below 14 knots, anti-skid is completely disabled. Below 35
knots, judicious braking is required to avoid flat spotting tires.
Anti-skid protection is bypassed when the ANTI SKID switch is OFF. Normally limited by the
anti-skid system, 3,000 psi hydraulic brake pressure is available and regulated only by pilot brake pedal
forces. When using brakes at high speed without anti-skid protection, there is a very small margin
between effective braking and blown tires. Any force greater than approximately 55 to 60 lbs applied
to the pedals (6° to 7° of pedal rotation) will likely result in blown tires with either the ANTI SKID
switch OFF or emergency brakes selected. The use of normal (anti-skid off) brakes at high speed
should be done with extreme caution. If braking without anti-skid is needed at high speeds, initially
apply very light brake pedal pressure and gradually increase as required.
Use of brakes without anti-skid at high speed can result in blown tires
resulting in loss of directional control. If practical, rollout speed should be
as slow as possible before applying brake pedal pressure.
NOTE
Hot brakes and/or melted wheel assembly fuze plugs can be expected
any time maximum effort braking is used at heavy gross weights with
or without anti-skid, e.g., aborted takeoff or heavy weight landing
(above 46,000 lb GW) with high taxi brake usage.
2.8.3.4.1 ANTI SKID Switch. The ANTI SKID switch, located on the lower left main instrument
panel, is used to manually disable the anti-skid system, e.g., for carrier operations or following an
anti-skid failure (ANTISKID caution displayed). The switch is lever-locked in the OFF position.
ON
Anti-skid system enabled for use with normal braking.
OFF
Anti-skid system disabled (SKID advisory displayed when the landing gear is
down).
2.8.3.4.2 Anti-skid BIT and the ANTISKID Caution. The anti-skid control unit performs two types of
BIT: initiated and periodic. IBIT is performed when power is initially applied to the anti-skid system:
(1) when the landing gear is lowered, (2) when the ANTI SKID switch is selected from OFF to ON
inflight, or (3) on the ground with the parking brake set. IBIT performs a complete test of the anti-skid
system 9 seconds after power is applied and runs for 4.5 seconds. With WonW, IBIT is inhibited with
the parking brake released, as wheel motion will cause a false BIT failure and brake pressure would be
dumped if brakes were applied. PBIT only performs a partial anti-skid test and runs whenever power
is applied and IBIT is not running.
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A1-F18EA-NFM-000
If an anti-skid failure is detected by either BIT, the ANTISKID caution will be displayed at BIT
completion. If an anti-skid failure is detected by PBIT, cycling the ANTI SKID switch will command
an IBIT and a more complete test of the system. When IBIT is running, the ANTISKID caution is
inhibited or is removed if previously displayed. If the ANTISKID caution returns after IBIT, the
ANTI SKID switch must be placed to OFF in order to isolate the failure and make sure that normal
braking (without anti-skid) is available.
For instance, assume the right wheel speed sensor has failed and an ANTISKID caution is displayed.
If the ANTI SKID switch is left ON during landing, touchdown protection circuitry will dump and
never restore brake pressure to both wheels. Normal braking will be lost. In this case, placing the ANTI
SKID switch to OFF will restore normal braking (without anti-skid), or pulling the EMERG BRK
handle will enable emergency braking (bypassing anti-skid).
• Do not cycle the ANTI SKID switch in response to an ANTISKID
caution immediately prior to landing. Cycling the ANTI SKID switch
removes the ANTISKID caution for up to 13.5 seconds as the system
performs IBIT even though the anti-skid system may still be failed
and, if the system is not failed, wheel motion at touchdown may cause
a false BIT failure and a dump of normal brake pressure when brakes
are applied.
• If the ANTI SKID switch is not placed to OFF with an ANTISKID
caution displayed, normal braking capability may be lost completely.
2.8.3.5 Emergency Braking. Emergency braking is enabled when either EMERG BRK handle is
pulled to detent. This action opens the emergency brake valve and applies backup hydraulic pressure
to the hydraulic servovalves. If available, backup pressure from HYD 2B is utilized through the aft
isolation and arming valves, which are open with WonW. If HYD 2 is failed completely, backup
pressure from both the brake and APU accumulators is used. Check valves are incorporated to prevent
the loss of accumulator pressure if HYD 2B is failed. With backup pressure applied, the servovalves
isolate HYD 2A pressure, if still available, so that the normal brake lines are unpressurized.
During emergency braking, backup pressure is applied through the left and right servovalves
proportional to the amount of pilot applied brake pedal force and is routed to the main wheel brakes
through the emergency brake lines. These lines bypass the anti-skid control valve, so the pilot must
regulate brake pedal force to prevent wheel skid.
Hydraulic accumulators and the brake accumulator pressure gauge are discussed in the Hydraulic
Power Supply System section.
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Figure 2-18. Emergency/Parking Brake Handle
2.8.3.5.1 EMERG BRK Handle (Front Cockpit). The EMERG BRK handle is combined with the
PARK BRK handle and is located on the lower left main instrument panel in the front cockpit. When
the handle is in the stowed, emergency position (horizontal), the ‘‘EMERG’’ label appears upright. See
figure 2-18. To select emergency brakes, the handle must be pulled to the detent while in the horizontal
position.
Stowed
(unmarked)
Emergency brake valve closed. Normal braking selected.
PULL
(to detent)
Emergency brake valve open. Emergency braking selected.
The position of the EMERG BRK handle is the only indication that
emergency braking is selected: no warning or caution is displayed. The
EMERG BRK handle(s) must be fully stowed in both cockpits to make
sure that normal braking with anti-skid is available.
Due to friction in the EMERG BRK handle mechanism, the handle may
not return to the fully stowed position unless positively pushed.
2.8.3.5.2 EMERG BRK Handle (Rear Cockpit Lots 21 thru 25). The EMERG BRK handle, located
on the lower left main instrument panel in the rear cockpit, is used to select emergency braking from
the rear cockpit. When stowed, the handle is oriented vertically.
Stowed
(unmarked)
Emergency brake valve closed. Normal braking selected.
PULL
Emergency brake valve open (to detent). Emergency braking selected.
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The position of the EMERG BRK handle is the only indication that
emergency braking is selected: no warning or caution is displayed. The
EMERG BRK handle(s) must be fully stowed in both cockpits to make
sure that normal braking with anti-skid is available.
Due to friction in the EMERG BRK handle mechanism, the handle may
not return to the fully stowed position unless positively pushed.
2.8.3.6 Parking Brake System. The parking brake is used to lock the main landing gear wheels when
the aircraft is parked. The parking brake is activated when the PARK BRK handle is rotated and
pulled to the locked position. This action places the emergency brake valve in the parking brake mode.
Backup hydraulic pressure from HYD 2B or the brake and APU accumulators is applied to the wheel
brake hydraulic servovalves and routed to the main wheel brakes through the emergency brake lines.
The PARK BRK caution will come on to alert the pilot that the parking brake is still set when both
throttles are advanced above about 80% N2 rpm (INS on).
2.8.3.6.1 PARK BRK Handle. The PARK BRK handle is combined with the EMERG BRK handle
and is located on the lower left main instrument panel in the front cockpit. From the stowed
(horizontal) position, the handle must be rotated 90° counterclockwise and pulled to the locked
position, in order to activate the parking brake. When the handle is in the vertical position, the
‘‘PARK’’ label appears upright. See figure 2-18. If emergency brakes are selected, the handle must be
returned to the stowed position before the parking brake can be activated. Rotating the handle 45°
counterclockwise releases the lock and allows the handle to return to the stowed (horizontal) position.
Stowed
(unmarked)
Parking brake released. Normal braking selected.
TURN/
PULL
Parking brake set.
Several aircraft systems utilize parking brake activation to enable or disable logic. A set parking
brake is used to enable anti-skid BIT logic, GEN TIE logic, and INS alignment and is used to trigger
the PARK BRAKE caution.
2.8.3.7 Main Wheel Anti-Spin. The anti-spin function stops main landing gear wheel rotation prior
to landing gear retraction. When the LDG GEAR handle is moved to the UP position, main landing
gear retract pressure is supplied to the anti-skid control valve. Normal brake pressure is blocked and
this anti-spin pressure is routed to the wheel brakes through the normal brake lines. Unlock and
retraction of the main landing gear is delayed until anti-spin pressure is applied and main wheel
rotation has stopped.
2.8.4 Launch Bar System. The launch bar is electrically controlled, hydraulically extended, and
mechanically retracted. With weight on the nose gear, placing the LAUNCH BAR switch to EXTEND
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A1-F18EA-NFM-000
energizes the launch bar control valve and routes HYD 2A pressure to unlock, lower, and hold down the
launch bar. The green L BAR advisory light indicates that the launch bar has been extended.
With the launch bar extended, returning the LAUNCH BAR switch to RETRACT deenergizes the
launch bar control valve, isolates HYD 2A pressure, and allows dual retract springs to mechanically
return the launch bar to the up and locked position. A launch bar proximity switch is energized when
the launch bar is fully retracted.
When the launch bar is fully extended it is held against the deck by HYD 2A pressure. Deck load
springs allow vertical movement of the launch bar during taxi over the catapult shuttle. When the
aircraft is placed in tension on the catapult, the launch bar is held captive in the extended position by
the shuttle. Once in tension, the LAUNCH BAR switch should be placed to RETRACT in order to
remove HYD 2A pressure from the launch bar. When the LAUNCH BAR switch is placed to
RETRACT, the green L BAR light should go out.
Failure to place the LAUNCH BAR switch to RETRACT prior to
catapult launch may result in launch bar hydraulic seal failure and
possible loss of HYD 2A.
At the end of the catapult stroke, launch bar/shuttle separation occurs and allows the retract springs
to return the launch bar to the up and locked position. When engaged, the launch bar uplock prevents
the launch bar from dropping to the deck due to g-loads during landing.
If the launch bar does not return to the up and locked position after catapult launch (launch bar
proximity switch not energized), the nose landing gear cannot be retracted. In this case, placing the
LDG GEAR handle UP will raise the main landing gear and leave the nose landing gear extended.
2.8.4.1 LAUNCH BAR Switch. The LAUNCH BAR switch, located on the lower left main instrument
panel in the front cockpit, is used to control the position of the aircraft’s launch bar. The switch is
spring loaded to the RETRACT position and is electrically held in the EXTEND position only if
weight is on the nose gear.
RETRACT
Launch bar control valve deenergized. Launch bar up.
EXTEND
Launch bar control valve energized. Launch bar unlocked and extended by HYD
2A pressure. Green L BAR advisory light on.
2.8.4.2 LB Circuit Breaker. The LB circuit breaker is located on the left-hand circuit breaker panel
above the left console. The LB circuit breaker provides a secondary means to raise the launch bar
following a launch bar malfunction. When pulled, the circuit breaker manually deenergizes the launch
bar control valve, removing HYD 2A pressure and allowing the retract springs to raise the launch bar.
Typically, this action would be required only if (1) the LAUNCH BAR switch failed in the EXTEND
position with weight on the nose gear or (2) the nose gear failed WonW after launch and the pilot failed
to place the LAUNCH BAR switch to RETRACT.
2.8.4.3 L BAR Warning/Advisory Lights. Two L BAR lights, one green and one red, are located on
the left warning, caution, and advisory lights panel. Both lights are controlled by the landing gear
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control unit (LGCU), based on inputs from the LAUNCH BAR switch, the launch bar proximity
switch, and various landing gear proximity switches.
The green L BAR advisory light is used to indicate that the launch bar has been extended. The
LGCU illuminates the green L BAR light when all of the following conditions are met: weight on the
nose gear, the LAUNCH BAR switch in EXTEND, the launch bar not up (launch bar proximity switch
not energized), and the red L BAR warning light not on.
The red L BAR warning light is used to indicate failure of the launch bar to retract after catapult
launch or a failure in the launch bar control system (proximity switch failure). The LGCU illuminates
the red L BAR light when one of the following sets of conditions are met:
1.
2.
3.
4.
Launch bar not up and weight off the left main gear.
Launch bar not up and left main gear not down.
LAUNCH BAR switch in EXTEND and weight off the left main gear.
LAUNCH BAR switch in EXTEND and left main gear not down.
The first set of conditions is the primary L BAR warning, e.g., the launch bar does not retract fully
after catapult launch. The other three sets of conditions provide a backup L BAR warning if one or
more of the proximity switches which control launch bar functioning fail.
2.8.5 Arresting Hook System. The arresting hook is always down-loaded by a nitrogen-charged
accumulator (arresting hook snubber) contained in the arresting hook retract actuator. Arresting hook
extension is therefore accomplished by mechanically releasing the arresting hook uplatch mechanism
(HOOK handle down) and allowing snubber pressure and gravity to extend the hook. The hook should
extend in less than 2 seconds. At touchdown, the arresting hook snubber controls hook bounce and
provides a hold down force for arresting cable engagement.
Arresting hook retraction is accomplished by raising the HOOK handle. This electrically opens the
aft isolation valve and the arresting hook selector valve, routing HYD 2B pressure to the arresting hook
retract actuator. HYD 2B pressure overcomes the snubber down-load pressure and raises the hook. The
arresting hook uplatch mechanism captures and locks the hook in the up position. The hook should
retract in less than 4 seconds. If HYD 2B pressure is lost, the arresting hook cannot be retracted.
2.8.5.1 HOOK Handle. The HOOK handle, located on the lower right main instrument panel in the
front cockpit, is used to control arresting hook extension and retraction.
Up
(unmarked)
Retracts the arresting hook utilizing HYD 2B pressure.
Down
(unmarked)
Unlocks the arresting hook uplatch mechanism and allows arresting hook snubber
pressure and gravity to extend the hook.
2.8.5.2 HOOK Light. The red HOOK light is located on the lower right main instrument panel
directly above the HOOK handle. The HOOK light comes on any time hook position does not agree
with HOOK handle position. The light comes on when the hook leaves the up and locked position and
remains on until the hook is fully extended (hook proximity switch energized). With WonW, the hook
will strike the ground before it reaches full extension, so the HOOK light will remain on.
In Lot 26 AND UP, there is a green HOOK light in the rear cockpit on the left warning, caution, and
advisory panel. This HOOK light illuminates when the hook is down.
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2.9 WING FOLD SYSTEM
The aircraft’s outer wing panels are designed to fold vertically to reduce the amount of deck space
occupied by the aircraft in the carrier environment. Each wing contains an independent wingfold
mechanism, which consists of two electric motors (one to lock/unlock the wings and one to spread/fold
the wings). During normal operation, the wings are spread, locked, unlocked, and folded in unison.
2.9.1 Wingfold Mechanism. Each wingfold mechanism contains a dc electric motor, which locks and
unlocks the wings, and an ac electric drive unit, which spreads and folds the wings. When the wings are
spread and locked, a locking bolt is electrically driven through the wingfold hinge, holding it in place.
When the wings are unlocked, a wing unlock flag (commonly called a beer can) protrudes from the
upper surface of the wing near the leading edge of the wingfold hinge, indicating that the locking bolt
is unstowed. The shaft of each beer can is painted red for easy identification. When the wings are
locked, the top of the beer can should be flush or near flush with the upper surface of the wing, and no
red should be showing.
Additionally, when the wings are folded, the ailerons are mechanically locked in the faired position
by a hook on the inboard aileron hinge, which engages an aileron locking pin. The aileron locking pin
is mechanically extended as the wings fold. The hook and locking pin are designed to prevent the
ailerons from blowing inboard over the TEFs when hydraulic power is not applied. If an aileron locking
pin should break, it is possible for the aileron to blow inward over the TEF. If this condition exists
during engine start, the TEF will retract into the aileron, damaging both surfaces.
If the wings are folded, note the position of the ailerons during the
preflight walk-around. If the aileron locking pins do not restrain the
ailerons in the faired position, make sure the ailerons are moved to a
faired or outboard position prior to engine start to preclude damaging the
ailerons and TEFs.
Each wingfold mechanism also contains a wing safety switch, which electrically prevents wingfold
movement. The safety switch is activated by a ″remove before flight″ pin inserted in the underside of
the wing near the wingfold hinge.
For ground crew operations (such as loading wingtip missiles), each wing can be manually unlocked,
folded, or spread. The beer cans can be manually extended by inserting a screwdriver into the wing
unlock motor (underside, leading edge). Once unlocked, the wings can be folded or spread by inserting
a speed handle into the electric drive unit (underside, trailing edge).
2.9.2 Wingfold Operation. With the wings folded, wing spread and lock is commanded by placing the
WINGFOLD switch to SPREAD. The SPREAD command is sent directly to the electric drive units to
spread the wings (there are no WonW or FCC interlocks). When each wing reaches the completely
spread position, power is removed to that electric drive unit, and that wing is automatically
commanded to lock. The WING UNLK caution will not be removed until both wings are locked (both
beer cans down). Once the wings are spread and locked, the ailerons will droop to the position
scheduled by the FCCs based on FLAP switch position.
The wings can be stopped in an intermediate position by placing the WINGFOLD switch to HOLD.
If the wings are spread, selecting HOLD unlocks the wings without folding, allowing full time NWS HI
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to be engaged for operations in the carrier environment. This function is useful when NWS HI is
desired but wingfold is not, e.g., dearming wingtip missiles after carrier arrestment.
With the wings spread and locked, wingfold is commanded by placing the WINGFOLD switch to
FOLD. In order for the wings to unlock, ground power must be applied or the aircraft must be WonW
(left main). The initial FOLD command electrically unlocks the wings (beer cans extended, WING
UNLK caution displayed) and fairs the ailerons. When the FCCs determine that (1) weight is on
wheels, (2) airspeed is less than 100 KCAS accelerating or 66 KCAS decelerating, (3) the ailerons are
faired, and (4) both wings are unlocked, the FOLD command is sent to the electric drive units to fold
the wings. When each wing reaches the completely folded position, power is removed to that electric
drive unit
2.9.3 WINGFOLD Switch. The WINGFOLD switch, located on the lower right main instrument
panel, is lever-locked in all three positions. The switch has a barrier guard to prevent inadvertent
actuation.
FOLD
(& unlock)
Unlocks the wings (WING UNLK caution displayed), fairs the ailerons, and, when
allowed by the FCCs, folds the wings.
HOLD
(& unlock)
Stops wing movement in an intermediate position. If spread, unlocks the wings.
SPREAD
(& lock)
Spreads and locks the wings. (WING UNLK caution removed when both wings are
locked).
Ensure the WINGFOLD switch is lever-locked in the SPREAD position
during takeoff checks. If the wings are commanded to unlock or fold
during a catapult shot, the wings will unlock, the ailerons will fair, the
wings may fold partially, and the aircraft will settle.
2.9.4 Wingfold Overheat Cutout Protection. The wingfold electric drive units are designed to meet
the following duty cycle requirements: two (2) fold-spread cycles followed by a twelve (12) minute
cooldown period. If wingfold operation exceeds this duty cycle, overheat cutout protection may
shutdown wingfold operation to prevent actuator damage. Once overheat cutout protection has been
activated, normal wingfold operation is not restored until actuator temperature drops within limits;
however, the wings can still be unlocked, folded, or spread manually.
2.10 FCS - FLIGHT CONTROL SYSTEM
The flight control system (FCS) is a fly-by-wire, full authority control augmentation system (CAS).
The FCS provides four basic functions: aircraft stability, aircraft control, departure resistance, and
structural loads management. Since the basic airframe is statically neutral to slightly unstable, a
primary function of the FCS is to maintain aircraft stability at all flight conditions. The FCS also
provides full authority control of the aircraft by implementing the basic flight control laws which
determine aircraft response to pilot inputs. Pilot inputs from the stick and rudder pedals send
electrical commands to two quad-redundant, digital flight control computers (FCC A and FCC B).
There is no mechanical linkage between the stick and rudder pedals and the flight control surfaces.
FCC software determines what commands are sent to the various flight control surfaces to exercise
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pitch, roll, and yaw control of the aircraft. Additionally, the FCS provides departure resistance by
either refusing to accept or by tailoring pilot inputs that would otherwise lead to an aircraft departure.
Lastly, the FCS provides structural loads management by limiting g-available to prevent an aircraft
overstress or by retracting flight control surfaces at airspeeds that would otherwise exceed the
structural limits of the airframe. See figure 2-19 for a functional diagram of the flight control system.
2.10.1 Flight Control Surfaces. The aircraft has 12 primary flight control surfaces including leading
edge flaps (LEFs), trailing edge flaps (TEFs), ailerons, twin rudders, horizontal stabilators, and
spoilers. LEFs, TEFs, ailerons, and stabilators can be moved both symmetrically or differentially for
pitch and roll control. Flight control surface deflection limits are shown in figure 2-20.
Pitch control is accomplished with symmetric stabilators and, in some conditions, with rudder toe-in
or rudder flare. Roll control is accomplished with combinations of ailerons, differential stabilators,
differential LEFs, and differential TEFs dependent on flight condition and CAS operating mode. The
twin rudders deflect symmetrically for directional control. There is no dedicated speedbrake surface.
Instead, a ″speedbrake function″ is provided by partial deflection of several of the primary flight
control surfaces.
Hydraulic power to all flight control surface actuators is supplied by HYD 1 and HYD 2. Stabilator
and TEF actuators are powered simultaneously by one HYD circuit from each system. All other
actuators are powered by a single primary HYD circuit, with backup hydraulic power available through
a hydro-mechanical switching valve. See the Hydraulic System section, specifically the Hydraulic Flow
Diagram, to determine which HYD circuits power each flight control surface actuator.
Surface
Deflection limits *
Aileron
25° TEU to 42° TED
Rudder
40° left or right
Stabilator
24° TEU to 20° TED
LEF
5° LEU to 34° LED
TEF
8° TEU to 40° TED
LEX Spoilers
0° or 60° TEU
* Tolerance ±1°, or ±3° for spoilers.
Figure 2-19. FCS Surface Deflections
2.10.1.1 Spoilers. The spoilers are mounted on top of the fuselage near the aft end of the LEX. The
spoilers are controlled by the FCCs and have two fixed positions: 0° (down) or 60° TEU. The 60° TEU
position is activated by the speedbrake function or when more than 15° TED stabilator is commanded
(forward stick) above 22° AOA to aid in recovery from high AOA.
2.10.2 FCCs - Flight Control Computers. Two flight control computers (FCC A and FCC B) provide
the computations which implement the aircraft’s flight control laws. A four-channel architecture is
used to provide FCS redundancy. Each FCC contains two individual central processing units (CPUs),
which each run one channel of the FCS. CH 1 and CH 2 are resident in FCC A, with CH 3 and CH 4
in FCC B.
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Figure 2-20. Flight Control System Functional Diagram
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Most inputs to the FCCs (rate gyros, accelerometers, air data sensors, stick and rudder pedal
position sensors) are quad-redundant, one input for each channel. Each of the four CPUs runs
independent and parallel flight control computations. Sensor inputs as well as CPU outputs are
continuously monitored by the FCCs for agreement. When there is disagreement, the erroneous signal
is discarded, if possible.
Rate and acceleration data are provided by two independent Attitude and Heading Reference Sets
(AHRS), one for each FCC. Each AHRS has two sets of ring laser rate gyros and two sets of
accelerometers, which provide four independent sources of pitch, roll, and yaw rate information, and
four independent sources of normal and lateral acceleration. The AHRS units have the capability to
provide attitude, heading, and longitudinal acceleration data, but it is not currently utilized. The
physical rate and acceleration sensors in each AHRS channel are not aligned with the aircraft’s pitch,
roll, and yaw axis. This raw sensor data is converted to the aircraft’s pitch, roll, and yaw axis by
microprocessors internal to each AHRS. As a result of this architecture, a single rate gyro failure in one
channel results in all three axis rates being unusable in that same channel (CAS P, R, Y in one channel
Xd out). Similarly, if any of the accelerometers fail, all acceleration data from that AHRS channel is
unusable (N ACC and L ACC in one channel will both be Xd out).
FCC channel outputs are transmitted to the appropriate flight control actuators and to other aircraft
systems such as the MCs. While FCC computations run in all four channels, all flight control actuators
are not commanded in all four channels. The stabilators and TEF actuators do receive command
signals from all four FCC channels. However, each aileron, rudder, spoiler, and LEF actuator only
receives command signals from two FCC channels, one from FCC A and one from FCC B. The
2-channel actuators on the left side of the aircraft receive inputs from CH 1 and CH 4 while the
2-channel actuators on the right side receive inputs from CH 2 and CH 3. This channel distribution can
be seen on the FCS format.
2.10.2.1 FCC Temperature Monitoring. FCC A contains a thermocouple which monitors the
temperature within the computer and provides a signal to FCC CH 1 and CH 2. If an over-temperature
condition is detected, the FCS HOT caution and caution light come on, and the ″Flight computer hot,
Flight computer hot″ voice alert annunciates. Additionally, FCC A indicates OVRHT on the BIT
status line. In this case, placing the AV COOL switch to EMERG provides emergency ram air cooling
to FCC A and the right TR via a dedicated FCS ram air scoop. FCC B also contains a thermocouple,
but does not set the FCS HOT cautions. The only indication of an over-temperature condition in FCC
B is a BIT status indication of OVRHT.
2.10.2.2 AV COOL Switch. The AV COOL switch is located on the lower right main instrument panel
outboard of the caution light panel.
NORM
FCS ram air scoop retracted.
EMERG
Deploys the FCS ram air scoop for emergency ram air cooling of FCC A, the right
TR, and other essential avionics.
Once deployed, the FCS ram air scoop cannot be retracted inflight.
2.10.3 FCS Redundancy and Survivability. Hydraulic redundancy is provided by distributing flight
control actuators among the four HYD circuits. This arrangement minimizes the probability of losing
multiple actuators due to catastrophic damage to any single actuator or its hydraulic lines. Following
a single HYD system failure, the other HYD system is capable of powering the entire FCS. Loss of
HYD 1 or HYD 2 in up and away flight does not affect aircraft control. However, in the takeoff and
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landing configuration, small but controllable roll and/or yaw excursions may be expected as hydraulic
switching valves cycle to their backup circuits.
The primary electrical power source for each FCC channel is a dedicated output from one of two
permanent magnet generators (PMGs). See the Electrical System section for FCC Electrical Redundancy. Should a power interruption occur to any single FCC channel, the FCC power supply
automatically switches to a ″keep alive″ circuit connected directly to the maintenance bus for 7 to 10
seconds. This makes sure that the FCCs have uninterrupted power to maintain full operation during
all predictable electrical bus switching transients.
For survivability, wiring for one channel from each computer is routed through the upper part of the
aircraft with wiring for the other through the lower part of the aircraft. This routing minimizes the
possibility of loss of any one flight control surface due to system failures or battle damage. If a
stabilator actuator fails due to multiple FCS or hydraulic failures, the FCS automatically reconfigures
to maintain 3-axis control and acceptable handling qualities by using the remaining surfaces. There is
no mechanical FCS reversion mode.
2.10.4 CAS Operating Modes. The control augmentation system (CAS) operates in two basic modes:
Powered Approach (PA) and Up-AUTO (UA). Mode selection is controlled by FLAP switch position
and airspeed. With the FLAP switch in HALF or FULL and with airspeed below approximately 240
KCAS, CAS implements flight control laws tailored for the takeoff and landing configuration (PA).
With the FLAP switch in AUTO, CAS implements flight control laws tailored for up and away flight
(UA). If the FLAP switch is left in HALF or FULL, the aircraft automatically transitions from PA to
UA when airspeed increases above approximately 240 KCAS. This is known as ″auto flap retract.″ In
this case, the amber FLAPS light comes on to alert the pilot to check FLAP switch position. The flight
control laws utilized in each mode are tailored to provide maximum maneuverability while maintaining
predictable handling qualities and departure resistance.
2.10.4.1 FLAP Switch. The FLAP switch, located on the lower left main instrument panel, is used to
select the CAS operating mode and to position the TEFs and aileron droop for takeoff and landing.
AUTO
Selects UA operating mode for up and away flight.
HALF
Selects PA operating mode for the takeoff and landing configuration. Sets TEF
deflection and aileron droop to 30° TED (WonW or at approach speed).
FULL
Selects PA operating mode for the takeoff and landing configuration. Sets TEF
deflection and aileron droop to 40° TED (WonW or at approach speed).
2.10.4.2 Flap Position Lights. Three flap position lights, two green and one amber, are located on the
lower left main instrument panel. The green HALF and FULL flap lights are used to indicate FLAP
switch position and are not indications of actual TEF/aileron position. The FCS format should be
referenced to determine actual LEF, TEF, and aileron position.
FLAPS
(amber)
FLAP switch in HALF or FULL and airspeed above 240 KCAS (auto flap retract),
abnormal flap condition (any flap is off or lacks hydraulic pressure), spin detected
by Spin Recovery System, or GAIN switch in ORIDE.
HALF
(green)
FLAP switch in HALF and airspeed below 240 KCAS.
FULL
(green)
FLAP switch in FULL and airspeed below 240 KCAS.
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2.10.5 Control Augmentation System (CAS).
2.10.5.1 Pitch CAS. Pitch CAS (P CAS) utilizes normal acceleration, pitch rate, and AOA feedback,
each scheduled based on aircraft flight conditions, to tailor aircraft response to pilot stick inputs and
to provide stabilator actuator commands. P CAS operates by comparing aircraft response to the pilot’s
longitudinal stick input, driving the stabilator actuators symmetrically until the difference is reduced
to zero.
In UA, with neutral longitudinal stick, comparing pilot input to aircraft response has the effect of
constantly trimming the aircraft to steady-state, hands-off 1g flight, essentially removing the
requirement for manual trim. In maneuvering flight, P CAS modifies aircraft response to stick inputs
creating the effect of changing stick forces to provide pilot cueing. Actual stick forces for a given stick
displacement do not change with flight condition. At high airspeeds, P CAS is a g-command system
requiring 3.5 pounds of stick-force-per-g. At medium airspeeds, P CAS acts as a hybrid pitch rate and
g-command system. Pitch rate feedback is used to increase apparent stick-force-per-g (heavier stick
forces) to cue the pilot that airspeed is decreasing and less g is available. At low airspeed, P CAS is
primarily an AOA command system using AOA feedback above 22° AOA to provide increasing stick
forces with increasing AOA. With large forward stick inputs, P CAS augments nose-down pitch rates
by flaring the rudders and raising the spoilers.
In PA, AOA and pitch rate feedbacks are used to augment inherent airframe pitch damping and
stability. P CAS nulls the difference between the commanded AOA and actual AOA. With neutral
longitudinal stick, P CAS maintains trim AOA. Unlike UA, pitch trim is required in PA to trim the
aircraft on-speed. Rudder toe-in is used to improve longitudinal stability and to aid aircraft rotation
during takeoff or bolter. Rudder toe-in is a function of AOA. At 0° AOA or with WonW, the rudders
are toed-in 40°. Rudder toe-in decreases linearly to 0° of toe at 12° AOA. Additional AOA feedback is
provided above 12° AOA which increases stick forces with increasing AOA to provide stall warning.
Pitch rate feedback helps maintain tight pitch attitude control during turns.
2.10.5.2 Roll CAS. Roll CAS (R CAS) schedules aileron, differential LEF, differential TEF, and
differential stabilator commands in response to lateral stick inputs to achieve the desired roll
characteristics. Roll rate feedback, scheduled based on aircraft flight conditions, is used to augment
inherent airframe roll damping. Differential LEFs and TEFs are only used in UA. The LEFs deflect
differentially up to 5° when below 25,000 feet and above Mach 0.6. Differential TEFs are not used
above 10° AOA or below -5° AOA. At high airspeeds, aileron, differential stabilator and differential
TEF travel are reduced to provide consistent roll rate response and to aid in preventing structural
loads exceedances. At low airspeeds, aileron and differential stabilator travel are reduced with
increasing AOA to minimize adverse yaw. Differential stabilator may also be limited due to pitch
commands which have priority over lateral commands.
With clean wing or A/A missile loadings (no wing tanks), maximum roll rate is limited to
approximately 225°/second. With A/G store or external fuel tank codes set in the armament computer
for any wing station and the pylon rack hooks closed for those stations, maximum roll rate is limited
to approximately 150°/second to avoid exceeding pylon structural load limits. If all stores are shown as
HUNG, roll rate limiting is removed; however, an R-LIM OFF caution appears on the DDI.
R CAS incorporates two features to reduce pitch-roll inertial coupling induced departures. Based on
pitch rate and Mach number, the first feature reduces the roll command when the pilot applies an
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excessive combined lateral/longitudinal stick input. The second feature limits the roll command when
the aircraft is already rolling and longitudinal stick is moved rapidly. This second feature is removed
at low altitude and high speed since available pitch rate does not result in significant pitch-roll inertial
coupling.
2.10.5.3 Yaw CAS. Yaw CAS (Y CAS) uses yaw rate and lateral acceleration feedback to provide
directional axis damping and to augment pilot commands to the twin rudder actuators. A rollingsurface-to-rudder interconnect (RSRI) adjusted by roll-rate-to-rudder crossfeed (scheduled with
AOA), and lateral acceleration feedback are used to minimize sideslip for roll coordination. To provide
departure resistance and enhanced maneuverability at high AOA, directional stability is augmented
utilizing INS pitch and roll attitudes along with the FCS sensors to synthesize sideslip and sideslip rate
feedback to the ailerons and differential stabilators. These lateral surfaces are used in this sense as
directional controllers by taking advantage of the strong yawing moments they produce at high AOA.
Below 13° AOA, rudder pedal deflections provide yaw by symmetric rudder deflection. At 25° AOA
and above, rudder pedal deflections no longer provide yaw control inputs but instead act entirely as a
roll controller (identical to lateral stick input) by commanding aileron and differential stabilator with
the RSRI commanding the required rudder deflection for roll coordination. Rudder pedal inputs are
summed with lateral stick inputs and this combined input is limited to a value equal to a maximum
lateral stick input. Therefore, applying pedal opposite to lateral stick cancels lateral stick inputs
proportional to the pedal input, e.g., full opposite pedal cancels a full lateral stick command resulting
in zero roll rate. Between 13° and 25° AOA, rudder pedal deflection gradually changes from pure yaw
control to pure roll control. This method of control provides enhanced departure resistance at high
AOA.
Some traditional directional control capability is returned at low airspeed and high AOA only when
the pilot applies lateral stick and rudder in the same direction. This feature starts becoming effective
only at airspeeds below approximately 225 KCAS, from 20° to 40° AOA, but is most effective at
approximately 170 KCAS and 34° AOA. Enabling this feature outside of these conditions would
compromise departure resistance. When this feature is enabled, the sum of lateral stick and rudder
pedal command is no longer limited to a value equal to a full lateral stick input. The excess roll
command is fed to the directional axis to command sideslip. For example, adding full rudder pedal with
a full lateral stick input provides a maximum roll and yaw command. Alternatively, adding lateral stick
to an existing full rudder pedal input has the same effect. The resulting aircraft motion is a highly
controllable nose-high to nose-low reversal.
At high airspeeds, symmetric rudder deflection is reduced and the rudders are toed in to avoid
exceeding vertical tail structural limits.
In PA mode, synthesized sideslip rate feedback augments aerodynamic directional damping and
stability.
2.10.5.4 Flap Scheduling. In UA, LEFs, TEFs, and aileron droop are scheduled as a function of AOA
and air data to optimize cruise and turn performance, to improve high AOA characteristics, and to
provide load alleviation (when required). In general, LEFs start to deflect as AOA increases above
approximately 3°, reaching full deflection (34° LED) by about 25° AOA. In general, TEFs start to
deflect above 2 to 3° AOA, are at full scheduled deflection (approximately 10 to 12° TED) from
approximately 6 to 15° AOA, and begin to retract as AOA increases further. In other words, TEFs are
deflected in the heart of the maneuvering envelope to produce more lift and are retracted at high AOA.
In UA, aileron droop is scheduled to 50% of TEF deflection at low AOA and to 0° at high AOA.
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In UA, flap scheduling is altered slightly based on the presence of wing tanks. With wing tanks
installed, TEF deflection is slightly lower at most flight conditions. LEFs and TEFs typically begin to
deflect at slightly slower Mach but follow the same trends as those mentioned above.
In PA, LEFs are scheduled as a function of AOA to maximize lift. TEFs are scheduled as a function
of airspeed for load alleviation but should be at maximum scheduled deflection at approach speed. In
PA, aileron droop is scheduled with TEF deflection. Following field takeoff or catapult launch,
TEF/aileron droop is latched for 10 seconds after the transition to WoffW. This feature is designed to
improve catapult launch characteristics by ensuring the flaps do not retract immediately after launch.
However, if approximately 190 KCAS is exceeded prior to expiration of the 10 second timer, the TEFs
and aileron droop do begin to retract for loads alleviation. LEF, TEF, and aileron droop scheduling are
shown in figure 2-21.
FCS Mode
Configuration
No Wing Tanks
UA
Wing Tanks
Status
LEF Position
TEF Position
AIL Droop
WonW
3° LED
2° TED
WoffW
Scheduled
with
M,AOA,Alt
Scheduled
with M,AOA
WonW
3° LED
4° TED
WoffW
Scheduled
with
M,AOA,Alt
Scheduled
with M,AOA
15° LED
Scheduled
with AOA
15° LED
Scheduled
with AOA
30° TED
30° TED
(on-speed)
40° TED
40° TED
(on-speed)
1° TED
50% of TEF
(<10° AOA),
0°
(>15° AOA)
2° TED
50% of TEF
(<10° AOA),
0°
(>15° AOA)
30° TED
30° TED
(on-speed)
40° TED
40° TED
(on-speed)
WonW
Flaps HALF
PA
WoffW
WonW
Flaps FULL
WoffW
Figure 2-21. Flap Schedules
2.10.6 Speedbrake Function. The aircraft is not fitted with independent speedbrake surfaces. A
″speedbrake function″ is provided to increase drag by partial deflection of several of the aircraft’s
primary flight control surfaces: ailerons, rudders, TEFs, and spoilers. The stabilators are commanded
to counter pitch transients during speedbrake extension and retraction. The full speedbrake function
can only be commanded in UA.
At subsonic speeds in UA, the speedbrake function flares the rudders and symmetrically raises the
ailerons TEU to approximately 95% of the capability of each surface at the given flight conditions.
This makes sure approximately 5% of surface authority is available for yaw and roll control. If needed,
rudder and aileron priority is given to yaw and roll commands. TEFs are also symmetrically lowered
to further increase drag and to counter the loss of lift caused by deflecting the ailerons TEU. The
spoilers are raised to the full up 60° position only when the speedbrake command reaches 75%. At
subsonic speeds, the stabilator is used to offset any pitch transients that occur due to the deflection of
all speedbrake surfaces except the spoiler. Delaying spoiler deflection until 75% allows the pilot to use
partial speedbrakes for speed modulation, while avoiding minor spoiler induced pitch transients.
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At supersonic speeds in UA, speedbrake surface deflections are changed. The rudders are not
deflected above 1.05 M due to vertical tail loads. The ailerons and TEFs are not deflected above 1.1 M
due to a lack of effectiveness. The spoilers are therefore deflected immediately upon speedbrake
actuation, since they are the only effective surface at these conditions. At supersonic speeds, the
stabilator is used to counter spoiler deployment. The speedbrake function is completely disabled above
1.5 IMN.
In UA, the speedbrake function is ramped out above 16° AOA or below -9° AOA to preserve
lateral-directional stability and between -3.0 to -1.5g for airframe loads.
In PA, the speedbrake function is disabled with WoffW. With WonW and the FLAP switch in HALF
or FULL, the speedbrake function only deploys the spoilers. While the spoilers can be deployed during
landing rollout or aborted takeoff, the drag increase is minimal, and rollout distance is not appreciably
decreased. With WonW and the FLAP switch in AUTO, full extension of the speedbrake function
commands 20° of rudder flare, 23° of TEU aileron, 7° of TED TEF, 60° of spoiler, and a 2° TED
stabilator change.
2.10.6.1 Speedbrake Switch. The speedbrake switch, located on the inboard side of the right throttle
grip, is used to enable/disable the speedbrake function, e.g., extend/retract the speedbrake surfaces.
The forward and center positions are detented, while the aft position is spring-loaded back to center
and must be held. In the F/A-18F (trainer configuration), the rear cockpit speedbrake switch has
override priority over the front cockpit switch.
Forward
(unmarked)
Retracts speedbrake surfaces (full retraction in 2 seconds).
Center
(unmarked)
Stops speedbrake surfaces at an intermediate position.
Aft
(unmarked)
Extends speedbrake surfaces (full extension in 2 seconds).
NOTE
• If the speedbrake switch is held or fails in the aft position for more
than 5 minutes, the speedbrake switch is declared failed, and the FCS
caution is set. If the switch is failed or is held in the aft position when
the FCS RESET button is pushed, the speedbrake surfaces are
retracted, Xs are set on the DEGD row of the FCS page, and the
speedbrake function is disabled for the remainder of the flight.
• If the front cockpit switch is held in the aft position during any FCS
RESET attempt, the speedbrake switch is declared failed; the speedbrake surfaces are retracted; and the speedbrake function is disabled
for the remainder of the flight. This allows the speedbrake surfaces to
be retracted before the 5 minute timer expires, if the front cockpit
switch is stuck in the aft position.
• In the F/A-18F (trainer configuration), if the rear cockpit switch fails
in the aft position, the 5 minute timer must expire before the
speedbrake surfaces can be retracted with the FCS RESET button.
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2.10.6.2 SPD BRK Light. The green SPD BRK light is located on the left warning, caution, and
advisory panel on the main instrument panel. The SPD BRK light comes on anytime the speedbrake
surfaces are not fully retracted.
2.10.7 G-Limiter Considerations. In order to understand what protection the aircraft’s g-limiter
provides, pilots must understand the difference between ″design limit-g″ and ″reference load factor (Nz
REF).″ See the Acceleration Limitations chart in the Operating Limitations chapter for a plot of Nz
REF versus gross weight, for g-limiter specifics, and for gross weight related g-restrictions.
2.10.7.1 Design Limit-g. The aircraft was designed to sustain a limit-g of +7.5g or -3.0g (symmetric)
only at or below its fighter design gross weight of 42,097 lb. At higher gross weights, design limit-g is
reduced to keep from exceeding the structural limitations of the airframe. An ″overstress″ is defined as
a g-level that exceeds the design limit-g at the aircraft’s current gross weight. Above 42,097 lb gross
weight, design limit-g is reduced by the aircraft’s relative gross weight (42,097/GW), such that the
positive design limit is +7.5g * (42,097/GW) and the negative design limit is -0.4 * (positive limit-g). At
the aircraft’s maximum gross weight (66,000 lb), design limit-g is only +4.8g or -1.9g.
Due to the increased airframe and pylon loads that accompany high-g rolling maneuvers, the aircraft
also has a design limit-g for abrupt full-stick rolls (FSR). Abrupt FSRs are defined as full lateral stick
in less than 1 second. The positive design FSR limit is +6.0g below 42,097 lb GW and 80% of the
symmetric design limit-g above 42,097 lb. The negative design FSR limit is -1.0g at all gross weights.
At 66,000 lb GW, the positive design FSR limit is only +3.8g.
2.10.7.2 Reference Load Factor (Nz REF). Reference load factor (Nz REF) is the value that the MC
uses to set the g-limiter when outside of the transonic g-bucket (described below). With increasing
gross weight, Nz REF is the same as design limit-g until the gross weight where +5.5g (-2.2g) is
available (57,405 lb GW). Above 57,405 lb GW, Nz REF is held fixed at +5.5g (-2.2g) in order to assure
that the pilot always has those g-levels available even if they would result in an overstress. Since the
g-limiter may not prevent an overstress at gross weights above 57,405 lb, the pilot must be responsible
for preventing an overstress in this gross weight region.
2.10.7.3 G-Limiter. The g-limiter essentially limits the amount of positive and negative g that can be
commanded by the pilot at a particular gross weight in order to prevent an aircraft overstress. Once the
pilot reaches the stick displacement required to attain the Nz REF g-limit, further stick inputs do not
increase g. This is commonly called ″being on the limiter.″ Once the stick is relaxed to the limit
displacement, g-control below Nz REF is regained. The g-limiter functions to maintain both the
positive and negative Nz REF limits.
During abrupt longitudinal stick inputs, g-limiter overshoots are not uncommon. G-limiter overshoots of up to +0.5g or -0.2g are allowed and do not constitute an over-g. An ″over-g″ is defined as a
g-level which exceeds the overshoot thresholds and sets MSP code 811 (positive exceedance) or 925
(negative exceedance). An over-g condition requires a postflight inspection to determine if an
″overstress″ occurred.
For rolling maneuvers commenced above the positive FSR limit, the g-limiter also provides some
protection. In this region, the g-limiter attempts to reduce commanded-g towards the positive FSR
limit to prevent an overstress. However, if the rate of lateral stick input exceeds the capability of the
g-limiter, an actual rolling overstress may result without setting an 811 MSP code (set only if the
symmetric over-g threshold is exceeded). The g-limiter treats rolling maneuvers with less than ¾ inch
lateral stick as symmetric maneuvers.
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A G-LIM 7.5G caution, accompanied by the FLIGHT CONTROLS, FLIGHT CONTROLS voice
alert, is set for any of the following: FUEL XFER, CAUT DEGD, MC2, SMS failure, or an invalid fuel
quantity. A G-LIM 7.5G caution indicates that the positive symmetrical command limit has been set
to +7.5g regardless of gross weight or stores loading. If the G-LIM 7.5G caution is set, the pilot must
limit commanded g-level to prevent an overstress.
Very high g-onset rates are possible with rapid aft stick movement, with
or without g-limiter override. A very high g-onset rate can cause immediate loss of consciousness (G-LOC) without the usual warning symptoms
of tunnel vision, greyout, and blackout. The effects of G-LOC may last 20
seconds or longer after the g level is reduced to near 1.0g.
2.10.7.4 G-Bucket. Due to the aerodynamic phenomenon known as transonic pitch-up, the g-limiter
incorporates a g-bucket designed to prevent an aircraft positive over-g during transonic deceleration.
In the g-bucket, the g-limiter reduces the positive command g-limit below Nz REF (figure 2-22). This
reduction is a maximum of 1.0g above 20,000 feet, and 1.7g below 15,000 feet. For example, if Nz REF
is +7.5g and altitude is ≤15,000 feet, the g-limiter only allows +5.8g to be commanded while in the
g-bucket. The symmetrical command limit is never reduced below +4.5g.
NOTE
• G-bucket reduction reduces maximum commandable-g.
• Magnitude of transonic pitch-up increases as rate of Mach change
increases. High drag loadings with idle power settings generally have
the largest transonic pitch-ups. High drag loadings (e.g., A/G stores)
have a g-bucket that extends into a lower Mach range.
• Largest measured transonic pitch-up was 2.2g for <15,000 feet. This
magnitude pitch-up was seen on both A/A and A/G loadings.
The 0.2g deep mini-g-bucket extension in the Mach 0.85 to Mach 0.94 range was added in the FCC
OFP to protect against over-g in that region.
The Mach range for the deeper part of the g-bucket is dependent on external stores configuration.
The deeper part of the g-bucket is entered at Mach 0.905 accelerating (with at least one wing tank or
A/G wing store) or at Mach 0.941 accelerating with no wing tanks or A/G stores. Regardless of stores
loading, when decelerating, the g-bucket is entered at Mach 1.045 and is exited at Mach 0.83.
Full stick roll (FSR) limits are reduced in the g-bucket to 80% of (Nz REF minus no more than a
1.0g reduction). For example, if Nz REF is +7.5g, Mach is 0.95, and altitude is ≤15,000 feet, the
g-limiter sets the FSR limit to 80% of (Nz REF minus 1.0) even though the bucket depth is 1.7g.
If the pilot wants to have maximum-g available during a turning maneuver (e.g., the merge), or avoid
the deeper part of the g-bucket, IMN should be 0.90 or less (with at least one wing tank or A/G store),
or 0.93 or less (clean or with A/A stores). Note that even though the 0.2g mini-g-bucket might be active,
flight testing has shown the g-level will be at or slightly above Nz REF.
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Figure 2-22. G-Limiter G-Bucket Reductions in Maximum Commandable G-Level
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2.10.7.5 G-Limiter Override. A g-limiter override feature can be enabled to allow a 33% increase in
the command g-limit for emergency use (allows a 10g command at 7.5g Nz REF). G-limiter override is
selected by momentarily pressing the paddle switch when the stick is near the full aft limit. When
g-limiter override is selected, a G-LIM OVRD caution is set along with a 927 MSP code. Override is not
disengaged until the stick is returned to near the neutral position.
2.10.7.6 Roll Rate Limiting. Roll rate limiting is enabled in R CAS when external wing tanks or A/G
stores are mounted on wing pylons (hooks closed). If any A/G store indicates HUNG, a R-LIM OFF
caution is set and roll rate limiting is removed. In this case, higher than normal roll rates are possible
and may exceed the structural limitations of the airframe/pylons if pilot-imposed lateral stick limits
are not applied.
2.10.8 Air Data Function. The air data function is provided by the FCCs and not a separate
computer. The FCCs receive input from pitot-static sensors, total temperature sensors, the angle of
attack probes, the standby altimeter barometric setting, and the mission computers. The FCC air data
function applies appropriate source error corrections to the air data sensor inputs and calculates
accurate true altitude, airspeed, Mach number, AOA, and outside air temperature (OAT). Computed
air data is used internally by the FCC control augmentation system (CAS) and is also supplied to the
MCs for IFF altitude reporting, weapon system calculations, and landing gear wheels warning, to the
FADECs for engine control, and to the ECS controller for ECS scheduling and fuel tank pressurization
and vent.
2.10.8.1 Pitot-Static/Total Temperature Probes. Two combined pitot-static/total temperature
probes are mounted on the left and right forward fuselage. Each probe contains one pitot pressure
source, two static pressure sources, and a total temperature sensor. One of the static pressure sources
from each probe is connected together and pneumatically averaged. This average static source is
provided to the left and right pressure transmitter sets along with the corresponding pitot pressure
source. The left pressure transmitter set provides pitot-static input to FCC CH 1 and 4 with the right
providing input to FCC CH 2 and 3. The FCC air data function corrects sensed pitot and static
pressures for position error to provide accurate true air data for FCC calculations, MC calculations, and
display in the HUD. The pitot pressure source and the second static pressure source from the left probe
are used to drive the standby flight instruments (altitude, airspeed, and VSI). The second static
pressure source from the right probe is unused.
Each pitot-static probe also contains an integral total temperature sensor. Each total temperature
sensor converts sensed temperature to an electrical signal. The output of the left total temperature
sensor is sent to FCC CH 2, with the right to FCC CH 4. The FCCs use total temperature to calculate
OAT. Each pitot-static/total temperature probe is electrically heated to prevent icing.
2.10.8.2 AOA Probes. Two AOA probes are mounted on the left and right forward fuselage. Each
probe mechanically measures local AOA by aligning with the airstream. An integral AOA transmitter
set converts the mechanical input to a two-channel electrical signal which is sent to the FCCs. The left
AOA transmitter set provides input to FCC CH 1 and 4 with the right to FCC CH 2 and 3. The FCC
air data function corrects the sensed local AOA to a true AOA and provides the output to the MC for
display on the HUD. FCS CH 4 supplies the AOA signal which drives the AOA indexer lights and the
approach lights.
It is possible to damage and jam an AOA probe such that it continues to send signals to the FCCs.
FCC software is designed to minimize flying qualities degradation in the event of a stuck/jammed AOA
probe. The FCCs incorporate an AOA estimator which is used to identify the good AOA probe if one
is damaged. If an AOA probe split is transient, the estimator is used to identify the good probe and no
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cautions are set. If the AOA probe split persists, an FCS caution is set, AOA is Xd in all four channels,
and the estimator is used for FCC calculations. See the HUD Symbology Degrades with Air Data
Function Failure paragraph and Part V for more details on AOA failures. Each AOA probe and AOA
probe cover are electrically heated to prevent icing.
2.10.8.3 PITOT ANTI ICE Switch. The PITOT ANTI ICE switch is located on the ECS panel on the
right console. This switch is used to power the electric heaters for the pitot-static/total temperature
probes, the AOA probes, and the AOA probe covers. All heaters are thermostatically controlled to
prevent damage to their corresponding sensors. With WonW, the thermostat set points are reduced to
prevent damage when cooling airflow is not provided.
ON
Pitot and AOA heaters on manually (WonW or WoffW).
AUTO
Pitot and AOA heaters on automatically with WoffW. Heaters off with WonW.
Failure of both AOA probe heaters in icing conditions may cause a sharp
uncommanded nose down attitude, uncontrollable by normal stick forces
or paddle switch actuation.
2.10.9 Flight Controls.
2.10.9.1 Stick. A traditional center mounted control stick is used to provide pitch and roll inputs to
the FCS. Since there is no mechanical linkage between the stick and the FCCs or the flight control
surfaces, stick feel is provided by two feel-spring assemblies and two eddy current dampers. The feel
spring assemblies provide a linear stick force versus stick displacement gradient in each axis. Two
4-channel position sensors, one in each axis, measure stick displacement and send longitudinal and
lateral stick commands to the FCCs proportional to stick displacement. Stick force and displacement
are listed in figure 2-23 for full stick travel. The eddy current dampers provide stick motion damping
in each axis. Additionally, the control stick is mass balanced to minimize longitudinal stick movement
resulting from accelerations normally experienced during catapult launch.
In the F/A-18F (trainer configuration), a control stick is also fitted in the rear cockpit and is
mechanically linked to the one in the front cockpit.
Stick
Pedal
Direction
Displacement
(in)
Force
(lbs)
Forward
Aft
Left/Right
Left/Right
2.5
5.0
3.0
1.0
20
37
13
100
Figure 2-23. Stick and Pedal Travel Limits
2.10.9.2 Rudder Pedals. Two rudder pedals (left and right) are used to provide directional inputs to
the FCS for yaw/roll control inflight or NWS control with WonW. Since there is no mechanical linkage
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between the rudder pedals and the FCCs or the flight control surfaces, rudder pedal feel is provided
by two feel-spring assemblies. The feel spring assemblies provide a linear pedal force versus
displacement gradient. Two 4-channel position sensors, one on each pedal, measure pedal displacement and send directional commands to the FCCs proportional to pedal displacement. Rudder force
and displacement are listed in figure 2-23 for full pedal travel. The rudder pedals are also used to
provide NWS commands and to actuate toe-operated wheel brakes.
In the F/A-18F (trainer configuration), two rudder pedals are also fitted in the rear cockpit but are
not mechanically linked to the rudder pedals in the front cockpit. Pedal inputs from either cockpit are
summed together and transmitted to the FCCs. A half pedal input from the front cockpit and a half
pedal input from the rear cockpit results in a full rudder pedal command to the FCCs. Similarly,
opposing rudder pedal inputs in each cockpit cancel each other.
2.10.9.2.1 RUD PED ADJ Lever. A RUD PED ADJ lever, located on the center pedestal in each
cockpit, is spring loaded to the up and locked position. When the lever is held down, the rudder pedals
are unlocked and can be moved forward and aft in ½ inch increments. Both pedals are spring loaded
to move aft and must be pushed forward to the desired position. Releasing the RUD PED ADJ lever
locks the pedals in the new position.
• Restrain the rudder pedals during adjustment. Unrestrained rudder
pedals may damage the rudder pedal mechanism.
• Ensure the rudder pedals are locked in position after adjustment.
Failure to lock the rudder pedals may result in uncommanded forward
rudder pedal movement inflight.
2.10.9.3 Stick Grip FCS Controls. The FCS controls located on the stick grip include the pitch and
roll trim switch, the NWS button, and the autopilot/NWS disengage switch. See figure 2-24. In the
F/A-18F (trainer configuration), the FCS controls on the rear cockpit stick grip are identical to those
in the front cockpit.
2.10.9.3.1 Pitch and Roll Trim Switch. The pitch and roll trim switch is located on the top right of
the stick grip. Movement of the pitch and roll trim switch electrically biases the FCCs and does not
reposition the stick.
Forward
Trims nose-down.
Aft
Trims nose-up.
Left
Trims left-wing-down.
Right
Trims right-wing-down.
Pitch and roll trim inputs can be made incrementally or held for faster trim rates. In UA, little if any
pitch trim is required due to the automatic trimming function provided by P CAS. In PA, pitch trim
is required to trim for on-speed AOA. Lateral trim is typically only required immediately after takeoff
or following changes in lateral weight asymmetry (fuel and/or stores). Pitch trim is not monitored for
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Figure 2-24. Stick Grip FCS Controls
runaway trim. However, roll trim is monitored for a stuck switch. If the trim switch is held or is stuck
in the left or right position for more than 40 seconds, the FCS caution is set and the roll trim function
of the switch is disabled for the remainder of the flight. Roll trim can be faded to zero by pushing and
holding the FCS RESET button or TO/TRIM button for approximately 4 seconds.
2.10.9.3.2 NWS Button. The undesignate/nosewheel steering button is located on the front of the
stick grip. The NWS button is used to engage NWS modes, as described in the NWS System
paragraphs in the Utility Hydraulic Functions section. The undesignate function of the NWS button
is described in the Weapon Systems Controls section.
2.10.9.3.3 Paddle Switch. The autopilot/NWS disengage switch, commonly called the ″paddle
switch,″ is located on the lower front of the stick grip. The paddle switch is used to disengage NWS with
WonW, to disengage all autopilot modes with WoffW, and to enable g-limiter override with WoffW. To
enable g-limiter override, the paddle switch must be momentarily pressed with the stick near the aft
limit.
2.10.9.4 RUD TRIM Knob. The RUD TRIM knob is located on the FCS panel on the left console in
the front cockpit only. Movement of the RUD TRIM knob electrically biases the FCCs and does not
reposition the rudder pedals. Rudder trim authority is ±10° and ±22.5° of rudder surface deflection in
UA and PA, respectively. PA rudder trim authority is set to allow zero pedal forces during a HALF flap,
single engine approach. Yaw trim is zeroed by mechanically centering the RUD TRIM knob when the
T/O TRIM button is pushed with either WonW or WoffW.
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2.10.9.5 T/O TRIM Button. The T/O trim button is located in the center of the RUD TRIM knob on
the FCS panel on the left console. With WonW, holding the T/O TRIM button pressed drives pitch
trim to 4° TEU stabilator, roll trim to neutral, and yaw trim to neutral by mechanically centering the
RUD TRIM knob. Depending on initial trim position the T/O TRIM button may need to be pressed
for up to 4 seconds. When these takeoff trim settings are reached, the TRIM advisory is displayed on
the LDDI for as long as the T/O TRIM button is held depressed. With WoffW, pressing the T/O TRIM
button for as long as 4 seconds drives roll trim to neutral, centers the RUD TRIM knob, but does not
affect pitch trim.
2.10.10 Yaw Rate Warning Tone. In UA (flaps AUTO), a yaw rate warning tone is provided to alert
the pilot of excessive yaw rate that may lead to an aircraft departure. The yaw rate warning tone is
generated by the FCCs and is initiated at 40°/second yaw rate with a 1 Hz pulse rate. The tone pulse
rate increases linearly as yaw rate approaches 60°/second, where the pulse rate remains constant at 10
Hz. There is no yaw rate warning tone in PA.
2.10.11 AOA Warning Tone. An AOA warning tone is provided to alert the pilot of excessive AOA
that may lead to aircraft settle and/or departure.
With flaps HALF or FULL, the AOA warning tone is triggered at 14° AOA with a 1 Hz pulse rate.
The tone pulse rate increases linearly as AOA approaches 35°, where the pulse rate remains constant
at 10 Hz.
With MC OFP H3E AND UP, with flaps AUTO, the AOA warning tone is triggered when the AOA
limits corresponding to the FLY lateral weight asymmetry are exceeded. A 500 ft-lb buffer is applied
to the lateral weight asymmetry threshold when triggering the tone to account for fuel slosh. If there
is an AOA failure, or more than one fuel quantity is invalid and/or weapon station indicates HUNG on
stations 2 − 10 (FLY value removed), the tone will not be triggered and the AOA TONE caution will
be displayed.
2.10.12 Spin Recovery System. The aircraft incorporates an automatic spin detection and recovery
system. A spin is declared when both of the following conditions are met: (1) airspeed is below
approximately 120 ±15 KCAS and (2) the yaw rate threshold is exceeded. The yaw rate threshold is
exceeded, for example, if a 15 to 20°/second yaw rate persists for approximately 15 seconds or a 50 to
60°/second yaw rate persists for approximately 2 seconds. For cases where the pilot is intentionally
commanding a high AOA roll (e.g., pirouette) the yaw rate threshold persistence is increased from 15
seconds to 25 seconds.
Immediately following a low-speed maneuver less than 77 KCAS (e.g., tail-slide), airspeed limits for
SPIN logic are opened to 180 KCAS or 12 seconds, whichever comes first. During this time, the normal
acceleration feedback gain is removed to avoid excess coupling, and spin mode arrows will be displayed
to aid recovery if yaw rate exceeds the threshold.
Once a spin has been detected, the spin recovery system places the SPIN MODE recovery displays
on both DDIs (figure 2-25), illuminates the amber FLAPS light, and drives the LEFs to 34° LED and
the TEFs to 4° TED. The displayed spin recovery arrow always indicates the proper direction for
anti-spin lateral stick inputs whether the spin is upright or inverted. Anti-spin lateral stick inputs are
aileron-into for upright spins and aileron-opposite for inverted spins. When lateral stick is placed with
the arrow, automatic spin recovery mode (ASRM) is engaged. With ASRM engaged, all CAS feedback
and control surface interconnects are removed, providing full aileron, rudder, and stabilator authority
for spin recovery. If the stick is neutral or is moved in the wrong direction, the SPIN MODE formats
remain displayed, the LEFs and TEFs remain deflected, but the FCS remains in CAS, and ASRM is
not engaged.
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Figure 2-25. SPIN Recovery Display
If ASRM is engaged during spin recovery, the spin arrow is removed and the FCS automatically
reverts to CAS when either of the following conditions are met: (1) airspeed is above approximately 245
KCAS or (2) the yaw rate threshold is no longer exceeded. After recovery, LEFs and TEFs return to
normal scheduling, and the SPIN MODE formats are replaced with the MENU page after 2 seconds.
NOTE
During highly oscillatory spins or spins that transition from upright to
inverted or from inverted to upright, the SPIN MODE displays may
disappear momentarily.
2.10.12.1 Spin Recovery Displays. When the spin recovery system detects an upright left spin or an
inverted right spin, a left spin arrow appears on both DDIs to indicate the proper direction of the
anti-spin lateral stick input.
SPIN MODE
STICK
LEFT
When the spin recovery system detects an upright right spin or an inverted left spin, a right spin
arrow appears on both DDIs to indicate the proper direction of the anti-spin lateral stick input.
SPIN MODE
STICK
RIGHT
When lateral stick is placed in the direction of the arrow, the word ″ENGAGED″ appears below the
words ″SPIN MODE″ on both DDIs to indicate that ASRM has been successfully engaged.
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ORIGINAL
A1-F18EA-NFM-000
SPIN MODE
is replaced by
SPIN MODE
ENGAGED
When the SPIN MODE formats appear on the DDIs, airspeed is always displayed in the upper left
corner, with altitude in the upper right and AOA in the lower center. See figure 2-23.
2.10.12.2 SPIN Switch. The SPIN switch is located on the right side of the main instrument panel.
The switch is guarded to prevent actuation. The SPIN switch was designed to allow for activation of
a manual spin recovery mode (MSRM). However, with all CAS feedback and control surface
interconnects removed, flight in MSRM will result in a departure and, once departed, will prevent
departure and/or spin recovery. SPIN arrow logic and ASRM functionality have been optimized and
thoroughly flight tested to produce accurate spin mode detection and positive spin recovery.
RCVY
Prohibited.
NORM
ASRM available when a spin is detected.
Selection of manual spin recovery mode (SPIN switch in RCVY) seriously degrades controllability, will prevent recovery from any departure
or spin, and is prohibited.
2.10.13 Stabilator Failure Control Law Reconfiguration. Stabilator reconfiguration consists of
additional control laws which augment baseline CAS control laws to compensate for the complete loss
of a single stabilator. Stabilator reconfiguration is automatically enabled following the detection of a
complete stabilator failure (3 or more FCS Xs in a single stabilator or a dual HYD circuit failure - HYD
1B/2A or 1A/2B). If hydraulics are intact, the failed stabilator is driven to 2° TEU and locked.
Following a dual HYD circuit failure, the failed stabilator must be driven to the locked position by
aiding airloads. If unaiding airloads are applied, actuator mechanization prevents the stabilator from
moving further away from the locked position.
The reconfigured control laws are designed to compensate for the loss of the pitch and roll
contribution of the failed stabilator. In the pitch axis, pitch commands to the remaining stabilator are
doubled to produce more pitching moment. In PA, rudder toe-in and rudder flare are also used to aid
the pitching moment capability of the remaining stabilator. In the roll axis, differential stabilator
commands are disabled. A stabilator-to-rolling-surface interconnect is used to compensate for the roll
generated by single stabilator movement. In PA, this interconnect is stabilator to aileron. In UA, it is
stabilator to aileron and differential TEF. At high airspeed in UA, differential LEFs are also used. In
the yaw axis, the baseline differential stabilator portion of the RSRI continues to be used to counter
the yaw generated by single stabilator movement.
2.10.14 GAIN ORIDE. GAIN ORIDE allows the pilot to select a set of fixed CAS gains when an FCS
malfunction prevents normal CAS gain scheduling (e.g., loss of AOA or pitot-static data). With the
GAIN switch in ORIDE, the FCCs use fixed values for speed, altitude, and AOA depending on the
position of the FLAP switch. These fixed gains cause the LEFs, TEFs, and aileron droop to be driven
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to the fixed positions shown in figure 2-26. GAIN ORIDE should generally provide acceptable handling
qualities at flight conditions which approximate the fixed gains. At flight conditions that deviate from
the fixed gains, a slight degradation in handling qualities should be expected. Refer to chapter 11 for
details. The aircraft stalls at a lower than nominal AOA since the LEFs are fixed. Transition to or from
the landing configuration should be performed at 180 KCAS. For best results, maintain on-speed AOA
during the approach and landing.
Flight with GAIN ORIDE selected is prohibited above 10° AOA or above 350 KCAS (flaps AUTO),
200 KCAS (flaps HALF), or 190 KCAS (flaps FULL) to ensure control system stability and to reduce
the potential for departure. When GAIN ORIDE is selected, the amber FLAPS light comes on along
with either the CRUIS advisory (flaps AUTO) or the LAND advisory (flaps HALF or FULL). Alpha
tone is disabled in GAIN ORIDE.
LEF
(°LED)
TEF
(°TED)
AIL Droop
(°TED)
AUTO
5
4
HALF
21
FULL
21
FLAP Switch
Fixed Gains
Mach
KTAS
Feet
°AOA
2
0.80
459
39,000
3.5
30
30
0.23
151
500
8.1
40
40
0.21
139
500
8.1
Figure 2-26. GAIN ORIDE Flap Positions and Gain Schedules
2.10.14.1 GAIN Switch. The GAIN switch, located on the FCS panel on the left console, is used to
select GAIN ORIDE. The switch is guarded in the NORM position to prevent inadvertent actuation.
ORIDE
Selects fixed CAS gains according to FLAP switch position.
NORM
Selects normal CAS gain scheduling.
2.10.15 FCS Failures. The FCS detects failures through three types of BIT: initiated (IBIT), periodic
(PBIT), and maintenance (MBIT). FCS IBIT is performed during Before Taxi Checks to run a
thorough test of the system prior to flight. PBIT is a less thorough test of the FCS and runs
continuously when other BITs are not running. MBIT is typically run by maintenance personnel and
is the most comprehensive test of the FCS.
FCS failures are annunciated by any or all of the following indications: the FCS caution, the FCES
caution light, the ″Flight controls, Flight controls″ voice alert, FCS format Xs, and/or BIT Logic
Inspection (BLIN) codes. FCS format Xs and BLIN codes identify the location and type of failure.
However, not all FCS related components/functions are covered by Xs on the FCS format matrix. For
such components/functions, valid BLIN codes may be the only indication of the location of the failure.
Therefore, until the nature of the failure is determined, BLIN codes that appear without Xs should be
treated with the same level of concern as those that do.
Typically, BLIN codes that have three digits or less are generated by PBIT, e.g., 341. Four digit and
five digit BLIN codes are generated by FCS IBIT, e.g., 4573 and 10165.
2.10.15.1 FCES Caution Light. The Flight Control Electronic Set (FCES) caution light is located on
the lower right caution lights panel. The primary purpose of the FCES caution light is to alert aircrew
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of critical FCS related failures when MC1 is failed. When MC1 is failed, the normal DDI FCS related
cautions are not generated and the FCS format is not available for troubleshooting. When MC1 is
operative, the FCES caution light is merely a secondary indication of an FCS related failure. The
specific FCS cautions which also trigger the FCES caution light are listed in figure 2-27.
2.10.15.2 FCS RESET Button. The FCS RESET button is located on the FCS panel on the left
console. This button is used to perform several FCS related functions. Following detection of FCS
related hardware and/or software failures (e.g., FCS Xs and/or BLIN codes), pressing the FCS RESET
button commands a reset of FCC failure detection circuitry. If the FCS related failure was momentary
and no longer exists, an FCS RESET (a) restores the failed actuator/component, (b) removes all FCS
failure indications (FCS caution, FCES caution light, and Xs; preflight BLIN codes only), and (c)
displays the RSET advisory for 10 seconds to indicate a successful reset. If the failure remains (a) the
failed actuator/component is not restored, (b) the FCS failure indications return, and (c) the RSET
advisory is displayed for 10 seconds to indicate an unsuccessful reset. In other words, the FCS
RESET button does not fix a detected failure; it merely allows components to be
restored and failure indications to be removed, if and only if the failure no longer exists.
Prior to takeoff (cycle to WoffW), a successful FCS RESET automatically clears all BLIN codes.
Inflight or post-flight, however, BLIN codes are not automatically cleared with a successful FCS
RESET in order to preserve this data for maintenance troubleshooting. Inflight and post-flight BLIN
codes can be cleared, if desired, by pushing the FCS RESET button simultaneously with the paddle
switch.
Additionally, the FCS RESET button is used in conjunction with the FCS BIT consent switch to
enter the FCS exerciser mode.
2.10.15.3 FCS Exerciser Mode. The FCS exerciser mode is incorporated to aid hydraulic system
warming during cold weather starts. The exerciser mode allows hydraulic fluid and hydraulic seals to
warm towards normal operating temperatures without making large surface movements. Large surface
movements with a cold hydraulic system can result in hydraulic seal damage, leaks, and loss of fluid.
On the ground, the FCS exerciser mode is initiated by simultaneously holding the FCS BIT consent
switch in the ON position while pressing the FCS RESET button. When initiated, the mode cycles the
stabilators, flaps, ailerons, and rudders through 20% of full travel for 10 cycles in 20 seconds. The
operation can be stopped prior to 20 seconds by pressing the paddle switch.
During cold weather starts, avoid activating any hydraulic actuated system for two minutes after
both engines are online. This allows hydraulic fluid to warm both systems and prevents hydraulic seal
damage and potential hydraulic leaks. If the aircraft has not flown within 4 hours with ambient
temperatures below -18°C (0°F), up to three selections of the FCS exerciser mode may be required in
order to obtain a successful FCS RESET (after the initial 2 minute warmup).
In standard or warm conditions, do not initiate the FCS exerciser mode
multiple times in an attempt to get a successful FCS RESET. In such
conditions, multiple initiations may excessively elevate hydraulic system
temperatures, increasing actuator and hydraulic pump seal wear and
potentially decreasing component life.
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2.10.15.4 FCS BIT Consent Switch. The FCS BIT consent switch is located above the right console
beneath the right canopy sill. The switch is used in conjunction with the FCS BIT option or the FCS
RESET button to initiate FCS IBIT or the FCS exerciser mode, respectively. See the FCS Initiated
BIT (IBIT) section at the end of chapter 2 for details.
ON
When held (for at least 2 seconds) during selection of the FCS option, initiates
FCS IBIT. When held during a press of the FCS RESET button, initiates FCS
exerciser mode.
OFF
FCS IBIT and FCS exerciser mode not selected.
2.10.15.5 FCS Related Cautions. FCS related cautions shown in figure 2-27 are described in the
Warning/Caution/Advisory Displays in Part V.
Associated Cockpit Indications
FCES Light
″Flight Controls,
Flight Controls″
Voice Alert
X
X
Caution
Master Caution
Light
AOA
Air Data
ATC Fail
AUTO PILOT
P CAS
R CAS
Y CAS
CHECK TRIM
CK FLAPS
FC AIR DAT
FCS
FCS HOT
FLAPS OFF
FLAP SCHED
G-LIM 7.5G
G-LIM OVRD
HYD 5000
NWS
R-LIM OFF
RIG
S/W CONFIG
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
Master Caution
Tone
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
Note 1
X
X
X
X
X
Note 2
Note 1
X
X
X
X
X
X
NOTES
1. The FLIGHT COMPUTER HOT, FLIGHT COMPUTER HOT voice alert and FCS HOT light on the
caution lights panel are activated when the FCS HOT caution is set.
2. Also displayed when any aileron, stabilator, or rudder actuator failed off (Xd out and a ″bold X″ over
the surface position on FCS Status Display).
Figure 2-27. FCS Related Cautions and Cockpit Indications
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Figure 2-28. FCS Status Display
2.10.16 FCS Status Display. When an FCS failure has occurred, the FCS status display (figure 2-28)
can be used to determine the location and type of failure. The FCS status display is selected by the FCS
option on the SUPT MENU. For FCS components/functions which are displayed in the matrix, an ″X″
is displayed in the failed channel(s) along with a corresponding BLIN code. For FCS components/
functions which are not displayed in the matrix, BLIN codes are the only indication of failure location.
A. Surface Position: In degrees from streamline for all surfaces. Tolerance is ±1° for all surfaces.
B. Surface Arrow: Direction of surface deflection relative to the surface hinge point except for
stabilators. For stabilators, surface arrow indicates trailing edge position.
C. Column of Xs: An entire FCC channel is failed due to a processor fault or loss of power.
D. Bold X across surface position: Surface failed; FCCs are no longer commanding movement of
that surface in any channel.
E. G-LIMX.XG: ″X.X″ is the current Nz REF value as calculated by the MC. This value is
decremented if in the transonic g-bucket. INVALID is displayed in this position if the interface
between FCC CH 1, FCC CH 3, and the MC is invalid. In this case, all data on the display is invalid.
During FCS IBIT, G-LIM0.0G is displayed in this position.
F. Bold X over G-LIMX.XG: Nz REF data from the MC is invalid or out of range, and Nz REF has
defaulted to +7.5g. This X also appears when gross weight is above 57,405 lb, indicating that Nz REF
has been set to +5.5g even though this may result in an overstress.
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SEE IC # 34
A1-F18EA-NFM-000
G. BLIN Codes. Up to eight FCC BLIN codes are displayed in octal format for each channel. The codes are
displayed in the order of occurrence. If the list exceeds eight, additional codes may be viewed with a memory
inspect of unit 14 or 15 with address 2253.
In-flight Memory Inspect (MI) of FCC (UNIT 14 or 15) addresses (ADDR)
greater than six digits long is prohibited since it may cause all four FCC channels
to shut down which will result in loss of aircraft control.
H. BLIN Code Channel: The FCC channel corresponding to the list of BLIN codes. The channel is
incremented from 1 thru 4 and back to 1 by selecting the BLIN option.
I. L/R XX.X: FCC air data function corrected true AOA for the left and right AOA probes (see AOA Select in
paragraph K below).
J. AOA XX.X: True AOA based on INS data. INS true AOA is displayed for reference only to aid the pilot in
determining which AOA probe is valid when one is damaged. INS true AOA is normally boxed, indicating that
an average of the left and right probes has been selected for display in the HUD and for use by the AOA indexer
lights and approach lights.
K. AOA Select: The AOA option is displayed only when GAIN ORIDE is selected. If one AOA probe is
damaged, the AOA option can be used to select output from the good probe for display in the HUD and for use
by the AOA indexer lights and approach lights. AOA probe selection does not affect the fixed gains used by the
FCCs in GAIN ORIDE.
NOTE
If a single probe is declared invalid (a two channel AOA failure), and that
probe is selected, the AOA indexer lights and the HUD AOA are blanked
immediately.
L. DEGD Xs: An FCC failure has occurred in the Xd channel that is not covered by other matrix Xs. BLIN
codes should be used to determine the degraded FCC channel function.
M. PTS Xs: The static or total pressure data is failed in the Xd channel. If a three channel PTS failure occurs
(three Xs), the FCC control laws use data from the remaining PTS channel. If a total PTS failure occurs (four
Xs), the FCCs use fixed PTS values. If a PTS failure clears, PTS Xs are removed automatically with or without
an FCS RESET attempt.
NOTE
With a four channel PTS failure, HUD airspeed and altitude are blanked.
N. AOA Xs: AOA data failed in the Xd channel. A three or four channel AOA failure sets four Xs (AOA Four
Channel failure). In UA, P CAS uses the AOA estimator for control law scheduling. If a UA failure clears, AOA
Xs are removed automatically with or without an FCS RESET attempt. In PA, P CAS uses a fixed 8.1° AOA
value, and R CAS uses the AOA estimator for control law scheduling. If a PA failure clears, AOA Xs are not
removed until an FCS RESET is attempted.
O. Sensor Xs (CAS P, R, or Y; N ACC, L ACC, STICK, or PEDAL): The corresponding sensor (rate gyros,
normal or lateral accelerometers, stick or pedal position) is failed in the Xd channel. A three or four channel
sensor failure sets four Xs (total sensor failure). However, for a three channel failure, the FCCs average the
remaining channel with the last channel that failed. For N ACC and L ACC only, the FCCs use a single channel
if the signal from the third failed channel exceeds 90% of full range. A single X for CAS P, CAS R, or CAS Y is
not possible. Any single gyro failure in P, R, or Y sets Xs in CAS P, CAS R, and CAS Y for that channel. Similarly,
an AHRS acceleration failure sets Xs in the N ACC and L ACC for that channel. A failed AHRS sets Xs in all
five rows for that channel.
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ORIGINAL W/IC 34
A1-F18EA-NFM-000
P. 1 2 3 4: Column legends for each FCC channel. FCC A contains channels 1 and 2 while FCC B
contains channels 3 and 4.
Q. Actuator Xs: The actuator is no longer commanded by the FCC in the Xd channel due to a
detected failure (for all actuators except spoilers). The actuator is still commanded by the other
operating channel(s).
R. SPOIL Xs: A single X is caused by a difference between commanded and actual position or by a
SOV over-current. If a two channel failure (two Xs) was caused by a difference between commanded
and actual position, the FCCs will continue to command the spoilers in both channels. This condition
is indicated by two Xs, a blanked surface position, and no bold surface position X. The vent surface
position value is always blanked.
S. Blank Surface Position: FCCs and/or MCs unable to report actuator position.
T. ROLL TRIM: With MC OFP H3E AND UP, roll trim is displayed in the right column with
WoffW. Arrows indicate trim direction. The roll trim value is dimensionless. Roll trim effects for the
same trim value are different at different airspeeds. The roll trim value provides a quantitative
measure of how much roll trim has been commanded.
U. PITCH TRIM: With MC OFP H3E AND UP, pitch trim is displayed in the left column with
WoffW. Arrows indicate trim direction. The pitch trim value is degrees AOA with flaps in HALF or
FULL and g-level with flaps in AUTO.
2.11 AFCS - AUTOMATIC FLIGHT CONTROL SYSTEM
The AFCS or autopilot provides three basic functions: pilot relief, coupled steering, and data link
control.
Different pilot relief modes are provided for the pitch and roll axes. Pitch-axis pilot relief modes
include barometric altitude hold (BALT), radar altitude hold (RALT), and flight path attitude hold
(FPAH). Roll-axis pilot relief modes include roll attitude hold (ROLL), ground track hold (GTRK),
ground track select (GSEL), heading hold (HDG), and heading select (HSEL).
Coupled steering modes allow the roll-axis to be coupled to a TACAN station (CPL TCN), to a
waypoint (CPL WYPT), to the azimuth steering line (CPL ASL), or to bank angle (CPL BNK).
Data link control modes include automatic carrier landing (ACL) and vector (VEC).
2.11.1 AFCS Mode Selection. Selection of the various autopilot (A/P) modes is accomplished from
the A/P sublevel of the CNI format on the UFCD. Before any autopilot mode can be selected, bank
angle must be less than 70°, pitch attitude must be less than 45°, and the A/P sublevel must be
displayed on the UFCD. The left column of the A/P sublevel displays the couple (CPL) option and the
pitch-axis pilot relief mode options: BALT, RALT, and FPAH. The right column displays the roll-axis
pilot relief mode options: ROLL, GTRK, and HDG. See figure 2-29.
An autopilot mode is enabled by selecting the corresponding option on the UFCD. Once selected, a
highlighted box appears around the option and the corresponding autopilot advisory appears on the
LDDI. If an option is not available, it is not displayed.
Once in the GTRK mode, subsequent selection of the GTRK option enables the GSEL mode (GSEL
replaces GTRK on the UFCD). Once in the HDG mode, subsequent selection of the HDG option
enables the HSEL mode (HSEL replaces HDG on the UFCD).
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-29. AFCS Controls and Indicators
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ORIGINAL
A1-F18EA-NFM-000
2.11.2 Basic Autopilot. The basic or default autopilot mode is FPAH/HDG. When any autopilot
mode is requested from the UFCD, the AFCS first engages FPAH/HDG and then engages the
requested mode. This makes sure that the AFCS is controlling both the pitch and roll axis whenever
an autopilot mode is engaged.
2.11.3 AFCS Mode Deselection. Autopilot modes can be disengaged by either reselecting (unboxing)
the UFCD option or by actuating the paddle switch. If an autopilot mode is unboxed on the UFCD, the
AFCS reverts to the basic autopilot mode in that axis and the AUTO PILOT caution illuminates. If the
stick is moved longitudinally with BALT or RALT engaged or laterally with CPL engaged, the AFCS
reverts to the basic autopilot mode in that axis as the AUTO PILOT caution illuminates. The basic
autopilot mode, however, cannot be disengaged (unboxed) from the UFCD and must, therefore, be
disengaged with the paddle switch. Paddle switch actuation is the only means to make sure that all
autopilot modes have been completely disengaged.
2.11.4 Pitch-Axis Pilot Relief Modes. Any of the pitch-axis pilot relief modes can be engaged in
conjunction with any coupled steering mode (except ACL) or with any roll-axis pilot relief mode.
However, only one pitch-axis mode can be selected at a time. If one pitch-axis mode is requested while
another is engaged, the AFCS switches to the requested mode.
2.11.4.1 BALT - Barometric Altitude Hold. When BALT is engaged, the baro-inertial altitude at the
time of engagement is captured and maintained. While the mode is engaged, this reference altitude
cannot be changed. Longitudinal stick inputs or trim changes disengage BALT, and the AFCS reverts
to FPAH and the selected roll-axis mode.
2.11.4.2 RALT - Radar Altitude Hold. RALT is not available above 5,000 feet AGL. When RALT is
engaged, the radar altitude at the time of engagement is captured and maintained. While the mode is
engaged, this reference altitude cannot be changed. Longitudinal stick inputs or trim changes
disengage RALT, and the AFCS reverts to FPAH and the selected roll-axis mode.
2.11.4.3 FPAH - Flight Path Angle Hold. When FPAH is engaged, the flight path angle at the time
of engagement is captured and maintained. This reference flight path angle can be changed by
longitudinal stick inputs (stick sensitivity similar to CAS) or by pitch trim changes (2°/sec in flaps
AUTO or 0.5°/second in flaps HALF or FULL). When pilot inputs cease, the flight path attitude at
release is captured and maintained.
2.11.5 Roll-Axis Pilot Relief Modes. Any of the roll-axis pilot relief modes can be engaged in
conjunction with any pitch-axis pilot relief mode. However, only one roll-axis mode can be engaged at
a time. If one roll-axis mode is requested while another is engaged, the AFCS switches to the requested
mode.
2.11.5.1 ROLL - Roll Attitude Hold. When ROLL is engaged, the roll attitude at the time of
engagement is captured and maintained. This reference roll attitude can be changed by lateral stick
inputs or roll trim changes (stick and trim sensitivity similar to CAS). When pilot inputs cease, the roll
attitude at release is captured and maintained.
2.11.5.2 GTRK - Ground Track Hold. When GTRK is engaged, aircraft response depends on the roll
attitude at the time of engagement. If roll attitude is less than ±5°, ground track is captured and
maintained. If roll attitude is greater than or equal to ±5°, roll attitude is captured and maintained.
While in GTRK, the aircraft responds to lateral stick or roll trim inputs (stick and trim sensitivity
similar to CAS). When pilot inputs cease, GTRK holds roll attitude (if greater than or equal to ±5°)
or ground track (if less than ±5°).
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ORIGINAL
A1-F18EA-NFM-000
2.11.5.3 GSEL - Ground Track Select. The desired ground track angle is selected by slewing the
command heading marker with the HDG/TK switch, located to the left of the MPCD. When GSEL is
engaged, the aircraft turns from the existing ground track through the smallest angle to the selected
ground track. While in GSEL, the aircraft responds to lateral stick inputs. However, the selected
ground track angle is not changed by stick inputs, so the aircraft returns to the selected ground track
angle upon stick release.
2.11.5.4 HDG - Heading Hold. When HDG is engaged, aircraft response depends on the roll attitude
at the time of engagement. If roll attitude is less than ±5°, magnetic heading is captured and
maintained. If roll attitude is greater than or equal to ±5°, roll attitude is captured and maintained.
While in HDG, the aircraft responds to lateral stick or roll trim inputs (stick and trim sensitivity
similar to CAS). When pilot inputs cease, HDG holds roll attitude (if greater than or equal to ±5°) or
magnetic heading (if less than ±5°).
2.11.5.5 HSEL - Heading Select. The desired heading is selected by slewing the command heading
marker with the HDG/TK switch, located to the left of the MPCD. When HSEL is engaged, the
aircraft turns from the existing heading through the smallest angle to the selected heading. While in
HSEL, the aircraft responds to lateral stick inputs. However, the selected heading is not changed by
stick inputs, so the aircraft returns to the selected heading upon stick release.
2.11.6 CPL - Coupled Steering Modes. The coupled steering modes couple the aircraft in the
roll-axis only. If the CPL option is selected, the AFCS disengages any currently engaged roll-axis pilot
relief mode. The AFCS has the ability to couple to the following sources: a waypoint, waypoint
courseline, or offset aimpoint (CPL WYPT); to a TACAN station or TACAN courseline (CPL TCN);
or to an auto sequence (CPL SEQ#). Refer to chapter 24 for detailed navigation steering information
on waypoint/OAP, auto sequential, and TACAN steering.
2.11.7 Coupled Data Link Modes. The AFCS can couple to data link commands in one of two modes:
ACL and VEC. With ACL boxed on the HSI format, the CPL P/R option appears on the A/P sublevel
when pitch/roll couple capability is available. Selecting the CPL P/R option couples the aircraft to
pitch and roll commands for a Mode 1 carrier approach. With VEC boxed on the HSI format, selecting
the CPL option couples the aircraft (roll-axis only) to data link steering commands. Refer to chapter
24 and the Tactical Manual for detailed information on the ACL and VEC modes.
2.11.8 AFCS Related Caution and Advisories. The AUTO PILOT caution and the following AFCS
related advisories are described in the Warning/Caution/Advisory Displays in Part V:
• BALT
• CPLD
• FPAH
• GSEL
• GTRK
• HDG
• HSEL
• RALT
• ROLL
2.12 WEAPON SYSTEMS CONTROLS
All of the primary controls for the aircraft’s weapon systems (weapons, sensors, and displays) are
located on the front cockpit throttles and stick, the rear cockpit throttles and stick (trainer configured
F/A-18F), or the rear cockpit hand controllers (missionized F/A-18F). This concept, hands on throttles
and stick (HOTAS), allows the aircrew to manipulate the weapon systems without removing the hands
from the aircraft’s primary flight controls. Additionally, the canopy sill DISP switch(es) and the rear
cockpit grab handle switches provide secondary controls for dispensing expendables from the ALE-47
self-protect system.
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ORIGINAL
A1-F18EA-NFM-000
Detailed descriptions of the functionality of the weapon systems controls are contained in the
Tactical Manual, F/A-18EA-TAC Series.
2.12.1 Stick Grip Switches/Controls (Front Cockpit). The weapon systems controls located on the
front cockpit stick grip are the A/A weapon select switch, the sensor control switch, the gun/missile
trigger, the A/G weapon release button, and the undesignate/NWS button. See figure 2-30.
Figure 2-30. Stick Grip Switches/Controls
2.12.1.1 A/A Weapon Select Switch. The A/A weapon select switch, located on the left side of the
stick grip, is spring loaded to the center/up position. The weapon select switch is used to select the
desired A/A weapon and can be used to enter A/A master mode. When the weapon select switch is
actuated while in NAV or A/G master mode, the A/A master mode is automatically entered, the RDR
ATTK format is automatically displayed on the RDDI, and the appropriate A/A weapon/radar format
is selected. Once in the A/A master mode, actuating the weapon select switch merely changes the
selected A/A weapon/radar format.
Center
Neutral
Forward
Selects AIM-7 and the corresponding radar format.
Down
Selects AIM-9 and the corresponding radar format.
Aft
Selects A/A GUN and the gun acquisition mode (GACG).
Inward
Selects AIM-120 and the corresponding radar format.
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ORIGINAL
A1-F18EA-NFM-000
Selection of an A/A missile initiates launch preparation of the priority missile, if more than one is
carried and brings up the corresponding radar format (e.g., scan volume presets). Subsequent selection
steps to the next available missile in the priority sequence.
2.12.1.2 Sensor Control Switch. The sensor control switch, commonly called the ″castle″ switch, is
located on the top center of the stick grip and is spring loaded to the center/up position. The castle
switch is used to assign throttle designator controller (TDC) priority to a particular cockpit display or,
once air combat maneuvering (ACM) mode functionality is enabled, to select a particular ACM mode.
Center
Neutral
Forward
Assigns TDC priority to the HUD in NAV or A/G master mode. In A/A
master mode, selects the boresight acquisition mode (BST) and enables
ACM mode functionality.
Forward (twice
within 0.5 seconds)
Selects/deselects EMCON.
Depress/release then
forward (within 1
second)
Assigns TDC priority to the UFCD in all master modes. If the top level
CNI, a CNI sublevel, or a data entry format is displayed, selects the last
displayed DDI format.
Left
Assigns the TDC to the LDDI in all master modes. With ACM mode
functionality enabled in A/A master mode, selects the wide acquisition
mode (WACQ).
Right
Assigns the TDC to the RDDI in all master modes. With ACM mode
functionality enabled in A/A master mode, selects the automatic acquisition mode (AACQ).
Aft
Assigns the TDC to the MPCD in all master modes. With ACM mode
functionality enabled in A/A master mode, selects the vertical acquisition
mode (VACQ).
When the TDC is assigned to a display which cannot accept TDC priority, automatic format
initialization occurs, typically selecting the format which is most commonly used on that particular
display (e.g., RDR ATTK on the RDDI). With TDC assignment, certain displays (e.g., RDR ATTK,
FLIR, etc.) perform a specific action (e.g., track, break track, etc.) when the castle switch is
subsequently bumped toward that display.
2.12.1.3 Trigger. The gun/missile trigger, located on the front of the stick grip, has two detented
positions.
First
Detent
Initiates strike camera automatic mode operation (based on the selected A/G
weapon) and if CVRS is running, commands HUD recording.
Second
Detent
In A/A master mode, fires the gun or selected A/A missile. Activates AGI protection. In A/G master mode, fires the gun or the laser, if either is selected. Commands the HUD event marker.
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ORIGINAL
A1-F18EA-NFM-000
2.12.1.4 A/G Weapon Release Button. The A/G weapon release button, commonly called the
″pickle″, is located on the top center of the stick grip to the left of the castle switch. The pickle is used
to command weapon release while in A/G master mode. The pickle also initiates AGI protection, strike
camera automatic mode operation, and, if CVRS is running, commands HUD recording and the HUD
event marker.
2.12.1.5 Undesignate/NWS Button. The undesignate/NWS button is located on the lower front of
the stick grip. In NAV or A/G master mode, the undesignate button undesignates all A/G designated
targets and commands the radar and FLIR to break lock, if either is tracking. If TDC priority is
assigned to the FLIR format, pressing the button twice in 1 second selects/deselects FLIR velocity
vector slave (VVSLV) pointing mode. In A/A master mode, the undesignate button creates a launch
and steering (L&S) target designation. Subsequent actuation steps target designation to the next
priority trackfile or to a second designated trackfile (DT2), if one exists. If ACM mode functionality is
enabled, the undesignate button exits the ACM mode and returns the radar to the previous search
mode.
2.12.2 Stick Grip Switches/Controls (Trainer Configured F/A-18F). In the F/A-18F (trainer
configuration), the front and rear cockpit stick grips are identical. However, the rear cockpit trigger
and A/G weapon release button are not functional. The rear cockpit A/A weapon select switch does not
automatically select A/A master mode. From the rear cockpit, A/A master mode must be entered by
actuation of the A/A master mode light. The front and rear cockpit control sensor switches are
functionally identical, including ACM mode selection. However, the rear cockpit TDC can be assigned
to a sensor different from the front cockpit TDC. The systems controlled by the stick grip
switches/controls respond to the last crewmember action taken from either cockpit.
2.12.3 Throttle Grip Switches/Controls (Front Cockpit). The weapon systems controls located on
the front cockpit throttle grips are the chaff/flare/ALE-50 dispense switch, the cage/uncage button, the
throttle designator controller (TDC), the radar elevation control, and the raid button. See figure 2-31.
Figure 2-31. Throttle Grip Switches/Controls (Front Cockpit)
2.12.3.1 Chaff/Flare/ALE-50 Dispense Switch. The chaff/flare/ALE-50 dispense switch is located
on the top inboard side of the right throttle. The switch is used to dispense expendables from the
ALE-47 and ALE-50 self-protect systems.
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ORIGINAL
A1-F18EA-NFM-000
Forward
Provides semi-automatic consent. Dispenses chaff singles (C/F mode).
Aft
Initiates the selected manual program. Dispenses flare singles (C/F mode).
Up
Dispenses a single ALE-50 decoy and provides transmit enable.
Down
Not functional
2.12.3.2 Cage/Uncage Button. The cage/uncage button is located on the rear inboard side of the
right throttle. In NAV master mode and A/G master mode (AUTO delivery), the cage/uncage button
is used to cage and uncage the velocity vector. Depending on master mode and TDC priority
assignment, the cage/uncage button can be used to (1) cage/uncage weapons, (2) command a radar
STT, (3) toggle between weapon modes, or (4) reset sequenced HARM targets.
(18E OFP) With the velocity vector caged, data such as the flight path/pitch ladder and steering
information is displayed near the center of the HUD despite large yaw rates and/or crosswind angles.
2.12.3.3 TDC - Throttle Designator Controller. The TDC is located on the front right side of the
right throttle. It is used to control the positioning of the acquisition cursor or the slewing of a particular
sensor or weapon (e.g., FLIR or Maverick). The TDC can be pressed for target designation. When held
pressed (action slew) or released (no-action slew), the fore-aft and left-right movement of the TDC
sends X-Y slew commands to the display or to the sensor/weapon to which TDC priority is assigned.
2.12.3.4 Radar Elevation Control. The radar elevation control is located on the front left side of the
right throttle. Momentary actuation changes the elevation of the radar antenna in 1,000 foot
increments at the range where the cursor is positioned on the RDR ATTK format. Press and hold
produces a faster elevation change.
Up
Raises the radar antenna/scan volume.
Center
Neutral
Down
Lowers the radar antenna/scan volume.
2.12.3.5 RAID Button. The RAID button is located on the left side of the left throttle. Depending on
master mode and TDC priority assignment, the RAID button can be used to (1) sequence between
available HARM targets, (2) toggle between narrow and wide field of view (FOV) for the FLIR or
Maverick, or (3) select/deselect the STT RAID or SCAN RAID radar modes.
2.12.4 Throttle Grip Switches/Controls (Trainer Configured Rear Cockpit). The rear cockpit
throttle grips contain the same weapon systems controls as those in the front cockpit, except no
chaff/flare/ALE-50 switch is installed. The systems controlled by the throttle grip switches/controls
respond to the last crewmember action taken from either cockpit.
2.12.5 Hand Controllers (Missionized Rear Cockpit Lots 21 thru 25). Two hand controllers are
installed in the rear cockpit, one on the front inboard section of each console. With the exception of
weapons release, the hand controllers are used to provide the same weapon systems control as stick and
throttle switches/controls do in the front cockpit. The right hand controller is a mirror image of the left
hand controller. Each contains a multi-function switch (MFS), a designator controller (DC) assignment
switch, a designator controller (DC), a radar elevation control, a chaff/flare dispense switch, and an
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ORIGINAL
A1-F18EA-NFM-000
undesignate button. Except for the chaff/flare dispense switches, the functions of the left and right
hand controllers are identical. See figure 2-32.
2.12.5.1 Multi-Function Switch (MFS). The MFS is located on the lower inboard side of each hand
controller. The MFS provides the same functionality as the front cockpit RAID and cage/uncage
buttons.
Forward
Sequences between available HARM targets.
Aft
Cages/uncages an A/G weapon.
Down
Selects RAID or changes FLIR/Maverick FOV.
Forward
Not functional
2.12.5.2 Designator Controller (DC) Assignment Switch. The DC assignment switch is located on
the top inboard side of each hand controller. It is used to assign left and right DC priority to the
displays in the rear cockpit. ACM modes cannot be selected by the rear cockpit DC assignment
switches.
Forward
Commands FLIR track/break lock if opposite DC is assigned to the FLIR format.
Aft
Assigns DC priority to the AUFCD.
Inboard
Assigns DC priority to the AMPCD
Outboard
Assigns DC priority to the outboard ADDI. Following DC assignment, commands
track/break lock on the outboard format.
Each DC is initially assigned to its corresponding ADDI. One but not both DCs can be assigned to
one of the center displays, AUFCD or AMPCD. If one DC is assigned to the AMPCD, the other is
forced back to its corresponding ADDI.
2.12.5.3 Designator Controller (DC). The DC is located on the top center of each hand controller.
The track and slew functions of the rear cockpit DCs are identical to the front cockpit TDC. While only
one hand controller DC can be used to designate at a time, both DCs may be used simultaneously on
their assigned formats.
If the front cockpit TDC and the rear cockpit DC are both assigned to the same format, control is
captured by the first controller actuated. If one controller is active when another is selected, the second
input is ignored. If both front and rear TDC/DCs are pressed simultaneously, the TDC/DC assignment
diamond, located in the upper right corner of the specific format, flashes. If both front and rear
TDC/DCs are being slewed simultaneously, the SLEW cue is displayed and flashed.
2.12.5.4 Radar Elevation Control. The radar elevation control is located on the top outboard side of
each hand controller. The front and rear cockpit radar elevation controls are functionally identical.
Control of the radar antenna is captured by the first elevation control actuated from either cockpit. If
one control is active when another is selected, the second input is ignored.
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-32. Hand Controllers (Sheet 1 of 2)
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-32. Hand Controllers (Sheet 2 of 2)
2.12.5.5 Chaff/Flare Dispense Switch. The chaff/flare dispense switch is located on the outboard
side of each hand controller. Dispense switch functionality differs between the left and right hand
controllers.
Left Hand Controller Forward
Initiates manual program 6. Dispenses chaff and flare singles (C/F mode)
Aft
Initiates manual program 6. Dispenses chaff and flare singles (C/F mode)
Right Hand Controller Forward
Provides semi-automatic consent. Dispenses chaff singles (C/F mode)
Aft
Initiates the selected manual program. Dispenses flare singles (C/F mode)
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ORIGINAL
A1-F18EA-NFM-000
2.12.5.6 Undesignate Button. The undesignate button is located on the lower front of each hand
controller. The front and rear undesignate buttons are functionally identical, except the rear cockpit
buttons do not command NWS.
2.12.6 Hand Controllers (Rear Cockpit LOT 26 AND UP). The functions of the left and right hand
controllers are not identical. The left hand controller contains the following switches: Countermeasures, Growth, Sensor Weapons Control, Designator Control Assignment, A/G Weapon Release, Left
Trigger, ECM, and Cage. The right hand controller contains the following switches: HARM, Display
Scroll/Toggle, FOV Wheel, Designator Control, A/A Weapon Release, A/A Weapon Select, Undesignate, and Right Trigger. See figure 2-32 Sheet 2. Refer to Tactical Manuals for complete Hand
controller Switch description.
2.12.7 ALE-47 DISP Switch. An ALE-47 DISP switch is located on the left canopy sill in both front
and rear cockpits. Either switch can be used to initiate manual program 6 or to dispense chaff and flare
singles (C/F mode).
2.12.8 Grab Handle Chaff/Flare Dispense Switches (Rear Cockpit LOTs 21 thru 25). The grab
handle dispense switches are located on the left and right ends of the center grab handle in the rear
cockpit. These switches provide a secondary means to dispense expendables from the ALE-47
self-protect system.
Forward
Provides semi-automatic consent. Dispenses flare singles (C/F mode)
Aft
Initiates the selected manual program. Dispenses flare singles (C/F mode)
Inboard
Initiates manual program 6. Dispenses chaff and flare singles (C/F mode)
Outboard
Initiates manual program 6. Dispenses chaff and flare singles (C/F mode)
2.12.9 Grab Handle Chaff/Flare Dispense Switches (Rear Cockpit LOT 26 and up). The grab
handle dispense switches are located on the left end of the left grab handle and the right end of the
right grab handle in the rear cockpit. Switch function is the same as described above.
2.13 ECS - ENVIRONMENTAL CONTROL SYSTEM
The environmental control system (ECS) utilizes engine bleed air to provide pressurization, heating
and cooling air to various aircraft systems. Warm air is provided for internal fuel tank pressurization
(LOT 23 and below), external fuel tank pressurization, canopy seal inflation, g-suit operation, radar
waveguide pressurization, windshield anti-ice and rain removal, gun gas purge, RECCE bay heating,
and on-board oxygen generating system (OBOGS) operation. Cold, dry conditioned air is provided for
avionics cooling. Warm and cold air are mixed to provide temperature controlled air for cabin heating,
cooling, and pressurization and windshield defog.
A liquid cooling system (LCS) is used to cool the radar transmitter. A digital ECS controller is used
to schedule ECS output, regulate system temperatures, monitor system health, and detect and isolate
faults. See foldout section for a schematic of the ECS. Items numbers, listed in ( ) next to ECS valves,
are maintenance and MSP code nomenclature.
2.13.1 Bleed Air Shutoff Valves. Engine bleed air is tapped from the final (seventh) stage of the
engine high pressure compressor. A primary bleed air pressure regulator and shutoff valve (Item 1), one
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A1-F18EA-NFM-000
on each engine, is used to regulate bleed air output pressure as well as to control flow application/
shutoff.
Both valves are electrically controlled by the BLEED AIR knob and are pneumatically actuated.
The primary bleed air shutoff valves failsafe to the closed position if either electrical power or air
pressure is lost. If an engine is shut down before placing the BLEED AIR knob to OFF, the
corresponding primary bleed air shutoff valve may not fully close, resulting in residual engine fumes
in the cabin on subsequent start of that engine.
Bleed air from each engine passes through a check valve, which prevents reverse flow, and is mixed
through a Y-junction prior to the secondary bleed air pressure regulator and shutoff valve (Item 2).
This valve is electrically controlled by the OFF and AUG PULL positions of the BLEED AIR knob and
is pneumatically actuated. The secondary bleed air shutoff valve fails to the open (safe) position if
either electrical power or air pressure is lost.
All three valves can be automatically commanded to the closed position by the bleed air leak
detection (BALD) system.
2.13.1.1 BLEED AIR Knob. The BLEED AIR knob, located on the ECS panel on the right console, is
used to select the engine bleed air source for the ECS system.
NORM
Commands both primary bleed air shutoff valves open, selecting bleed air from both
engines.
L OFF
Commands the left primary bleed air shutoff valve closed, selecting bleed air from the
right engine only.
R OFF
Commands the right primary bleed air shutoff valve closed, selecting bleed air from the
left engine only.
OFF
Commands all three bleed air shutoff valves closed, isolating the ECS. Closes the ECS
auxiliary duct doors.
AUG
PULL
Commands the secondary bleed air shutoff valve closed, opens the ECS air isolation
valve, and allows APU compressor air to operate the ECS.
2.13.1.2 L or R BLD OFF Cautions. The L or R BLD OFF cautions indicate that the corresponding
primary bleed air shutoff valve(s) are commanded closed. The cautions are not an indication of actual
valve position. The L and/or R BLD OFF cautions are displayed in the following circumstances:
a.
b.
c.
d.
e.
BLEED AIR knob in L OFF, R OFF, or OFF (L, R, or both cautions).
ENG CRANK switch in L or R (L or R caution, respectively).
BALD system detects a leak in one or both bleed air systems (L, R, or both cautions).
FIRE switch in TEST A or TEST B (both cautions).
Over pressurization in one or both bleed air systems (both cautions).
If a bleed air shutoff valve has been commanded closed other than by ENG CRANK, the BLEED
AIR knob must be cycled to OFF and back to NORM to reopen the valve(s).
2.13.2 Bleed Air Subsystem. Engine bleed air downstream of the secondary bleed air shutoff valve
is routed to the primary heat exchanger and three valves. The primary heat exchanger is used for first
stage cooling of hot engine bleed air. The ECS air isolation valve directs bleed air to the air turbine
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starter control valves (ATSCV) for crossbleed start and accepts air from the APU compressor for
alternate (AUG PULL) ECS operation. On the ground and during slow speed flight, the ejector shutoff
valve is opened to direct bleed air to the ram air ejectors to induce cooling airflow through the primary
and secondary heat exchangers. The warm air temperature control valve is used to mix uncooled engine
bleed air with primary heat exchanger output air to regulate the temperature of air in the warm air
manifold.
2.13.3 Primary Heat Exchanger. The primary heat exchanger is located near the base of the right
vertical tail and is used to reject heat from engine bleed air to ram air from one of two inlets.
During medium to high speed flight, ram air cooling is provided by the main ram inlet which draws
air from the engine intake. During slow speed flight and ground operations, ram air cooling is provided
by an auxiliary ram inlet which draws free stream air from the top of the fuselage. A ram air ejector in
the ram air exhaust duct is used to induce more airflow when required, such as on the the ground,
during slow speed flight, and during windshield anti-ice operation with either throttle above IDLE.
2.13.3.1 ECS Auxiliary Duct Doors. Two ECS auxiliary duct doors are located on the upper surface
of the fuselage forward of the base of the vertical tails and immediately in front of the primary and
secondary heat exchangers. The ECS auxiliary duct doors are electrically actuated and open into the
free stream, closing the main ram air inlets and exposing the auxiliary ram air inlets.
On the ground, the ECS auxiliary duct doors are open with the BLEED AIR knob in any position
except OFF. Inflight, the doors are positioned based on Mach number and throttle setting. The doors
are always open below Mach 0.33 and are always closed above Mach 0.40. Between Mach 0.35 and Mach
0.40, the doors are open if either throttle is near MIL (THA greater than 25°).
The ECS auxiliary duct doors should operate symmetrically. If either ECS auxiliary duct door is not
in the commanded position for greater than 8 seconds, an ECS DR advisory is displayed along with an
ECS BIT indication of DEGD. If the door(s) return to the commanded position, the ECS DR advisory
is removed.
2.13.3.2 Ram Air Exhausts. The ECS ram air exhausts are located on the upper fuselage just aft of
the leading edge of the vertical tails. There is one exhaust for the primary heat exchanger (right side)
and one for the secondary heat exchanger (left side). These exhausts discharge the heated ram air
overboard. The exhausts have been redesigned to a five swept stack configuration to prevent
overheating of the aft fuselage structure.
2.13.4 Warm Air Subsystems. Air leaves the primary heat exchanger at a greatly reduced temperature. Warm air from the primary heat exchanger is routed directly to the following systems: internal
fuel tank pressurization (LOT 23 and below), external fuel tank pressurization, canopy seal, g-suit,
radar waveguide pressurization, and OBOGS.
The ECS controller modulates the warm air temperature control valve (Item 21) to mix air from the
primary heat exchanger with uncooled engine bleed air to regulate the temperature in the warm air
manifold. The warm air manifold supplies air for windshield anti-ice and rain removal, gun gas purge,
RECCE bay heating, ECS turbine anti-ice, and cabin heating.
2.13.5 Air Conditioning System (ACS) Pack. The ACS pack provides cold, dry conditioned air for
cabin, FLIR, and avionics cooling. The primary components of the ACS pack are a compressor, turbine,
condenser, reheater and water extractor.
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The ECS controller modulates airflow through the ECS flow modulator valve (Item 4) to control
ECS flow and the speed of the ECS compressor and turbine. Compressor discharge air is directed to
the secondary heat exchanger for additional cooling. Cooled air from the secondary heat exchanger is
directed to the ECS turbine and is used as another source of OBOGS air inflight.
Expansion through the ECS turbine drives the compressor and greatly reduces the temperature of
the airflow, typically to below freezing. The ECS controller regulates the ECS turbine output
temperature and prevents ECS icing by adding warm air through the anti-ice add heat valve (Item 51).
Output temperature is regulated during low altitude operations where ECS icing would be likely.
The condenser, reheater and water extractor remove water from the cold conditioned air. With the
large heat load of the APG-79 radar, the condenser can run at sub−freezing conditions. During
prolonged operation of the APG-79 in high humidity conditions with warm fuel temperatures, periodic
automatic deicing of the condenser occurs. During de-icing, a short term increase in cabin supply air
temperature may be noticed.
In LOT 26 and up, at altitudes above 40,000 feet the water removal system may be bypassed by
opening the water extractor bypass valve (Item 171) in order to improve ECS operating efficiency. If
this valve fails to close after descending below 37,000 feet, water removal capability is degraded. Water
or ice pellets may be blown into the cabin and erratic system behavior (flow/pressure surges) may
result. An ECS ICING caution may also occur.
2.13.6 Secondary Heat Exchanger. The secondary heat exchanger is located near the base of the left
vertical tail and is used to reject heat from ECS compressor air to ram air. Operation of the secondary
heat exchanger is identical to operation of the primary heat exchanger. However, excess ECS system
moisture is sprayed onto the secondary heat exchanger to increase system cooling. This moisture may
be seen exiting the secondary heat exchanger exhaust duct when the throttles are advanced during
ground operations.
2.13.7 Avionics Cooling Fans. Two avionic cooling fans augment ECS cooling of avionics. The
avionics ground cooling fan is located in the nosewheel well. The aft avionics cooling fan is located in
door 108. These fans normally provide primary avionics cooling on deck, and also provide contingency
avionics cooling in flight. Fan activation is a function of bleed air pressure, which varies with engine N2
rpm (i.e. throttle position). Fan operation is described in the individual ECS mode descriptions.
2.13.7.1 Aft Cooling Fan Shutoff Valve. The aft cooling fan shutoff valve, when closed, secures flow
from the aft avionics cooling fan into the ECS and also prevents ECS backflow to the aft avionics
cooling fan. If this valve fails (indicated by MSP 863) while in the open position then, depending on
system pressure, ECS backflow can overspeed the fan in the wrong direction. This may cause a
structural failure of the fan and result in collateral fragment damage to adjacent systems. ECS
backflow is most likely to occur when ECS MAN mode is selected and ECS flow is commanded by
default to maximum.
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Selection of ECS MAN mode is prohibited. Selecting ECS MAN mode
while the aft cooling fan shutoff valve is open may cause the fan to
overspeed, resulting in a catastrophic fan failure and potential loss of
OBOGS.
2.13.8 ECS Operating Modes. The ECS has three operating modes: AUTO, MAN (manual) (PROHIBITED), and OFF/RAM. Each mode is selected by the corresponding position of the ECS MODE
switch. On the ground only, APU compressor air may be used instead of engine bleed air to run the ECS
and cool the avionics (BLEED AIR knob in AUG PULL).
2.13.8.1 ECS AUTO Mode. ECS AUTO mode is the normal operating mode of the ECS. The ECS
controller modulates ECS output to provide the required airflow to the cabin and the avionics. Cabin
airflow is scheduled as a function of ram air temperature and throttle setting, with the highest flow
delivered at hot and cold temperature extremes. Cabin flow may decrease at IDLE. ECS flow to the
avionics is dependent on WonW status, throttle setting, Lot number, and radar configuration. For all
aircraft configurations, more air is provided to avionics inflight.
With APG−73 radar installed, WonW and both throttles at IDLE, the avionics ground cooling fan
and the aft avionics cooling fan energize. With APG−79 radar installed, WonW and both throttles at
IDLE, only the avionics ground cooling fan energizes. The fan(s) provide the primary source of avionics
cooling on deck. The ECS provides a second source of ground avionics cooling but at a fixed, low flow
rate with the remainder of the flow going to the cabin.
With WonW and at least one throttle advanced to approximately 74% N2 rpm, or with WoffW, the
fan(s) secure. Once secured, the fan(s) will not reenergize until both throttles are retarded below
approximately 70% N2 rpm. With WonW and fans secured, the ECS provides all avionics cooling and
controls to the cabin and avionics airflow schedules. The ECS controller schedules avionics airflow
based on the temperature of the air being delivered (warmer air requires higher flow to maintain
constant cooling).
The ECS controller modulates airflow output to meet scheduled airflow requirements. ECS output
temperature is scheduled by the ECS controller to meet avionics cooling requirements and to prevent
ECS turbine icing. Temperature is regulated by adding warm air as required. The controller divides the
airflow between the cabin and avionics. Inflight, if ECS output is inadequate for demand, cabin airflow
is normally given priority. The avionics typically receive that portion of the ECS output which is not
used by the cabin or FLIR cooling systems.
Cabin temperature is controlled by gradually mixing warm air with cold conditioned air, according
to the position of the CABIN TEMP knob. Cabin temperature is normally selectable in the range of
35 to 135° F between the CABIN TEMP knob positions of COLD and HOT, respectively (60 to 160°
F with the DEFOG handle in the HIGH position).
NOTE
In ECS AUTO mode, cabin temperature may take 1 to 2 minutes to
stabilize following a large movement of the CABIN TEMP knob.
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2.13.8.2 ECS MAN Mode.
Selection of ECS MAN mode is prohibited. Selecting ECS MAN mode
while the aft cooling fan shutoff valve is open may cause the fan to
overspeed, resulting in a catastrophic fan failure and potential loss of
OBOGS.
2.13.8.3 ECS OFF/RAM Mode. ECS OFF/RAM mode is used to terminate normal ECS operation
following a major ECS malfunction. Conditioned ECS air is terminated, the cabin ram air scoop is
deployed, and if inflight the aft avionics cooling fan energizes.
NOTE
If ECS OFF/RAM mode is selected inflight, the AV COOL switch
should be placed in EMERG to deploy the FCS emergency ram air
scoop and maximize the emergency avionics cooling available.
Partial cabin pressurization is provided, but only if the CABIN TEMP knob is above the full COLD
position. Cabin airflow and pressurization is provided by ram air from the cabin ram air scoop. Ram
air supply varies with inflight dynamic conditions (airspeed and altitude), and cabin airflow and
pressurization vary similarly.
In ECS OFF/RAM mode, avoid operations above 25,000 feet MSL in
order to prevent decompression sickness (DCS). Normal cabin pressurization is not provided. Partial cabin pressurization is available under
certain circumstances, but may be lost insidiously without warning.
Inflight, cabin temperature is no longer automatically controlled. The CABIN TEMP knob
manually controls addition of heated air according to the CABIN TEMP knob. Temperature control
in between the full COLD and full HOT positions is difficult and imprecise, and ECS response to
commanded temperature changes is slow and nonlinear.
NOTE
In ECS OFF/RAM mode, changes in the position of the CABIN
TEMP knob should be held for 30 seconds due to a slower temperature
response.
With the CABIN TEMP knob not in full COLD, cabin temperature varies with throttle position and
flight condition. System flow/temperature cycling can be expected with the CABIN TEMP knob in full
HOT since overtemperature protection operates intermittently.
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With WonW and both throttles at IDLE, the avionics ground cooling fan and the aft avionics cooling
fan provide the only source of avionics cooling. Above approximately 74% N2 rpm, both fans secure,
and avionics equipment is deprived of all cooling (AV AIR HOT caution).
NOTE
During ground operations in the ECS OFF/RAM mode, the throttles
should be kept below 70% N2 rpm whenever possible in order to
preserve avionics cooling. If either throttle must be advanced above
approximately 74% N2 rpm more than momentarily or if an AV AIR
HOT caution is present, placing the BLEED AIR knob to OFF
reenergizes both fans and provides avionics cooling.
Inflight, the aft avionics cooling fan and the FCS emergency ram air scoop (with AV COOL in
EMERG) provide emergency avionics cooling.
NOTE
With the FCS emergency ram air scoop extended, avionics cooling
inflight is maximized by maintaining altitude below 25,000 feet MSL
and airspeed between 200 to 300 KCAS.
2.13.8.4 AUG PULL. With both generators online and the APU running, selecting AUG PULL (up
on the BLEED AIR knob):
a.
b.
c.
d.
e.
overrides APU automatic shutdown
closes the secondary bleed air shutoff valve.
opens the ECS air isolation valve.
shuts down the aft avionics cooling fans.
and directs APU air to the ECS for cabin and avionics cooling.
2.13.8.5 ECS MODE Switch. The ECS MODE switch, located on the ECS panel on the right console,
is used to select the ECS operating mode.
AUTO
Selects ECS AUTO mode, the normal ECS operating mode. Provides automatic
cabin and avionics airflow, automatic temperature scheduling, and full cabin pressurization.
MAN
PROHIBITED
Selection of ECS MAN mode is prohibited. Selecting ECS MAN mode
while the aft cooling fan shutoff valve is open may cause the fan to
overspeed, resulting in a catastrophic fan failure and potential loss of
OBOGS.
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OFF/RAM
Selects ECS OFF/RAM mode, which terminates conditioned ECS airflow. Provides
ram air for cabin airflow and pressurization depending on flight conditions, and
manual temperature control.
2.13.8.6 CABIN TEMP knob. The CABIN TEMP knob, located on the ECS panel on the right
console, is used to control the temperature of air delivered to the cabin. In ECS AUTO mode, clockwise
rotation of the CABIN TEMP knob linearly increases cabin temperature. In ECS OFF/RAM mode, the
CABIN TEMP knob controls the cabin add heat valve directly, producing a nonlinear temperature
response.
NOTE
• In ECS AUTO mode, cabin temperature may take 1 to 2 minutes to
stabilize following a large movement of the CABIN TEMP knob.
•
In ECS OFF/RAM mode, changes in the position of the CABIN
TEMP knob should be held for 30 seconds due to a slower temperature response.
2.13.8.7 Cabin Louvers/Foot Air Outlet. Three sets of louvers are provided to direct airflow within
the cabin, one on either side of the main instrument panel and one at the base of the center console
behind the control stick. The left and right louvers have controls for elevation and azimuth. The center
louver has a single control for elevation and can be closed completely by moving the lever to the full
aft position. When the center louver is closed, airflow through the left and right louvers is increased.
Maximum aircrew cooling is provided by pulling the DEFOG HANDLE full aft, closing the center
louver, and directing the side louvers towards the body. A fixed foot air outlet directs airflow to the
base of the cabin.
2.13.8.8 Windshield Defog Outlets Fixed windshield defog outlets direct cabin air onto the inner
windshield to remove and inhibit fog. Windshield fogging can occur during rapid environmental
changes, such as high rates of descent and high humidity.
2.13.8.9 DEFOG Handle. The DEFOG handle, located on the right console outboard of the ECS
panel, controls the division of airflow between the windshield defog outlets, the three cockpit louvers,
and the foot air outlet. The DEFOG handle mechanically controls the position of the cabin air/defog
diverter valve.
HIGH
Directs all cabin airflow to the windshield defog outlets to maximize defog, and
increases the temperature range controlled by the CABIN TEMP knob in ECS AUTO
mode.
NORM
Equally divides cabin airflow between the cabin louvers/foot air outlet and the windshield defog outlets. Provides adequate windshield defog for most conditions.
LOW
Directs all cabin airflow to the cabin louvers and the foot air outlet to maximize cabin
cooling.
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2.13.9 Cabin Pressurization. The cabin is pressurized using airflow from the cabin heating and
cooling system. Cabin pressurization is controlled by the CABIN PRESS switch and automatic
operation of the cabin pressure regulator. Cabin pressure altitude is displayed on a cabin pressure
altimeter.
Cabin pressure is controlled by the cabin pressure regulator which regulates exit airflow to maintain
a pressure/altitude schedule. The cabin is unpressurized from sea level to an aircraft altitude of 8,000
feet. Between 8,000 and 24,500 feet aircraft altitude, cabin pressure is maintained at a constant 8,000
feet. Above 24,500 feet, cabin altitude increases slowly to approximately 14,500 feet at 35,000 feet
aircraft altitude and 20,000 feet at 50,000 feet aircraft altitude. A rule of thumb for cabin altitude above
24,500 feet aircraft altitude is aircraft altitude x 0.4.
A cabin safety and dump valve is incorporated to limit cabin pressure if the pressure regulator fails.
When the CABIN PRESS switch is placed to DUMP or RAM/DUMP, the cabin safety and dump valve
opens to release pressure to the cabin pressure regulator, reducing cabin pressure to ambient.
In ECS OFF/RAM mode, avoid operations above 25,000 feet MSL in
order to prevent decompression sickness (DCS). Normal cabin pressurization is not provided. Partial cabin pressurization is available under
certain circumstances, but may be lost insidiously without warning.
2.13.9.1 CABIN PRESS Switch. The CABIN PRESS switch, located on the ECS panel on the right
console, is used to control cabin pressurization. The switch is lever-locked in the NORM position.
NORM
Automatically regulates cabin pressurization according to the cabin pressure schedule
using the cabin pressure regulator (ECS AUTO mode).
DUMP
Dumps cabin pressurization. Normal ECS airflow to the cabin and the avionics is not
affected.
RAM/
DUMP
Dumps cabin pressurization. Terminates all ECS airflow to the cabin by closing the
cabin flow valve (Item 9). Cabin airflow is provided by the cabin ram air scoop. Avionics airflow is provided by a simplified control scheme.
2.13.9.2 Cabin Pressure Altimeter. A cabin pressure altimeter, located on the center console in the
front cockpit and the lower left instrument panel in the rear cockpit, displays the current cabin
pressure altitude.
2.13.9.3 Cabin Pressurization Warning System (CPWS). The CPWS pressure switch monitors cabin
pressure and aircraft relays monitor related controls to warn aircrew of potentially hazardous cabin
pressurization conditions.
2.13.9.3.1 CABIN Caution Light. The yellow CABIN caution light is located on the lower right
caution lights panel. The light illuminates when cabin pressure altitude is above 21,000 +/- 1,100 feet.
The light may not extinguish until cabin pressure altitude is below 16,500 feet.
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• CABIN light may appear with normal cabin pressurization when
aircraft altitude is above 47,000 feet MSL. If altitude is maintained,
aircrew should continuously monitor physiological condition.
• DCS may be experienced when operating with cabin pressure altitude
above 25,000 feet even with a working oxygen system. Symptoms of
DCS include pain in joints, tingling sensations, dizziness, paralysis,
choking, and/or loss of consciousness.
NOTE
There is no corresponding DDI caution for the CABIN caution light.
2.13.9.3.2 CK ECS Caution Light. The yellow CK ECS caution light is located on the lower right
caution lights panel. The light illuminates when the position of cabin pressurization related controls
will inhibit cabin pressurization.
NOTE
There is no corresponding DDI caution for the CK ECS caution light.
2.13.10 Windshield Anti-ice and Rain Removal. The windshield anti-ice and rain removal systems
use the same air nozzle to direct warm air over the external windshield in order to improve pilot
forward visibility in icing/raining conditions. Warm air is provided by mixing engine bleed air with
output from the primary heat exchanger. Windshield air nozzle orientation is intended to affect (in
flight) an area roughly 20 inches to the left and 9 inches to the right of centerline at design eye level
and below. System operation is controlled by the WINDSHIELD switch.
2.13.10.1 WINDSHIELD Switch. The WINDSHIELD switch, located on the right console outboard of
the ECS panel, is used to select either windshield anti-ice or rain removal. The switch is lever-locked
to the OFF position.
ANTI
ICE
Delivers a high flow rate of 290 ±20°F air to the external surface of the windshield.
OFF
Terminates anti-ice/rain removal airflow.
RAIN
Delivers a low flow rate of 270 ±20°F air to the external surface of the windshield.
2.13.11 Anti-g System. The anti-g system delivers air pressure to the g-suit proportional to sensed
load factor. A button in the anti-g valve allows the aircrew to test system operation by manually
inflating the anti-g suit. The system incorporates a pressure relief valve to prevent over-pressurization.
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2.13.12 ECS RESET and AV COOL Switch.
2.13.12.1 ECS RESET. If the ECS is DEGD, selecting the ECS RESET option from the BIT/
HYDRO-MECH display commands an ECS controller software reset and may restore normal
functionality following a transient fault.
Resetting the ECS controller at mid to high power settings may cause
uncomfortable cabin pressure surges and ear pain.
Resetting the ECS controller results in the loss of system overtemperature protection for 10 to 70 seconds.
2.13.12.2 AV COOL Switch. The AV COOL switch, located on the lower right instrument panel,
extends the spring−loaded FCS ram air scoop located on the right side of the forward fuselage for
emergency cooling of FCC A, the right TR, and one AHRS unit. The AV COOL switch should be placed
in EMERG if an FCS HOT caution is displayed or during ECS OFF/RAM mode operation as a
preventative measure. Once extended, the FCS emergency ram air scoop cannot be retracted in flight.
EMERG Extends the FCS emergency ram air scoop to provide direct ram air cooling of FCC A,
the right TR, and one AHRS unit, and supplemental cooling to other avionics.
NORM
FCS emergency ram air scoop not extended. The switch is spring−loaded to the NORM
position.
2.13.13 LCS - Liquid Cooling System. The LCS is a closed loop system normally used to transfer
heat from the radar transmitter to fuel and/or ambient air. When fuel temperatures are warmer than
the liquid coolant (typically at low fuel levels and high ambient temperatures), the LCS is used to
transfer heat from the fuel to ambient air. The LCS contains a liquid coolant pump, a liquid coolant/air
heat exchanger, and an LCS ground cooling fan (all three located in the left LEX) and two liquid
coolant/fuel heat exchangers, and in LOT 26 and up, two liquid coolant/ECS air heat exchangers.
The LEX liquid coolant/air heat exchanger has one air inlet located on the bottom of the LEX, and
two air exhausts. Only one exhaust path is commanded open at any given time. During LCS ground
cooling fan operation, the exhaust on the bottom of the LEX is commanded open. Inflight, only the air
exhaust on the top of the LEX can be commanded open.
LCS ground cooling fan air is the primary cooling source for liquid coolant on deck. During ground
operations, the liquid coolant pump and LCS ground cooling fan are commanded on when power is
applied to the radar (RADAR knob in STBY, OPR, or EMERG).
During ground operations, when feed tank fuel temperatures exceed 40°C (30°C in LOT 26 and up),
the liquid coolant pump and LCS ground cooling fan may also be commanded on if the RADAR knob
in OFF in order to provide LCS cooling of the fuel system. In LOT 25 and below, the bypass valve,
which allows liquid coolant to the liquid coolant /fuel heat exchangers, is solely controlled by the liquid
coolant temperature, so the amount of fuel cooling provided (if any) depends on the temperature of
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each fluid. In LOT 26 and up, the bypass valve is controlled by the ECS controller and will only allow
liquid coolant to the liquid coolant/fuel heat exchanger with the RADAR knob in OFF.
NOTE
In LOT 26 and up, the RADAR knob must be in OFF in order to
provide any postflight LCS fuel cooling.
In either case, placing the RADAR knob to OFF removes the radar as a heat source and should
extend ground operating time.
Inflight, the LCS ground cooling fan secures, the lower air exhaust path closes, and the upper air
exhaust path is controlled by the mission computer. The upper air exhaust path opens when feed tank
fuel temperatures exceed 40°C (30°C in LOT 26 and up), ram air is cooler than fuel, and AOA is below
15°C. In LOT 25 and below, heat not removed by the LEX liquid coolant/air heat exchanger is removed
by the fuel system through the liquid coolant/fuel heat exchangers. In LOT 26 and up, heat not
removed by the LEX liquid coolant/air heat exchanger is removed by a combination of the liquid
coolant/fuel and liquid coolant/ECS air heat exchangers.
In LOT 26 and up, provisions have been incorporated to limit the risk of fire following a major liquid
coolant leak. When LCS low pressure is detected, the liquid coolant pump secures and check valves
limit coolant leakage. Since the SDC controls the liquid coolant pump, an SDC reset secures the pump,
resulting in a radar OVRHT (APG−73) or LOFLOW (APG−79) indication, and the radar stops
transmitting.
NOTE
In LOT 26 and up, an SDC reset temporarily secures the liquid coolant
pump. Before selecting SDC RESET, the RADAR knob should be set
to STBY until the SDC reset is complete and normal radar cooling
capability is restored.
2.13.14 ECS Related Warnings, Cautions, and Advisories. The following ECS related warnings,
cautions, and advisories are described in the Warning/Caution/Advisory Displays in Part V:
D
D
D
D
D
D
D
D
L BLEED and R BLEED warning lights
BLEED AIR LEFT (RIGHT) voice alert
AV AIR HOT caution
L or R BLD OFF cautions
CABIN caution light
CK ECS caution light
ECS ICING caution
EXT TANK caution
D
D
D
D
D
D
D
FCS HOT caution
FLIR OVRHT caution
GUN GAS caution
TK PRES LO caution
TK PRES HI caution
WDSHLD HOT caution
ECSDR advisory
If ECS troubleshooting has been done while on deck, check for an 8A6
MSP code after takeoff. An 8A6 MSP code is set by a system flow
modulating regulating valve failure. This failure can potentially lead to
an unexpected loss of cabin pressurization.
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2.14 OXYGEN SYSTEMS
2.14.1 On Board Oxygen Generating System (OBOGS). OBOGS provides oxygen rich breathing gas
to the aircrew while either engine is operating. Engine bleed air is cooled and routed through the
OBOGS inlet air shutoff valve to the OBOGS concentrator. The breathing gas is routed from the
concentrator to a cockpit plenum, where the temperature is stabilized and a limited supply is stored for
peak flow demands. From the plenum, the breathing gas flows through the pilot services panel oxygen
disconnect, through the seat survival kit, to the aircrew regulators and masks.
A leak or break in the breathing gas system anywhere from the pilot
services panel oxygen disconnect to the mask will prevent either normal
OBOGS breathing gas or emergency oxygen from being delivered to the
mask in its intended concentration. Leaks and breaks may be difficult to
locate and/or verify. If hypoxic symptoms are experienced or an OBOGS
system degrade occurs, immediate descent below 10,000 feet cabin
altitude is required to prevent severe or incapacitating hypoxia.
The OBOGS concentrator is powered by the left 115 volt ac bus. Two molecular sieve beds in the
OBOGS concentrator remove most of the nitrogen from the engine bleed air. The nitrogen is dumped
overboard while the remaining output of oxygen rich breathing gas is supplied to the aircrew.
2.14.1.1 OBOGS Monitor. The CRU-99/A solid state oxygen monitor is located on the left side of the
seat bulkhead in the front cockpit and is powered by the left 28 volt dc bus. The monitor continuously
measures oxygen concentration in the OBOGS breathing gas and provides a discrete signal to activate
the OBOGS DEGD caution if the oxygen concentration falls below a predetermined level.
Loss of electrical power to the OBOGS monitor prevents reporting of
OBOGS DEGD conditions.
The monitor performs a power-up BIT during a 2 minute warm-up period and conducts a periodic
BIT every 60 seconds. No indication is provided if power-up or periodic BIT pass.
Preflight BIT of the monitor is accomplished by using either the pneumatic BIT plunger or the
electronic BIT pushbutton. Refer to figure 2-33. Pressing up and holding the pneumatic BIT plunger
for 15 to 65 seconds tests the operation of the OBOGS monitor by diverting cabin air into the monitor
to create a low oxygen concentration condition.
After initiating OBOGS pneumatic BIT, failure to ensure plunger is fully
extended may result in contamination of OBOGS plenum which may lead
to hypoxic conditions.
Momentarily pressing and releasing the electronic BIT pushbutton tests the monitor electronically.
Successful completion of either test activates the OBOGS DEGD caution, which clears automatically
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Figure 2-33. OBOGS Monitor
after BIT is complete and OBOGS resets to normal operation.
Successful OBOGS monitor BIT is required prior to flight. A BIT failure
indicates that there is no protection against inadequate oxygen concentration or hypoxia due to a degraded OBOGS monitor. Good breathing
gas flow alone does not ensure adequate oxygen concentration.
2.14.1.1.1 OBOGS DEGD Caution. An OBOGS DEGD caution is set by the OBOGS monitor when
oxygen concentration is below a predetermined level. The OBOGS DEGD caution threshold is always
above cabin air conditions, in order to provide a physiological safety margin.
NOTE
Oxygen concentration may drop below the predetermined OBOGS
DEGD level if breathing gas flow is unlimited. Removing the mask
without placing the OXY flow knob to OFF, system leaks, and/or loose
aircrew hose connections can overwhelm system capacity and may
result in an OBOGS DEGD caution.
After a total loss of bleed air, OBOGS breathing gas flow will be available until the residual gas
within the system is depleted. The residual oxygen concentration may be sufficient to keep the OBOGS
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DEGD caution from illuminating, but loss of system pressure will ultimately lead to inadequate
pressure and an abrupt inability to breathe.
Low mask flow or increased breathing resistance may occur without an accompanying OBOGS
DEGD caution, which indicates a potential system degradation that may result in oxygen levels below
physiological requirements.
When cabin altitude is above 10,000 feet, OBOGS DEGD caution
procedures shall be executed for low mask flow, increased breathing
resistance, or OBOGS DEGD cautions of any duration.
A single brief appearance of the OBOGS DEGD caution immediately following placing the OBOGS
control switch to ON, turning the OXY FLOW knob to ON, or after donning or removing the mask is
normal. Certain malfunctions affecting the OBOGS system also may momentarily cause OBOGS
DEGD cautions. A momentary OBOGS DEGD caution may clear from the DDI so rapidly that the
aircrew may be unable to determine the cause of the MASTER CAUTION light/tone.
Repeated unexplained MASTER CAUTION lights/tones may be an
indication of OBOGS system degradation.
Even with the OBOGS DEGD caution displayed, OBOGS breathing gas under normal flow is of
higher quality (i.e., oxygen concentration, partial pressure and purity) than cabin air. Once cabin
altitude is below 10,000 feet, aircrew may elect to conserve emergency oxygen by resetting the
emergency oxygen release tab, then either removing the mask and breathing cabin air, or returning the
OXY FLOW knob and the OBOGS control switch to ON and breathing through the mask.
Pure oxygen accelerates recovery from hypoxia. Emergency oxygen shall
be used whenever hypoxic symptoms are recognized.
2.14.1.2 Breathing Regulator. The aircrew torso mounted breathing regulator reduces both normal
and emergency oxygen system operating pressures to breathing pressure levels. The regulator delivers
undiluted OBOGS breathing gas or emergency oxygen to the aircrew at positive pressure, the limits of
which increase automatically with altitude. It interfaces with the hose assembly, which connects with
the seat survival kit oxygen disconnect.
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2.14.1.3 OBOGS Control Switch. The OBOGS control switch, located on the left console in the front
cockpit, is used to control electrical power to the OBOGS concentrator and the OBOGS inlet air shutoff
valve.
ON
Supplies electrical power and engine bleed air to the OBOGS concentrator.
OFF
OBOGS system off.
2.14.1.4 OXY FLOW Knob. The OXY FLOW knob, located on the left console in both cockpits, is
used to control the supply of OBOGS breathing gas to each aircrew’s mask.
ON
OBOGS flow supplied to the mask.
OFF
OBOGS flow secured.
It is possible to place the OXY FLOW knob in an intermediate position between the ON and OFF detents, which may result in a reduced
flow of breathing gas. The OXY FLOW knob should always be fully
rotated to the ON or OFF detent position.
2.14.2 Emergency Oxygen. Emergency gaseous oxygen is contained in a bottle in the seat survival
kit. The bottle is connected into the OBOGS supply hose as it passes through the kit. From this point,
the emergency oxygen and OBOGS breathing gas share a common path to the aircrew mask.
A leak or break in the breathing gas system anywhere from the pilot
services panel oxygen disconnect to the mask will prevent either normal
OBOGS breathing gas or emergency oxygen from being delivered to the
mask in its intended concentration. Leaks and breaks may be difficult to
locate and/or verify. If hypoxic symptoms are experienced or an OBOGS
system degrade occurs, immediate descent below 10,000 feet cabin
altitude is required to prevent severe or incapacitating hypoxia.
A pressure gauge is visible on the inside left front of the survival kit. The bottle provides
approximately 10−20 minutes of oxygen. Oxygen duration decreases with lower altitude.
Under less than optimum conditions (low altitude, heavy breathing,
loose−fitting mask, etc.), as few as 3 minutes of emergency oxygen may be
available.
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The emergency oxygen supply is activated automatically upon ejection. The emergency oxygen
supply may be activated manually by pulling the emergency oxygen green ring on the outside of the left
thigh. The emergency oxygen supply may be deactivated at aircrew discretion by pushing down on the
release tab immediately forward of the green ring.
With emergency oxygen selected, the OXY FLOW knob(s) shall be
placed to OFF. If not secured, OBOGS system pressure may prevent
emergency oxygen from reaching the breathing regulator. Additionally,
the OBOGS control switch should be placed to OFF to backup the OXY
FLOW knob.
2.15 FIRE DETECTION, FIRE EXTINGUISHING, AND BLEED AIR LEAK DETECTION SYSTEMS
The fire detection system contains dual-loop fire detectors and three FIRE warning lights. The fire
extinguishing system contains a READY/DISCH light and a fire extinguisher bottle. The two systems
provide engine bay, AMAD bay, and APU bay fire warning, engine and APU emergency shutdown, and
selective fire extinguishing capability. The fire extinguisher bottle is located in the aft fuselage
between the engines. The bottle contains a nontoxic gaseous agent which provides a one-shot
extinguishing capability.
Electrical power from the 28 vdc essential bus is required to operate the fire detection and
extinguishing systems. The systems can operate on battery power alone with the BATT switch ON.
A separate dry bay fire suppression (DBFS) system is incorporated to automatically detect and
extinguish a fire or explosion in the dry bays below fuel tanks 2, 3, and 4.
2.15.1 FIRE Lights. Two FIRE warning lights, one for each engine/AMAD bay, are located on the
upper left and right sides of the main instrument panel. The warning lights come on when a fire
condition is detected in the respective engine/AMAD bay. The left FIRE light indicates a fire condition
in the left engine/AMAD bay. The right FIRE light indicates a fire condition in the right engine/
AMAD bay.
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Each light is a pushbutton, which is guarded to prevent inadvertent actuation. Pushing the left or
right FIRE light arms the fire extinguisher bottle (FIRE EXTGH READY light on) and closes the
corresponding engine feed shutoff valve, the crossfeed valve, and the crosscooling valve.
Because the engine VEN/start pumps are fuel lubricated, pushing a FIRE
light prior to throttle OFF may damage the corresponding pump. To
reduce the likelihood of damage, FIRE lights should only be pressed as
directed by NATOPS (following throttle OFF for actual emergencies or
as specifically delineated for an FCF A profile).
If a FIRE light is pressed, the pushbutton stays in and approximately 1/8 inch of yellow and black
stripes is visible around the outer edges of the light.
2.15.2 APU FIRE Light. The APU FIRE warning light is located on the main instrument panel
inboard of the right FIRE light. The APU FIRE light comes on when a fire condition is detected in the
APU bay. The APU FIRE light is also a pushbutton. Pushing the APU FIRE light arms the fire
extinguisher bottle (FIRE EXTGH READY light on) and secures fuel to the APU. If the APU FIRE
light is depressed, the pushbutton stays in and approximately 1/8 inch of yellow and black stripes is
visible around the outer edges of the light.
2.15.3 FIRE Warning Voice Alerts. When the left, right, and/or APU FIRE lights are illuminated, the
ENGINE FIRE LEFT, ENGINE FIRE RIGHT, or APU FIRE voice alerts, respectively, are also
activated. If more than one FIRE light comes on at the same time, the voice alert priority is LEFT,
RIGHT, then APU.
2.15.3.1 FIRE Warning Lights (F/A-18F). The rear cockpit left, right, and APU FIRE warning lights
are advisory only. These lights are not pushbuttons, and they do not arm the fire extinguisher bottle
or shut down the engines or APU.
2.15.4 FIRE EXTGH READY/DISCH Light. The FIRE EXTGH READY/DISCH light is located on
the MASTER ARM panel on the left side of the main instrument panel. The top half of the light is
yellow and is labeled READY. The bottom half of the light is green and is labeled DISCH. When the
fire extinguisher bottle is armed (left, right, or APU FIRE light pressed), the yellow READY light is
illuminated.
The FIRE EXTGH READY/DISCH light is also a pushbutton. When the READY light is on,
pushing the light discharges the fire extinguisher bottle into the selected engine/AMAD/APU bay(s).
The FIRE EXTGH READY/DISCH light does not latch like the FIRE lights, which means a signal is
sent to discharge the fire extinguisher bottle ONLY when the light is held in the fully pressed position.
When the fire extinguisher bottle has discharged or pressure has been lost, the green DISCH light
comes on. Therefore, it is good practice with an engine/AMAD bay fire to hold the READY/DISCH
light pressed until the green DISCH light comes on. The fire extinguisher bottle should discharge
within 5 seconds.
For an inflight APU fire, discharge of the fire extinguisher bottle is delayed approximately 10
seconds from when the APU FIRE light is pushed. It is not necessary to hold the READY/DISCH light
for those 10 seconds. However, if the DISCH light does not come on 10 seconds after the APU FIRE
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and READY/DISCH lights have been pushed, the READY/DISCH light should be pushed and held
until the DISCH light does come on.
If more than one FIRE light is pressed, the fire extinguisher bottle may not discharge and, if it does,
the concentration of the extinguishing agent sent to the selected bays may be insufficient to extinguish
both fires.
2.15.5 APU Fire Extinguishing System. The APU fire extinguishing system is automatically activated with WonW and must be manually activated with WoffW. If an APU fire condition is detected
with WonW, the automatic function secures fuel to the APU, arms the fire extinguisher bottle, and
after 10 seconds discharges the fire bottle. Discharge of the bottle is delayed for 10 seconds to allow the
APU time to spool down before extinguishing agent is introduced.
Manual activation is accomplished by pushing the APU FIRE light and then the FIRE EXTGH
READY light. Like automatic activation, discharge of the bottle is delayed for 10 seconds after the
APU FIRE light is pushed. If an APU FIRE condition is detected with WonW, manual activation
should be performed to backup the automatic system.
Since the fire extinguishing system requires 28 vdc essential bus power,
the fire extinguisher bottle may not be discharged if the BATT switch is
turned off during the 10 second delay time.
2.15.6 Engine/AMAD Fire Extinguishing System. The engine/AMAD fire extinguishing system
must be manually activated. Manual activation is accomplished by lifting the guard and pressing the
corresponding FIRE light and then the FIRE EXTGH READY light. Pushing the FIRE light secures
fuel to the engine at the engine feed shutoff valve and isolates the left and right fuel systems by closing
the crossfeed and crosscooling valves. When the FIRE EXTGH READY light is pushed, the fire
extinguisher bottle is discharged without delay into the corresponding engine/AMAD bay.
Fire testing indicates that the probability of extinguishing a fire and
preventing relights is greatly increased by immediately discharging the
fire extinguisher.
2.15.7 FIRE Detection System Test. Each of the three FIRE warning lights contains four individual
light bulbs. Light bulb integrity can be tested during a LT TEST with ac power applied. If a
malfunction exists in a fire detection loop associated with the APU FIRE light, the APU FIRE voice
alert does not annunciate and none of the four individual bulbs in the light illuminate. If a malfunction
exists in a fire detection loop associated with either FIRE light, the corresponding ENGINE FIRE
LEFT/RIGHT voice alert does not annunciate and only the individual bulb (or bulbs) associated with
the malfunctioning sensor does not come on. Care must be taken to detect bulbs that are not on in the
FIRE lights during the loop test.
A successful test of the FIRE detection system should illuminate all four bulbs in each of the three
FIRE lights and should annunciate the ENGINE FIRE LEFT, ENGINE FIRE RIGHT, and APU
FIRE voice alerts.
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2.15.8 Bleed Air Leak Detection (BALD) System. The BALD system is designed to protect the
aircraft from damage resulting from a bleed air leak. The system contains a BALD controller and 11
detector sensing elements routed along the bleed air distribution lines (ducts, valves, and heat
exchangers). If a bleed air leak is detected, the system attempts to isolate the affected ducting by
automatically closing the appropriate bleed air shutoff valve(s). A leak detected upstream of the
secondary bleed air shutoff valve closes only the appropriate primary bleed air shutoff valve. A leak
detected downstream of the secondary bleed air shutoff valve closes all three valves (secondary and
both primaries).
When a leak is detected, the BALD system sends commands directly to the appropriate bleed air
shutoff valves, to the BLEED warning lights, and to the ACI, triggering the BLEED AIR LEFT
(RIGHT) voice alerts. When the bleed air shutoff valve(s) are commanded closed, the appropriate L
and/or R BLD OFF caution is displayed. The L BLEED and/or R BLEED warning lights extinguish
as soon as bleed air is removed from the leaking duct and may not be on long enough to be recognized
by the aircrew.
Automatic functioning of the BALD system may extinguish the L(R)
BLEED warning lights prior to aircrew recognition and may not trigger
the appropriate voice alerts. In this case, cycling the BLEED AIR knob to
remove the L and/or R BLD OFF cautions reintroduces hot bleed air to
the leaking duct. If the sensing element was damaged by the leak,
automatic shutdown and isolation capability may be lost. Extensive
damage and/or fire may result.
The BALD controller also sends a separate command to the SDC which sets an MSP code for the
appropriate sensing element.
A bleed air leak can be verified by MSP codes 953, 954, 955, 956, 957, 958, 959, 960 or 961 (code
determines leak location). An overpressure condition is indicated by MSP code 833 with no bleed air
leak codes.
Sustained high power, high speed operation at low to medium altitude
may result in dual BLEED warnings and loss of OBOGS, cabin pressurization, and g-suit due to secondary bleed air pressure regulator bay
overheat.
2.15.8.1 Bleed Air Leak Detection System Test. The BALD system is tested by the FIRE test switch
in conjunction with the FIRE detection system. Actuation of the FIRE test switch tests the BALD
system sensors and circuitry. The BALD controller turns on the L and R BLEED warning lights,
annunciates the BLEED AIR LEFT/RIGHT voice alerts, and commands the bleed air shutoff valves
closed setting the L and R BLD OFF cautions. The warnings and cautions indicate that the test has
successfully passed.
The L BLEED and R BLEED warning lights go out when the FIRE test switch is released to NORM,
but the L and R BLD OFF cautions remain until the BLEED AIR knob is cycled through OFF to
NORM with ac power applied.
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2.15.8.2 FIRE Test Switch. The FIRE test switch, located on the forward left console, is used to
initiate a test of the FIRE detection and BALD systems. Operation of the FIRE test switch requires
28 vdc essential bus power. The switch is spring-loaded to the NORM position.
TEST A
Initiates a test of loop A of the fire detection and BALD systems.
NORM
Provides normal fire and bleed air leak detection.
TEST B
Initiates a test of loop B of the fire detection and BALD systems.
A successful test of the FIRE detection and BALD systems should illuminate all four bulbs in the
left, right, and APU FIRE lights, both L BLEED and R BLEED warning lights, and should annunciate
all of the following voice alerts in order: ″ENGINE FIRE LEFT, ENGINE FIRE RIGHT, APU FIRE,
BLEED AIR LEFT, BLEED AIR RIGHT″ (each repeated twice).
The BALD controller is sensitive to the duration of FIRE test switch actuation. If the switch is not
held in the TEST A or TEST B position for at least 2 seconds, the BALD controller may set a false
MSP code. Additionally, the BALD controller requires 3 seconds between TEST A and TEST B to
successfully reset. If the switch does not remain in NORM for at least 3 seconds, the BALD controller
may not successfully initiate the bleed air warnings and may set a false MSP code.
2.15.9 DBFS - Dry Bay Fire Suppression System. An active DBFS system is incorporated in the
center dry bays under fuel tanks 2, 3, and 4 to automatically extinguish any fires or explosions which
are detected in these areas. The system consists of fourteen optical fire detectors, seven extinguishing
units and a control unit. The control unit integrates system operation and performs BIT. If the DBFS
system fails BIT, a BIT advisory appears, and DBFS BIT status indicates DEGD.
The DBFS system is totally automatic and requires no aircrew action. The system is armed and
capable of suppressing a fire/explosion in the dry bays when the LDG GEAR handle is up and at least
one generator is on line. If a fire/explosion is detected, the controller discharges all extinguishers,
flooding all dry bays with an inert gas (BAY DISCH caution set). If a fire condition is still detected
after 3 seconds, a BAY FIRE caution is set to alert the aircrew that the fire was not extinguished. If
the fire condition ceases, the BAY FIRE caution resets. With the LDG GEAR handle down, the system
still gives a fire warning (BAY FIRE caution) but is not capable of extinguishing a dry bay fire.
2.16 ENTRANCE/EGRESS SYSTEMS
2.16.1 Canopy System. The cockpit is enclosed by a clamshell type canopy. The main components
of the canopy system are an electromechanical actuator, which provides powered and manual operation
of the canopy, and a cartridge actuated thruster with associated rocket motors, which provides
emergency jettison. When closed, the canopy is latched in place by three hooks on the bottom of each
side of the canopy frame and two forward indexer pins on the lower leading edge of the canopy frame.
When the canopy is closed, the latch hooks and indexer pins engage fittings along the canopy sill, and
the canopy actuator rotates the canopy actuation link over-center, locking the canopy. A mechanical
brake in the canopy actuator motor provides a redundant lock. An inflatable seal, installed around the
edge of the canopy frame, retains cockpit pressure when the canopy is locked. A rain seal is installed
outboard of the pressure seal to divert rain water away from the cockpit. See figure 2-34.
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A high voltage (100,000 volt) static electrical charge may build up in flight
and be stored in the canopy. If possible, make sure that ground crew
discharge the static electricity prior to egress. Otherwise, avoid direct
contact with the outside of the windscreen and canopy to prevent
electrical shock.
Taxiing with the canopy at an intermediate position can result in canopy
attach point damage and failure. Do not open or close the canopy with the
aircraft in motion.
2.16.1.1 Canopy Operation. During normal operation, the canopy is electrically actuated using
either the internal CANOPY switch or external canopy switch. The canopy actuator is powered by the
maintenance bus, which is powered directly by the battery in the absence of ac power (BATT switch
ON or OFF). On battery power, at least five open/close cycles should be available. With the canopy
open, the CANOPY switch must be held to lower the canopy to the rails, slide it approximately 1.5
inches forward, and lock it in place. With the canopy closed and locked, selecting OPEN on the
CANOPY switch (WonW) automatically unlocks and opens the canopy to the full up position. The
switch does not have to be held. Whether the canopy is opening or closing, selecting HOLD stops the
canopy at its present position.
If electrical power is not available, the canopy can be manually operated using either an internal or
external crank system. The canopy can be jettisoned using one of the internal CANOPY JETT
handles.
2.16.1.1.1 CANOPY Switch (Internal). The internal CANOPY switch is located beneath the right
canopy sill in the front cockpit. In LOT 26 and up, the rear cockpit CANOPY switch is located on the
lower right portion of the instrument panel. The CANOPY switch is spring loaded to the HOLD
position and is solenoid-held in the OPEN position only with WonW. The solenoid can be overridden
at any time by placing the switch to HOLD. With WoffW, the switch must be held in the OPEN
position to raise the canopy. Opposing position control commands between the front cockpit and rear
cockpit switches result in a fail-safe OPEN command.
OPEN
Unlocks and/or raises the canopy.
HOLD
Stops the canopy at any point during the open or close cycle.
CLOSE
Lowers and, if held, closes and locks the canopy (CANOPY caution out when
closed and locked).
2.16.1.1.2 Canopy Switch (External). The external canopy switch is located inside the external
power receptacle door (door 9) on the left side of the aircraft below the canopy and LEX. The switch
provides electrical operation of the canopy from outside the cockpit. The switch has the same positions
and operates identically to the internal CANOPY switch, except that the OPEN position is not
solenoid held. After AFC 366, the switch is guarded to prevent inadvertant actuation.
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Figure 2-34. Canopy Controls
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2.16.1.1.3 Manual Canopy Handcrank (Internal). The internal manual canopy handcrank is stowed
in a clip beneath the left canopy sill. The canopy can be manually opened or closed by inserting the
handcrank into the crank socket immediately above the stowage clip. Approximately 70 counterclockwise turns are required to fully open the canopy. Clockwise cranking closes the canopy. A cable is
provided to prevent loss of the handle if dropped.
2.16.1.1.4 Manual Canopy Actuation Fitting (External). The external manual canopy actuation
fitting, a 3/8 inch drive socket on the left side of the aircraft below the canopy, is used to manually
operate the canopy. Inserting a 3/8 inch drive tool in the socket and then turning counterclockwise
approximately 35 turns opens the canopy. Turning the drive tool clockwise closes the canopy.
2.16.1.2 Canopy Jettison System. For canopy jettison, a cartridge initiated thruster is utilized to
unlatch the canopy by moving it 1.5 inches rearward, after which two canopy frame mounted rocket
motors fire to rotate the canopy up and aft, clear of the ejection seat path. The thruster, which provides
attachment for the canopy actuator link during normal canopy operation, is activated by pulling the
ejection seat firing handle or internal canopy jettison handle(s) (figure 2-34). The canopy can be
jettisoned closed, open, or in any intermediate position.
2.16.1.2.1 CANOPY JETT Handle. The CANOPY JETT handle is black and yellow striped and is
located on the left inboard canopy sill just aft of the instrument panel in the front cockpit. Pressing the
button on the tip of the handle unlocks the handle. Pulling the handle aft initiates the canopy jettison
sequence. A ‘‘remove before flight’’ pin is used to manually secure the CANOPY JETT handle between
flights.
2.16.1.2.2 CANOPY JETT Handle (F/A-18F). The rear cockpit CANOPY JETT handle is black and
yellow striped and is located on the left console. Pressing the button on the forward tip of the handle
unlocks the handle. Pulling the handle up initiates the canopy jettison sequence. A REMOVE
BEFORE FLIGHT pin is used to manually secure the CANOPY JETT handle between flights.
2.16.2 Boarding Ladder. A five-step boarding ladder (figure 2-35), stowed under the left LEX,
provides access to the cockpit and the top of the aircraft. Ladder extension and retraction can be
accomplished only from outside the cockpit, either manually or by the ladder remote release button
(after AFC 366) or switch (before AFC 366).
The ladder is extended manually by releasing the latch on the stow assist handle on the ladder’s left
rail (allowing it to drop slightly) and while supporting the ladder, rotating the stow assist handle to
vertical (releasing the remaining two mechanical uplocks on the underside of the LEX). The ladder
rotates down to the extended position. The stow assist handle is then secured. The drag brace locks
when extended to its full length to provide longitudinal stability for the ladder. Lateral stability is
provided by the V-shaped side brace attached to the side of the fuselage.
The LEX is narrow and highly sloped. Use caution to avoid loss of
footing.
NOTE
The ladder is not visible from the cockpit.
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Normal cockpit egress is accomplished by grasping the canopy sill firmly with both hands, leaning
outboard and, using the ladder marking decals as a guide, stepping over the LEX toward the ladder.
The first step is approximately 15 inches below the leading edge of the LEX.
The ladder is stowed by detaching the rigid side brace connection from the fuselage. Pulling the
collar on the drag brace down permits the telescoping drag brace to unlock and compress as the
boarding ladder is rotated up and aft to the stowed position. The latches are manually engaged and
locked by pushing them full up until locked flush with the forward beam. If necessary, the stow assist
handle can be used to assist in stowing the ladder by releasing the handle and pushing the ladder to
the stowed position and pushing the stow assist handle to the closed (stowed) position and releasing it.
With electrical power on the aircraft, a LADDER caution comes on whenever the proximity switch in
the aft portion of the ladder well is not actuated. With the ladder stowed and the 3 latches locked, the
LADDER caution goes out.
The ladder is extended remotely by opening the ground power receptacle door and holding the
ladder remote release switch (lower switch) to the down (DEPLOY) position. The three latches are
opened by a battery powered actuator. The ladder drops to the open position while being restrained
from free falling by a dampening strut (drop time approximately 3 to 4 seconds). All other procedures
for securing the ladder are the same as manual opening.
2.16.2.1 Ladder Remote Release Switch (Before AFC 366). The external ladder remote release
switch is located inside the external power receptacle (door 9) on the left side of the aircraft below the
canopy and LEX. The switch is lever locked and spring loaded to the UP position. Switch positions are:
UP
Removes electrical power from the unlock actuators which allows the ladder to lock
when it is manually stowed.
DEPLOY
Applies electrical power to the three unlock actuators which unlocks the actuators
so the ladder can free fall. The switch must be pulled out and down and held in
DEPLOY until the actuators are unlocked.
Due to the close proximity of the ladder remote release switch and the
external canopy switch, positive switch identification is required to
prevent inadvertently lowering the canopy and injuring personnel egressing the aircraft.
2.16.2.2 Ladder Remote Release Button (After AFC 366). The guarded external ladder remote
release button is located inside the external power receptacle (door 9) on the left side of the aircraft
below the canopy and LEX. Pressing and holding the button applies electrical power to the three
unlock actuators which unlocks the actuators so the ladder can free fall.
2.16.3 Ejection Seat. The SJU-17 (V)1/A, 2/A, and 9/A, and SJU-17A (V)1/A, 2/A, and 9/A NACES
(Navy Aircrew Common Ejection Seat) are ballistic catapult/rocket systems that provide the pilot with
a quick, safe, and positive means of escape from the aircraft. See Ejection Seat, foldout section, for
ejection seat illustrations. The seat system includes an initiation system which, after jettisoning the
canopy and positioning the occupant for ejection, fires the telescopic seat catapult. Canopy breakers on
the top of the seat give capability of ejecting through the canopy. As the seat departs the aircraft and
the catapult reaches the end of the stroke a rocket motor on the bottom of the seat is fired. The thrust
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-35. Boarding Ladder
of the rocket motor sustains the thrust of the catapult to eject the seat to a height sufficient for
parachute deployment even if ejection is initiated at zero speed, zero altitude in a substantially level
attitude.
NOTE
Safe escape is provided for most combinations of aircraft altitude,
speed, attitude, and flight path within the envelope of 0 to 600 KCAS
airspeed and 0 to 50,000 feet.
2.16.3.1 SJU-17(V) 1/A, 2/A, and 9/A, and SJU-17A (V)1/A, 2/A, and 9/A NACES Seat. Timing of
all events after rocket motor initiation is controlled by the electronic sequencer which utilizes altitude,
acceleration, and airspeed information to automatically control drogue and parachute deployment and
seat/man separation throughout the ejection seat’s operational envelope. In the event of partial or total
failure of the electronic sequencer, a 4-second mechanical delay initiates a barostatic release unit which
frees the occupant from the seat and deploys the parachute between 14,000 and 16,000 feet MSL if the
ejection occurred in or above this altitude range. The emergency barostatic release unit operates
immediately after the 4-second delay if the ejection occurred below 14,000 feet MSL. An emergency
restraint release (manual override) system provides a backup in the event of failure of the barostatic
release unit. The seat is stabilized and the forward speed retarded by a drogue chute attached to the
top and bottom of the seat. The parachute deployment rocket is automatically fired to withdraw the
parachute from deployment bag. Full canopy inflation is inhibited until the g forces are sufficiently
reduced to minimize opening shock. There are 5 modes of operation. See figure 2-36 for parameters
that determine the mode of operation and the corresponding parachute deployment and drogue chute
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-36. SJU-17 and SJU-17A Ejection Modes
release times. At high altitude the drogue chute deploys to decelerate and stabilize the seat. The seat
falls drogue retarded to 18,000 feet MSL where the drogue is released, the main parachute is deployed,
and seat/man separation occurs. At medium altitude, (between 18,000 and 8,000 feet MSL), and at low
altitude (below 8,000 feet MSL) parachute deployment is automatically delayed from 0.45 to 2.90
seconds (depending upon airspeed and altitude) after first seat motion to allow the drogue chute to
decelerate and stabilize the seat.
The main parachute is a 21 foot aeroconical canopy type, stored in a headbox container on top of the
ejection seat. The parachute is steerable and contains water deflation pockets which aid in dumping air
from the canopy after landing in water. The seat drogue chute is stored in a separate container on top
of the drogue deployment catapult. The seat contains controls for adjusting seat height and for locking
and unlocking the inertia reel shoulder restraint straps. A survival kit is installed in the seat pan.
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A1-F18EA-NFM-000
2.16.3.2 SEAWARS - SEAWATER Activated Release System. SEAWARS is a seawater activated
system that automatically releases the parachute from the crew member. When the sensing-release
units are immersed in seawater, cartridges are fired which allow the crew member to separate from the
parachute.
2.16.3.3 Ejection Control Handle. The ejection control handle, located between the crewman’s legs
on the front of the seat pan, is the only means by which ejection is initiated. The handle, molded in the
shape of a loop, can be grasped by one or two hands. To initiate ejection, a 20 to 40 pound pull removes
the handle from its housing, and a continued pull of 30 to 60 pounds is required to pull both sears from
the dual initiators. Either of the initiators can fire the seat. After ejection, the handle remains attached
to the seat. The ejection control handle safes the ejection seat safe/armed handle.
2.16.3.4 Ejection Seat SAFE/ARMED Handle. To prevent inadvertant seat ejection, an ejection seat
safe/armed handle is provided. The handle, forward on the right seat armrest, safeties the seat when
rotated up and forward, and arms the seat when it is rotated aft and down. The safe/armed handle is
locked when placed to either of these two positions and the handle must be unlocked by squeezing a
locking lever within the handle cutout before changing positions. When in the armed position the
visible portion of the handle (from the occupant’s vantage point) is colored yellow and black with the
word ARMED showing. In the safe position, the visible portion is colored white with the word SAFE
showing. The seat is safe only when the word SAFE is entirely visible on the inboard side of the
SAFE/ARM handle and the handle is locked in the detent. Placing the handle to the SAFE position
causes a pin to be inserted into the ejection firing mechanism to prevent withdrawal of the sears from
the dual seat initiators.
2.16.3.4.1 CK SEAT Caution. The CK SEAT caution light is located on the caution light panel and
repeats the DDI CHECK SEAT caution. The caution comes on when the right throttle is at MIL or
above, weight is on wheels, and the ejection seat is not armed.
2.16.3.5 Shoulder Harness Inertia Reel. Pilot shoulder harness restraint is provided by a dual strap
shoulder harness inertia reel mounted in the seat below the parachute container. The dual inertia reel
shoulder straps connect to the parachute risers which in turn are buckled to the seat occupant’s upper
harness. The inertia reel locks when the reel senses excessive strap velocity. Manual locking and
unlocking of the reel is controlled by the shoulder harness lock/unlock handle on the left side of the
seat bucket. During ejection a pyrotechnic cartridge is fired to retract the shoulder harness to position
the seat occupant for ejection.
2.16.3.6 Shoulder Harness Lock/Unlock Handle. The shoulder harness lock/unlock handle on the
left side of the seat bucket has two positions. To operate, the handle must be pulled up against spring
pressure, moved to the desired position, and released.
FORWARD
(locked)
The inertia reel prevents the reel straps from being extended and ratchets any
slack in the straps back into the reel.
AFT
(unlocked)
The reel allows the pilot to lean forward, but the inertia portion of the reel continues to protect by locking the reel when it senses excessive strap velocity. Once
locked, the pilot can normally lean forward again after a slight release in pressure
on the reel straps.
2.16.3.7 Leg Restraint System. A leg restraint system is located on the front of the ejection seat. The
function of the system is to secure the occupant’s legs to the seat during ejection. The system consists
of two adjustable leg garters, a restraint line, and a snubber box for each leg. One garter is worn on the
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-37. Leg Restraint System
thigh and one on the lower leg. The restraint lines are routed through the garter rings and the snubber
box as shown in figure 2-37. One end of each restraint line is secured to the cockpit floor and the other,
after being routed through the snubber box and both garter rings, is secured to the seat just outboard
of the snubber box by a releaseable pin. During ejection, the slack in each line is taken up and the
tension builds up to finally separate the lines at the tension rings in the leg lines. At man/seat
separation, the pins on the other end of the lines are released by the time release mechanism. The pins
are also released when the manual override handle is pulled. Both the lower garter and thigh garter
contain a quick release buckle which disconnects the ring through which the leg restraint line runs,
permitting the pilot to egress from the aircraft wearing both upper and lower garters. In addition, toe
clips are installed on the tops of the rudder pedals to prevent contact between the toes and the
instrument panel during ejection.
2.16.3.7.1 Leg Restraint Snubber Release Tabs. The leg restraint lines are adjusted to give the pilot
more leg movement by pulling inboard the leg restraint snubber release tabs (figure 2-37) and
simultaneously pulling the leg restraint lines forward through the snubber box.
2.16.3.8 Seat Survival Kit (SKU-10/A). The SKU-10/A survival kit is used with the SJU-17 and
SJU-17A ejection seats. This survival kit, which fits into the seat bucket, is a contoured rigid platform
which contains an emergency oxygen system and a fabric survival rucksack (figure 2-38). A cushion on
top of the platform provides a seat for the aircrew.
The rigid platform forms a hard protective cover to the survival package and oxygen system and is
retained in position in the seat bucket by brackets at the front and lugs secured in the lower harness
locks at the rear. Attached to the lugs are two adjustable lap belts with integral quick release fittings.
A flexible oxygen and communication hose is installed in the left aft side of the upper kit to provide
a connection to the aircrew for aircraft oxygen and communication. An emergency oxygen cylinder,
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-38. Survival Kit
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ORIGINAL
A1-F18EA-NFM-000
pressure reducer, and associated pipe work are mounted on the underside of the platform. A green
manual emergency oxygen operating handle is mounted on the left side of the platform and a pressure
gage is on the inside face of the left leg support. The emergency oxygen can be activated manually by
pulling the green emergency oxygen handle upwards. The green emergency oxygen handle can be reset,
shutting off the flow of emergency oxygen, by pushing downward on the button on the front end of the
emergency oxygen handle assembly. The emergency oxygen is automatically activated during ejection
by a lanyard connected between the floor and the survival kit. An AN/URT-33A locator beacon is
located in a cutout in the left leg support. The beacon is actuated during ejection by a lanyard
connected to the emergency oxygen lanyard.
The survival rucksack is retained to the underside of the rigid platform by 5 fabric straps and a
double cone and pin release system. The package accommodates a life raft and survival aids. Two
yellow manual deployment handles are mounted on the aft surface of the kit. Pulling either handle
enables the aircrew to deploy the raft and survival package after man/seat separation. The life raft
inflates automatically on survival package deployment and is attached to the survival package with a
line. If the survival kit must be deployed after water entry, a snatch pull on the red manual activation
handle near the CO2 bottle is required to inflate the life raft.
2.16.3.9 Manual Override Handle. A manual override handle permits releasing the pilot’s lower
harness restraints and the leg restraint lines for emergency egress and permits resuming part of the
ejection sequence (man/seat separation and main parachute deployment) in the event of sequencing
failure during ejection. The manual override handle, on the right side of seat bucket and just aft of the
ejection seat safe/arm handle, is actuated by pressing a thumb button on the forward part of the handle
and rotating the handle up and aft. If the manual override handle is actuated on the ground or in the
air before ejection, survival kit attachment lugs and leg restraint lines are released, the inertia reel is
unlocked, and the ejection seat safe/armed handle automatically rotates to the SAFE position. During
ground emergency egress, after the manual override handle is pulled and the parachute riser fittings
are released, the pilot is free to evacuate the aircraft with the survival kit still attached. If the manual
override handle is actuated after ejection but before man/seat separation occurs, the following events
take place: release of survival kit attachment lugs, negative-g strap (SJU-17 (V)1/A, 2/A, and 9/A only),
leg restraint lines, and inertia reel straps; firing of the manual override initiator cartridge; firing of the
barostatic release unit; and firing of the parachute deployment rocket, which deploys the parachute.
The ejection seat safe/armed handle automatically rotates to the SAFE position whenever the manual
override handle is actuated.
Pulling the manual override handle automatically rotates the ejection
seat safe/armed handle to the SAFE position, releases the survival kit
attachment lugs and leg restraint lines, and unlocks the inertia reel. If
this is done inflight, the aircrew will be unable to eject.
2.16.3.10 Seat Bucket Position Switch. The seat bucket position switch is on the left side of the seat
bucket, forward of the shoulder harness lock/unlock handle. The forward switch position lowers the
seat bucket, the aft position raises the seat bucket. The center off position, to which the switch is spring
loaded, stops the seat bucket. The maximum vertical travel of the seat bucket is 5.1 inches for the
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ORIGINAL
A1-F18EA-NFM-000
SJU-17 (V)1/A, 2/A, and 9/A, and 6.1 inches for the SJU-17A (V)1/A, 2/A, and 9/A. The actuator
should not be operated over 1 minute during any 8 minute period.
• To prevent increased risk of thigh slap or leg contact injuries, aircrew
with a buttock-to-knee length greater than 25.5 inches should not use
either of the two forward backpad positions. Aircrew with buttock-toknee length between 24.6 and 25.5 inches should not use the full
forward backpad position.
• Actuation of the seat bucket position switch with the leg restraints
under the seat bucket may result in an inadvertent ejection.
Actuation of seat bucket position switch with lap belts, shoulder harness,
and/or leg restraints outside or under seat bucket may damage ejection
seat, leg restraints, and/or Koch fittings.
2.16.3.11 Backpad Adjustment Mechanism (SJU-17A (V)1/A, 2/A, and 9/A). The backpad adjustment mechanism handle is on the seat bucket adjacent to the top left hand side of the backpad and is
connected to the backpad by a linkage. The backpad has three positions, full-forward, middle, and fullaft, which give a total forward/aft adjustment of 1.6 inches. When the handle is in the full-up position,
the backpad is full-aft, and when the handle is full-down, the backpad is full-forward. To move the
backpad, the adjustment handle is moved within a quadrant until a spring-loaded plunger engages in
one of the three detent positions in the quadrant. Set the backpad for personal comfort and best access
to flight controls during initial strap-in and prior to flight.
2.16.4 Ejection Seat System (F/A-18F). The ejection seats in the F/A-18F are ejected at opposite
divergent angles to one another. The rear seat diverges to the left while the forward seat diverges to the
right. The amount of divergence is influenced by the weight of the aircrew and the speed of the ejection.
The heavier the aircrew and the faster the speed the less the resulting divergent angle. In addition, a
sequencing system is installed to allow dual ejection initiated from either cockpit or single (aft) seat
ejection initiated from the rear cockpit. A command selector valve is installed in the rear cockpit to
control whether ejection from the rear cockpit is dual or single.
2.16.4.1 EJECTION MODE Handle (F/A-18F). The EJECTION MODE handle is located on the right
side of the main instrument panel in the rear cockpit. The EJECTION MODE handle is used to select
the desired ejection sequence to be initiated from the rear cockpit, or provide for single ejection for solo
flight. Positioning is accomplished by pulling out while turning to the desired position. The SOLO
position requires the use of a collar to hold the handle in that position. To release from AFT
INITIATE, pull then turn clockwise.
NORM
(vertical)
Single rear seat ejection when initiated from the rear cockpit. Dual ejection (rear
seat first) when initiated from the front cockpit.
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ORIGINAL
A1-F18EA-NFM-000
AFT
INITIATE
(horizontal)
Dual ejection (rear seat first) when initiated from either cockpit.
SOLO
(45° CCW)
Front seat ejection only when initiated from the front cockpit. Front seat ejection
is immediate. Rear seat ejection only when initiated from the rear seat. Rear seat
ejection is immediate.
• SOLO mode shall NOT be selected when both seats are occupied. If
SOLO mode is selected when both seats are occupied, simultaneous
ejection initiation may result in a collision between seats.
• SOLO mode shall be selected when the aircraft is being flown solo.
Alternate selection when flying solo results in ejection of unoccupied
seat and possible collision with the front cockpit seat.
When selecting NORM or SOLO from AFT INITIATE, the handle must
be pulled before rotation or damage to the command selector valve may
result.
2.16.4.2 SEAT CAUT MODE Switch (F/A-18F). The SEAT CAUT MODE switch is located in the
rear cockpit above the command selector valve. The switch position changes the operation of the CK
SEAT caution for solo or dual flight.
NORM
CK SEAT caution is activated by either seat remaining safed. Switch is spring
loaded to this position.
SOLO
CK SEAT caution is activated only by the front seat remaining safed. Switch must
be pinned to remain in this position.
2.16.4.3 CK SEAT Caution (F/A-18F). The CK SEAT caution light is located on the caution light
panel, and repeats the DDI CHECK SEAT caution display. The caution is displayed when the right
throttle is at MIL or above, weight is on wheels, and the cockpit ejection seat is not armed with the rear
cockpit NORM/SOLO switch set to SOLO or either ejection seat is not armed with the NORM/SOLO
switch set to NORM.
2.17 EMERGENCY EQUIPMENT
2.17.1 Jettison Systems. The jettison systems consist of the emergency jettison system and the
selective jettison system.
2.17.1.1 Emergency Jettison. Emergency jettison is performed by pushing the EMERG JETT
button with either the LDG GEAR handle UP or with WoffW. When activated, the emergency jettison
system jettisons all stores, launchers, and racks from the BRU-32 racks on the six wing pylon stations
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ORIGINAL
A1-F18EA-NFM-000
(2, 3, 4, 8, 9, and 10) and the centerline station (6). Emergency jettison is sequential by station pairs:
3 and 9, 2 and 10, 4 and 8, then 6. There is a 100 msec, ±25 msec, delay before the first set of stations
is jettisoned and in between each subsequent set.
2.17.1.2 EMERG JETT Button. The EMERG JETT button is located on the left side of the main
instrument panel and is black and yellow striped. The button must be pressed and held during the
entire jettison sequence. The EMERG JETT button is used to initiate emergency jettison with the
LDG GEAR handle UP or with WoffW.
The EMERG JETT button must be pressed for 500 msec to make sure all
stores are jettisoned.
If the EMERG JETT button has been pushed on the ground prior to
takeoff and remains stuck in, emergency jettison is activated as soon as
the aircraft goes WoffW. The only cockpit indication of this condition is
SMS BIT status DEGD and MSP 082 (Emergency Jettison Switch Failed
On).
2.17.1.3 Selective Jettison. Selective jettison is performed using the SELECT JETT knob, in
conjunction with the JETT STATION SELECT buttons, and jettisons stores in a safe condition. The
stores or the launchers/racks (with any attached stores) can be jettisoned from the centerline and wing
stations, and the missiles can be jettisoned from the fuselage stations.
Selective jettison requires ARM conditions satisfied and all the landing gear up and locked. ARM
conditions are satisfied with WoffW, LDG GEAR handle UP, MASTER ARM switch in ARM, and
SIM mode unboxed. ARM status can be confirmed on the STORES page. All the landing gear up and
locked can be confirmed by the absence of the LDG GEAR handle warning light/landing gear warning
tone with the LDG GEAR handle UP.
Selective jettison of the centerline and wing stations requires station(s) selection by the JETT
STATION SELECT buttons and STORES or RACK/LCHR selection by the SELECT JETT knob.
Selective jettison of a fuselage station missile requires R FUS MSL or L FUS MSL selection by the
SELECT JETT knob.
With all the requirements met, selective jettison is performed by pressing the JETT center
pushbutton in the SELECT JETT knob.
2.17.1.3.1 JETT STATION SELECT Buttons. The JETT STATION SELECT buttons are on the left
edge of the instrument panel below the emergency jettison button. The buttons are labeled CTR, LI,
RI, LM, RM, LO and RO. Pressing a button turns on an internal light and selects a weapon station for
jettison. The JETT STATION SELECT buttons are also used in the backup A/G weapon delivery
modes for weapon selection; refer to A1-F/A-18EA-TAC (Series).
2.17.1.3.2 SELECT JETT Knob. The SELECT JETT knob on the left vertical panel has rotary
positions L FUS MSL, SAFE, R FUS MSL, RACK/LCHR, and STORES. L FUS MSL and R FUS
MSL select either fuselage missile for jettison. The RACK/LCHR and STORES positions select what
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ORIGINAL
A1-F18EA-NFM-000
is to be jettisoned from the weapon station(s) selected by the JETT STATION SELECT buttons. The
JETT center pushbutton activates the jettison circuits provided ARM conditions are satisfied and all
the landing gear are up and locked. The SAFE position prevents any selective jettison.
2.17.1.3.3 AUX REL Switch. The AUX REL switch, on the lower instrument panel, is used to enable
jettison of hung stores or store and rack/launcher combinations from BRU-32 racks on stations 2, 3, 4,
6, 8, 9, and 10. A need to use the AUX REL switch is indicated by a hung indication on the DDI after
a selective jettison or a normal weapons release is attempted. Placing the switch to ENABLE selects
the auxiliary release function. AUX RELEASE has the same requirements as those listed above for
selective jettison. Jettison is initiated by selecting RACK/LCHR or STORES on the selective jettison
knob, selecting the hung store station by pressing the appropriate station jettison select button, and
pressing the JETT center pushbutton of the selective jettison knob. The SMS provides a jettison signal
to fire the auxiliary cartridge in the BRU-32 rack on which the hung store or store and rack/launcher
combination is loaded. After the cartridge is fired, the store or rack/launcher is gravity dropped with
the store in a safe condition. This switch is also used with some weapons for a second normal release
attempt after these weapons have been hung during a first normal release attempt. Refer to
A1-F18EA-TAC-Series for these weapons and procedures.
2.17.2 Warnings/Cautions/Advisories. The warning/caution/advisory system provides visual indications of normal aircraft operation and system malfunctions affecting safe operation of the aircraft.
The lights are on various system instruments and control panels in the cockpit. The red warning lights
normally indicate a systems malfunction that could be a severe hazard to further flight, and may
require immediate action. Caution lights and displays normally, but not always, indicate malfunctions
that require attention but not immediate action. After the malfunction has been corrected, warning
and caution lights and caution displays go out. Advisory lights and displays indicate safe or normal
conditions and supply information for routine purposes. Warning, caution and advisory displays are
NVG compatible. Caution and advisory displays appear on the left or right DDI and the MPCD,
depending on the number of displays in operation. The advisory displays start at the bottom of the
display and are preceded by ADV. The caution displays, in larger characters than the advisory displays,
appear immediately above the advisory displays. The caution lights, located on the caution lights panel
and the instrument panel, are yellow lights. The advisory lights, scattered throughout the cockpit(s),
are green. Lights that have been lit on the caution lights panel flash when overheated to prevent light
damage.
2.17.2.1 MASTER CAUTION Light. A yellow MASTER CAUTION light, on the upper left part of the
instrument panel, comes on when any of the caution lights or caution displays come on. The MASTER
CAUTION light goes out when it is pressed (reset). An audio tone is initiated whenever the MASTER
CAUTION light comes on. The tone is of 0.8 second duration and consists of a 0.25 second sound
followed by a 0.15 second sound of higher pitch, followed by one repetition of these sounds. The tone
does not repeat unless the original condition causing the tone clears and recurs 5 seconds after the first
tone, regardless of whether or not the MASTER CAUTION is reset. Additional cautions sound the
tone, regardless of whether or not the MASTER CAUTION is reset, providing about 5 seconds have
elapsed since the previous caution. Pressing the MASTER CAUTION when it is unlighted causes the
uncorrected caution and advisory displays to reposition to the left and to a lower level, provided there
is available space vacated by corrected caution and advisory displays. To restack the cautions and
advisories when the MASTER CAUTION is lighted, the MASTER CAUTION must be pressed twice:
first, to turn off the MASTER CAUTION light and second, to reposition the caution and advisory
displays. A reset MASTER CAUTION light (and tone) comes on if there is at least one uncorrected
caution present when weight is on the wheels and both throttles are moved beyond approximately 80%
rpm if both throttles were below 80% for at least 60 seconds.
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ORIGINAL
A1-F18EA-NFM-000
2.17.2.2 MASTER CAUTION Light (F/A-18F). A yellow MASTER CAUTION light, on the upper
instrument panel comes on whenever the MASTER CAUTION light in the front cockpit comes on.
The rear cockpit MASTER CAUTION light goes out whenever the front cockpit MASTER CAUTION
is reset.
2.17.2.3 Dimming and Test Functions. There are no provisions for testing the caution and advisory
displays and each DDI contains its own display dimming controls. The warning/caution/advisory lights
are dimmed by the warning/caution lights knob and are tested by the lights test switch. The following
lights can be dimmed by the warning/caution lights knob, but once in the dimmed lighting range
cannot be varied in intensity: MASTER CAUTION light, landing gear handle warning, L BAR
warning, HOOK warning, L BLEED warning, R BLEED warning, APU FIRE warning, left and right
engine FIRE warning.
2.17.3 Voice Alert System. For certain critical warnings and cautions, voice alert transmissions are
sent to the aircrew’s headset. The message is repeated twice; for example, APU FIRE, APU FIRE. The
voice alert requires no reset action on the pilot’s part and the alert is not repeated unless the original
condition ceases for 5 seconds or more and then recurs. For cautions with voice alert, the voice alert
replaces the master caution tone; however, the master caution tone backs up the voice alert system and
provides a tone if the voice alert system malfunctions. FIRE, APU FIRE, L BLEED, and R BLEED
warning lights are not backed up by the master caution tone. Voice alert is the only audio warning for
these problems. With dual generator failure, the following voice alert warnings operate from battery
power: APU FIRE, ENGINE FIRE LEFT (RIGHT), and BLEED AIR LEFT (RIGHT). All voice alert
cautions, and the master caution tone are inoperative on battery power during dual generator failure.
Once a voice alert has been activated, it cannot be interrupted by a higher priority voice alert. All
voice alerts play until completed. The primary radar low altitude warning (WHOOP, WHOOP), is
repeated at the lowest priority until reset or disabled by the pilot. With MC OFP 20X AND UP OR
H3E AND UP, the BINGO voice alert is repeated every 30 seconds until the BINGO setting is
adjusted.
NOTE
With an MC1 failure (LOTs 21-24), the voice alert does not sound
when the aircraft descends below the altitude set by the primary low
altitude warning setting.
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ORIGINAL
A1-F18EA-NFM-000
CAUTION
VOICE ALERT
IFF 4
MODE 4 REPLY
FCS
FCS HOT
FLIGHT CONTROLS
FLIGHT COMPUTER HOT
L (R) OVRSPD
L (R) EGT HIGH
L (R) FLAMEOUT
L (R) OIL PR
L (R) STALL
L (R) ENG
L (R) ENG VIB
ENGINE LEFT (RIGHT)
FUEL LO
FUEL LOW
BINGO
BINGO
WARNING
VOICE ALERT
ALTITUDE
ALTITUDE
L (R) BLEED AIR
BLEED AIR LEFT (RIGHT)
L (R) FIRE
ENGINE FIRE LEFT (RIGHT)
APU FIRE
APU FIRE
2.17.4 Terrain Awareness Warning System (TAWS) (MC OFP 18EA and H2E AND UP). The terrain
awareness warning system alerts the aircrew of a controlled flight into terrain (CFIT) condition during
all mission phases. The system operates any time that the navigation mission computer (MC1) and
TAMMAC digital mapping set (DMS) are functional. TAWS functions as a safety backup system and
not as a performance aid. TAWS has been designed to eliminate false warnings, minimize nuisance
warnings, and generate consistent aircrew response in all aircraft master modes. Five possible voice
warnings are provided to indicate the correct initial response to an impending CFIT condition, and a
visual cue is provided to indicate the recovery direction of pull, or in some instances, to command an
increase in turn rate. All TAWS warnings should be treated as though an imminent flight into terrain
condition exists. Pilot response to a TAWS warning should be instinctive and immediate.
TAWS uses data from the following inputs: FCC, INS, RADALT, GPS, and digital terrain elevation
data (DTED). DTED resides in the DMS as part of TAMMAC and is used to provide the
forward-prediction capability that protects against flight into rising terrain. The TAWS option is
reached by pressing MENU-HSI-DATA-A/C as shown in figure 2-39. The TAWS option boxes
automatically at start-up.
When a DMS is not installed in the aircraft or is not operational, protection from CFIT events is
provided by the Ground Proximity Warning System (GPWS). BIT may be initiated on the DMS by
pressing the appropriate pushtile of the BIT display. The BIT can take up to 185 seconds to complete.
During the BIT, TAWS is not operational. Therefore, the GPWS algorithm is used to determine the
presence of possible CFIT events. There is no capability for pilot selection of GPWS if DMS is
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-39. HSI-DATA-A/C Controls
operational. The GPWS algorithm runs continuously with outputs being overwritten if TAWS is
operational. This prevents erroneous values during an unexpected transition from TAWS to GPWS.
2.17.4.1 TAWS Modes. TAWS has two operational modes: TAWS-with-DTED, and TAWS- withoutDTED. These modes switch automatically depending upon the available sensor data and flight phase.
When the aircraft position (latitude and longitude) is accurately known and DTED for the local area
has been loaded onto the DMS (during the theater load process), TAWS is in the TAWS-with-DTED
mode and provides protection against varying terrain ahead of the aircraft. When the aircraft position
is not accurately known, DTED for the local area is unavailable, or TAWS determines the aircraft is
in a landing phase, TAWS transitions to the TAWS-without-DTED mode and provides protection
against flight into level or descending terrain as GPWS does.
When operating over the ocean, DTED does not exist and TAWS will be in the TAWS-withoutDTED mode. However, there is no degradation in protection because the ocean is relatively flat. As the
aircraft approaches the coast or islands, DTED may be available (depending upon the theater load)
and TAWS will automatically switch back to the TAWS-with-DTED mode.
Operation of TAWS in the TAWS-without-DTED mode is still an improvement over GPWS as
TAWS incorporates a more robust performance model and additional input sensor redundancy.
2.17.4.2 TAWS Operation. TAWS incorporates signal processing that determines a best estimate of
aircraft position and altitude (AGL and MSL). TAWS protection algorithm continuously computes
two recovery trajectories: Vertical Recovery Trajectory (VRT) and Oblique Recovery Trajectory
(ORT). VRT is the standard GPWS-like recovery: roll to wings-level, if needed, and pull to recover.
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ORT assumes that you maintain the current bank angle and pull to recover (increase turn rate). Both
computed trajectories include the following assumptions:
a. Pilot Response Time is the time from issuance of a TAWS warning to the time that the pilot
actually initiates recovery. Pilot Response Time is set at 1.3 seconds.
b. Roll Recovery Phase is the time necessary to roll the aircraft to near wings-level. This assumes
at least ½ lateral stick will be used for bank angles less than 70° and at least ¾ lateral stick
will be used for bank angles ≥ 70°.
c. G-Onset Phase is the time required to pull to the target recovery g. The target recovery g is
80% of the instantaneous g available, or 5g, whichever is less. The g-onset phase assumes that
rapid aft stick motion will be used (full deflection within ¾ second). In addition, TAWS
assumes that throttles will be moved to MAX if below corner speed and to IDLE if above
corner speed.
d. Dive Recovery Phase is the remainder of the trajectory until terrain clearance is achieved.
TAWS assumes a terrain clearance of 50 ft.
When TAWS senses that the aircraft is in the landing configuration, the recovery assumptions must
change since the desire is to land. TAWS defines the landing phase as below 500 ft AGL, less than 200
KCAS, landing gear down and locked, and more than one minute since a waveoff or takeoff. In the
landing phase, TAWS protects against landings of greater than the structural limit of the landing gear
(1584 fpm). To allow this, TAWS switches to TAWS-without-DTED and provides a warning when the
landing is predicted to exceed the structural limit of the landing gear.
TAWS provides protection against gear-up landings. When the aircraft is below 200 KCAS, below
150 ft AGL, more than one minute since waveoff or takeoff, and the landing gear is not down and
locked, a TAWS warning is provided.
2.17.4.3 TAWS Warnings. TAWS provides clear, unambiguous, and directive aural and visual cues
to the aircrew. Aural warnings provide the aircrew with a wake-up call and correct initial response
while visual warnings provide the aircrew with correct follow-on recovery information.
2.17.4.3.1 Voice Warnings. TAWS uses the ACI to provide aural cues to the aircrew. The aural cues
are distinct from any other cues that the aircrew may receive. The TAWS voice alert warnings are:
″Roll−Left...Roll−Left″, ″Roll−Right.....Roll−Right″, ″Pull−Up...Pull−Up″, ″Power...Power″, and
″Check Gear″. Each of these warnings is issued at a level 3−6 dB above the present voice alerts. The
TAWS voice warnings provide a wake−up call to the aircrew and indicate the most appropriate initial
response for the given aircraft state, not necessarily the only required response. The aural cue repeats
until the warning condition is cleared. TAWS aural warnings have priority over all current aural tones.
Note that TAWS requires a -1018 or greater ACI load (or MIDS equivalent) which is capable of
generating the ″Roll Left/Right″ warnings. Earlier ACI loads were only capable of generating ″Roll
Out″ warnings and are not desired for use with TAWS.
A ″Roll Right...Roll Right″ warning is issued when a roll to the right is the correct initial response.
A ″Roll Left...Roll Left″ warning is issued when a roll to the left is the correct initial response.
A ″Power...Power″ warning is issued when the roll requirement conditions have not been met and
adding power is the correct initial response. This occurs when the aircraft is below 200 KCAS, the AOA
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Figure 2-40. TAWS HUD Visual Recovery Warning - Pull Up (VRT)
is above 8.5° for PA configuration (or 18° AOA for up and away configuration) and the throttle is not
already at MAX. The correct response to this warning is to select MAX afterburner.
A ″Pull Up...Pull Up″ warning is issued when the above conditions have not been met and pulling up
is the correct response or when the ORT is the recovery trajectory.
When a warning is given to protect against a gear-up landing, the following aural cues may be heard:
″Pull Up...Pull Up″ followed two seconds later by ″Check Gear″ when the gear handle is in the UP
position and a gear-up landing condition has been assessed (repeated every 4 seconds).
″Check Gear″ repeated every 8 seconds when the gear handle is down and a gear up landing condition
has been assessed.
2.17.4.3.2 Visual Warnings. A visual recovery arrow is provided in the center of the HUD and HUD
format on the DDI. The recovery arrow indicates the direction of recovery. The visual warning is
displayed when a CFIT condition is present and is removed when the CFIT condition is cleared.
TAWS visual recovery cues are designed to be used in conjunction with TAWS voice warnings.
There are several voice warning/visual recovery cue combinations. When the arrow points UP in the
HUD (i.e., along the lift vector), a longitudinal pull is the correct response and an aural ″Pull Up...Pull
Up″ is heard. This is a VRT recovery if the aircraft is close to wings level, or it is an ORT (increased
turn rate) recovery if the aircraft is banked such that the TAWS algorithm assessed that an increased
turn rate would provide the quickest recovery from an impending CFIT condition. Figures 2-40 and
2-41 depict these two situations. Both situations require a longitudinal pull as the correct response,
however, the first case (VRT) depicts a dive recovery while the second case (ORT) depicts a recovery
requiring an increase in turn rate by increasing g when already in an established angle of bank.
When the arrow points anywhere other than UP in the HUD (i.e., not along the lift vector, but
perpendicular to the horizon), it may be accompanied by either a ″Roll Left (Right)″...″Roll Left
(Right)″ or ″Pull Up...Pull Up″ voice warning. The voice warning indicates the correct initial response,
then the aircrew should roll or pull as required to place or maintain the TAWS recovery arrow straight
up in the HUD (i.e., along the lift vector). For example, if a ″Roll Left...Roll Left″ voice warning is
issued with an accompanying HUD recovery arrow displayed in the HUD that is perpendicular to the
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Figure 2-41. TAWS HUD Visual Recovery Warning - Pull Up (ORT)
horizon, the correct response is to roll left to align the lift vector with the HUD recovery arrow and then
perform a dive recovery. If a ″Pull Up...Pull Up″ voice warning is issued with an accompanying HUD
recovery arrow displayed in the HUD that is perpendicular to the horizon, the correct response is to
apply g along the current lift vector and then, referencing the HUD recovery arrow, roll to align the lift
vector with the HUD recovery arrow and perform a dive recovery. Figure 2-41 depicts a situation in
which a Roll or Pull Up aural warning could be issued. If a ″Roll Right...Roll Right″ aural warning was
issued, a roll to the right would be the correct initial response and then a dive recovery would be
continued with a longitudinal pull. If a ″Pull Up...Pull Up″ aural warning was issued, a longitudinal pull
would be the correct initial response and then a roll to the right to align the lift vector with the HUD
recovery arrow followed by a longitudinal pull for a dive recovery would be the follow-on recovery
procedure.
2.17.4.4 ACI Configuration Check. There are different ACI configurations in the F/A-18 aircraft.
Two TAWS software loads exist: -1016 and -1018. The -1016 ACIs can command these four aural cues:
″Roll Out...Roll Out″, ″Pull Up...Pull Up″, ″Check Gear″, and ″Power...Power″. The -1018 and greater
ACIs added logic for replacing the ″Roll Out...Roll Out″ aural cue with ″Roll Left...Roll Left″ and ″Roll
Right...Roll Right″. Due to the variety of possible combinations, on cold start power-up, the MC
commands a ″Roll Left...Roll Left″ to the ACI to determine if it is in the -1018 configuration. If there
is no response, the MC commands a ″Roll Out...Roll Out″ to the ACI to determine if the ACI is in the
-1016 configuration. If no response is received the second time the MC attempts to command a ″Roll
Out...Roll Out″, the TAWS/GPWS voice warnings will not be heard, however the visual arrow will still
be present when a warning is issued.
2.17.5 GPWS - Ground Proximity Warning System. GPWS is designed to backup the pilot by
providing an alert of impending controlled flight into terrain (CFIT). GPWS provides warnings of
potentially unsafe maneuvering flight conditions such as excessive bank angles, excessive sink rates,
gear up landings, floor altitude violations, and altitude loss during recovery. The system is operational
as long as MC1, radar altimeter, and air data systems are ON and functional. The GPWS algorithm
operates in the background of the OFP with no cockpit indications until an actual CFIT warning is
required. The system provides distinctive aural and visual warning cues only, to alert and direct
recovery from an impending CFIT condition. The pilot maintains full control of the aircraft for
recovery.
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Figure 2-42. TAWS HUD Visual Recovery Warning (VRT)
2.17.5.1 GPWS Sensors/Modes. GPWS is a look-down system with no forward-looking capability.
GPWS uses the radar altimeter as the primary source of terrain clearance information and the FCC air
data function, GPS, and INS as backup altitude sources when radar altitude is invalid. Radar altitude
is considered invalid by GPWS above 4,950 feet AGL or at a pitch or angle of bank greater than 50°.
With valid radar altitude data, GPWS calculates terrain slope from inputs from the INS and the radar
altimeter. Over descending terrain, GPWS assumes the terrain descends indefinitely (until the system
senses a change in terrain slope). This mechanization allows for maximum protection while minimizing
nuisance warnings.
For the first 5 seconds after radar altitude becomes invalid (as indicated by a flashing ″B″ in the
HUD or RALT Xd out on the UFCD), GPWS provides no CFIT protection. After 5 seconds, the system
enters ″COAST″ mode for a period of up to 2 minutes. While in COAST mode, GPWS calculates an
estimate of the aircraft current height above terrain. COAST mode can only be enabled while the
aircraft is not transonic and was over flat terrain (defined as slope less than 2°). CFIT warnings can still
be generated while in COAST mode. If the aircraft was transonic or was not over flat terrain when
radar altitude data was lost, GPWS transitions into the BYPASS mode. In the BYPASS mode, no
CFIT warnings are generated. Full protection is resumed from both modes when valid radar altitude
data is restored.
2.17.5.2 Altitude Required For Recovery Calculations. GPWS calculations for altitude required for
recovery include the loss of altitude due to persistency timers, pilot reaction time, time to roll wings
level, target g-onset rate, and steady state dive recovery time. GPWS pilot reaction time varies
depending on flight conditions but is a minimum of 0.5 second in the GPWS LAT envelope (±30° AOB,
0 to 30° dive, 450 to 560 KCAS). Pilot reaction time is reduced in the GPWS LAT envelope, where pilot
situational awareness is typically good, in order to reduce false warnings. Time to roll wings level is
based on a ½ to ¾ lateral stick displacement roll at 1g. Target g-onset rate is 80% of the available
g-onset rate up to (1) 5g/sec (less than 400 KCAS or greater than 30° AOB) or (2) 6g/sec (greater than
400 KCAS and less than 30° AOB). Steady state dive recovery time is based on a target sustained-g of
80% of g-available up to (1) 5g (less than 400 KCAS or greater than 30° AOB) or (2) 6g (greater than
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400 KCAS and less than 30° AOB). Regardless of which category applies, these g-onset
rates and sustained-g levels require an aggressive pilot response.
2.17.5.3 CFIT Protection Provided.
Above 150 feet AGL Above 150 feet AGL, GPWS continuously calculates the altitude required to recover. A CFIT
warning is issued if the altitude required to recover plus a variable safety buffer and an added terrain
clearance altitude is greater than the current altitude above terrain. The terrain clearance altitude
varies between 30, 50, and 90 feet, depending on flight conditions.
Below 150 feet AGL Below 150 feet AGL, GPWS transitions to provide warnings of CFIT conditions related to takeoff
and landing. These warnings are based on (1) the time since a WoffW transition (takeoff or T&G) or
a waveoff and then (2) a combination of landing gear position, airspeed, altitude, and sink rate. GPWS
defines a waveoff as 1000 fpm rate of climb for more than 5 seconds while below both 500 feet AGL and
200 KCAS. If the following sets of conditions are valid for greater than 0.3 seconds when the aircraft
altitude is less than 150 feet, a CFIT warning is provided. The CFIT warning is cancelled when the
condition no longer exists for 0.3 seconds.
1. Less than 60 seconds after WoffW or a waveoff:
a. Floor Altitude - less than 90 feet AGL and greater than 250 KCAS.
b. Takeoff Sink Rate - less than 150 feet AGL, less than 250 KCAS, greater than 300 fpm sink.
2. More than 60 seconds after WoffW or a waveoff:
a. Floor Altitude - less than 90 feet AGL and greater than 200 KCAS.
b. Check Gear - less than 150 feet AGL, less than 200 KCAS, descending, and landing gear not
down.
c. Landing Sink Rate - less than 150 feet AGL, less than 200 KCAS, landing gear down, and an
excessive sink rate. The allowable sink schedule varies from a maximum of 2,040 fpm to a
minimum of 1,488 fpm based on altitude and GW.
d. Bank Angle - less than 150 feet AGL, less than 200 KCAS, greater than 45° AOB for one
second.
Below 150 feet AGL, GPWS does not directly account for the recovery
capabilities of the aircraft. Therefore, recovery may not be possible
following a warning under extreme flight conditions.
2.17.5.4 GPWS Warning Cues. GPWS provides distinctive, clear, unambiguous and directive visual
and aural cues to the aircrew for each potential CFIT condition.
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Figure 2-43. GPWS HUD Roll Warning Cues
2.17.5.4.1 GPWS HUD Recovery Arrow. The GPWS visual warning cue is a steady arrow located in
the center of the HUD. See figure 2-43. The HUD recovery arrow is always perpendicular to the horizon
and points in the direction of pull required for recovery. The HUD recovery cue is displayed
simultaneously with all voice warnings except CHECK GEAR. The HUD recovery arrow remains
displayed until GPWS calculates that a CFIT condition no longer exists.
2.17.5.4.2 GPWS Voice Commands. Refer to figure 2-44 for GPWS aural warning cues.
Voice commands automatically transition to the appropriate command for the current stage of
recovery (e.g., ROLL OUT transitions to PULL UP when AOB becomes less than 45°). The voice
commands are terminated when the appropriate recovery maneuver is initiated (e.g., a PULL UP is
initiated within 0.5 g of the GPWS calculated target-g).
• In addition to following the voice commands, additional pilot action
may be required to avoid an unrecoverable situation (e.g., aft stick
with a POWER call or power addition/subtraction with a PULL UP
call.)
• GPWS voice alerts are delayed if other voice alerts are currently being
transmitted.
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GPWS Warning
Condition
Aural Cue
Repetition Rate
Excessive bank angle
″ROLL LEFT (RIGHT), ROLL LEFT (RIGHT)″
2 seconds
Excessive take-off
sink rate
″POWER, POWER″
2 seconds
Excessive landing
sink rate
″POWER, POWER″
2 seconds
Gear-up landing
″CHECK GEAR″
8 seconds
ALDR or floor
altitude
″POWER, POWER″ for airspeed <210 KCAS and
AOB ≤45°
2 seconds
″ROLL LEFT (RIGHT), ROLL LEFT (RIGHT)″
for AOB >45°
2 seconds
″PULL UP, PULL UP″ for all other flight
conditions
2 seconds
Figure 2-44. GPWS Aural Cues
2.17.5.5 Areas of Limited CFIT Protection. Areas where CFIT protection is considered limited are
as follows:
1. In the COAST mode (5 to 120 seconds outside the valid RALT envelope).
2. Over rising terrain of greater than 2° slope (GPWS is inhibited to prevent nuisance warnings).
3. Within the GPWS LAT envelope where allowable pilot reaction times have been reduced (±30°
AOB, 0 to 30° dive, 450 to 560 KCAS).
4. Below 150 feet AGL in the landing phase (less than 200 KCAS) where warnings are designed only
to prevent hard landings.
At certain high speed, high gross weight conditions, overriding the
g-limiter may be required for recovery from dives greater than 50° and
will likely be required for dives between 10 and 25°.
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2.17.5.6 Areas of No CFIT Protection. Areas of no protection are as follows:
1. Loss of air data or RALT, INS, or MC1 failed or off (non-AMCD aircraft), or either MC1 or MC2
failed or off (AMCD aircraft).
2. Less than 6 seconds after WonW.
3. Less than 5 seconds or greater than 120 seconds outside the valid RALT envelope.
4. Transonic flight (0.95 to Mach 1.04) outside the valid RALT envelope.
5. For 1.5 seconds after a break X is displayed.
6. After a waveoff until exceeding 1,000 fpm for 5 seconds.
7. Dives greater than 50° after 2 minutes above 5,000 feet AGL.
2.18 INSTRUMENTS
Refer to foldout section for cockpit instrument panel illustration. For instruments that are an
integral part of an aircraft system, refer to that system description in this section.
2.18.1 Standby Attitude Reference Indicator. The standby attitude reference indicator is a selfcontained electrically driven gyro-horizon type instrument. It is normally powered by the right 115
volts ac bus. If this power fails it is automatically powered by an inverter operating off the essential 28
volts dc bus. An OFF flag appears if both power sources fail or the gyro is caged. During caging the gyro
initially cages to 4° pitch and 0° roll regardless of aircraft attitude. After 3 to 5 minutes, the indicator
reads 0° pitch and 0° roll. Power should be applied for at least 1 minute before caging. The indicator
displays roll through 360°. Pitch display is limited by mechanical stops at approximately 90° climb and
80° dive. As the aircraft reaches either stop, the gyro tumbles 180° in roll. A needle and ball are at the
bottom of the instrument. A one needle width turn is 90° per minute.
2.18.2 Standby Airspeed Indicator. The standby airspeed indicator displays airspeed from 60 to 850
KIAS. It operates directly from left pitot and static pressure.
2.18.3 Standby Altimeter. The standby altimeter is a counter-pointer type. The counter drum
indicates altitude in thousands of feet from 00 to 99. The long pointer indicates altitude in 50-foot
increments with one full revolution each 1,000 feet. A knob and window permit setting the altimeter
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Figure 2-45. Angle of Attack Indexer
to the desired barometric setting. This setting is also used by the flight control computers. The standby
altimeter operates directly from left static pressure.
2.18.4 Standby Rate of Climb Indicator. The standby rate of climb indicator displays vertical speed
on a scale from 0 to ±6,000 fpm and operates directly from left static pressure.
2.18.5 Standby Magnetic Compass. A conventional aircraft magnetic compass is mounted on the
right windshield arch in the front cockpit and in the rear cockpit in Lots 21 thru 25.
2.18.6 Angle Of Attack Indexer. The angle of attack indexer is mounted to the left of the HUD. It
displays approach angle of attack (AOA) with lighted symbols; corresponding AOA indications are
shown on the HUD (see figure 2-45). The indexer operates with the landing gear down and locked and
weight off the gear. The lighted symbol(s) flash if the arresting hook is up and the hook bypass switch,
on the left vertical panel, is in CARRIER. The symbols will not flash with the arresting hook up and
the hook bypass switch in FIELD. The switch is solenoid held to FIELD and automatically goes to
CARRIER when the arresting hook is lowered or aircraft power is removed. The AOA indexer knob on
the HUD controls dimming of the symbols. All symbols light when the lights test switch on the interior
lights control panel is held to TEST.
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2.19 AVIONICS SUBSYSTEM
The avionics subsystem combines the integration and automation needed for operability with the
redundancy required to ensure flight safety and mission success. Key features of the system include
highly integrated controls and displays, inertial navigation set with carrier alignment capability, and
extensive built in test capability. The avionics subsystems operate under the control of two mission
computers with primary data transfer between the mission computers and the other avionics
equipment via the mux buses.
2.19.1 Mission Computer System. The mission computer system consists of two digital computers
(MC1 and MC2) which are high speed, stored program, programmable, general purpose computers
with core memory. Computer de-selection is made with the MC switch on the MC/HYD ISOL panel.
With non-AMCD aircraft, the two mission computers interconnect with the primary avionics
equipment on the avionics multiplex (mux) buses. MC1, referred to as the navigation computer,
performs processing for navigation, control/display management, aircraft built in test (BIT), status
monitoring operations and backup for MC2. MC2, referred to as the weapon delivery computer,
performs processing for air-to-air combat, air-to-ground attack, and tactical control/display.
There are six avionics mux bus channels with redundant paths (X and Y) for each channel.
The mission computer:
1. Computes and controls the data sent to the cockpit displays,
2. Computes missile launch and weapon release commands,
3. Provides mode control and options for various avionics systems,
4. Generates BIT initiate signals to and equipment operational status from various avionics
systems.
With AMCD aircraft, the front and rear DDIs are driven directly by the MC over a high speed
interface bus, not by avionics mux bus commands. The HUD is driven directly by redundant
connection to either MC. MC1 drives the front and rear LDDIs and HUD while MC2 drives the front
and rear RDDIs and HUD. When an MC is off or non-functional, the displays driven by that MC show
a green square in the center of the display. Each MC provides the same level of functionality in the
single MC backup mode of operation.
On AMCD aircraft, with both MCs inoperative and the left generator operative, the SDC provides
a limited HUD format on the front MPCD/UFCD, prevents the FADEC and ECS controller from going
into default mode operation, and provides left/right ATS cautions when necessary. See Chapter 25
Backup/Degraded Operations for a description of SDC Backup Mode.
AMCD II mission computers are used on aircraft with the aft cockpit 8 x 10 display installed. The
8 x 10 display receives its primary signal through MC2, and is inoperative with computers prior to
AMCDII. The AMCD II computer provides digital video color capability to the 8 x 10 display via the
Fiber Channel Network Switch (FCNS) and High Speed Video Network (HSVN).
The computers receive inputs for navigational data and steering command computations from the
inertial navigation system, electronic flight control system, multipurpose display group, TACAN, and
backup attitude and the navigation system. The computers control display symbology and information
presented to the pilot by the multipurpose display group.
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2.19.1.1 MC Switch. The MC switch, located on the outboard edge of the aft left console, is used to
manually turn OFF either of the two mission computers, MC1 or MC2.
1 OFF
Removes power to MC1.
NORM
Both MC1 and MC2 are powered with ac electrical power available.
2 OFF
Removes power to MC2.
On non-AMCD aircraft with WonW, both engines will spool up from ground idle to flight idle, if
power is removed from MC1 (MC1 switch in 1 OFF).
2.19.1.2 Mission Data Entry. Mission data (date and flight number) can be manually loaded into the
mission computer for data recorder documentation. Data is entered by performing the following:
1. On the DDI - Press MENU, MUMI, then ID.
2. On the UFCD - Enter Julian Date (DATE).
3. On the UFCD - Enter Flight Information (FLT).
Mission data can be manually loaded into the mission computer through the Memory Unit Mission
Initialization (MUMI) display or automatically loaded into the mission computer through the Data
Storage Set (DSS). The DSS consists of the Memory Unit (MU) and the Memory Unit Mount (MUM)
and provides memory storage for aircraft parameters, maintenance data, and avionics initialization
data. The DSS receives, stores, retrieves, and transmits data with the mission computer.
2.19.1.2.1 Mission Initialization. The MU provides the capability to load the following mission
initialization files: HARM, RADAR, MU ID, TACAN, WYPT/OAP, Combined Interrogator Transponder (CIT), Sequential Steering (S/S), data link/ID, Overlay Controlled Stores (OCS), bomb wind
data, and ALR-67 display resets. The S/S file can have a 15 point sequence consisting of Geographic
Reference Points (GEOREF), GPS waypoints, and almanac data initialization files. Loading is done at
aircraft power up or when MUX communication is lost for more than 1 second and regained. If MUX
communication is not regained, a MU LOAD caution is displayed and an AV MUX error message is
displayed on the MUMI display. Manual loading may be done using the MUMI display.
2.19.1.2.2 MUMI - Memory Unit Mission Initialization Format. The MUMI format (see figure 2-46)
is accessible from the SUPT MENU and with WonW provides a visual indication of mission
initialization files loaded from the MU. If the MU directory indicates that no user files are present, the
MU ID displays NO IDENT. When the MU directory indicates a user file is present, MC1 displays the
option. When the option is selected and the file is being read by MC1, the option is boxed. If the read
is successful, the file is loaded and the option is unboxed. When a file is present and errors have
resulted from reading the file, the following occurs:
1. The MU ID displays NO IDENT.
2. The applicable load error is displayed (HARM, RDR, TCN, WYPT, S/S, OCS, GPS WYPT, GPS
ALM, ALR 67, WIND, DL13, or CIT).
3. MC1 sends the appropriate maintenance code to the SDC.
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Figure 2-46. MUMI Display
4. If WonW, an MU LOAD caution is displayed on the DDI.
2.19.1.2.3 CRYPTO Switch. Setting the intercommunications amplifier control CRYPTO switch to
the ZERO position sends an erase signal to the MU. This causes the MU to erase all data stored
between predetermined memory locations.
2.19.1.2.4 Erase and Hold Data. The erase controller (EC) within MC1 provides the capability to
automatically or manually erase, or inhibit erasing, of classified data contained in the MU, SMS, MC1,
and MC2.
NOTE
With AMCD aircraft, during an ERASE, the DDI controlled by the
MC undergoing the erase will flash STANDBY, then briefly display a
green square, then flash STANDBY until the erase is complete.
When the EC determines classified mission initialization files have been read from the MU, the EC
classified data management system is activated. When activated, MC1:
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1. Displays the HOLD and ERASE options on the MUMI display.
2. Displays the CDATA advisory.
3. Sends applicable maintenance code(s) to the SDC.
2.19.1.2.5 Automatic Erase. The MU, SMS, MC1, and MC2 automatically erase classified data
when all of the following criteria are met.
1. Airspeed is less than 80 KCAS.
2. Left and right engine THA less than 29°.
3. Transition from WoffW to WonW.
4. Pilot does not select erase inhibit (HOLD) or MC SUSPEND options.
NOTE
Automatic erase can be inhibited by selecting the HOLD pushbutton
option.
5. HOLD boxed with MU displayed prevents automatic erase of the MU.
6. HOLD boxed with ALL displayed prevents automatic erase of all units (MU, armament
computer, MC1, and MC2).
The EC commands the MU and the armament computer, then MC1 and MC2 to erase. The MC
ERASE IN XX SEC countdown timer starts (60 seconds). During the countdown, an MC SUSPEND
pushbutton option is displayed. The MC SUSPEND option may be toggled between boxed (selected)
and unboxed (deselected). When the timer reaches zero and the MU and armament computer have
finished erasing, the decision to continue erasing the remainder of MC1 and MC2 depends on the MC
SUSPEND option being deselected (unboxed); when deselected, the remaining erase of MC2 and MC1
is completed.
NOTE
With AMCD aircraft, during an ERASE, the DDI controlled by the
MC undergoing the erase will flash STANDBY, then briefly display a
green square, then flash STANDBY until the erase is complete.
Automatic erase is also initiated by pilot ejection. The state of the HOLD options is ignored during
pilot ejection.
2.19.1.2.6 Manual Erase. Manual erase is a two pushbutton process and is initiated by pressing the
ERASE pushbutton on the MUMI display. When the ERASE pushbutton is pressed, the option to
proceed with the erasure (ERASE) and the option to cancel the erase (CNX) replaces the HOLD and
ERASE options. Selecting the second ERASE option initiates erasure. While erase is in progress,
ERASE is boxed and erasing proceeds the same as automatic erase. When erasing is complete, the
ERASE pushbutton unboxes. While erase is in progress one of the following is displayed on the MUMI
display:
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1. ERASING - erasing of unit is in progress.
2. COMPLETE - erasing of unit is complete.
3. FAILED - unit failed to erase.
NOTE
With AMCD aircraft, during an ERASE, the DDI controlled by the
MC undergoing the erase will flash STANDBY, then briefly display a
green square, then flash STANDBY until the erase is complete.
When erase fails, the MC1 retains the MUMI ERASE and HOLD pushbutton options and displays
the ERASE FAIL caution on the DDI. When erasing is complete, MC1 removes the ERASE, HOLD,
and MC SUSPEND pushbutton options from the MUMI display, removes the CDATA advisory from
the display, and resets the applicable maintenance code(s).
2.19.1.2.7 Backup Erase Controller. If MC1 fails, backup erase capability is provided by MC2 by
providing an ERASE option on the HSI format. When the ERASE pushbutton is depressed, the option
to proceed with the erasure (ERASE) and the option to cancel the erasure (CNX) are provided.
Selecting the second ERASE option initiates erasure. While ERASE is in progress, the ERASE option
is boxed; however, additional cuing is not provided. In backup mode, pilot ejection is the only
automatic erase provided. With AMCD aircraft, the MUMI display for the MC being erased disappears
while the erase is in progress.
2.19.2 Master Modes. There are three master modes of operation: navigation (NAV), air-to-air
(A/A), and air-to-ground (A/G). Controls, displays, and the avionics equipment operation are tailored
as a function of the master mode selected. The navigation master mode is entered automatically when
power is applied to the aircraft, when the air-to-air or air-to-ground modes are deselected, when the
landing gear is lowered, when the SPIN mode activates, or when the aircraft has WonW and the THA
is greater than 27°. The A/A master mode is entered either by pressing the A/A master mode button
alongside the left DDI or by selecting an A/A weapon with the A/A weapon select switch on the control
stick. The A/G master mode is selected by pressing the A/G master mode button. The selection is
performed by the stores management set (SMS), and the SMS identifies the selected master mode to
the mission computer.
2.19.2.1 Steering Information. The sources of steering information available in the NAV master
mode are waypoint, TACAN, instrument landing system, and data link. The data link modes available
in the NAV master mode are vector and automatic carrier landing. TACAN and waypoint steering are
mutually exclusive; selecting one automatically deselects the other. Data link, ILS, and TACAN (or
waypoint) steering can be provided simultaneously. The ACL mode is selectable only in the NAV
master mode and the vector mode is available in all master modes. Steering information is used by the
Automatic Flight Control System to provide coupled steering options.
2.19.3 Cockpit Controls and Displays. The cockpit controls and displays which are used for
navigation operation are on the multipurpose display group.
2.19.4 Multipurpose Display Group. The multipurpose display group consists of the right and left
digital display indicators (DDIs), the multipurpose color display (MPCD), 8 x 10 display, the digital
map set (DMS), the head-up display (HUD), the CRS (course) set switch, the up front control display
(UFCD) and the HDG/TK (heading/ground track) set switch. The multipurpose display group
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A1-F18EA-NFM-000
presents navigation, attack, and aircraft attitude displays to the pilot. The multipurpose display group
converts information received from the mission computer system to symbology for display on the DDIs,
the MPCD, the UFCD, and the HUD. The HUD camera records the outside world and HUD
symbology. The left and right DDIs, 8 x 10 display and the MPCD contain pushbuttons for display
selection and selection of various equipment operating modes. The UFCD is an active matrix liquid
crystal display with an infrared (IR) touchscreen for operator inputs. Refer to Part VII for the
operation of each component.
2.19.4.1 CRS Set Switch. The course set switch, located on the main instrument panel on the video
record panel, manually sets the desired course on the HSI display.
2.19.4.2 HDG/TK Set Switch. The heading/ground track set switch located on the main instrument
panel on the video record panel, manually sets the heading marker on the desired heading/ground track
on the HSI display.
2.19.4.3 DDIs. The left and right DDI (LDDI/RDDI) are physically and functionally interchangeable, giving the ability to display desired information on either indicator. With non-AMCD aircraft, the
DDIs also provide symbology to the HUD. The left indicator is used primarily for stores status, built
in test status, engine monitor, caution, and advisory displays. The right indicator is normally used for
radar and weapon video displays.
(Non-AMCD aircraft) The DDIs are NVG compatible and display three colors (red, yellow, and
green) for stroke information. A monochrome version of the digital map can be selected on the left DDI.
Either DDI can provide raster generation for the HUD.
NOTE
It is possible that a transient condition may cause the displays to
blank or provide an erroneous display on the left or right DDI, or
HUD. The problem may be cleared by manually cycling the power to
the left or right DDI.
(AMCD Aircraft) The DDIs have full color capability in all display modes and are NVG compatible.
2.19.4.3.1 Brightness Selector Knob - Non-AMCD Aircraft. Placing this rotary knob to OFF
prevents the DDI from operating. Placing the knob to NIGHT provides a lower brightness control
range and no automatic contrast control. Turning the knob to DAY provides a higher brightness
control range.
2.19.4.3.2 BRT Control Knob - Non-AMCD Aircraft. This knob varies the intensity of the DDI
presentation.
2.19.4.3.3 Brightness Knob - AMCD Aircraft. Placing this rotary knob to OFF prevents the DDI
from operating. When turned on, rotating the knob clockwise increases display brightness, while
rotating the knob counterclockwise decreases display brightness.
2.19.4.3.4 GAIN Control - AMCD Aircraft. This three-position rocker switch affects the existing
gray-scale, shifting the background brightness up or down, with no impact on displayed symbology.
The center position is off, while the up arrow increases background brightness and the down arrow
decreases background brightness. The up and down arrows have momentary and scroll functionality,
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A1-F18EA-NFM-000
depending on how long they are held. Current values are temporarily displayed when switches are
pressed.
2.19.4.3.5 CONT Control. This knob (non-AMCD aircraft), or three-position rocker switch (AMCD
aircraft), varies the contrast between symbology and the dark background on any level of brightness.
With AMCD aircraft, when the contrast rocker is pressed, a graphical and numeric representation of
the current setting is momentarily displayed on the DDI.
2.19.4.3.6 DDI Pushbuttons. There are 20 pushbuttons on each DDI which are used to select the
function and the mode for proper indicator display.
2.19.4.3.7 MENU Formats. There are two MENU formats (figure 2-47), TAC (tactical) and SUPT
(support) through which display selections can be made. The two menu formats can appear on either
the DDI, the UFCD, 8 x 10 display or MPCD.
The TAC MENU is indicated by the word TAC appearing just above the MENU option. The TAC
MENU is selected by actuating the MENU option on any format other than the TAC MENU format.
The SUPT MENU is indicated by the word SUPT appearing just above the MENU option. The SUPT
MENU is selected by actuating the MENU option on the TAC MENU. With MC OFP H2E AND UP,
the current time of day, consisting of two minute digits and two second digits appear in place of the
MENU legend (cold start default). Time of day is also turned on/off with the TISM option on the
Engine Format. MENU pushbutton functionality is unchanged.
The TAC MENU allows selection of the following formats: AZ/EL, HUD, RDR ATTK, STORES,
FLIR, NFLIR, DL13, HARM, A/G WPN, JSOW, SA, IMAGE, CAS, EW, and TGT DATA. The SUPT
MENU allows selection of the following formats: ADI, HSI, HMD, NETS, GPS, MIDS, ROE/IFF
PROG, BIT, CTT, MUMI, CHKLST, ENG, FCS, UFC BU, FPAS, and FUEL.
Some of the options on the MENU formats are conditional and are not always displayed. NFLR,
FLIR, LST, and CAM are listed only if the equipment is communicating with the mission computer.
HARM DSPLY is displayed when HARM is on board and CLC communicating. A/G missile display
(WEDL DSPLY, MAV DSPLY, etc.) is displayed when the MC has determined from the armament
control processor set that a weapon station has been selected which contains one of these weapons.
In Non-AMCD Aircraft, if the navigation computer (MC1) is not on, only HSI is displayed on the
SUPT menu. If the weapon delivery computer (MC2) is not on, the SUPT menu remains unchanged
and the TAC menu does not display the STORES, FLIR/LST/CAM or NFLR, AWW-13, HARM, or
A/G displays. If both mission computers are off or not communicating with the display, the DDI,
MPCD, and UFCD display only a flashing STANDBY in the center of the screen.
In AMCD Aircraft, if either mission computer is off or failed (MC Backup Mode), the options
available on the displays are the same regardless of which computer is failed. MC1 is no longer referred
to as the navigation computer and MC2 is no longer referred to as the weapons computer. If MC1 is
failed or not communicating, the left DDIs in both cockpits will either flash STANDBY or have a green
square present in the middle of the displays. If MC2 is failed or not communicating with the display,
the right DDIs in both cockpits will either flash STANDBY or have a green square present in the
middle of the displays. The 8 x 10 display will not be capable of displaying format symbology or video.
The following options are available in MC Backup mode: HUD, EW, FCS, MIDS, MUMI, ENG, UFC
BU, FUEL, ADI, and HSI. If A/A master mode is selected, the RDR ATTK, SA, and AZ/EL formats
are available as well. No A/G displays are supported in MC Backup mode. If both mission computers
are off or failed, a backup HUD format (driven by the SDC) is displayed on the UFCD and MPCD in
the front cockpit.
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A1-F18EA-NFM-000
Figure 2-47. MENU Format
2.19.4.3.8 Electronic Attitude Display Indicator (EADI). The electronic attitude display indicator is
available for display on the left or right DDI as an alternative to the attitude display on the HUD
(figure 2-48). The EADI display is selected by selecting the ADI option on the SUPT MENU. The pitch
ladder is displayed in 10° increments. A small circle is displayed on the ball to represent the zenith and
a circle with an inscribed cross is displayed to represent the nadir. A turn indicator which displays FCS
yaw rate is provided below the ball. A standard rate turn (3°/second) is indicated when the lower box
is displaced so that it is under one of the end boxes.
Selecting the INS or STBY options at the bottom of the display determines the source of attitude
information used to generate the display. Upon power-up with WonW, the EADI attitude initializes to
STBY (STBY boxed), thus using the standby attitude reference indicator for attitude source
information. Selecting the INS option (INS boxed) uses attitude information provided by the INS.
Selection of INS or STBY on the EADI does not change the source of attitude data for the HUD.
Airspeed and altitude are displayed in boxes at the top left and right. Altitude source is displayed
to the right of the altitude box and the vertical velocity is displayed above the altitude box. When ILS
is selected the deviation needles are displayed in reference to the waterline symbol. The ILS needles
are yellow when COLOR is selected on the Attack display.
2.19.4.3.9 Shaped Attitude Display Indicator. When the ADI option is selected on the SUPT menu,
a shaped attitude display indicator is available on the MPCD, UFCD, or DDI and 8 x 10 display
(AMCD aircraft). The pitch ladder is displayed in 10° increments, with ±30° in pitch being displayed
when the gear is up. With the gear down, the pitch ladder is displayed in 5° increments with ±15° in
pitch being displayed. A turn indicator is provided below the ball. A standard rate turn (3°/second) is
indicated when the lower box is displaced so that it is under one of the end boxes.
On the MPCD and 8 x 10 display, the shaped ADI is black above the horizon to represent sky, and
green below the horizon to represent the ground. On the UFCD, the shaped ADI is unshaded above the
horizon (sky) and shaded green below the horizon (ground).
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-48. Electronic Attitude Display Indicator
Selecting the INS or STBY options at the bottom of the display determines the source of attitude
information used to generate the display. Upon power-up with WonW, the ADI attitude initializes to
STBY (STBY boxed), thus using the standby attitude reference indicator for attitude source
information. Selecting the INS option (INS boxed) uses attitude information provided by the INS.
Selection of INS or STBY on the ADI does not change the source of attitude data for the HUD.
Airspeed and altitude are displayed in boxes at the top, altitude source is displayed to the right of
the altitude box and vertical velocity is displayed above the altitude box. When ILS is selected the
deviation needles are displayed in reference to the waterline symbol. The ILS needles are yellow when
COLOR is selected on the attack display.
2.19.4.4 MPCD - Multipurpose Color Display. The MPCD is an NVG compatible digital display
capable of providing any MENU selectable format except the video on the A/G radar display (figure
2-50). The MPCD drives itself using information received on the MUX.
The BRT knob controls the overall video and symbology brightness and also acts as the power
control for the MPCD and UFCD; a detented OFF position at the extreme counterclockwise position.
The CONT knob adjusts the video contrast of the MPCD display. The day/night mode is controlled
by the day/night/NVG switch on the interior lights panel.
2.19.4.4.1 Standby Indication. A flashing STANDBY indication is provided in the center of the
display on initial power-up and when there is invalid mux communication with the MC. If the
STANDBY condition persists for a few seconds, cycling power to the MPCD may return the MPCD
and UFCD to normal operation.
2.19.4.4.2 Brightness and Contrast Controls. The brightness control is used to adjust the overall
brightness of the MPCD display surface, allowing symbology and video to be adjusted. The full
counterclockwise position of the brightness control knob shuts off power to both the MPCD and the
UFCD. Selecting ON powers both the MPCD and the UFCD.
The contrast control is used to adjust video contrast.
2.19.4.4.3 Symbology Rocker Switch. The symbology (SYM) brightness rocker switch is used to
adjust the brightness of the symbology without affecting the brightness of the MPCD video.
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ORIGINAL
A1-F18EA-NFM-000
Symbology change feedback in provided in the left center of the display by a digit from 0 to 9. This
feedback is provided while the rocker switch is pressed and for 5 seconds after the rocker switch is
released.
2.19.4.4.4 DDI Formats on MPCD. Every display format is available on the L or R DDI. A/G radar
symbology and monochrome video is available on the MPCD. A/G video is not available.
2.19.4.4.5 MPCD Formats. With non-AMCD aircraft, if two different map formats are provided in
the front and rear cockpits, only one displays the map; the other provides only symbology. The priority
of the format which displays the map is HSI, followed by SA.
With AMCD aircraft, if two different formats are requesting MAP video underlay, only one displays
the video underlay; the other provides only symbology. MAP video display priority is determined by
two things: 1) which display surface is requesting the MAP underlay and 2) which format is requesting
the MAP underlay. Highest display surface priority is given to the MPCD followed by the DDIs. The
SA format is given the highest priority, followed by the HSI format, for formats requesting the MAP
underlay. Once the MAP is displayed for any format, (SA, HSI), any display surface with that same
format selected will also have the MAP underlay displayed. The MAP underlay is only supported on
the SA or HSI formats.
2.19.4.4.5A 8 x 10 Display. On aircraft LOT 26 AND UP, the 8 x 10 display replaces the aft MPCD.
The 8 x 10 display has a full color Active Matrix Liquid Crystal Display surface and 30 pushbuttons
used for operator inputs. Only 20 buttons are functional, and they duplicate the MPCD/DDI menu
layouts. The top row outboard buttons and top four buttons on each side are not used. The brightness
control switch adjusts the video brightness. The full counterclockwise position, a detented OFF
position, removes power from the 8 x 10 display and the aft cockpit UFCD. Two rocker switches
provide Gain and Contrast control. A numeric value displayed near the switches is used to adjust the
gain and/or contrast to a desired level. Pushbuttons are backlit, controlled by the INST PNL knob on
the aft cockpit interior lights panel. The following formats and related sublevel formats are available:
TAC Menu, SUPT Menu, HSI, EW, SA, TFLIR and BIT. Figure 2-49 shows symbology placement and
map coverage for the 8 x 10 display. See figure FO - 4, foldout section for aft cockpit arrangement with
the 8 x 10 display installed.
The 8 x 10 display is connected directly to MC2. The computer recognizes whether an 8 x 10 display
or an MPCD is installed and uses the corresponding software that matches the configuration at
startup. This affects the UFCD controls and functions in addition to the 8 x 10 display. When MC2 is
offline or during initialization, the 8 x 10 display does not show a format or process bezel inputs. A
flashing STANDBY indication is provided in the center of the 8 x 10 display when MC2 is not
communicating properly with the 8 x 10 display. A standby display pattern (approximately a 1- by
1-inch green square) is present when the Image Transfer Bus (ITB) has lost symbology, video, or a
combination, and the selected format is affected. When interface between MC2 and 8 x 10 display is
lost, only symbology is displayed (via ITB). There is no pushbutton feedback, video control, or BIT
reporting when interface is degraded. The day/night mode of the 8 x 10 display (and rear UFCD) is set
via the DAY/NITE/NVG switch on the forward cockpit interior lights panel through MC2.
In the Situational Awareness (SA) format, the cursor (captains bars) may be slewed over the entire
8 x 10 display surface. Because this area is larger than the DDIs/MPCD, if another display is showing
the same format and the cursor on the 8 x 10 display is slewed beyond the center area, the cursor on
the DDI/MPCD is limited. When this happens, the cursor on the DDI/MPCD displays flashes
continuously. When a SA format has DC priority assigned, the initial cursor position is determined by
the cockpit that initiates the slewing action.
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-49. 8 x 10 Display
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ORIGINAL
A1-F18EA-NFM-000
2.19.4.5 HSI Display Symbology. Basic HSI symbology such as the compass rose, ground track
pointer, lubber line (for magnetic heading), true airspeed readout, ADF bearing pointer, groundspeed
readout, and aircraft symbol are not described. These symbols are shown in figure 2-50.
Display format lines, just below the top row of pushbutton labels, indicate the formats being
displayed on the left, center, right and UFCD displays in both the cockpit (left side of HSI display) and
rear cockpit (right side of HSI display).
Radar target and GEOREF symbols are described in A1-F18EA-TAC-Series. The following
paragraphs describe unique navigation symbology. Refer to part VII for a description of how these
symbols are integrated with the navigation system:
1. Waypoint/OAP data. Data for the current steer to waypoint/OAP is displayed on the upper right
corner of the HSI. Waypoint/OAP data consists of bearing, range, and TTG (time-to-go) up to
8:59:59 based on distance and ground speed. When a waypoint/OAP or offset to the OAP is
designated (becomes a target), this data relates to the target. When a waypoint is a waypoint that
was transferred from the GPS, an ID code is displayed under the waypoint data.
2. TACAN data. TACAN data is displayed on the upper left corner of the HSI. TACAN data
consists of bearing, range (slant range), TTG (based on distance and present ground speed), and the
station identifier.
3. Waypoint/OAP symbology. Waypoint/OAP symbology consists of a waypoint/OAP symbol and a
bearing pointer and tail. The waypoint/OAP symbol indicates the position of the selected
waypoint/OAP relative to the aircraft symbol. The waypoint/OAP bearing pointer and tail are
displayed inside the compass rose and indicate bearing to the selected waypoint/OAP. Waypoint/
OAP symbology is displayed whether or not waypoint/OAP steering is selected. When the selected
waypoint/OAP is outside the HSI range scale, the waypoint/OAP symbol does not appear, but the
bearing pointer and tail appears. When a waypoint/OAP is designated, the waypoint/OAP symbol
and circle inside the pointer change to a diamond shape. The offset symbol appears when steering
is to an OAP. The offset symbol indicates the position of the offset relative to the OAP.
4. TACAN symbology. TACAN symbology consists of a TACAN symbol, and TACAN bearing
pointer and tail. The TACAN symbol indicates the position of the TACAN station relative to the
aircraft symbol. The TACAN bearing pointer and tail are located outside of the compass rose and
indicate bearing to the TACAN station. When the TACAN station is outside the HSI range scale, the
TACAN symbol does not appear but the bearing pointer and tail appear. When TACAN range
becomes invalid the TACAN symbol is not displayed.
NOTE
TACAN symbology displayed inside of the compass rose is filtered to
prevent excessive movement of TACAN symbols due to RF
interference. However, the ″fly−to″ needle displayed in the HUD with
TACAN steering selected is not filtered and represents the raw data
received by the TACAN. As a result, for brief periods of time, the
HUD and HSI may display conflicting information regarding aircraft
position with respect to the selected TACAN course line. Aircrew
should use HUD displayed TACAN information when conducting
TACAN approaches.
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-50. MPCD Controls and HSI Symbology
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ORIGINAL
A1-F18EA-NFM-000
5. Heading select/ground select marker and readout. The heading select/ground select marker is
maneuvered along the periphery of the compass rose using the HDG/TK switch. The digital readout
of the selected heading is located on the lower left corner of the HSI. The heading select/ground
select marker and digital readout are part of the heading select/ground track select mode of the
autopilot.
Course line arrow and readout. The course line arrow indicates the selected course to the
waypoint/OAP or TACAN station. The course is selected using the CRS switch. The digital readout
of the selected course is displayed on the lower right corner of the HSI. The course line arrow is not
displayed when TACAN range is invalid.
6. TDC assignment symbol. The TDC assignment diamond is displayed on the upper right corner
of the HSI. This symbol indicates that the TDC is assigned to the HSI. The TDC assignment
diamond indicates control is assigned to both cockpits. Other symbols indicate cockpit (d) or rear
cockpit (e) TDC control and SLEW control. SLEW is done by actuating the sensor control switch
AFT while in the NAV or A/G master mode. The word SLEW is displayed in the TDC assignment
diamond position when the SLEW option is active.
7. Coupled steering symbology. CPL and the source of the steering information is displayed on
either side of the aircraft symbol in the center of the HSI display whenever the flight control system
is coupled in azimuth to a steering source. Steering source can be WYPT, TCN, or SEQ#. The couple
cue flashes for 10 seconds and is removed if the steering signal is lost or becomes invalid.
8. Sequential steering lines. The sequential steering lines are displayed when a sequence is entered
and when one of the sequence options (SEQ1, SEQ2, SEQ3, or SEQL) is boxed. The sequential
steering lines are available for display in all HSI modes and range scales. Sequential steering lines
are not displayed at power up with WonW and are removed when: magnetic heading is invalid,
aircraft position is invalid, or map slew is selected.
9. Time of day. Zulu time of day (ZTOD) or local time of day (LTOD) are displayed on the lower left
corner of the HSI. For aircraft that pass the FIRAMS real time clock power up BIT, ZTOD does not
need to be entered. For aircraft that do not pass the FIRAMS real time clock power up BIT, ZTOD
must be entered.
10. Groundspeed required. Groundspeed required appears below the current groundspeed readout.
Groundspeed required indicates the groundspeed required to a target based on entered ZTOD, time
on target (TOT), and the target.
11. Elapsed time (ET)/countdown (CD) time. ET and CD time are displayed on the lower right
corner of the HSI, however, only one of the timers can be displayed at a time. ET or CD timer must
be selected to be displayed. ET initializes to zero minutes and seconds and CD time initializes to six
minutes and zero seconds.
12. Aircraft heading. Aircraft heading is indicated on the compass rose. Aircraft heading and bearing
data can be selected as either magnetic or true. With true heading selected, the letter T appears
below the lubber line and the word TRUE appears below the selected scale readout. There is no
indication when magnetic heading is selected.
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ORIGINAL
A1-F18EA-NFM-000
Figure 2-51. HUD Controls
2.19.4.6 HUD - Head-Up Display. The HUD is on the center main instrument panel. The HUD is
used as the primary flight instrument, weapon status, and weapon delivery display for the aircraft
under all conditions. (Non-AMCD aircraft) The HUD receives attack, navigation, situation, and
steering control information from the left or right DDI symbol generators (under mission computer
control) and projects symbology on the combining glass for head-up viewing. (AMCD aircraft) The
HUD receives all symbology from MC1 or MC2. The HUD is electrically interfaced with the UFCD.
The HUD has NVG compatible raster display capability to allow it to display NFLR video. The
controls for the HUD are below the UFCD and are described in the following paragraphs. See figure
2-51.
2.19.4.6.1 HUD Symbology Reject Switch. The HUD reject switch is located on the HUD control
panel on the center main instrument panel. This switch is used to control the amount of symbology
displayed on the HUD.
NORM
Displays full HUD symbology.
REJ 1
Removes the Mach, g, and peak-g indications, the bank angle scale and pointer, the airspeed and altitude boxes, energy caret (landing gear down), and the ground speed
required cue.
REJ 2
Removes REJ 1 symbology and the heading scale, current heading caret, command
heading marker, NAV/TCN range, and the ET, CD, LTOD or ZTOD timer.
2.19.4.6.2 HUD Symbology BRT Control Knob. This knob is used to turn on the HUD and then
varies the display intensity.
2.19.4.6.3 HUD Symbology Brightness Selector Switch. This is a two-position toggle switch with
positions of DAY and NIGHT. Placing the switch to DAY provides maximum symbol brightness in
conjunction with the HUD symbology brightness control. With the switch set to NIGHT, a reduced
symbol brightness is provided in conjunction with the HUD symbology brightness control.
2.19.4.6.4 Black Level Control Knob. The black level control knob, located on the HUD control
panel, adjusts the NFLR video plus or minus ½ a shade of gray per increment when rotated.
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A1-F18EA-NFM-000
2.19.4.6.5 HUD Video Control Switch. The video control switch, located on the HUD control panel,
enables NFLR video display on the HUD with selectable polarity (white hot/black hot).
W/B
Selects white hot/black hot polarity.
VID
Displays NFLR video in the HUD, if available.
OFF
NFLR video off.
2.19.4.6.6 BAL Control Knob. The balance control, located on the HUD control panel, adjusts the
stroke brightness relative to the raster brightness. Rotating the switch from 12 o’clock towards the VID
position holds the brightness of the video (as set by the brightness control switch) and reduces the
brightness of the stroke symbology. The opposite is true when rotating the switch toward the SYM
position.
2.19.4.6.7 AOA Indexer Control Knob. This knob controls the brightness of the indexer lights.
2.19.4.6.8 ALT Selector Switch. The ALT selector switch, located on the HUD control panel, is used
to select the primary altitude source for display on the HUD and for use in the mission computer
(weapon systems calculations).
BARO
Selects barometric altitude.
RDR
Selects radar altitude.
2.19.4.6.9 ATT Selector Switch. The ATT selector switch, located on the HUD control panel, is used
to select the primary attitude source used for display in the HUD and in MC and FCC computations.
INS
Functions identically to the AUTO position.
AUTO
Selects filtered INS data as the primary attitude source. The INS automatically reverts
to gyro mode, using unfiltered data if its processor fails. The MC automatically selects
the standby attitude reference indicator for attitude information if the INS fails completely.
STBY
Selects the standby attitude reference indicator. The FCCs no longer use INS data and
the HIAOA advisory is displayed.
2.19.4.6.10 HUD Symbology. The following paragraphs describe HUD symbology as related to basic
navigation, steering (direct, great circle, courseline, and ILS), navigation target designation, advisories
and landing, see figure 2-52. Refer to part VII for a description of how these symbols are integrated into
the navigation system. Also, refer to section VII for unique ACL data link symbology. Refer to
A1-F18EA-TAC-Series, for symbology concerning the A/A and A/G master modes, weapons, RWR and
the data link vector mode.
1. Heading. The aircraft magnetic/true heading is indicated by the moving 30° heading scale. The
actual aircraft heading is directly above the caret/T symbol. The moving heading scale provides
trend information during turns. As the aircraft turns right, the scale moves from right to left.
Magnetic or true heading may be selected. Magnetic heading is indicated by a caret below the
heading scale. True heading selection is indicated by a T appearing below the current heading.
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2. Airspeed. Calibrated airspeed from the FCC is provided in the box on the left side of the HUD.
The tops of the airspeed and altitude boxes are positioned at the aircraft waterline, which is 4° up
from the optical center of the HUD.
HUD airspeed should normally read less than 50 knots while sitting still
on the ground. A reading of more than 50 knots on the ground may
indicate an Air Data failure.
3. Altitude. The altitude presented in the box on the right side of the HUD may be either barometric
altitude or radar altitude depending on the setting of the altitude switch on the HUD control panel.
When the altitude switch is in the BARO position, barometric altitude is displayed. When the
altitude switch is in the RDR position, radar altitude is displayed and is identified by an R next to
the altitude. If the radar altitude is invalid, barometric altitude is displayed and a B next to the
altitude flashes to indicate that barometric altitude is being displayed rather than radar altitude. An
X displayed next to the barometric altitude indicates that the altitude value may be inaccurate. The
ten thousand and thousand digits are 150% size numbers. The hundred, ten, and unit digits are
120% size numbers, except that below 1,000 feet they are 150% size.
4. Barometric setting. The barometric setting used by the air data function in the FCC is the value
set in the standby altimeter. When the barometer setting is changed on the standby altimeter, the
barometric setting is presented below the altitude on the HUD to provide a head-up baro-set
capability. The display remains for 5 seconds after the change is made. In addition, the baro-set
value is displayed and flashed for 5 seconds when the aircraft descends below 10,000 feet at an
airspeed less than 300 KCAS.
5. Angle of Attack. True angle of attack in degrees is displayed at the left center of the HUD. AOA
values displayed on the HUD are filtered and may slightly lag actual true AOA. Therefore, it may
be possible to trigger the AOA tone slightly prior to seeing the applicable limit AOA in the HUD.
The HUD AOA is generally driven by the FCS using the AOA probes. However, around 42° AOA
(and -9°), the HUD reverts to an MC-computed AOA based on INS data and winds. The HUD AOA
flashes indicating that the displayed HUD AOA may be inaccurate because the INS computed winds
have not been updated in the past three minutes. Low AOA, small bank angles, and small rates are
required to update the wind. With MC OFP H3E AND UP, the pitch trim AOA value is displayed
next to the ATC HUD advisory location while trimming and for two seconds after trimming with
WoffW and flaps HALF or FULL. The value is displayed with or without ATC engaged but is not
displayed with autopilot engaged.
6. Mach number. The aircraft Mach number is displayed immediately below the angle of attack.
7. Aircraft g. Normal acceleration of the aircraft is displayed immediately below the Mach number.
8. Peak aircraft g. A peak positive g indication is displayed on the HUD below the normal g when
a threshold of 4.0g is exceeded. The peak positive g display can be removed by cycling the clutter
reject switch to one of the reject positions.
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Figure 2-52. HUD Symbology (Sheet 1 of 2)
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Figure 2-52. HUD Symbology (Sheet 2 of 2)
9. Bank angle scale. A bank angle scale and pointer are displayed at the bottom of the HUD for bank
angle reference up to 45°. At bank angles in excess of 47°, the bank angle scale pointer is limited at
45° and flashes.
10. Velocity vector. The velocity vector provides an outside world reference with regard to actual
aircraft flight path. The velocity vector represents the point towards which the aircraft is flying
(aircraft flight path). With a functioning INS, the velocity vector is driven by INS attitude and
velocities. If GPS data is valid, a ″hybrid″ GPS vertical velocity correction is used to correct errors
in the INS vertical velocity loop regardless of the INS mode switch position (CV, GND, NAV, or
IFA). If GPS data is not available for use in the hybrid correction, a VVEL advisory is displayed.
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With a VVEL advisory displayed, sustained climbs and descents, such as
penetration from the marshal stack, can result in uncued (no cautions)
vertical velocity errors and a possible inaccurate velocity vector position.
Error magnitudes increase at slower airspeeds and lower altitudes. Errors
of up to 3° (actual flightpath 3° below the displayed velocity vector) have
been observed in the landing configuration. Three minutes of level flight
may be required to allow the INS to correct the vertical velocity errors.
NOTE
The VVEL advisory will be displayed if a GPS is not installed or if
masking prevents GPS positioning of a sufficient quality to ″aid″ the
INS vertical velocity loop.
The position of the velocity vector is limited to an 8° radius circle centered at the HUD optical
center. If the velocity vector reaches this limit during high angle of attack flight or large yaw and/or
drift angles, it flashes rapidly to indicate that it does not accurately indicate flight path (velocity vector
is HUD limited).
With GPS operating, if the INS velocity data becomes unreliable, the mission computer utilizes GPS
information. If INS velocity data becomes unreliable the mission computer utilizes FCC air data
function information and the last available wind data to compute the velocity vector and this degraded
velocity vector is indicated by a slow flashing of the symbol. In the NAV master mode, the velocity
vector may be caged to the vertical center line of the HUD by the cage/uncage switch on the throttle.
When it is caged, a ghost velocity vector is displayed at the true velocity vector position if that position
is more than 2° from the caged position. The flight path/pitch ladder and steering information are
referenced to the caged position. The ghost velocity vector flashes when limited. The flight path/pitch
ladder is referenced to the waterline symbol when the velocity vector is caged.
The velocity vector is automatically fixed at the horizon with WonW and ground speed less than 80
knots, and its status cannot be changed until WoffW. However, the velocity vector’s selected
caged/uncaged status does not change during touch-and-go landings if the ground speed remains above
80 knots.
11. Flight path/pitch ladder. The vertical flight path angle of the aircraft is indicated by the position
of the velocity vector on the flight path/pitch ladder. The horizon and flight path/pitch angle lines
represent the horizon and each 5° of angle between ±90°. Positive pitch lines are solid and are above
the horizon line. Negative pitch lines are dashed and are below the horizon line. The outer segments
of the lines point toward the horizon. Each line is numbered and the numbers rotate with the lines
so that inverted flight can easily be determined. To aid in determining flight path angle when it is
changing rapidly, the pitch lines are angled toward the horizon at an angle half that of the flight path
angle. For example, the 50° pitch line is angled 25° toward the horizon. In level flight, the pitch lines
are not angled. The zenith is indicated by a circle and the nadir is indicated by a circle with an X
in it. Aircraft pitch angle can be determined by comparing the tops of the altitude and airspeed
boxes (which represent the aircraft waterline) with the pitch ladder when the wings are level, but the
flight path/pitch ladder normally rotates about the velocity vector and determination of pitch angle
may be difficult at high roll angles.
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12. Vertical velocity readout. This value is displayed above the altitude box and indicates vertical
velocity in feet per minute. This is displayed in the NAV master mode only. Descent is indicated by
a minus sign.
13. HUD landing symbology. When any two landing gear are down, the Mach number, g, and peak
g are deleted and an AOA bracket, extended horizon bar, waterline symbol, and energy caret appear.
The center of the AOA bracket represents the optimum approach AOA. The bracket moves lower
with respect to the velocity vector as AOA increases and moves higher as AOA decreases. When the
energy state of the aircraft is in equilibrium, the energy caret points to the ″right wing″ of the
velocity vector and the aircraft neither accelerates or decelerates. With an energy deficit, the energy
caret moves lower with respect to the velocity vector and the aircraft decelerates; with excess energy,
the energy caret moves higher and the aircraft accelerates.
14. Waypoint/OAP, mark point, TACAN, or target data. Waypoint/OAP and mark data consists of
range (horizontal), and the steer-to point identifier (W, O, or M) and number located on the lower
right corner of the HUD. TACAN data consists of slant range and a Morse code identifier located
on the lower right corner of the HUD. When a steer-to point is designated, range remains displayed
and the steer-to point identifier changes to TGT.
15. Coupled steering symbology. While coupled steering is engaged CPL SEQ#, CPL WYPT, CPL
TCN, CPL BNK, CPL ASL, CPL HDG, or CPL P/R appears on the right side of the HUD display
above the navigation data.
16. ILS symbology. When ILS steering is selected, an azimuth deviation bar (localizer) and elevation
deviation bar (glideslope) appear on the HUD.
17. ZTOD, LTOD, ET, and CD time. The ZTOD, LTOD, ET, or CD time is displayed on the lower
left corner of the HUD. These timers are mutually exclusive. Only one timer is available for display
on the HUD at a time. When the FIRAMS real time clock power up BIT passes, ZTOD does not need
to be entered, but when the FIRAMS real time clock power up BIT does not pass, ZTOD must be
entered. ET initializes to zero minutes and seconds. CD initializes to 6 minutes and zero seconds.
18. Command heading marker. When waypoint/OAP or TACAN direct great circle steering is
selected, the command heading marker is displayed just below the heading scale.
19. Steering arrow and dots. When waypoint/OAP or TACAN course line steering is selected, the
steering arrow and dots appear on the HUD.
20. Required ground speed cue. When steering is engaged to the target in a sequence, the required
ground speed cue appears under the airspeed box.
21. Target designation symbology. When a target is designated, a target designation symbol
(diamond) appears below the heading scale indicating target heading. Another target designation
symbol (diamond) appears indicating the target line of sight (LOS).
2.19.4.6.11 HUD Symbology Degrades. The avionics suite has built in redundancy with two mission
computers for data management and two DDIs (LOTs 21-24), or two MCs (LOTs 25 and up), for
symbol generation. Likewise, if the attitude select switch is in the AUTO or INS position, back up data
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sources are automatically selected to provide HUD symbology when failures are detected. Refer to
figure 2-53, for the HUD displays discussed below.
a. HUD Symbology Degrades with INS Failure. When a failure occurs in the INS, HUD
bank angle, velocity vector, pitch ladder, and heading indications can be expected to be
impacted. With GPS operating, the mission computer utilizes GPS information for the velocity
vector. If INS attitude is valid but INS velocities are not valid the mission computer
automatically uses the INS attitude and GPS velocities to position a non-flashing velocity
vector. With a degradation of the air data function (probe or pressure transmitter set damage
or failure) calibrated airspeed, barometric altitude, indicated Mach number, and vertical
velocity indications may be impacted.
When the INS experiences a total shutdown (dump) with the attitude select switch in AUTO
or INS, or if the attitude switch is deliberately placed in standby, a stationary waterline symbol
replaces the velocity vector indicating that the standby attitude reference indicator is now
providing attitude data. This failure is normally accompanied by the MASTER CAUTION
light, tone, and INS ATT caution. Place the attitude select switch in the STBY position,
crosscheck the HUD against standby instruments, and attempt an in-flight alignment.
Due to the tendency of the standby attitude reference indicator to precess, it is suggested that
flying in instrument meteorological conditions (IMC) using the ARI as a primary attitude
reference be minimized. A partial IFA (In-Flight Alignment) is always recommended whenever
possible to recover the INS attitude platform.
b. HUD Symbology Degrades with Air Data Function Failure. An air data function
failure in the FCC results in loss of associated data from the HUD display as shown in figure
2-53. Such a failure also inhibits operation of cruise flight Automatic Throttle Control and
disables the altitude signal used for IFF altitude reporting. An air data function failure may
affect cabin air flow and cabin air temperature.
The pressure transmitter set can produce erroneous signals without cautions or advisories if
the pitot tube or AOA probes receive damage. As the air data function degrades, loss of some
or all of the following data from the HUD may occur:
(1) Calibrated airspeed or barometric altitude. The loss of calibrated airspeed and/or barometric altitude data results in activation of the landing gear handle warning light and tone
with the gear UP. Aircrew action is to reference the applicable standby airspeed or altitude
indicator and then silence the tone.
(2) Angle of Attack. Loss of AOA in three or more FCC CHs causes AOA to be removed from
the HUD.
(3) Vertical velocity indicator. Pilot action on loss of the vertical velocity indication is to check
that the aircraft is in the NAV master mode and to reference the standby vertical velocity
indicator.
(4) Mach number. Pilot action on loss of the Mach number indication is to reference the
standby airspeed indicator.
If an AOA probe becomes jammed (does not move), the FCC continues to receive valid signals until
the pilot executes a maneuver that causes the reading between the AOA probes to differ more than 15°
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Figure 2-53. HUD Symbology Degrades
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in UA or 5.5 to 15° in PA, depending on sideslip. HUD displayed airspeed may be inaccurate without
annunciation if a pitot tube is damaged.
A jammed, blocked, or damaged pitot tube/AOA probe may not be
annunciated if system errors are not large enough to set a caution. Be
alert for unannunciated pitot static and AOA errors during flight in icing
conditions or if damage is suspected after a bird strike or IFR basket
impact during inflight refueling.
Air data inputs from the MC are used by the INS to help smooth or dampen pitch ladder and
velocity vector position. A complete air data function failure does not immediately affect the pitch
ladder/velocity vector, but these displays eventually degrade. If subtle damage to the AOA probe is
suspected, the pilot should make a crosscheck of airspeed with a wingman if possible. The standby
airspeed indicator receives signals from the left pitot static probe, so it is accurate if only the right
probe is damaged. AOA checks with a wingman should be made in landing configuration if a jammed
AOA probe is suspected. Crosschecking in cruise configuration may give a satisfactory crosscheck, but
the probe may be bent in such a way that AOA anomalies are accentuated on landing configuration.
Landing with automatic throttle control (ATC) may be affected. If damage is suspected, ATC during
landing is not recommended.
When AOA is declared invalid (e.g., AOA Four Channel failure), the HUD AOA display and AOA
bracket are removed and the AOA indexer lights and approach lights are inoperative. GAIN ORIDE
provides fixed gains to the FCS and allows the pilot to select, through the FCS status display, either
the left or right probe. The center (INS) AOA value allows the pilot to compare AOA values to select
the undamaged probe. Once selected, this probe drives the HUD AOA display, AOA bracket, AOA
indexer, and approach lights. If the incorrect probe is selected, the information provided to the pilot
and LSO may be in error but has no impact on the flight control system as the gains are fixed. Notify
the LSO that a single probe has been selected.
2.19.4.6.12 HUD Advisory Data Symbology. The displays in figure 2-52 show some of the advisories
that can appear on the HUD in the NAV master mode. The advisories are associated with nose wheel
steering and approach power compensator. Although the advisories are shown on the gear down
display, most of them can appear on the basic HUD display. Refer to Part VII for description of data
link system and advisories.
The automatic throttle control/nosewheel steering advisories are displayed above the distance
display whenever the ATC or the NWS is engaged. If the ATC is disengaged by any means other than
actuation of the ATC engage/disengage switch, the advisory is flashed for 10 seconds before it is
removed from the display or, if a pilot attempt to engage ATC is not successful, ATC is flashed for 10
seconds and removed.
2.19.4.6.13 HUD BIT Checks. The HUD has two methods of built-in tests: manually initiated and
automatic test. Refer to BIT-Status Monitoring Subsystem for the procedures and displays used for
the HUD BIT checks.
2.19.4.7 CRS Set Switch. The course set switch manually sets the desired course on the HSI display.
2.19.4.8 HDG/TK Set Switch. The heading/ground track set switch manually sets the heading
marker on the desired heading/ground track on the HSI display.
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2.19.5 Up Front Control Display (UFCD). The UFCD is on the main instrument panel below the
HUD in the front cockpit. In the rear cockpit, the UFCD is located below the MPCD in LOTs 21 thru
25, and above the MPCD or 8 x 10 display in LOT 26 AND UP. The UFCD is an active matrix liquid
crystal display with an IR touchscreen used for data entry inputs consisting of digits or NEWS, and
control of the CNI systems (autopilot modes, IFF, TACAN, ILS, data link, radar beacon, UHF/VHF
radios, and ADF), see figure 2-55. In addition, the touchscreen can be used as a multi-function display
for display formats, including video. The UFCD is used in conjunction with the two DDIs, the 8 x 10
display and the MPCD to enter navigation, sensor, and weapon delivery data. UFCD option selections
and inputs are transmitted directly to the MPCD and on to the mission computers. (The mission
computers pass these inputs to the control converter (CC) for CNI equipment control). The UFCD is
NVG compatible. In the F/A-18F, the front and rear cockpit UFCDs operate independently. When
different formats are being displayed, the only data common to both UFCDs is radio channel and
frequency information. Both cockpit UFCDs present the results of changes to radio channels or
frequencies at the same time, regardless of which cockpit performed the change. The cockpit not
performing data entry does not see touch highlights as the data is entered, only the result of the data
entry. If both pilot and WSO enter digits on the keypad for the same option, both entries are accepted,
with the second entry overwriting the first. When pilot and WSO are on the same data entry or CNI
format, asterisks are provided in the top left and right of the scratch pad.
(LOT 26 AND UP) The aft UFCD is electrically controlled through the 8 x 10 display Off/On/
Brightness knob when it is installed. The aft UFCD flashes STANDBY when the MC1 communication
to the UFCD is disrupted. A standby display pattern in the the center of the UFCD display surface,
similar to the DDI pattern, indicates degraded image processing from MC1. When the Mono video
connection from the 8 x 10 display is lost or degraded, the aft UFCD will not display anything. When
video synchronization is lost or degraded, the aft UFCD may display symbology that is not coordinated
with the video. If the interface between MC1 and the aft UFCD is lost, the touch screen capability and
brightness/contrast control will not work. When MC1 is inoperative, the aft UFCD is not capable of
displaying format symbology or analog video.
When the pilot or WSO touch a keypad option on the touch screen a highlight appears indicating
that the option has been selected. Figure 2-55 shows an example of a selected option. When a new
format is selected on the touchpad, the highlight does not remain on the new format.
Keypad options use a first finger in mechanization. Only one option can be selected at a time. If two
or more selections are attempted at one time, none of them are considered valid.
Some formats initialize with data in the scratchpad, e.g., the COMM sublevel of the CNI format
initializes with the comm frequency in the scratchpad. When data entry is started, the digits in the
scratchpad are blanked, allowing data entry. If a comm frequency is being entered the decimal remains
displayed, allowing frequency entry in relation to the decimal point to be viewed.
The UFCD uses a double clear mechanism. The first selection of the CLR keypad option removes the
last digit that was entered in the scratchpad. The second selection of the CLR option removes all the
digits which have been entered in the scratchpad. If the scratchpad is flashing due to the 10 second
timer (no entries in the previous 10 seconds), pressing CLR stops the flashing and performs the
previously described CLR function. If the scratchpad is flashing due to an error, pressing CLR stops
the flashing and removes all digits from the scratchpad.
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Uncommanded option selections may occur with a malfunctioning
UFCD. This can include uncommanded autopilot selection or deselection. In order to reduce the flight safety risk, a UFCD that self-selects
uncommanded options should be secured in flight and only turned on if
required for safety-of-flight functions (radio, IFF or navigation selections). The malfunction may or may not be present when turned back on,
but can be expected to reoccur during flight if operational. In F/A−18F
aircraft, an unaffected UFCD may remain on and in use with no adverse
effect. A frozen unresponsive display is the result of another unrelated
failure mode that may occur and can be corrected by cycling power to the
MPCD/DECD.
NOTE
Securing the UFCD will reset the MPCD/DECD and cause a transitory
display loss.
2.19.5.1 UFCD Data Entry. The UFCD has two data entry protocols, Standard Data Entry (SDE)
and Fast Data Entry (FDE).
2.19.5.1.1 Standard Data Entry. With MC OFP 18E, SDE digits are entered using the keyboard
followed by the enter (ENT) key. In many cases, the option for which the data is being entered must
first be selected prior to entering digits. Selection of the ENT option causes the digits in the scratchpad
to flash once, confirming that the ENT has been accepted. The majority of the data entry displays use
SDE protocol and are recognized by the presence of the ENT key.
2.19.5.1.2 Fast Data Entry. With MC OFP 18E some data, and with MC OFP H1E AND UP all data
are entered using the keyboard followed by selecting the option directly. This protocol does not use the
enter (ENT) key. With MC OFP 18E, MAN replaces ENT. With MC OFP H1E AND UP, NEWS
replaces ENT. See figure 2-54.
2.19.5.1.3 Data Entry Using the Shifted Keypad. The N-E-W-S option is provided on all UFCD
data entry displays with MC OFP H1E AND UP. The shifted keypad provides the ability to enter a
negative sign, a decimal point, degrees, minutes, and seconds symbols, and North (N), East (E), West
(W), and South (S) entries for latitude and longitude entries. With MC OFP H2E+ AND UP, latitude
and longitude can be entered to ten thousandths of a minute, or hundredths of a second for increased
precision.
2.19.5.2 UFCD CNI Function. The aircraft powers up with all CNI systems off and the top level CNI
format as the default UFCD display. See figure 2-55. This format is the central point for controlling all
CNI systems. For Communication-Identification Equipment, see Chapter 23. For Navigation Equipment, see Chapter 24.
The following top level options are not related to CNI systems: DDI, EW, and FLR. The DDI option
provides the last selected DDI display on the UFCD. The EW and FLR options provide the specified
display format on the UFCD.
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Figure 2-54. Alphanumeric Entry Format
The top level CNI display does not use an ENT function, the MAN option is used to turn to a
previously entered manual radio frequency.
2.19.5.3 UFCD STANDBY Indication. A flashing STANDBY indication is provided just above the
center of the UFCD display surface when there is no valid mux communication between the MC and
MPCD. The STANDBY indication is superimposed over any display format. When the STANDBY
indication is displayed, power to the MPCD should be cycled to attempt a reset. If the STANDBY
condition clears before power can be recycled, the UFCD screen blanks and the previously displayed
format reappears.
2.19.5.3.1 Aft UFCD MUX FAIL. When the 8 x 10 display is installed, the aft UFCD generates a MUX
FAIL BIT indication when MC1 loses communication with the UFCD. This message occurs only when
the 8 x 10 display is turned on and powering the aft UFCD. Using periodic BIT monitoring, the MC1
tries to reset the aft UFCD through the 8 x 10 display communication if an error is detected.
2.19.5.4 RALT - Radar Altimeter Function. The RALT function indicates clearance over land or
water from 0 to 5,000 feet. Operation is based on precise measurement of time required for an
electromagnetic energy pulse to travel from the aircraft to the ground and return. A warning tone and
visual warnings are activated when the aircraft is at or below a selectable primary or secondary low
altitude limit. The primary/secondary radar low altitude warnings are reset by setting the low altitude
index (primary), or UFCD selected altitude (secondary) to an altitude below the present altitude or by
climbing above the previously set limit. The warning tone can be disabled in either cockpit by pressing
the flashing RALT option on the UFCD Low Altitude Warnings format or UFCD DDI format, turning
off the radar altimeter, setting the warning value below current radar altitude, or by climbing above the
warning altitude. When disabled, the tone cannot be triggered until after being reset as described.
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Figure 2-55. Up Front Control Display (UFCD)
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The radar altimeter consists of a receiver-transmitter and individual transmitting and receiving
antennas. The receiver-transmitter produces the energy pulses, transmits the energy to the ground,
receives the reflected signal, and processes the data for display as altitude by the head-up display unit
(HUD).
Indicators and controls used with the electronic altimeter set are the left or right DDI on the
instrument panel (for BIT checks), the ALT switch, UFCD (for secondary low altitude warning), and
the head-up display. Radar altimeter BIT is initiated from the BIT display.
Pressing the emission control (EMCON) switch on the UFCD inhibits operation of the radar
altimeter. EMCON is toggled on or off each time the EMCON switch is pressed. When EMCON is on
the letters E, M, C, O, N are displayed either in the left side option boxes (DDI display) or in the
scratchpad of the top level CNI display and on the HUD.
2.19.5.4.1 Low Altitude Warning Tone. When the aircraft descends below the primary low altitude
set in the UFCD, a ‘‘Whoop, Whoop’’ warning tone is heard in the aircrew’s headset and the word
ALTITUDE is displayed on the HUD. The warning tone, when initiated by the primary radar low
altitude warning, is repeated at the lowest priority until reset or disabled.
Numbers ending in zero are valid entries for both the Low Altitude
Warning and COMM 1. Due to the close proximity of the COMM 1 and
RALT options, it is possible to inadvertently change the low altitude
warning setting for the radar altimeter when attempting to use COMM 1
fast data entry.
A barometric low altitude and secondary radar low altitude warning function are enabled by entering
the appropriate altitude, up to a maximum of 25,000 (BARO) and 5,000 (RADAR), on the UFCD. The
barometric low altitude and secondary radar low altitude warning provide a single voice alert warning
“ALTITUDE, ALTITUDE” when the aircraft descends through the selected altitude. Refer to Part VII
for information on entering altitude. The barometric low altitude warning function does not affect the
operation of the radar altimeter low altitude warning function.
2.19.5.5 UFCD Controls. A description of UFCD switches follows. Refer to Part VII for operating
instructions for CNI equipment.
2.19.5.5.1 COMM 1 and 2 Channel Knobs. The COMM 1 and 2 channel knobs permit independent
selection of up to 20 preset channels on radios 1 and 2. Guard (G), Manual (M), Ship Maritime (S), and
SINCGARS Cue (C) channels are also available on the continuously rotatable knobs.
2.19.5.5.2 COMM VOL Knobs. The VOL knobs independently control radio volume for COMM 1 and
2. The OFF position (detent) is at the full counterclockwise position. In the F/A-18F, both sets of knobs
must be in the OFF position for power to be removed from the radios. When a radio is powered, the
respective COMM option is corner highlighted. When a radio is actively receiving, a half intensity
highlight is shown on the upper half of the COMM option (figure 2-55).
2.19.5.5.3 ID (IDENT) Pushbutton. The ID pushbutton commands an IFF identification/position
squawk (IDENT), for modes 1, 2, and 3 (if enabled).
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2.19.5.5.4 UFCD BRT Knob. The BRT knob adjusts the overall brightness of the UFCD display, both
symbology and video. Turning the knob counterclockwise to the OFF position removes power from the
UFCD.
2.19.5.5.5 UFCD CONT Knob. The CONT knob adjusts video contrast on the UFCD.
2.19.5.5.6 UFCD SYM Knob. The SYM knob is used to adjust the brightness of the UFCD symbology
without affecting the brightness of the UFCD video.
2.19.5.5.7 EMCON Pushbutton. If the UFCD is displaying a DDI format the letters E M C O N are
displayed in the left side option boxes. If the UFCD is displaying the top level CNI format, EMCON
is displayed in the scratch pad.
2.19.6 SDC - Signal Data Computer. The signal data computer (SDC) operates under mission
computer control and records aircraft fatigue strain data, engine parameters when out of tolerance
conditions occur, fuel information, and aircraft and target parameters when targets are designated and
weapons are delivered. It includes fuel transfer controls and gaging capabilities, incorporates ground
support equipment fuel transfer and gaging fault isolation functions, and provides interface for
multiple sensors and controls. It provides analog-to-digital conversion of aircraft parameters. In
addition, BIT fail indications are stored in the SDC to be displayed by the maintenance status panel
(MSP) for readout by maintenance personnel after the flight, or on the EFD for readout during the
flight.
The RESET SDC option is available from the SUPT MENU/FUEL format. This option is used to
reset the SDC by momentarily removing power. When selected, the RESET portion of the option
legend is boxed, and remains boxed until the SDC reestablishes AVMUX communication or 15 seconds
after the option was selected. The RESET SDC option is removed from the FUEL format if the CSC
is not communicating on the AVMUX.
With AMCD aircraft, the SDC is used to compensate for periods of blank cockpit displays and
interrupted avionics multiplex (AVMUX) bus communication occurring when both advanced MCs are
off-line or in initialization. The SDC will take temporary control of AVMUX busses 1 and 6 during the
ground start and inflight conditions. While acting as the backup bus controller, the SDC transmits to
the forward MPCD and forward UFCD a limited HUD display of essential flight information. Refer to
chapter 25. Additional data transfers to the left and right Full Authority Digital Engine Controllers
(FADECs) and the Environmental Control System (ECS) Controller sustains standard operation. The
SDC also enables display of the Left and Right Air Turbine Starter (L ATS and R ATS) cautions.
2.19.7 CVRS - Cockpit Video Recording System. LOTs 21−24 CVRS contains two video tape
recorders, a HUD camera, and two over the shoulder cameras. The system has the capability to record
either front DDI, the MPCD, the HMD, the aft MPCD, the aft UFCD, or the HUD (in color), but in
limited combinations. Headset audio is also recorded as long as the KY−58 encryption function is
inactive. The switches used to operate CVRS are located on the VIDEO RECORD panel in each
cockpit.
The CVRS in LOT 25 is identical to that for LOTS 21−24 with the exception of the two
over−the−shoulder cameras which are removed and the DDIs are recorded using direct video outputs.
In LOTs 26 and up without the solid state recorder, the CVRS consists of two video tape recorders
and a HUD camera. No over the shoulder cameras are installed. The system has the capability to direct
record various combinations of either front DDI, aft DDI, MPCD, aft MPCD or 8 x 10 display, the
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HMD, and the HUD in color. Headset audio is also recorded as long as the KY−58 encryption function
is inactive. The switches to operate the CVRS are located on the VIDEO RECORD panel in each
cockpit.
With AMCD aircraft with solid state recorder (SSR) installed (after AFC 443), the CVRS consists
of one SSR with removable memory module (RMM). The system has the same record options and
cockpit interface as the 8 mm CVRS.
2.19.7.1 CVRS Video Tape Recorders/Solid State Recorder (SSR). The two CVRS video tape
recorders (VTRs) are located behind the ejection seat in the F/A−18E and behind the rear cockpit
ejection seat in the F/A−18F. Each VTR provides a minimum of 2 hours recording time on removable
8 mm video tape cartridges.
The SSR is located in the avionics bay behind the ejection seat in the F/A-18E and behind the rear
cockpit ejection seat in the F/A−18F. VTR1 and VTR2 record options are the same as CVRS. The SSR
provides a minimum of 3 hours recording time on a removable memory module (RMM).
2.19.7.1.1 SECURE ERASE Button. The SECURE ERASE button is guarded and located on the
right hand forward vertical console. Pressing the SECURE ERASE button erases the information
stored in the RMM.
2.19.7.1.2 SSR Advisories. The following SSR related advisories are described in the Warning/
Caution/Advisory Displays in part V: RMMCD, RMMFL.
2.19.7.2 HUD Camera. The HUD camera, positioned immediately in front of the HUD, records a
color image similar to what the pilot sees through the HUD, e.g., symbology superimposed on a picture
of the outside world. This combined image is made available for recording.
2.19.7.2.1 HUD Event Marker. The HUD event marker is a small black marker generated by the
HUD camera and positioned in the upper left corner of the video signal sent to the VTR for recording.
The event marker is displayed when the weapon release pickle button is pressed or the trigger is
squeezed to the second detent and remains displayed until pickle button/trigger release. When
reviewing CVRS video post−flight, the event marker is a useful training aid, used to determine the
timing and duration of pickle button and trigger actuations (e.g., shot validation).
2.19.7.2.2 HUD Camera BIT Button. The BIT button, located on the front face of the HUD
assembly, is used to initiate a BIT of the HUD camera. The GO and NO GO BIT status balls, normally
black, are used to determine the BIT status of the HUD camera. When the BIT button is pressed, BIT
status is indicated by a green GO ball or an orange NO GO ball.
2.19.7.3 Over the Shoulder Cameras (LOTs 21−24). Two over the shoulder cameras, one mounted
on each side of the canopy frame approximately 30 inches from the instrument panel, are oriented to
record a picture of the L and RDDIs. The image from each camera is made available to a VTR for
recording.
2.19.7.3.1 Over the Shoulder Camera BIT Buttons (LOTs 21−24). Each camera contains a BIT
button located on its front face. Press and hold of the BIT button initiates a BIT of that camera. BIT
status is indicated by a green LED (go) or a red LED (fail).
2.19.7.4 CVRS Control Switches. The switches used to operate CVRS are located on the VIDEO
RECORD panel on the lower left instrument panel in the forward cockpit, and the aft cockpit for LOTs
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21−25 aircraft. (See figures 2-56 thru 2-64) They are located below the center display in the aft cockpit
of LOT 26 AND UP aircraft. The aft cockpit switches override the front cockpit switches for any
selection on LOT 26 AND UP aircraft. The VTR 1 and VTR 2 sources are grouped to nominally allow
front cockpit recording on VTR 2 (includes HUD) and allow aft cockpit recording on VTR 1 (includes
8 x 10 display).
2.19.7.4.1 CVRS Mode Switch. The CVRS mode switch is used to select manual or automatic video
recording.
MAN
CVRS records continuously (according to the inputs
selected by the two VTR selector switches).
OFF
CVRS recording off.
AUTO
CVRS records automatically only when in the A/A or A/G master modes (according to the inputs selected by the two VTR selector switches).
NOTE
The MODE switch must be in the OFF position to remove power from
the VTRs. Failure to place the MODE switch in the proper position
prior loss of aircraft power will keep the 8mm tapes from dethreading
and prevent the removal of the tapes from the VTRs until aircraft
power is reapplied.
2.19.7.4.2 CVRS RDCR ON Light. The green RDCR ON light, located on the left warning/caution/
advisory lights panel on the upper left instrument panel, comes on when CVRS power is on. An
illuminated light does not mean that the VTR or SSR is recording.
2.19.7.4.3 VTR Selector Switches. The VTR selector switches are used to select the video input
source for recording on the two VTRs.
Figure 2-56. Forward CVRS Control Panel (LOTs 21-22)
Forward CVRS (LOTs 21-22)
RDDI
Selects right over−the−shoulder camera video.
LDDI
Selects left over−the−shoulder camera video.
HUD
Selects HUD video camera.
UFCD
Selects aft UFCD direct video.
MPCD
Selects MPCD direct video.
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Figure 2-57. Forward CVRS Control Panel (LOTs 23-25 Before AFC 445)
Forward CVRS (LOTs 23-25 Before AFC 445):
HUD
Selects HUD video camera.
LDDI
Selects left over−the−shoulder camera video.
HMD
Selects HMD camera video.
RDDI
Selects right over−the−shoulder camera video .
Figure 2-58. Aft CVRS Control Panel (LOTs 21-22) and LOTs 23-25 Before AFC 445.
Aft CVRS (LOTs 21-22 and LOTs 23-25 Before AFC 445):
FWD
Selects front cockpit video. Default start−up position.
AFT/ MAN
Selects rear cockpit video. Turns CVRS on if system is off.
Electrically held in this position.
MPCD
Selects aft MPCD direct video.
UFCD
Selects aft UFCD direct video.
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Figure 2-59. Forward CVRS Control Panel (LOTs 23-24 After AFC 445)
Forward CVRS (LOTS 23-24 After AFC 445)
HMD
Selects HMD video camera.
LDDI
Selects left over-the-shoulder camera video.
RDDI
Selects right over-the-shoulder camera video.
HUD
Selects HUD video camera.
LDIR
Selects LDDI direct video.
RDDI
Selects right over-the-shoulder camera video.
NOTE
When both VTR 1 and VTR 2 have RDDI selected, the actual
recording occurs on VTR 2 only.
Figure 2-60. Forward CVRS Control Panel (LOT 25 After AFC 445
Forward CVRS (LOT 25 After AFC 445):
HMD
Selects HMD video camera.
LDDI
Selects LDDI direct video.
RDDI
Selects RDDI direct video.
HUD
Selects HUD video camera.
LDDI
Selects LDDI direct video.
RDDI
Selects RDDI direct video.
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Figure 2-61. Aft CVRS Control Panel (LOT 23-25 After AFC 445)
Aft CVRS Control Panel (LOT 23-25 AFTER AFC 445):
FWD
Selects front cockpit video. Default start−up position.
AFT
Selects rear cockpit video.
Electrically held in this position.
UFCD
Selects aft UFCD direct video.
MPCD
Selects aft MPCD direct video.
HMD
Selects HMD video camera.
Figure 2-62. Fwd CVRS Control Panel (LOT 26 AND UP)
Forward CVRS Control Panel (LOT 26 AND UP):
HMD
Selects HMD video camera.
LDDI
Selects LDDI direct video.
RDDI
Selects RDDI direct video.
HUD
Selects HUD video camera.
RDDI
Selects RDDI direct video.
MPCD
Selects MPCD direct video.
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Figure 2-63. Aft CVRS Control Panel (LOTs 26-29 Before AFC 445)
Aft CVRS Control Panel (LOTs 26-29 Before AFC 445):
CNTR
Selects center display direct video. Electrically held in this position.
FWD
Selects front cockpit video. Default start−up position.
LDDI
Selects LDDI direct video. Electrically held in this position.
FWD
Selects front cockpit video. Default start−up position.
RDDI
Selects RDDI direct video. Electrically held in this position.
Figure 2-64. Aft CVRS Control Panel (LOTs 26-29 After AFC 445 AND LOT 30 AND UP)
Aft CVRS Control Panel (LOTs 26-29 After AFC 445 AND LOT 30 AND UP):
CNTR
Selects center display direct video. Electrically held in this position.
FWD
Selects front cockpit video. Default start−up position.
LDDI
Selects LDDI direct video. Electrically held in this position.
HMD
Selects HMD video camera.
FWD
Selects front cockpit video. Default start−up position.
RDDI
Selects RDDI direct video. Electrically held in this position.
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NOTE
In LOT 26 AND UP F−model aircraft both aft VTR switches must be
in the FWD position AND the forward MODE switch must be in the
OFF position to remove power from the VTRs. Failure to place all 3
switches in the proper position prior loss of aircraft power will keep
the 8mm tapes from dethreading and prevent the removal of the tapes
from the VTRs until aircraft power is reapplied.
2.19.7.5 VTR2 Override Function. CVRS incorporates a VTR2 override function designed to make
sure that HUD camera video is recorded when an A/G weapon is released, the gun is fired, or an A/A
missile is launched. In A/A or A/G master modes, when the pickle button is pressed or the trigger is
squeezed to the second detent, VTR2 automatically switches from the selected VTR2 source to the
HUD camera. VTR2 records HUD video from pickle button/trigger actuation to pickle button/trigger
release plus a set overrun time before reverting to the video source selected by the VTR2 selector
switch. Overrun times are 5 seconds for AIM−9 launch and gun firing and 10 seconds for AIM−7/
AIM−120 launch and A/G weapon release. In the A/G master mode, if FLIR video is displayed on a
DDI, the UFCD, or the MPCD, VTR2 does not switch to record HUD camera video.
2.19.8 Fast Tactical Imaging Set (FTI-II). The FTI-II (AN/AVX-4) provides the F/A-18F (LOT 24
AFTER AFC 395) with near real-time capability to capture, view, send, and receive cockpit display
information as still-frame images in either air-to-air or air-to-ground mode. FTI-II captures any of 5
single-source video inputs from one of the following displays: aft LDDI, aft RDDI, HUD, aft MPCD,
and forward HMD. The system sends parallel video to the forward and aft DDIs for image viewing
when the monochrome map is selected for display. An additional output signal is sent to VTR1 for
recording. Transmitted and received imagery is encrypted using the KY-58 secure speech system and
the ARC-210 radio.
2.19.8.1 FTI-II Major Components.
2.19.8.1.1 Remote Switching Control (RSC). The RSC is installed on the aft cockpit left hand
console aft of the volume control panel. The RSC has two lines of 24 characters each and six buttons,
and is used for programming the FTI-II and executing commands.
2.19.8.1.2 Digital Imaging Processor (DIP). The DIP is installed in the forward cockpit, right hand
console just aft of the KY-58 control panel. The DIP has a removable compact flash card that stores
full resolution images. The DIP contains a CPU, communications protocol, LPEG software, wavelet
software, operating system software, frame grabber function, and a Merlin card for auto detection and
down-scale conversion.
2.20 TACTICAL AIRCRAFT MOVING MAP CAPABILITY (TAMMAC)
The TAMMAC avionics subsystem provides a moving map capability to enhance operational
effectiveness and survivability and addresses supportability/obsolescence issues facing existing moving
map and data storage systems currently deployed.
TAMMAC effectively replaces 3 separate Weapons Replaceable Assemblies (WRAs), including the
AN/ASQ-196 DVMS which consists of two WRAs; the Digital Map Computer (DMC) and the Digital
Memory Unit (DMU), and the AN/ASQ-194 DSS which is a single WRA.
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(LOT 26 AND UP) Independent maps and related formats are available in both cockpits with
Digital Video Map Computer (DVMC) installed, allowing independent map functionality and video
routing. The DVMC provides 5 outputs through two channels to various aircraft displays; however,
each map channel output is not available on every display. DVMC Channel 1 (using MC1) has two
analog video outputs. One output is viewable on the front LDDI, MPCD and rear LDDI. The other
channel 1 output is used for the aft UFCD. DVMC Channel 2 (using MC2) has two analog video
outputs and one digital color fiber optic output. One of the analog outputs is viewable on the front and
rear RDDIs, and 8 x 10 display. The other analog output is used for Wrap-Around-Test (WAT) and
removes video capability from the front UFCD. Channel 2 digital (fiber optic) is only available on the
8 x 10 display. When the Channel 2 digital output is in use on the 8 x 10 display, the analog outputs
are disabled. The DVMC only produces digital or analog on Channel 2 at any time, not both
simultaneously.
The 8 x 10 display processes analog video (via MC1 or MC2) or digital video, via Fiber Channel
Network Switch (FCNS) 2 along with the symbology received from MC2 to create a composite format
display image. FLIR mono video and map mono video (Wrap-Around-Test only) are the only analog
videos routed to the 8 x 10 display. The aft UFCD video displays are routed through the 8 x 10 display
for processing, and include sensor, weapon, and TAMMAC DVMC. Digital color map video on the 8 x
10 display is routed from the TAMMAC DVMC via the High Speed Video Network (HSVN) and
FCNS2. If FCNS2 fails, no digital color video is available on the 8 x 10 display. Map video routing is
shown in figure 2-65.
The TAMMAC subsystem consists of a new MU-1119/A Advanced Memory Unit (AMU), a new
CP-2414/A digital map computer and a High Speed Interface Bus (HSIB) which connects the two. The
DMC is a functional replacement for the existing DVMS. Data Transfer Devices (DTDs) used in
conjunction with TAMMAC for transferring data to and from ground-based stations are Personal
Computer Memory Card International Association (PCMCIA) cards or PC cards. Ground-based
stations that process data for TAMMAC include the Joint Mission Planning System (JMPS) and the
Automated Maintenance Environment (AME). Theater data is installed on the map loading card using
JMPS.
The AMU is a functional replacement for the existing DSS, more commonly referred to as the
Memory Unit (MU). The AMU contains two PC card receptacles, one for maintenance (ground
support) operations and one for mission (pilot) operations. This unique configuration allows maintenance and mission data to be separated both pre- and post-flight which reduces logistics coordination
and facilitates operational readiness. Three types of PC cards are used in the TAMMAC subsystem
operation: 1) Mission Card, 2) Maintenance Card, and 3) Map Loading Card.
2.20.1 TAMMAC Status Monitoring. TAMMAC status monitoring functions are based on the
existing MU and DVMS status monitoring functionality. Some of the existing cautions, advisories, BIT
mechanizations, and MSP codes that satisfied the requirements for the MU and DVMS were not
changed. However, additional status monitoring functionality was added to accommodate changes to
the MC/AMU/DMC interfaces including, but not limited to, the PC cards and the High Speed
Interface Bus (HSIB).
All existing display references to the MU and DMC are unchanged. The TAMMAC AMU functionality
discussed here equates to the MU nomenclature on all existing displays. The TAMMAC DMC equates
to the DMC on all existing displays. The only operator change to the status monitoring BIT displays
is the addition of the AMU MAINT option.
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Figure 2-65. Video Display Routing
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Figure 2-66. AMU Maintenance Format
The MC verifies the AMU and DMC software configuration IDs are compatible with the MC
software.
2.20.2 AMU Maintenance Format Options and Display Information. The AMU maintenance
format contains two relay mode options, MAP LOAD and MBIT. MAP LOAD provides access to
sublevel formats used to upload map theater data to the DMC. MBIT is used for troubleshooting and
fault isolation.
Information displayed on the AMU maintenance format is limited to the MU OFP configuration ID.
The OFP CONFIG identifies the OFP version currently installed in the AMU. See figure 2-66.
2.20.3 Map Theater Data Loading. Map theater data loading can include either updates to an
existing theater load or a new theater load. In both cases, the map loading cards are processed on JMPS
and loaded in the DMC nonvolatile mass memory using the same procedure. The number of map
loading cards is contingent on the size of the update or new theater load. The number of map loading
cards can be as few as one or as many as seven.
Map theater data loading is controlled by the map loading format. The map loading format is
accessed by selecting the MAP LOAD option (PB 11) on the AMU maintenance format as shown in
figure 2-67. The map loading format contains 3 relay mode options; LOAD, ABORT, and RTN.
2.20.4 Map Loading Format Options. The LOAD option (PB 11) is used to initiate a theater load
when a map loading card is installed in the AMU maintenance card receptacle and the AMU door is
closed. Once the process is initiated, the LOAD option is removed from the format. Multiple card
theater loads/updates require the insertion of another map card when prompted by the DDI display.
Installing another map loading card and closing the AMU door continues the theater loading process.
This procedure is repeated until all theater data is loaded. Map loading cards can be loaded in any
sequence.
The ABORT option (PB 13) is used to abort a theater load in progress. It is removed from the
display prior to and after completion of a load. To initiate an abort, the operator presses PB 13 which
changes the option to ABORT ENABLE. Selection of ABORT ENABLE executes the abort process.
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Figure 2-67. Map Loading Format
The ABORT ENABLE legend is displayed for 3 seconds following selection of the ABORT option. If
the operator does not select the ABORT ENABLE option within three seconds, the ABORT legend is
redisplayed. The intent of this two-step abort process is to preclude any inadvertent operator initiated
aborts from being performed. This abort process is permanent and the entire load process will be
terminated.
Aborting a theater map load causes the aborted theater to be deleted. No maps in CHRT, DTED,
or CIB will be available if a theater load is aborted. Previous theater maps are automatically deleted
upon loading a new theater.
The RTN option (PB 15) is used to return the AMU maintenance format if a theater load is not in
progress. If a theater load is in progress, the RTN option is removed from the display.
2.20.5 Map Loading Format Status Information. Status information displayed on the map loading
format provides the operator with on-line instructions and associated feedback necessary to perform a
successful theater load. The information presented is grouped into several status fields; THEATER,
STATUS, CARD/STATUS, CARD, and OPER.
The THEATER status field contains; theater identification, theater update revision letter, and the
theater update version number that is stored on the map loading card currently being loaded. If at least
one card is not installed in the AMU (to initiate the load), the field is blank. After a map loading card
is loaded the theater identification information remains displayed as other cards are loaded.
The status field below the THEATER status field contains the overall status of the loading process.
This field contains one of the following status indications:
HALTED - Indicates the load process has been halted (between successive cards).
LOADING - Indicates the load process has been initiated and/or is in progress.
ABORTED - Indicates the load process has been aborted.
NO DMS COMM - Indicates HSIB communications with the DMC have failed.
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DMS FULL - Indicates the DMC nonvolatile mass memory is full.
LOAD ERROR - Indicates the load process has failed.
WRONG CARD - Indicates the card installed in the AMU maintenance card receptacle is not a Map
loading card.
COMPLETE - Indicates the load process has successfully been completed.
CARD/STATUS fields contain status information regarding the PC cards used in the loading
process. The CARD field indicates the card ID number(s) in the theater load card set. The maximum
number of card IDs that can be displayed is seven. The card ID number(s) displayed is dependent on
which order the cards are loaded. The STATUS field below the CARD field contains the actual load
status of the card number directly above it. Once the card is installed and the load process initiated,
the field contains one of the following status indications:
L - Indicates the card is currently being loaded.
F - Indicates the card has failed to load properly.
C - Indicates the card has been successfully loaded.
If none of the above conditions exists, the STATUS field will be blank. If any card fails to load
properly resulting in an ″F″ status, the operator has the option of reinserting the card in an attempt to
obtain a successful load.
The CARD field contains the load status of the card that is currently installed. The card ID number
will be displayed followed by the percent complete (%) for the card. The percentage will be displayed
in 1% increments.
The OPER field contains instructions for the operator. The field contains one of the following
instructional status indications. If any card has yet to be installed for loading, the OPER field will be
blank.
INIT LOAD - Indicates the AMU is ready to start the load process and the LOAD option needs to be
selected.
CLOSE DOOR - Indicates the AMU door needs to be closed.
REMOVE CARD - Indicates the installed card has been successfully loaded and needs to be removed.
INSERT CARD - Indicates another card is required to complete the load process.
2.20.6 Map Loading Interruptions. Interruptions to the map theater data loading process can occur
as a result of several events: loss of power to the MC, DMC, or AMU, operator aborts, or inadvertent
transfers out of AMU relay mode.
If the AMU experiences a power loss greater than 5 seconds, or if the DMC experiences a power loss
of any duration, or if the operator initiates an abort, the interruption in the loading process results in
a nonrecoverable abort and the load process cannot be recovered without reloading all the cards.
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If power is reapplied to the AMU within 5 seconds, then the load process can be recovered with
minimum impacts. Once the operator reselects the map loading format, the status of the load prior to
the interruption is reflected on the format status fields. If a card was in the process of being loaded
when the interruption occurred, its status is blank indicating it has not been loaded. Selecting the
LOAD option reinitializes the load process following this type of interruption.
2.20.7 AMU/PC Cards Cautions and Advisories. The AMU has the ability to trigger 3 caution and
5 advisory messages. The caution messages are: MU LOAD, ERASE FAIL, and S/W CONFIG. The
advisory messages are: Maintenance Card Advisory (MNTCD), Mission Card Advisory (MSNCD),
Classified Data Advisory (CDATA), AMU Full advisory (AMU FL), and the BIT advisory.
The MU LOAD caution is generated when the AMU door is open; if the AMU fails; if the AMU
declares a card interface fail; if the AMU is mux fail or not ready; if the mission card is improperly
formatted, not installed, or is declared failed by the AMU, if the initialization data is not downloaded,
if an incorrect checksum is calculated. The MU LOAD caution is disabled while the AMU is in relay
mode or the aircraft is in flight.
The ERASE FAIL caution is generated when the AMU has failed to erase its internal RAM memory
buffer following a classified data transfer.
The S/W CONFIG caution is generated if the AMU and MC software are not compatible. When an
AMU OFP checksum failure occurs, the AMU OFP software configuration ID displayed on the S/W
configuration BIT sublevel format indicates XXXXXXXX.
The MNTCD advisory is generated when the AMU door is open, if the maintenance card is not
installed or properly formatted, or if the AMU declares a maintenance card failure. The advisory only
displays with WonW and clears in flight.
The MSNCD advisory is generated when the AMU door is open, or if there is an AMU/PC card
interface fail, or if a down load of data is incorrect, or there is a checksum failure with the data
downloaded, or if the mission card is not installed, or if the mission card is not properly formatted, or
if there is a mission card failure. The advisory only displays with WonW and clears in flight.
The BIT advisory is generated when the AMU is degraded or an AMU RAM classified erase failure
occurs.
The CDATA advisory is generated when the mission card contains classified data. It is removed
when a successful classified data erase of all avionics has been performed, or a successful classified data
erase of all avionics except the mission card is performed and the ERASE (MU HOLD) option has been
selected on the MUMI format.
The MU FL advisory used for the existing MU is changed for the AMU. The MU FL advisory
indicates a data wraparound has occurred on the maintenance card and the corresponding MSP code
(809) is set. When the MC determines there is not enough memory on the maintenance card to perform
the next sequential write operation, it begins overwriting previously recorded data.
DFIRS data download requests do not cause a data wraparound to occur. If there is insufficient
memory available, based on the current sequential write address pointer, the DFIRS data download
request is not executed.
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2.21 COUNTERMEASURES DISPENSING SYSTEM
2.21.1 ALE-47 Countermeasures Dispensing Set. The ALE-47 countermeasures dispensing set can
be manually actuated or can use information from various Electronic Warfare (EW) systems to
generate countermeasures dispensing programs. Refer to A1-F18EA-TAC series for details on ALE-47
operation and displays.
2.21.1.1 DISPENSER Switch. The DISPENSER switch, located on the center pedestal, is used to
control power to the ALE-47 system and to enable the BYPASS dispensing mode.
BYPASS
Selects the BYPASS mode for ALE-47 operation.
ON
Powers ALE-47. Enables the ALE-47 sublevel on the EW format.
OFF
ALE-47 off
2.21.1.2 ALE-47 Advisories. The D LOW advisory is displayed when expendable loadouts drop to the
BINGO level set on the ALE-47 sublevel of the EW format. The D BAD advisory is displayed when a
dispense misfire occurs.
2.21.2 ALE-50 Decoy Dispensing Set. The ALE-50 decoy dispensing set provides an expendable
towed RF countermeasures capability. The system includes a multi-platform launch controller
(MPLC) and a removable dispenser with 3 expendable decoys. Refer to A1-F18EA-TAC series for
details on ALE-50 operation and displays.
2.21.2.1 JAMMER Switch. The JAMMER switch, located on the center pedestal, is used to control
power to the ALE-50 system and to provide a secondary means to sever a deployed towline.
CUT
Severs a deployed towline.
ON
Powers ALE-50. With WonW, initiates start-up BIT. Enables the ALE-50 sublevel
on the EW format.
OFF
ALE-50 off
2.21.2.2 ALE-50 BIT Anomalies. There are currently three anomalies related to ALE-50 BIT, which
can corrupt decoy inventory or damage the MPLC. These anomalies have not yet been corrected, so
care should be taken when ALE-50 BIT is run.
1. When a full ALE-50 dispenser has been installed between flights, a full ALE-50 start-up BIT
must be run with WonW, in order to inventory all decoys. If ALE-50 BIT is not run WonW, the
system defaults to the last known inventory (which is two if one was dispensed on the previous
flight).
2. If the JAMMER switch is placed to OFF while the ALE-50 start-up BIT or IBIT is running, the
decoy inventory can be corrupted and/or the SEVER caution can be inhibited (e.g., no cockpit
indication of a failed decoy is provided).
3. If the ALE-50 BIT is run post-flight after one or more decoys has been dispensed in flight, the
system may arc and damage the MPLC.
Therefore, to avoid the potential loss of ALE-50 operability, aircrew should perform the following
procedures on flights where ALE-50 use is anticipated. Run the ALE-50 start-up BIT and/or IBIT
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with WonW prior to flight. Do not place the JAMMER switch to OFF while ALE-50 BIT is running.
If this occurs, place the JAMMER switch back to ON and run another start-up BIT. Turn the
JAMMER switch OFF in flight prior to landing or, at a minimum, do not cycle the JAMMER switch
OFF then ON or initiate ALE-50 IBIT when WonW post-flight.
2.21.2.3 SEVER Caution. The SEVER caution is set under the following conditions: (1) the hook or
landing gear is lowered with an ALE-50 decoy deployed, (2) a deployed decoy fails (AUTO DEPLOY
option not selected), or (3) the signal was sent to sever a towline but the squib did not fire.
2.22 BIT-STATUS MONITORING SUBSYSTEM
The BIT/status monitoring subsystem, provides the aircrew with a simple display of system status.
Most information is derived from BIT mechanizations within the avionics sets and from non-avionic
built in tests implemented in the computer software for other aircraft subsystems.
The subsystem monitors engine and airframe operational status for unit failures and caution/
advisory conditions when the mission computer system is operating. When the mission computer
system detects a caution/advisory condition, it commands display of the applicable caution or advisory
message on one of the DDIs. If the mission computer system detects a unit failure, it commands the
subsystem to store the applicable maintenance code. The mission computer displays the subsystem
BIT results on one of the DDIs.
Non-BIT equipment status includes configuration ID numbers and INS terminal data.
2.22.1 FIRAMS - Flight Incident Recorder and Aircraft Monitoring Set. The FIRAMS consists of a
signal data computer, a data storage set, an engine fuel display, and a maintenance status panel. The
FIRAMS monitors selected engine, airframe, avionic, non-avionic, fuel gauging and consumable
signals. It also performs conversion of sensed measurements, provides real time clock function, outputs
discrete and analog data to associated equipment, communicates with the mission computer, displays
fuel quantities and engine parameters, and performs fuel system health monitoring. FIRAMS also
provides nonvolatile storage for flight incident, maintenance, tactical and fatigue data, and bulk data
input of tactical mission planning data.
2.22.2 DFIRS - Deployable Flight Incident Recorder Set. The DFIRS system consists of the signal
data recorder (SDR), the data transfer interface unit, and the pyrotechnic release system. The SDR
consists of the flight incident recorder memory, beacon, battery, and antenna. DFIRS is contained in
a deployable aerodynamic airfoil located on the top of the fuselage between the rudders. The DFIRS
system stores up to 30 minutes of flight data and, when activated, deploys the SDR along with a rescue
beacon in an airfoil. The SDR is deployed upon pilot ejection or on ground impact. The data stored on
the flight incident recorder (FIR) is gathered by the mission computer from aircraft systems. DFIRS
records flight data, cautions, advisories, and spin data. The FIR memory wraps around to the
beginning when the end of memory is reached. Only the last 30 minutes of each flight is retained. The
MC controls the rate and the type of data that is stored. DFIRS data recording starts when both
throttles are advanced past the vertical, launch bar is lowered, ground speed exceeds 50 knots, or
WoffW and airspeed is over 80 knots. DFIRS recording stops 1 minute after WonW, both throttles less
than the vertical, and the ground speed less than 50 knots. All data with SPIN mode activated are
automatically recorded. A DFIRS DWNLD option is available on the engine display with WonW.
Selecting this option downloads the DFIRS data to the MU for easier retrieval.
2.22.3 Avionics BIT. In most instances, two types of BIT are mechanized, periodic and initiated.
Periodic BIT begins functioning at equipment power application. It provides a failure detection
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Figure 2-68. Flight Aids Reversion Mechanization
capability that is somewhat less than that provided by initiated BIT in that it does not interfere with
normal equipment operation.
Two forms of BIT derived data are supplied to the MC. One form is validity information associated
with selected data. The second form is equipment failure information which identifies failed
assemblies. The MC uses these two forms of BIT data to implement reversion operation and advisories
for the aircrew as well as equipment status displays for both the aircrew and maintenance personnel.
2.22.3.1 Reversion. When the BIT equipment determines that a function has exceeded a predetermined threshold, the data derived from that function is immediately indicated as not valid. The MC,
upon receiving this indication, reverts to the next best available source. This source is, in many cases,
as accurate as the original source. This reversion is maintained as long as the data remains invalid from
the primary source.
Figure 2-68 illustrates this concept for the flight aids. For each unit in the primary path, there is at
least one alternate source of data for reversion. The aircrew is provided appropriate display cueing only
when a reversion results in some loss of capability or performance. If the ALT switch is in RDR and
the radar altimeter fails, the MC removes the displayed radar altitude, replaces it with barometric
altitude, and replaces the ″R″ cue with a flashing ″B″ cue. If altitude is lost from the FCC and the
altitude switch is in BARO, the MC removes the displayed altitude from the HUD. These examples
illustrate 3 forms of degraded mode advisories: (1) reversion to an alternate data source of equivalent
accuracy with no cueing; (2) reversion to an alternate data source of lesser accuracy with cueing; (3) and
removal of displayed data when no acceptable alternate source is available. Refer to Part VIII for
further discussion on weapon system reversions.
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2.22.3.2 Equipment Status Displays. Equipment status displays (BIT, caution, and advisory)
provide the aircrew with continuous status of the avionics equipment and weapons. A cue to check
equipment BIT status is the appearance of the BIT advisory display on the caution/advisory. The
display is normally on the left DDI. A MENU selectable top level BIT format displays the status of
failed, NOT RDY, or OFF systems of all avionics equipment that interface with the MC. When the BIT
control display is selected on another display, the BIT advisory is removed until another BIT failure
occurs. For aircraft with the 8 x 10 display installed, messages appear as ACNTR. Messages displayed
as a function of equipment status are listed in figure 2-69.
Weapon and stores status is displayed primarily on the stores display (selected from the menu
display). When the BIT display indicates a stores management system (SMS) failure, the affected
stations and degree of failure are identified on the stores display as described in A1-F18EA-TACseries.
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STATUS
MESSAGE
NOT RDY
APPLICABLE SYSTEM
MESSAGE DEFINITION
All systems except MC1
Equipment OFF, not installed, or initializing.
OFF
RDR, MPCD, ACNTR, UFCD, CAM,
IFF, RALT, BCN, ILS, TCN, COM1,
COM2, D/L
Equipment OFF.
IN TEST
All systems except MC1, MC2, RWR
Initiated BIT in progress.
SF TEST
FLIR, NFLR, SMS, RDR, MPCD,
ACNTR, UFCD, RALT, WPNS, LTDR,
DFIRS, GPS ALE-47, ALE-50, DBFS,
ECS, FADEC
Self test in progress (cannot be operator
terminated).
All systems
Initiated BIT completed without failure.
All systems except MC1, MC2
Failure detected; equipment operation
degraded.
DEGD
+
OVRHT
LDT, FLIR, NFLR, SMS, MPCD,
ACNTR, UFCD, CAM, CSC, FCSA,
FCSB, INS, ASPJ, RWR, LTDR,
DFIRS, ALE-50, FADEC
Detected failure and overheat.
OVRHT
LDT, FLIR, NFLR, SMS, MPCD,
ACNTR, UFCD, CAM, CSC, FCSA,
FCSB, INS, ASPJ, RWR, LTDR,
DFIRS, ALE-50, FADEC
Overheat.
GO
DEGD
MUX FAIL CLC, FLIR, NFLR, SMS, RDR, LDDI,
RDDI, MPCD, ACNTR, CSC, MC2,
FCSA, FCSB, INS, COM1, COM2, D/L,
ASPJ, AISI, SDC, MU, LTDR, DMC,
DFIRS, GPS, ALE-47, ALE-50, ECS,
FADEC, UFCD (Aft)
Equipment is not communicating on
AVMUX and on/off discrete is set to
on.
RESTRT
All systems except MC1, MC2, RWR,
FADEC
Reinitiate BIT; equipment did not respond to BIT command, remained in
BIT too long and was terminated by
MC.
OP GO
NFLR, SMS, COM1, COM2, WPNS,
MU, DFIRS, GPS, ALE-47, ALE-50,
FADEC
Non critical BIT failure detected.
PBIT GO
All systems except MC1, MC2, RWR,
FADEC
Initiated BIT has not been run since
ground power-up and PBIT is not reporting any failures.
No indication (blank) adjacent to the equipment legend indicates that initiated BIT has not been
run on the equipment and that the periodic BIT has not detected any faults. LDDI, RDDI, MPCD,
and EFD have unique degraded messages of ALDDI, ARDDI, AMPCD/ACNTR, AEFD, and
AUFCD in the F/A-18F to allow distinguishing BIT status failures for aft cockpit displays.
Figure 2-69. Equipment Status Messages
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Figure 2-70. Caution/Advisory Displays
2.22.3.2.1 Cautions and Advisories. Cautions and advisories are displayed on the left DDI except
when the left DDI is used for BIT display or weapon video (figure 2-70). When the left DDI is off or
failed, or when the LDDI is used for BIT or weapon video, cautions and advisories are displayed on the
center display. If the left and center displays fail or are turned off, the right DDI displays the cautions
and advisories. Cautions and advisories automatically move to the center display when BIT is selected
on the LDDI. Caution displays appear as 150%-size letters compared to the normal message symbology
size. Cautions are displayed as they occur beginning in the lower left portion of the DDI display and
sequence to the right up to 3 displays across. The fourth caution reindexes to the left edge above the
first caution. A dedicated caution display automatically replaces the HSI display if the number of
cautions exceeds 3 lines. Advisory displays appear as 120%-size letters on a single line beneath the
caution displays. The advisories are preceded by an ADV- legend and the individual advisories are
separated by commas. A caution or advisory is removed when the condition ceases. If there is a caution
or advisory displayed to the right of the removed caution or advisory the display remains blank.
Pressing the MASTER CAUTION light when the light is out repositions the remaining cautions and
advisories to the left and down to fill the blank displays. When a caution occurs, the MASTER
CAUTION light on the main instrument panel illuminates and the MASTER CAUTION tone or a
voice alert is heard in the headset. The MASTER CAUTION light is extinguished by pressing the light.
Refer to Warning/Caution/Advisory Displays in chapter 12 for the display implications and corrective
action procedures.
2.22.3.3 BIT Initiation. In addition to displaying equipment BIT status, the BIT top level and nine
sublevel displays (figure 2-71) are used to command initiated BIT. Those avionics set groups identified
by the legends on the top level display periphery have an initiated BIT capability. BIT may be initiated
for all operating units simultaneously except for some BIT that cannot be performed in flight. Figure
2-71 shows which initiated BIT are not allowed in flight. Additional steps are required to test the INS
and FCS. BIT for individual units within groups may be initiated through the BIT sublevel displays.
Pressing BIT returns to the BIT top level display. Pressing STOP or MENU when BIT is in progress
terminates initiated BIT. Performance of BIT assumes that the required electrical and hydraulic
power is applied to the equipment tested.
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Figure 2-71. BIT Control Display (Sheet 1 of 2)
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Figure 2-71. BIT Control Display (Sheet 2 of 2)
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2.22.3.3.1 All Equipment. Simultaneous initiated BIT of all equipment installed is performed by
selecting AUTO on the BIT top level display. Equipment group, acronym and status are displayed at
the display options. Equipment group status indicates the lowest operating status reported by any unit
in the tested group. Individual system status results other than GO, PBIT GO, IN TEST, SF TEST,
and OP GO are displayed with a system acronym in the center of the display. If the equipment list is
too long to be displayed on one page, a PAGE pushbutton is displayed. Pressing PAGE displays the
remainder of the list that is on page 2. Pressing PAGE when page 2 is displayed returns page 1.
2.22.3.3.2 Equipment Groups. Initiated BIT of entire equipment groups is performed by selecting
SELBIT (SELBIT option becomes boxed) on the BIT top level display and the desired equipment
group pushbutton. One or more groups can be selected. Another way to select a group is to press the
group pushbutton (with the SELBIT option not boxed) on the BIT top level display and then ALL on
the group sublevel display. See figure 2-71.
2.22.3.3.3 Individual Units. Initiated BIT of an individual unit is performed by pressing the
equipment group pushbutton on the BIT top level display which contains the desired unit. The display
changes to a group sublevel display. Individual units from the group can then be tested by pressing the
pushbutton adjacent to the desired acronym. System status for all systems in the group is displayed on
the center of the display. Some systems require additional aircrew BIT input.
2.22.3.4 System BIT Steps. The following includes certain initiated BIT which require steps in
addition to pressing one of the buttons on the BIT display and reading the BIT status messages after
the test is complete. Figure 2-71 shows which initiated BIT are not allowed in flight.
2.22.3.4.1 FCS Initiated BIT (IBIT). For the FCS to enter IBIT the FCS BIT consent switch must be
held ON. This action prevents inadvertent IBIT initiation inflight for reasons of flight safety.
Control surfaces move during FCS IBIT with hydraulic power applied.
To prevent personnel injury or equipment damage, make sure personnel
and equipment are kept clear of control surfaces.
NOTE
• With the wings folded, both ailerons are Xd out, but no aileron BLIN
codes should be displayed. Even with wings folded, there are aileron
functions tested that may reveal FCS failures via valid BLIN codes.
• For FCS IBIT to start, the FCS BIT consent switch must be held for
at least 2 seconds. If not held for the required time, FCS A and FCS B
will indicate RESTRT on the BIT status line. If RESTRT is displayed, select STOP on the FCS-MC sublevel display and then repeat
the initiation procedure.
• The FCS will not enter IBIT if the throttles are above 14° THA or
NWS is engaged.
• Do not operate any FCS related switches or move the stick or rudder
pedals while FCS IBIT is running, as this may produce false failure
indications
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NOTE
• With the wings folded, a BIT status of GO will only be displayed for
approximately 2 seconds before reverting to a DEGD indication. BIT
status will return to GO when the wings are spread and locked.
• If the FCS IBIT fails, FCS A and FCS B will indicate DEGD on the
BIT status line. Note surface Xs and/or BLIN codes and contact
maintenance personnel for disposition.
1. Select MENU-SUPT/BIT/FCS-MC on right DDI.
2. While simultaneously holding FCS BIT consent switch to ON, select the FCS pushbutton on the
FCS-MC sublevel display.
3. Release FCS button and FCS BIT consent switch when FCSA and FCSB BIT status indicates IN
TEST. FCS IBIT requires approximately 1 minute.
NOTE
If IN TEST remains on the FCS BIT display longer than 2 minutes,
select STOP and actuate the paddle switch to exit FCS IBIT.
2.22.3.4.2 Preflight FCS Initiated BIT. The fly-by-wire flight control system uses redundant
hardware to provide continued safe operation after component failures. The level of redundancy
designed into the system was set by component failure rates, failure mode effects, aircraft mission time,
and survivability considerations. The ability to provide safe operation is fundamentally based on the
principle that there are no undetected (e.g., latent) failures prior to flight which would compromise
system redundancy. It is not possible to have an in-flight periodic BIT (PBIT) which can detect all
degradations in a fly-by-wire system. Many redundant pathways can be tested only by setting system
conditions that would be unsafe to establish in flight (e.g., verification of the ability to shut off an
actuator). Preflight FCS initiated BIT was designed to provide those tests and ensure the full
redundancy of the flight control system is available prior to flight. Without running preflight FCS
initiated BIT and performing the necessary maintenance, latent failures present in the system can
result in unsafe conditions should additional failures occur in flight.
2.22.3.4.3 Preflight FCS Initiated BIT Operation. Preflight FCS initiated BIT consists of a series of
tests which verify the integrity of the flight control system processors, actuators, sensors, and cockpit
interfaces.
Preflight FCS initiated BIT begins by testing lower level functions. If preflight FCS initiated BIT
detects a fault at this level which affects higher level functions, it halts and reports the fault(s). If
preflight FCS initiated BIT did not halt at this point, false BLIN codes would be generated on higher
level functions which depend upon the failed lower level function for their operation. If preflight FCS
initiated BIT detects a fault in a subsystem (e.g., left stabilator) testing of the failed subsystem is
discontinued, and testing of unrelated subsystems (e.g., rudders, trailing edge flaps, etc.) continues.
Since testing is not complete, preflight FCS initiated BIT must be run again after maintenance actions
in order to complete all tests.
2.22.3.4.4 Preflight FCS Initiated BIT PASS/FAIL. A successful preflight FCS initiated BIT results
in a GO indication on the DDI. An unsuccessful preflight FCS initiated BIT indicates a system
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degradation. A preflight FCS initiated BIT never sets an X on the DDI FCS status page (MENU-FCS)
since preflight FCS initiated BIT only sets BLIN codes. Launching in a degraded state (e.g., with BLIN
codes) places the aircraft in a situation where a portion of the flight control system is operating without
the normal redundancy.
2.22.3.4.5 Repetition of Preflight FCS Initiated BIT. If an aircraft fails preflight FCS initiated BIT
(e.g., BLIN codes present after preflight FCS IBIT) maintenance should be called to troubleshoot the
system. After completing troubleshooting, a successful preflight FCS IBIT is necessary to make sure
the system is fully operational. Except for cold weather operation, preflight FCS IBIT failure is
indicative of a component degradation, e.g., hydraulic or electrical components are out of tolerance, or
a cable conductor is intermittent (broken wire, loose connector pin, etc.).
2.22.3.4.6 FCS Exerciser Mode. In cold weather, actuator components do not respond normally until
hydraulic fluid temperature increases. Exerciser mode should be used to expedite system warm-up.
During exerciser mode, a number of PBIT actuator monitors are ignored to prevent generation of
nuisance BLIN codes. In cold weather it is appropriate to re-attempt preflight BIT after running
exerciser mode. Exerciser mode should not be used as a method to clear BLIN codes in normal start-up
temperature conditions. BLIN codes cleared in this manner could be associated with hydraulic
contamination or sticking control valves which could appear again in flight with catastrophic results.
Repeatedly running exerciser mode in normal and hot weather environments may lead to hydraulic system overheat.
2.22.3.4.7 Running Preflight FCS Initiated BIT After Flight. A good (no codes) preflight FCS IBIT
on the previous flight is no assurance against latent failures on the next flight. Electronic components
have a propensity to fail on power application. Damage can occur during deck handling or maintenance
activity not associated with the flight controls. The only insurance is to run preflight FCS initiated BIT
prior to flight.
2.22.3.5 SMS Initiated BIT. Safeguards have been built into the weapon system mechanization to
allow SMS initiated BIT to be performed on the ground with weapons loaded and cartridges installed.
During initiated BIT, weapon release signals and associated circuitry are not exercised unless all of the
following interlocks are satisfied simultaneously: MASTER ARM switch to ARM, armament safety
override in override, weapon load codes on stores processor set to zero, and no weapon ID detected on
any weapon station. SMS initiated BIT should not be attempted until the above interlocks are in a safe
condition. The SMS initiated BIT should be successfully completed within 180 seconds of initiation.
2.22.3.6 INS Initiated BIT. To perform initiated BIT, the INS must be in the TEST mode and a
ground/carrier selection must be made to indicate where the BIT is being accomplished. When the
BIT/SELBIT/NAV/ALL, BIT/NAV/INS, or BIT/AUTO is actuated, a status message of GND/CV?
appears next to the INS legend in the status display area. At the same time, GND and CV button labels
appear along the bottom of the display. These options allow entry of where the initiated BIT is
performed, e.g., on the ground or on a carrier. A third option, INS LONG (INS long test), identifies if
the platform slew portion of the INS test is desired. The platform slew test, which adds an additional
26 minutes to the normal 12 minutes INS BIT time, is only performed when the INS exhibits degraded
navigation performance during a flight and normal BIT routines do not detect any malfunction. This
option should be selected prior to selecting the ground or carrier option.
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1. Check parking brake set.
2. For ground initiated BIT insure waypoint zero is local latitude/longitude.
3. Select MENU/BIT/SELBIT/NAV or MENU/BIT/NAV/INS or MENU/BIT/AUTO or MENU/
BIT/NAV/ALL on the right DDI and TEST on the INS mode switch.
4. Select INS LONG (if required) and GND or CV on DDI, and start clock. At successful completion
of test BIT, display status message reads GO. Maximum time for INS initiated BIT is 12 minutes
and maximum time for INS initiated BIT and platform slew test is 45 minutes.
2.22.3.7 AUTO BIT. If the AUTO button is pressed, BIT are initiated in parallel for all equipment
turned ON and whose interlocks are satisfied. The test pattern associated with the DDI and HUD is
not displayed when the AUTO option is used. Approximately 2.5 minutes are required for all AUTO
BIT except FCS and INS.
1. Check power applied to all systems requiring BIT and check required interlocks in safe condition.
2. Select MENU/BIT/AUTO on DDI.
a. All systems read GO after required test period. GO indication is provided when system check
is complete and OK. Other messages may be displayed if malfunctions are detected.
3. If FCS test required, perform FCS Initiated BIT above while substituting the AUTO button for
the FCS button in the procedure. Insure the procedural warnings and notes are observed and that
the AUTO button and FCS BIT consent switch are held simultaneously to initiate test.
4. If INS test required, perform INS initiated BIT above while substituting the AUTO button for
the INS button in the INS Initiated BIT procedure.
2.22.3.8 Cockpit Displays Initiated BIT.
2.22.3.8.1 DDI/HUD Initiated BIT. Operator participation is required to detect failures and isolate
faults in the display equipment. The BIT/DISPLAYS/DDI-HUD option starts the MC generated test
patterns on the DDI and HUD immediately after each indicator BIT is concluded. The test pattern can
be compared on the 3 displays for similarity and individually for concentricity, intensity level, and
alphanumeric clarity. Options are tested by actuating all the buttons. A circle appears adjacent to the
button when the functional test is successfully completed.
With AMCD aircraft, IBIT takes approximately 20 to 90 seconds to complete (MC allows a
maximum of 200 seconds). During this interval, the displays cycle through a series of raster test
patterns (i.e., small square of raster, full screen of raster, small square again, and then a larger square
of raster). Immediately after the BIT for each display is concluded, the MC generated test pattern is
displayed on each of the displays. See figure 2-72. The STOP option on the test pattern terminates the
CRT test. The DDIs alternate color of the test pattern between green, red, and yellow.
Pressing DDI-HUD initiates BIT on two different equipment display groups. The following
procedure can be used to test one or both of the display groups by performing the appropriate parts
of the procedure.
1. Select BIT/DISPLAYS/DDI-HUD
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a. DDI and HUD displays go blank momentarily, flash IN TEST, and display a test pattern.
b. Check DDI test patterns are steady and in focus. See figure 2-72.
c. HUD test pattern flickers but remains on.
2. DDI and HUD displays - CHECK
a. Display commonality
b. Display concentricity
c. Proper intensity
d. Check right DDI pushbuttons (20) starting with the top left button on the horizontal row.
Circle is displayed next to each pushbutton after it is pressed.
e. Check BIT status messages for GO on the BIT displays. F/A-18F lists front and rear indicator
results separately.
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Figure 2-72. MPCD and UFCD Test Patterns
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Figure 2-73. EFD Test Pattern
2.22.3.8.2 EFD Initiated BIT. The EFD test pattern is initiated by performing EFD BIT using
pushbutton sequence BIT/DISPLAYS/EFD and observing the test pattern display following the
completion of EFD BIT. The EFD test pattern may be observed using the following procedure.
1. Select BIT/DISPLAYS/EFD
2. EFD - observe test display
3. Select STOP or MENU to terminate test pattern
While the test pattern is displayed, selecting the MODE button or pulling the BINGO knob toggles
between the two test patterns displayed in figure 2-73. Rotating the BRT knob does not change display
intensity. Rotating the BINGO knob increments or decrements the number displayed in the lower
right corner from 1 to 12, depending on the rotation direction (clockwise or counterclockwise).
2.22.3.8.3 MPCD Initiated BIT. The MPCD receives information directly from the MC for display
and also processes the information provided for display on the UFCD. The MPCD test pattern is
initiated by doing MPCD BIT using pushbutton sequence BIT/DISPLAYS/MPCD and observing the
test pattern display following the completion of MPCD BIT. The MPCD test pattern may be observed
using the following procedure.
1. Select BIT/DISPLAYS/MPCD
2. MPCD and UFCD - observe test displays
3. Select STOP or MENU to terminate test pattern
The BIT format can be displayed on the MPCD or UFCD during DDI/HUD BIT and on the DDI
during MPCD BIT. On F/A-18F, the MPCD option initiates BIT on both the cockpit and rear cockpit
MPCD. During MPCD BIT the message MPCD IN TEST is displayed on both the MPCD and UFCD.
When IBIT is completed, the test pattern (figure 2-72) is displayed on both the MPCD and UFCD.
2.22.3.8.4 8 x 10 Display Initiated BIT. The 8 x 10 display executes initiated BIT by selecting
ACNTR on the Display BIT format. When selected, the 8 x 10 display completes IBIT and displays a
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Figure 2-74. UFCD Test Pattern
test pattern on both the 8 x 10 display and the aft UFCD. The 8 x 10 display test pattern is the same
as the DDI test pattern. The mission computer removes the IBIT command after 10 seconds to allow
viewing of the test pattern on both displays.
2.22.3.8.5 UFCD Initiated BIT. UFCD test patterns are initiated by selecting BIT/DISPLAYS/
UFCD and observing the test pattern display following the completion of UFCD BIT. The UFCD test
pattern (figure 2-74) is obtained using the following procedure.
1. Select BIT/DISPLAYS/UFCD.
2. UFCD - Observe test display
3. Select STOP or MENU to terminate test pattern.
When BIT is initiated, UFCD IN TEST is displayed until BIT is completed. On F/A-18F, selecting
UFCD initiates BIT on both cockpit and rear cockpit displays. The UFCD test pattern is not displayed
if IBIT is not initiated on the UFCD option.
2.22.3.8.6 Radar Altimeter Initiated BIT. If BIT is initiated during RADALT time-in, the status on
the BIT display is NOT READY. If the BIT is initiated after time-in is complete, the display is GO
(indicating the radar altimeter is operating correctly), RESTRT (the BIT was not completed within
the design time limits), or DEGD (a WRA fail signal exists).
2.22.3.8.7 STOP Button. The STOP button allows the aircrew to stop initiated BIT at any time. BIT
is also stopped by pressing MENU, although MENU is not available with the DSPL/EPI/EFD/UFCD
BIT test pattern displayed. When the STOP (or MENU) button is pressed, any test in progress stops
and the equipment returns to normal operation. Exceptions to this are the radar and SMS power-on
BIT and the COMM 1/2, D/L, and TACAN BIT. The radar and SMS power-on BIT cannot be
terminated and indicate SF TEST when the MC detects the system is in BIT without having been
commanded to do so. The same is true of the COMM 1/2, D/L, and TACAN equipment which does a
canned non-interruptable BIT sequence. The mission computer terminates initiated BIT for any
equipment that it determines has taken too long to complete the test.
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2.22.3.9 Hydro-mechanical (HYDRO MECH) Initiated BIT. With SELBIT boxed on the top level
BIT format, selecting HYDRO MECH initiates ECS BIT if aircraft is WonW.
FADEC status is also provided on the HYDRO MECH display by L and R ENG A and B status
indications.
2.22.3.10 BLIN Codes. BIT Logic INspection (BLIN) codes are octal readouts identifying FCS
failures and can be read from the FCS status display. The following procedures may be used to display
and record BLIN codes. Channel 1 BLIN codes are displayed. Pressing the BLIN button displays the
next channel (1, 2, 3 and 4) BLIN codes.
1. On DDI - PRESS MENU-SUPT/FCS/BLIN
2. DDI BLIN codes - RECORD BY CHANNEL
3. Press BLIN button to view next channel BLIN codes
2.22.4 Non-Avionic BIT. Non-Avionic BIT is implemented in selected hydro-mechanical subsystems
primarily for the purpose of displaying subsystem status in the cockpit(s) (cautions and advisories)
and/or providing fault detection and fault isolation information for maintenance personnel. This status
data is provided to the status monitoring displays by the signal data computer which interfaces with
the following hydro-mechanical areas:
1. Engine/Secondary Power
2. Electrical
3. Hydraulics and landing/arresting gear
4. Fuel
5. Environmental control system and liquid cooling system
6. Controls/mechanisms/miscellaneous
The hydraulic system pressure cautions are interfaced directly by both mission computers, providing
redundancy for safety of flight.
2.22.4.1 Equipment Status Displays. NABIT cautions and advisories are displayed in the same
manner as avionics cautions and advisories.
2.22.5 Status Monitoring Backup. MC2 provides backup status monitoring if MC1 fails. It provides
an MC1 caution on the DDI indicating that MC1 has failed.
NOTE
With non-AMCD aircraft, if MC1 fails, all DDI cautions and advisories
are lost except MC1, HYD 1A, HYD 1B, HYD 2A, and HYD 2B. TAC
MENU loses the SA option and SUPT MENU displays only the HSI
option.
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Figure 2-75. CONFIG Display (Sample)
NOTE
With AMCD aircraft, if MC1 fails, all DDI cautions and advisories are
available.
2.22.6 Non-BIT Status. Equipment status derived by means other than BIT include DDI configuration display ID numbers and INS terminal data.
2.22.6.1 CONFIG Display Country ID Code. The country identifier code USN is displayed under the
CONFIG legend.
2.22.6.2 CONFIG Display. The ID numbers of the current operational flight program (OFP) loads for
the radar, stores management system, CLC, INS, mission computers, communication system control,
flight control computer, FLIR, GPS, HMD, SDC, MU, LDT, DMC, MPCD, DDI, EW equipment,
DFIRS, FADECs, and ECS controller can be determined by selecting the configuration display (see
figure 2-75). The configuration display is selected by the following procedure:
1. Select BIT from the SUPT MENU.
2. Select CONFIG.
With the configuration display selected, the current ID numbers are displayed to the right of the
equipment acronym. Refer to figure 2-75 for an example.
2.22.6.2.1 MC CONFIG Caution. An MC CONFIG caution indicates MC1 and MC2 OFP loads are
incompatible.
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2.22.6.2.2 S/W CONFIG Caution. With the exception of JHMCS, a S/W CONFIG caution indicates
MC1 and MC2 OFP loads are not concurrent releases (incompatible) or an avionic equipment
processor OFP is incompatible with MC OFPs. The incompatible OFP(s) are indicated by a line drawn
through the OFP ident. If the MC OFPs are incompatible a line is drawn through both MC OFP idents.
For JHMCS, a S/W CONFIG caution and a line through the HMD S/W configuration line on the
configuration display indicates a mismatch in the BuNo in the Magnetic Compensation Data file and
the BuNo in the MC, or failure to load the initialization file into the EU.
2.22.6.2.3 OVRD Button. The override option allows the pilot to override the software configuration
logic when the software country ID codes do not agree with the aircraft country ID codes.
2.22.6.3 INS Terminal Data. INS terminal data can be obtained if at least one update has been
performed after flight with the parking brake on. Terminal data is displayed by selecting the following
options in sequence: MENU, BIT, MAINT, INS, and POST. Note the PER (performance error rate)
and navigation time on the FLIGHT 1, POST 1 display (see figure 2-76). Select the POST option again
and note the velocity on the FLIGHT 1, POST 2 display. Turn the INS mode selector knob to OFF.
With GPS operating, if the aircraft is flown with IAF selected, the performance error rate does not
include the time flown in the AINS mode.
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Figure 2-76. INS Postflight Data Display
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2.23 JOINT HELMET MOUNTED CUEING SYSTEM (JHMCS)
The JHMCS allows the aircrew to target and employ existing short range missiles (SRMs) and High
Off-Boresight (HOBS) weapons, such as the AIM-9X, and cue the radar, and other sensors. When
using JHMCS to employ HOBS weapons, the aircrew can slave/acquire and shoot targets beyond the
gimbal limits of the aircraft radar and designate ground targets. The main display provides a
monocular 20° field of view that is visible in front of the pilot’s right eye.
The main components of the JHMCS include the helmet mounted displays, electronics unit,
HMD/AHMD off/brightness knobs, aft cockpit Boresight Reference Unit (BRU), cockpit units,
magnetic transmitter units, and seat position sensors in each cockpit. The JHMCS aircraft-integrated
components can be flown with or without the helmet system.
Increased weight and forward CG of the helmet will increase neck strain
during high or sustained g flight maneuvers.
NOTE
All aircrew shall receive simulator or dedicated ground training on
JHMCS helmet controls and displays prior to flight with JHMCS.
2.23.1 Helmet Mounted Display (HMD)/Aft Helmet Mounted Display (AHMD). Each HMD consists
of the helmet, Helmet Display Unit (HDU), Helmet-Vehicle Interface (HVI), and a universal connector
which connects the HDU to the helmet.
2.23.1.1 Helmet Display Unit (HDU). The HDU includes a CRT, Magnetic Receiver Unit (MRU),
camera, auto-brightness circuitry, uplook reticles, and visor. Aircrew can remove the HDU and
configure the helmet to accommodate the AN/AVS-9 night vision goggle system.
Ensure the HMD/AHMD OFF/BRT knob(s) are OFF before removing the HDU, and store the HDU
in the JHMCS stowage bag on the right bulkhead. See figure 2-78.
• To keep water out of the HDU on aircrew helmets and prevent the
possibility of electric shock, ensure the HDU cover is installed on the
helmet.
• To prevent damage to the HDU, do not expose the HDU to a
temperature exceeding 50°C (122°F) operationally or in storage.
2.23.1.2 Helmet-Vehicle Interface (HVI) Connectors. The HDU is connected to the aircraft by the
HVI, which consists of 3 connectors. These connectors are the Quick Disconnect Connector (QDC),
In-Line Release Connector (IRC), and Helmet Release Connector (HRC). The upper HVI is the
portion of the HVI from the helmet to the QDC. The lower HVI is the portion of the HVI below the
QDC that is installed in the aircraft.
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The upper HVI is routed under the survival vest (if worn) through the JHMCS bundle flue on the
torso harness. The QDC is seated in a Quick Mounting Bracket (QMB) attached to the lower left hand
leg strap of the torso harness. See figure 2-77.
The JHMCS upper HVI (UHVI) must be properly routed through the
JHMCS bundle flue under the survival vest and the QDC secured in the
QMB to ensure that no entanglement exists with the oxygen hose.
Misrouting of the JHMCS UHVI may allow the QDC to rub against the
oxygen hose disconnect causing unintentional oxygen/communications
disconnect in-flight.
The QDC is the primary disconnect for ejection, and both normal and emergency ground egress. The
QDC can be manually disconnected for normal egress by pushing a button on top of the QDC and
separating the top half. During an ejection or emergency egress the QDC is disconnected via an aircraft
mounted lanyard when a force of 15 to 25 pounds is applied. When the QDC is not connected, the
aircraft QDC should be properly stowed in its receptacle. When the QDC is not properly connected and
the system is on, an HMD/AHMD advisory is generated.
The JHMCS QDC must be properly attached to the aircrew torso harness
QDC mounting bracket to avoid possible death or severe injury during
ejection.
• Low voltage is present on the exposed QDC pins when the HMD/
AHMD OFF/BRT knob(s) are not in OFF. To prevent a minor
electrical shock from contact with exposed pins, ensure the HMD/
AHMD OFF/BRT knobs are OFF whenever the QDC is disconnected
or connected.
• To prevent damage to the QDC and aircraft components, ensure the
aircraft QDC is properly stowed in its receptacle when not mated to
the aircrew’s QDC.
NOTE
Ambient cockpit temperatures at or below 0°C (32°F) may cause
inadvertent HMD/AHMD advisories during preflight. Warming of the
QDC and quick mount bracket should remove the failure indication if
temperature is the cause.
The IRC is a back-up disconnect which functions in the event of QDC failure. The IRC is attached
to the left aft console and requires a force of 80 to 120 pounds to disconnect.
The HRC allows the cable to disconnect should the helmet be lost during ejection. The HRC
connects to the left shoulder harness and requires a force of 80 to 120 pounds to disconnect.
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HELMET RELEASE
CONNECTOR (HRC)
INLINE RELEASE
CONNECTOR (IRC)
JHMCS
BUNDLE FLUE
QUICK
DISCONNECT
CONNECTOR (QDC)
Figure 2-77. JHMCS Upper HVI Routing
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ORIGINAL
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2.23.2 Electronics Unit (EU). The EU contains the main system CPU, LOS module, graphics
processor/display drive (one for each HMD), and low voltage power supply. The CPU controls system
bus interfacing, display list generation, BIT, and other system functions. The LOS module calculates
helmet LOS while the graphics processor/display drive processes the display list and generates the
helmet display. The MC interfaces with the EU via the MUX bus. It is located in the rear cockpit.
2.23.3 Cockpit Unit (CU). The CU contains the system high voltage power supply for helmet
displays. On the F/A-18E, it is located in the upper equipment bay, and on the F/A-18F, both CUs are
located in the aft crew station.
2.23.4 Magnetic Transmitter Unit (MTU). The MTU is used to generate a magnetic field used to
determine HMD/AHMD position/orientation by the HMD MRU receiving the magnetic field and then
sending the received signal to the EU. One MTU per cockpit is mounted on the canopy frame aft of the
pilot/WSO’s left shoulder.
The MTU is energized when the HMD/AHMD is turned on. Warm-up time for the MTU is 15 to 20
minutes. System accuracy may drift up to 0.5° if the HMD/AHMD is aligned before MTU warm-up
is completed. An additional 5 to 10 minutes should be added to the MTU warm-up time if operating
in extremely cold temperatures (e.g., -40°C).
NOTE
To maintain system accuracy, run initial HMD/AHMD alignment, or
an additional HMD/AHMD alignment, after system is warmed-up.
2.23.5 Boresight Reference Unit (BRU). On the F/A-18F, the BRU is located on top of the
instrument blast shield and dust cover. An alignment cross is provided inside the BRU to permit coarse
and fine alignment of the AHMD to the aircraft reference. See figure 2-79.
2.23.6 Seat Position Sensor (SPS). The SPS is a linear potentiometer which indicates ejection seat
height to the JHMCS. It is mounted to the aft right side of the ejection seat. There is one SPS for each
ejection seat. This seat position information allows the JHMCS to compensate for disruption of the
magnetic field in the cockpit as the metal in the seat changes position when the seat is raised or
lowered.
2.23.7 HMD/AHMD OFF/BRT Knobs. The front cockpit HMD OFF/BRT knob is located on the spin
recovery panel. This knob removes and applies power to the HMD, and adjusts HMD display
brightness. See figure 2-78.
A BRU/HMD OFF/BRT stacked knob, located on the aft cockpit INTR LT control panel, removes
and applies power to the BRU/AHMD, and adjusts BRU/AHMD display brightness. See figure 2-78.
2.23.8 HUD Video Record Panel. A switch has been added to the HUD video record panel so that
HMD or LDDI can be selected while the other switch is used to select HUD or RDDI for video
recording.
2.23.9 Cautions/Advisories. When the BuNo in the Magnetic Compensation Data file does not
match the BuNo in the MC, or if the MC fails to download the initialization file to the EU, the MC sets
the SW CONFIG caution and displays a line through the HMD S/W configuration line on the
configuration display. An HMD/AHMD advisory is reported if the QDC is not properly secured to
Quick Mounting Bracket (QMB) or is disconnected, the HDU is not properly connected, or the coarse
alignment is invalid or has not been performed.
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Figure 2-78. HMD Controls
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ORIGINAL
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Figure 2-79. Boresight Reference Unit
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Figure 2-80. Displays BIT Sublevel
2.23.10 Configuration Check. When the JHMCS system is turned on, the MC provides the EU with
the aircraft make, model, and tail number.
2.23.11 Built-In Test (BIT). The JHMCS BIT system includes automated start-up BIT (SBIT) and
initiated BIT (IBIT), and displays a BIT status message. See figure 2-80 for the DISPLAYS BIT
sublevel display.
2.23.11.1 Start-up BIT (SBIT). When the HMD system is turned on, SBIT starts automatically and
the internal software is loaded in the EU. SBIT cannot be stopped until it is completed. PBIT GO or
DEGD, as appropriate, is displayed when SBIT is completed.
2.23.11.2 Initiated BIT (IBIT). IBIT is performed when the HMD (PB 11) option is selected in either
cockpit on the DISPLAYS BIT sublevel display. ENTERING IBIT flashes on both HMDs, and an
initiated BIT is performed on both helmets. When IBIT is complete, a series of four test patterns,
which are automatically changed each second, are displayed on the HMD/AHMD. See figure 2-81. The
test patterns are displayed until the STOP (PB 10) option is selected. If the ALL (PB 6) option is
selected, IBIT and HMD test patterns are performed.
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ORIGINAL
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Figure 2-81. HMD/AHMD Test Patterns
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ORIGINAL
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2.23.11.3 Status Messages. Refer to the following for status messages and associated descriptions:
MESSAGE
DESCRIPTION
STATUS
MUX FAIL
Equipment ready discrete is high but the EU is
not communicating on either MUX bus to the
MC
NOT RDY
Equipment ready discrete is low and the EU is
not communicating on either MUX bus to the
MC
IN TEST
Initiated BIT in progress
RESTRT
Re-initiate BIT, EU did not respond to the
IBIT command or IBIT did not complete
within 30 seconds
DEGD
EU has detected a failure which degrades system performance
OVRHT
EU has reported a component as overheated
DEGD+OVRHT
EU has detected a failure and EU has reported
a component as overheated
GO
EU responded with no failures.
OP GO
EU has detected a failure which does not degrade system performance
PBIT GO
EU responded with no failures prior to performing IBIT
2.23.11.4 BIT Recording On The Memory Unit. The MC records any failure or degrade reported by
the EU to the memory unit for fault reporting and isolation.
2.23.11.5 Overheat Condition. The system has the capability to detect an equipment overheat
condition. When an overheat condition is detected the system is automatically shut down to prevent
equipment damage.
2.23.12 JHMCS Alignment. The JHMCS must be boresighted (aligned) with the aircraft prior to
every flight. Selecting the ALIGN (PB 20) option on the HMD format boxes ALIGN, selects coarse
alignment mode, and displays the FINE (PB 1) alignment option after the coarse align function is
complete. See figure 2-82.
The forward and aft helmets are aligned independently. Selecting the ALIGN (PB 20) option on the
HMD format from the aft cockpit initiates aft HMD align. When in aft coarse or fine align mode, the
MC assigns the right hand controller undesignate and Designator Control (DC) switches to the HMD.
Both forward and aft HMD alignments function identically with the exception that the aft helmet is
aligned to the BRU mounted on top of the instrument blast shield and dust cover.
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2.23.12.1 Coarse Alignment. An alignment cross is displayed on the HUD/BRU and on the
HMD/AHMD. See figure 2-82. The pilot moves the HMD to superimpose the alignment cross on the
HMD over the alignment cross on the HUD. Once aligned, the cage/uncage switch is pressed and held
until ALIGN OK is displayed on the HMD. When coarse alignment is complete, fine alignment is
automatically selected. Fine alignment (PB 1) can also be manually selected.
The WSO moves the HMD to superimpose the alignment cross on the HMD over the alignment
cross on the BRU, figure 2-82, sheet 2. Once aligned, the undesignate switch on the right hand
controller is pressed and held until ALIGN OK is displayed on the HMD. When coarse alignment is
complete, fine alignment is automatically selected. Fine alignment (PB 1) can also be manually
selected.
2.23.12.2 Fine Alignment. When the FINE (PB 1) option is boxed, an alignment cross is displayed
in the HUD, and two alignment crosses are displayed on the HMD in the vicinity of the HUD
alignment cross. See figure 2-83. The display indicates which axis is being aligned. If azimuth and
elevation is indicated (FA DXDY), the pilot moves the TDC either left or right to align in azimuth, up
or down to align in elevation. The pilot presses and releases the cage/uncage switch when satisfied with
the quality of the azimuth and elevation alignment. This causes the display to toggle to roll alignment
mode. With the roll axis indicated (FA DROLL), TDC inputs to the left or right are used to rotate the
HMD alignment symbols to align with the HUD alignment cross. The pilot presses and releases the
cage/uncage switch when satisfied with the quality of the roll alignment. Pressing and releasing the
cage/uncage switch continues to toggle between these two modes until the pilot deselects FINE to
return to coarse alignment or exits alignment.
Upon entering fine alignment, automatically, or if commanded by the WSO selecting the FINE
option, the EU indicates which axis is being aligned. If the azimuth and elevation axis is indicated, the
WSO uses the DC to move the crosses up/down and left/right to align with the cross displayed on the
BRU, (figure 2-83 sheet 2). When satisfied with the alignment, the WSO presses and releases the
undesignate switch on right hand controller at which time the EU automatically switches to roll
alignment. The WSO uses the DC to rotate the cross so that it aligns with the cross displayed on the
BRU. When satisfied with the quality of the alignment the WSO presses and releases undesignate
switch on right hand controller.
2.23.12.3 Alignment Exit. Alignment is exited whenever ALIGN is deselected (unboxed), an A/A
weapon is selected, MENU is selected, TDC priority is reassigned, ACM mode is selected, or the master
mode is changed. This removes the alignment cross from the HUD, removes the FINE option, unboxes
ALIGN, and returns the cage/uncage function to the previously assigned system.
2.23.12.4 Alignment Verification. When the HMD is in normal mode a cross is displayed on the
HUD at the reported HMD LOS. See figure 2-84. If the reported HMD LOS is outside the HUD FOV,
the cross flashes at the HUD FOV limit.
2.23.13 HMD/AHMD Symbology
2.23.13.1 HUD Symbology Replicated on the HMD. The HMD layout essentially replicates the
HUD layout. Window locations, format, and occlusion level on the HMD are as identical to the HUD
locations, format, and occlusion level as practical.
2.23.13.2 HUD Symbology Not Replicated on the HMD. Some of the symbology on the HUD is
either not required on the HMD or would be disorienting if the information was presented. The
following paragraphs describe the items on the HUD which are not replicated on the HMD.
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2.23.13.2.1 Aircraft Attitude Data. Some HUD data only provides the pilot usable information
when presented along the aircraft boresight. HMD data is not always presented along the aircraft
boresight. For this reason, the aircraft pitch ladder, horizon bar, water line indicator, and velocity
vector are not displayed on the HMD/AHMD.
The HMD/AHMD does not provide adequate attitude information and
should not be used as a primary flight instrument.
2.23.13.2.2 Ground Proximity Warning System. In addition to aircraft attitude data, the HMD/
AHMD does not replicate the Ground Proximity Warning System (GPWS) arrow. However, when the
GPWS is activated, ALTITUDE is displayed in the HARM window of the HMD/AHMD.
2.23.13.2.3 Landing Aid Symbology. Landing aid symbology is not displayed on the HUD.
2.23.14 Navigation Master Mode. If present, the MC displays the A/A L&S with the TD box and its
associated TLL, or the A/G designation with the TD diamond and its associated TLL. However, if both
the A/A L&S and the A/G designation are present, only the A/A L&S TLL is displayed.
2.23.14.1 NAV Master Mode TDC Priority. When in NAV master mode, the TDC can be assigned
priority to the HMD by pressing the castle switch forward. This is indicated by an open aiming cross
with a dot in the center being displayed. If TDC priority is removed from the HMD, the dot is removed
from the center of the aiming cross.
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Figure 2-82. Coarse Alignment (Sheet 1 of 2)
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ORIGINAL
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Figure 2-82. Coarse Alignment (Sheet 2 of 2)
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2.23.15 Mission Computer Failure. With non-AMCD aircraft, in the event of an MC1 failure, MC2
provides back-up symbology for the HMD. MC2, at a minimum, provides airspeed, altitude, selected
A/A weapon and count, and the L&S and AIM-9 LOS. MC2 continues to slave the radar to the HMD
LOS when in HACQ mode. MC2 also continues to slave the AIM-9 to the HMD LOS when the AIM-9
is selected. Activation of the uplook reticles are also maintained during back-up.
With non-AMCD aircraft, in the event of an MC2 failure, MC1 provides back-up symbology for the
HMD. MC1, at a minimum, provides airspeed, altitude, selected A/A weapon and count. Slaving of the
radar and AIM-9 is suspended.
With AMCD aircraft, the HMD is not supported in the MC backup mode. If either MC fails, no
symbology is displayed on either helmet and the remaining operating MC functions as if no helmets are
installed.
2.23.16 Electronic Unit Failure. In the event of an EU failure which does not allow any symbology
to be displayed on the HMD, the MC provides the radar boresight symbol and AIM-9 FOV symbol on
the HUD. If the LOS is still valid, the MC continues to slave the radar or AIM-9 to the HMD LOS. If
the LOS is invalid, the MC reverts to the current no-helmet mechanization for slaving weapons,
sensors, and HOTAS.
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Figure 2-83. Fine Alignment (Sheet 1 of 2)
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Figure 2-83. Fine Alignment (Sheet 2 of 2)
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Figure 2-84. Alignment Verification
2.23.17 Helmet Tracker Failure. If the EU reports that the helmet tracker is failed, or the helmet
LOS is no longer valid, the MC discontinues slaving sensors and weapons to the HMD LOS. The MC
replaces aircraft boresight for the helmet LOS to the radar and AIM-9, and deselects slaving for the
A/A FLIR. Additionally, the MC removes any item from the HMD which is tied to the HMD LOS. The
radar boresight and AIM-9 FOV symbol are restored to the HUD. The MC also restores VACQ mode
and the HOTAS function to access the VACQ function. The MC continues to display information on
the HMD which is not tied to the HMD LOS.
2.23.18 Helmet Not Installed. When the MC determines that the HMD is not on, or the EU is not
responding to the MC via the mux bus, the MC reverts to the current no-helmet mechanization for
slaving weapons, sensors, HOTAS, and HUD display.
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CHAPTER 3
Servicing and Handling
3.1 SERVICING
Refer to A1-F18EA-NFM-600.
I-3-1 (Reverse Blank)
ORIGINAL
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CHAPTER 4
Operating Limitations
4.1 LIMITATIONS OF THE BASIC AIRCRAFT
All operating limitations listed in this section are based on the following assumptions • 480 gallon EFTs are model -1013, -1015, or -1017
• F/A-18E aircraft LOT 23 and up, and F/A-18F aircraft LOT 23 and up, with SUU-79B/A
wing pylons and AFC 315
4.1.1 Engine Operation Limitations. During normal engine operation, engine parameters (e.g., N1,
N2, and EGT) are maintained within limits by the FADEC. See figure 4-1 for engine operation
limitations.
Limitations
N2 (%)
N1 (%)
EGT
(°C)
Nozzle
(%)
Oil Press (psi)
Transient (MIL/MAX)
102
103
976
−
−
952
50 to 100
100
100
932
0 to 45
MAX
Steady state
80 to 150 (warm oil)
MIL
Ground IDLE
≥ 61
≥ 32
250 to 590
77 to 83
Start
≥ 10
−
871
−
35 to 90 (warm oil)
• Min 10 within 30 sec
• 180 max after 2.5 min
Figure 4-1. Engine Operation Limitations
4.1.1.1 Engine Vibration Limitations. Engine vibration limitations are:
1. FAN VIB:
1.6 ips max
2. CORE VIB:
2.2 ips max
4.1.2 CG Limitations. 16.8 to 31.8% MAC
4.1.3 Airspeed Limitations. The airspeed limitations for the basic aircraft (with or without pylons)
in smooth or moderately turbulent air with the landing gear retracted and flaps in AUTO are shown
in figure 4-2. Subsystem related airspeed limitations are shown in figure 4-3.
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ORIGINAL
A1-F18EA-NFM-000
Figure 4-2. Basic Aircraft Airspeed Limitations
Subsystem
Position/Action
Airspeed/Groundspeed
Extension/Retraction
300 KCAS
Extended
400 KCAS
Extension/Retraction/Extended
250 KCAS
Emergency Extension
170 KCAS
HALF/FULL
250 KCAS
Nose Gear
195 KGS
Main Gear
210 KGS
Wingfold
Spread/Fold
60 knots
Canopy
Open
60 knots
Refueling Probe
Landing Gear
Trailing Edge Flaps
Tires
Figure 4-3. Subsystem Airspeed Limitations
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ORIGINAL
A1-F18EA-NFM-000
4.1.4 Gross Weight and Lateral Weight Asymmetry Limitations. See figure 4-4 for gross weight and
lateral weight asymmetry limitations.
GW Limit (lb)
Asymmetry Limit1 (ft-lb)
Field Takeoff/In Flight
66,000
29,0002,3 (FLY)
Field Landing/FCLP/T&G
50,600
33,0004,5 (LAND)
Catapult
66,000
29,0002 (FLY) not to exceed
30,8004,5 (LAND)
Carrier Landing/Barricade
44,000
30,8004,5 (LAND)
Condition
1.
2.
3.
4.
5.
FLY and LAND values for MC OFP H3E AND UP, if unavailable, calculate using notes 2-5.
Calculate using station 2-10 stores/pylons/external fuel.
Add internal wing fuel split with FUEL XFER caution.
Calculate using station 1-11 stores/pylons/external fuel.
Add internal wing fuel split.
Figure 4-4. Gross Weight and Lateral Weight Asymmetry Limitations
NOTE
• Field landing/FCLP/T&G, and Carrier Landing/Barricade limits are
ground loads limits and apply to touchdown only. All in-flight phases,
including approach to landing must meet the in-flight lateral asymmetry restrictions in figure 4-4.
• Catapult limits are derived from both flying quality limits and ground
load limits. Ensure neither limit is exceeded during catapult launch.
4.1.4.1 Lateral Weight Asymmetry Calculations. Field takeoff and in-flight lateral asymmetry is
calculated by using the weight of asymmetric stores on stations 2 thru 10. Field landing events lateral
asymmetry is calculated by using the weight of asymmetric stores on stations 1 thru 11. For catapult
operations, lateral weight asymmetry is calculated for stations 1 thru 11 and stations 2 thru 10. For
asymmetric catapult launch endspeed (asymmetric weight board designation) and catapult launch
lateral trim calculations, use the asymmetry of stations 2 thru 10. For field takeoff and in-flight
calculations, the weight of asymmetric internal wing fuel can be ignored unless the FUEL XFER
caution is displayed due to asymmetric internal wing fuel. In addition, for catapult and landing events,
all asymmetric internal wing fuel weight (regardless of FUEL XFER caution) is included due to
landing gear structural load limitations.
With MC OFP H3E AND UP, lateral weight asymmetry is calculated by the MC and displayed on
the CHKLST page. Two values are displayed, FLY and LAND, and indicate thousands of ft-lb. The
values are displayed on the heavy side of the aircraft and it is possible that one value will be displayed
on one side of the aircraft and the other value on the other side. The FLY value is calculated using the
weight of asymmetric stores on stations 2 thru 10. The LAND value is calculated using the weight of
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ORIGINAL
A1-F18EA-NFM-000
asymmetric stores on stations 1 thru 11. Both values include internal wing fuel imbalances regardless
of FUEL XFER caution status. The values will flash when one fuel quantity is invalid or one weapon
station indicates HUNG and the values will be removed if more than one fuel quantity is invalid,
and/or weapon station indicates HUNG.
The FLY value is used for field takeoff, catapult, and in flight. The LAND value is used for landing
events or catapult. For asymmetric catapult launch endspeed (asymmetric weight board designation)
and catapult launch trim calculations, use the FLY value. In flight, it is possible for the FLY value to
exceed the value in figure 4-4 due to small fuel imbalances. A manual lateral asymmetry calculation
should be performed in accordance with figure 4-4. If AOA TONE caution is present and/or lateral
weight asymmetry calculations are flashing, calculate aircraft lateral asymmetry manually.
* The weight of an asymmetric missile or pod on stations 1 or 11 is not to be included in asymmetry
limitations calculations for field takeoff or in flight, but is included in landing events lateral weight
asymmetry calculations. For catapult operations, lateral asymmetry must be calculated for stations 1
thru 11 and stations 2 thru 10.
Figure 4-5. Asymmetric Stores Limitations
NOTE
• If a FUEL XFER caution is displayed and the internal wing tanks are
unbalanced, the weight of asymmetric internal wing fuel must be used
in calculating total weight asymmetry in all phases of flight. Completely split internal wing tanks (one full and one empty) have the
potential of reaching 14,000 ft-lb of lateral weight asymmetry.
• The following is a rule of thumb for calculating internal wing tank
asymmetry: pounds of fuel split x 10 (e.g., a 200 pound fuel split
equates to approximately 2,000 ft-lb of asymmetry).
• With a symmetrically loaded aircraft, release of any single store, with
the exception of a full midboard 480 gallon tank, will not exceed the
lateral weight asymmetry limitations. Release of dissimilar stores in
the normal SMS release sequence may exceed the lateral weight
asymmetry limitation.
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ORIGINAL
A1-F18EA-NFM-000
4.1.5 AOA Limitations - Flaps AUTO. Refer to figure 4-6 for AOA limitations with flaps in AUTO.
4.1.5.1 AOA Limit Display. With MC OFP H3E AND UP, the current positive AOA limit is
displayed on the CHKLST page. The limit is displayed below the aircraft and only when a limit exists.
For every set of positive/negative AOA limits, only the positive AOA limit is displayed. The AOA limit
will be removed if the FLY lateral weight asymmetry value is removed. The AOA limit of 14° will be
displayed with flaps in HALF or FULL regardless of the lateral weight asymmetry AOA limit.
AOA Limitations - Flaps AUTO1
Lateral Weight
Asymmetry
(1,000 ft-lb)
Subsonic
Supersonic
Unrestricted
Unrestricted
(> +15° Half lateral stick or half rudder
pedal inputs only)
≤6
> 6 to ≤ 8
> 8 to ≤ 12
> 12 to ≤ 29
Low or Slow
High and Fast
(≤ 20k ft or ≤ 250 (> 20k ft and > 250
KCAS)
KCAS)
Unrestricted
≤ +30°
-6 to +15°
and
Single axis inputs only2
-6 to +15°
and
Single axis inputs only2
Notes:
(1) Rolling maneuvers up to abrupt, full stick (full stick in less than 1 second) are authorized within the
AOA and acceleration limitations specified in figure 4-7.
(2) In ‘‘Single axis inputs only’’ regions, avoid rolling or yawing the aircraft while changing longitudinal stick
position. It is acceptable to pull, stop, then roll or to pull and counter any roll-off induced by the heavy
wing under g.
(3) With MC OFP H3E AND UP, AOA tone in flaps AUTO is based on the limits in this figure and
computed lateral asymmetry. The tone indicates that the AOA lateral asymmetry limit has been
exceeded. AOA shall be reduced to stay within the limits of figure 4-6.
Figure 4-6. AOA Limitations - Flaps AUTO
4.1.6 Acceleration Limitations. With flaps in AUTO, the acceleration limitations for the basic
aircraft (with or without empty pylons) in smooth air with the landing gear retracted are shown in
figure 4-7. In moderate turbulence, reduce deliberate accelerations 2g below that shown in figure 4-7
to minimize the potential of an aircraft over-g.
Acceleration limits during landing gear extension/retraction, or with landing gear extended, are 0.0
to +2.0g (symmetrical) and +0.5 to +1.5g (rolling).
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ORIGINAL
A1-F18EA-NFM-000
NOTES
1. At any specific gross weight, the G-limiter will attempt to limit command g to the levels shown
above, up to 57,405 lb GW. Above 57,405 lb GW, the G-limiter is fixed at +5.5.
2. Overshoots up to +0.5g or -0.2g do not constitute an over-g.
3. Above 57,405 lb, an over-g will occur if the pilot solely relies on the G-limiter.
4. The aircraft structural carriage g envelope is based on the product of the maximum normal
acceleration limits of +7.5g and -3.0g at a weight of 42,097 lb. As aircraft gross weight increases
above 42,097 lb, Nz must be decreased so that the maximum NzW allowable is not exceeded. Refer
to sheet 3 for an example of NzW correction for these calculations. Nz limits less than +2.0 and -2.0
need not be corrected for NzW.
5. See External Stores Limitations, A1-F18EA-TAC-020 (NWP 3-22.5-F/A18E/F VOL. IV) for
additional acceleration limitations which may apply when carrying stores. Unless otherwise noted,
Nz store limitations are based on an aircraft gross weight of 42,097 lb. As aircraft gross weight
increases above 42,097 lb, Nz must be decreased so that the maximum NzW allowable is not
exceeded. Refer to sheet 3 for an example of NzW correction for these calculations. Over-g
protection is not provided by the G-limiter for additional Nz restrictions due to store carriage.
Over-g due to store limitations will not trigger an over-g MSP code but aircraft over stress may
result. Additional store restrictions should be closely monitored by aircrew. Nz limits between +2.0
and -2.0 need not be corrected for NzW.
Figure 4-7. Acceleration Limitations - Basic Aircraft
(with or without empty pylons) (Sheet 1 of 3)
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ORIGINAL
A1-F18EA-NFM-000
‘
NOTES
1. At any specific gross weight, the G-limiter will attempt to limit command positive g to the levels
shown above, up to 57,405 lb GW with lateral stick inputs. Above 57,405 lb GW, the G-limiter is
fixed at +4.4. For negative rolling maneuvers beyond -1 g, the G-limiter will not reduce command
g in response to lateral stick input. Negative load factor when rolling shall be closely monitored.
2. G-limiter overshoots up to +0.5g or -0.2g do not constitute an over-g.
3. Above 57,405 lb, an over-g will occur if the pilot solely relies on the G-limiter.
4. The aircraft structural carriage g envelope is based on the product of the maximum normal
acceleration limits of +7.5g and -3.0g at a weight of 42,097 lb. As aircraft gross weight increases
above 42,097 lb Nz must be decreased so that the maximum NzW allowable is not exceeded. Refer
to sheet 3 for an example of NzW correction for these calculations. Nz limits between +2.0 and -2.0
need not be corrected for NzW.
5. See External Stores Limitations, A1-F18EA-TAC-020 (NWP 3-22.5-F/A18E/F VOL. IV) for
additional acceleration limitations which may apply when carrying stores. Unless otherwise noted,
Nz store limitations are based on an aircraft gross weight of 42,097 lb. As aircraft gross weight
increases above 42,097 lb, Nz must be decreased so that the maximum NzW allowable is not
exceeded. Refer to sheet 3 for an example of NzW correction for these calculations. Over-g
protection is not provided by the G-limiter for additional Nz restrictions due to store carriage.
Over-g due to store limitations will not trigger an over-g MSP code but aircraft over stress may
result. Additional store restrictions should be closely monitored by aircrew. Nz limits between +2.0
and -2.0 need not be corrected for NzW.
Figure 4-7. Acceleration Limitations - Basic Aircraft
(with or without empty pylons) (Sheet 2)
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ORIGINAL
A1-F18EA-NFM-000
Figure 4-7. Acceleration Limitations - Basic Aircraft
(with or without empty pylons) (Sheet 3)
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ORIGINAL
SEE IC # 34
A1-F18EA-NFM-000
4.1.7 Limitations with Flaps HALF or FULL. Refer to figure 4-8 for flaps limitations when HALF or
FULL.
Parameter
Limitation
AOA
0 to 14° (AOA tone) (1)
Bank angle
90° max
15° max during flap selection (HALF or FULL from
AUTO) with a HI AOA advisory
Acceleration
Symmetrical
0.0 to +2.0g
Rolling
+0.5 to +1.5g
NOTE:
1. Transitory excursions above 14° may be seen during catapult launch.
Figure 4-8. Limitations with Flaps HALF or FULL
4.1.8 Refueling Limitation. Maximum refueling pressure, inflight or on the ground, is 55 psi.
4.1.9 Prohibited Maneuvers.
Environmental 1. Flight in lightning or thunderstorms.
Systems 1. Takeoff with a FADEC DEGD indication (dual channel line outs).
2. Takeoff with a FCS A or FCS B DEGD.
3. Pulling any FCS circuit breaker inflight except as directed by NATOPS.
4. Use of RALT mode below 500 feet AGL.
5. Use of the auto sever/redeployment function of the ALE-50.
6. Landing with autopilot modes engaged except for the following:
a. Mode 1 ACL.
b. Field landings with FPAH/ROLL.
7. Takeoffs and landings while using any laser eye protection (LEP) devices.
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ORIGINAL
SEE IC # 34
A1-F18EA-NFM-000
8. Selection of MAN with the ECS MODE switch.
Selection of MAN with the ECS MODE switch while the aft cooling fan
shutoff valve is open may cause the fan to overspeed resulting in a
catastrophic fan failure potentially leading to loss of OBOGS.
9. Flight in RVSM controlled airspace with any of the following:
a.
b.
c.
d.
e.
f.
g.
h.
i.
Mach > 0.92
AOA > 15°
Load factor > 2g
Initial engagement of BALT when above Mach 0.90
BALT not engaged when level at assigned altitude (During formation flight, BALT engagement is only required by the flight leader.)
Refueling probe extended
GAIN ORIDE selected
Suspected AOA or pitot-static probe damage (e.g., birdstrike, refueling basket strike)
Baro altitude displayed with an ″X″
10. In-flight Memory Inspect (MI) of FCC (UNIT 14 or 15) addresses (ADDR) greater than six
digits long.
In-flight Memory Inspect (MI) of FCC (UNIT 14 or 15) addresses
(ADDR) greater than six digits long may cause all four FCC channels to
shut down which will result in loss of aircraft control.
Departure/Spin 1. Zero airspeed tailslides.
2. Intentional departures/spins.
3. Yaw rates over 40° second (yaw tone).
4. Holding roll inputs (lateral stick or rudder pedal) past 360° of bank angle change.
5. Inflight selection of RCVY on the SPIN switch.
Selection of manual spin recovery mode (SPIN switch in RCVY) seriously degrades controllability and prevents recovery from any departure
or spin.
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ORIGINAL W/IC 34
SEE IC # 34
A1-F18EA-NFM-000
Fuel and Engine Oil 1. Zero g except transient (over 2 seconds between +0.2 and -0.2g).
2. Negative g for more than 10 seconds (30 seconds required between negative g maneuvers).
Loads 1. Field full stop, FCLP, or T&G with lens settings greater than 3.25°.
2. Carrier arrestment or T&G with lens settings greater than 4.0°.
3. Pushing beyond -1g above 700 KCAS and below 10,000 feet MSL.
4. Holding lateral stick inputs past 180° of bank angle change when pushing between 0.0 and -1.0g.
5. Abrupt, full aft stick inputs (full aft stick in less than 0.5 seconds) with less than 3,500 pounds of
fuel.
6. AUX release of external fuel tanks or aerial refueling store.
7. Lateral and directional trim limitations. Use of lateral and/or directional trim in
combination with lateral stick inputs above certain airspeeds may result in the exceedance of
aircraft structural load limitations. Following are four conditions that require trim due to aircraft
imbalance and their associated limitations. These limitations are in addition to existing
limitations for the basic aircraft and any store peculiar limitation. Air to ground limitations are
for any configuration with a wing EFT or wing carried air to ground store.
Condition A: Roll off tendency > 5°/sec without lateral trim or
FLIR pod loading only on one side of the aircraft without a centerline 480 EFT
Condition B: Lateral weight asymmetry for air-to-air weapons > 6,000 ft-lb and ≤ 14,000 ft-lb
Condition C: Lateral weight asymmetry for to air-to-ground weapons > 6,000 ft-lb and ≤29,000
ft-lb
Condition D: Lateral weight asymmetry for to air-to-air weapons > 14,000 ft-lb
Flight Restrictions
A
B
C
D
Airspeed > 500KCAS
Symmetric Nz = -1.2 to +7.5g
X
X
X
X
Airspeed > 600 KCAS or IMN > 1.1 below 24,000 ft MSL
Airspeed > 500 KCAS above 24,000 ft MSL
1/2 Lateral Stick Deflection
X
X
Airspeed > 600 KCAS or IMN > 1.05
non-abrupt 1/2 lateral stick deflections
X
All airspeeds except for powered approach, 1/2 lateral stick input
X
These limitations assume that the aircraft is configured with ECP 6171 or IAFC 362
(aft fuselage stiffener).
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ORIGINAL W/IC 34
SEE IC # 34
A1-F18EA-NFM-000
Without this modification there are additional limitations that are defined below:
Configuration
Restriction
For aircraft with asymmetry due to air to air weapons ≤ 8,000 ft-lb
IMN between 0.9 and 1.04
Symmetric Nz = -3.0 to +7.5g
Asymmetric Nz = -1.0 to +5.0g
For aircraft with roll off tendency >5°/sec without lateral trim
IMN between 0.9 and 1.04
For aircraft with a FLIR pod only on one side of the aircraft without a Symmetric Nz = -3.0 to +5.0g
Asymmetric Nz = -1.0 to + 1.5g
centerline tank
For aircraft with lateral weight asymmetry > 8,000 ft-lb due to air-to-air
stores
For aircraft with lateral weight asymmetry > 6,000 ft-lb due to air-toground stores
8. Any rudder input when aircraft load factor is less than -1.2g.
Flutter 1. Flight without LAU-127 wingtip launcher rails. The launchers on both stations 1 and 11 shall be
either: (a) LAU-127A/A or B/A (with power supply and nitrogen bottle installed), or (b) LAU127C/A (with power supply and High Pressure Pure Air Generator (HiPPAG) unit installed).
Flying Qualities 1. Single-ship takeoffs with 90° crosswind component over 30 knots.
2. Section takeoffs with any of the following conditions:
a. 90° crosswind component over 15 knots.
b. Asymmetric loading over 9,000 ft-lb not including wingtip missiles or pods.
c. Dissimilar loading except pylons, FLIR, LDT, fuselage missiles, wingtip missiles or pods,
CVERS, MERS, or TERS.
3. Flight with GAIN ORIDE selected above 10° AOA or above 350 KCAS (flaps AUTO), 200 KCAS
(flaps HALF), or 190 KCAS (flaps FULL).
With GAIN ORIDE selected (fixed FCS gains), the aircraft is uncontrollable above approximately 450 KCAS.
4. Single-ship landings with 90° crosswind component over 30 knots.
5. Section landings with 90° crosswind component over 15 knots.
6. Aerobraking on landing rollout with crosswind greater than 5 knots, pitch attitude greater than
10°, airspeed less than 80 KCAS, GAIN ORIDE selected, FCS AIR DAT caution or FLAP
SCHED caution.
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ORIGINAL
A1-F18EA-NFM-000
4.2 EXTERNAL STORES LIMITATIONS
External stores limitations for external fuel tanks are provided in figures 4-9 thru 4-11. Figure 4-12
contains Air Refueling Store Carriage Limits. Figures 4-9 thru 4-12, in conjunction with the Tactical
Manual, A1-F18EA-TAC-020 (NWP 3-22.5-F/A18E/F Vol IV) define the stores limitations for all
(external fuel tanks and other) external stores. Only those pylons and stores shown in the Tactical
Manual may be carried and/or released.
Fuel in Tank
LIMITATIONS
(weight/tank)
Airspeed
(KCAS/IMN)
Load
Factor
(g)
≤ 500 lb
>500 lb
LON
Symmetrical
LON
Rolling
LON (1)
LON
Lateral Stick Deflection
LON (1)
LON
Rudder
LON
Field Landing, FCLP,
T&G, Field Arrest
Any fuel state
Catapult
Less than or equal to 100 lb or greater than 2,700 lb
Carrier Arrestment,
Carrier T&G
Less than or equal to 800 lb
NOTE:
1. Lateral stick inputs are restricted to checks to neutral only to terminate rolls. Abrupt full lateral stick deflection authorized for roll
initiation.
Simplified Limitations Schematic
> 500 lb Fuel
D Roll termination only to neutral. No opposite lateral stick check allowed.
Figure 4-9. Station 6 480-Gal External Fuel Tank Carriage Limits
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ORIGINAL
A1-F18EA-NFM-000
SUU-79B/A (Rev B) Wing Pylons, AFC-315 to Lot 22 aircraft
LIMITATIONS
(weight/tank)
Fuel in Tank
Airspeed
(KCAS/IMN)
Load
Factor
(g)
Lateral
Stick
Deflection
>500 lb
≤ 500 lb
635/1.6 above 15,000 ft MSL
570 KCAS at or below 15,000 ft MSL
635/1.6
Symmetrical
LON at or below 500 KCAS
-3.0 to +6.0 from 500 KCAS up to 550 KCAS
-1.0 to +4.0 above 550 KCAS (1)
Rolling
LON (rolling) at or below 400 KCAS
-1.0 to +2.0 above 400 KCAS and below 470 KCAS (1)
+1.0 to +2.0 above 470 KCAS (1)
At or below
15,000 ft MSL
LON at or below 470 KCAS
Smooth inputs of up to 3/4 inches above 470 KCAS
Above 15,000 ft MSL
LON at or below 570 KCAS
Smooth inputs of up to 3/4 inches above 570 KCAS
Rudder
LON
Smooth inputs above 500 KCAS
Field Landing, FCLP,
T&G, Field Arrest
Any fuel state
Catapult
Less than or equal to 100 lb or greater than 2,700 lb
Carrier Arrestment,
Carrier T&G
Less than or equal to 800 lb
NOTE:
1. Nz restriction due to carriage of 480 EFT need not be corrected for an aircraft weight greater than 42,097 lb (NzW).
Simplified Limitations Schematic
> 500 lb Fuel
≤ 500 lb Fuel
D Max Airspeed 570 KCAS/1.6 IMN
D LON at or below 400 KCAS
D Single Axis Control inputs above 400 KCAS
(roll then pull)
D 4g Max above 500 KCAS
D Max Airspeed 635 KCAS/1.6 IMN
Note: The simplified limitations are presented as a rule-of-thumb for ease of memorization.
The actual NATOPS limits are shown in the upper portion of the table.
Figure 4-10. Station 4/8 480-Gal External Fuel Tank Carriage Limits (Rev B Pylons)
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ORIGINAL
A1-F18EA-NFM-000
SUU-79A/A (Rev A) Wing Pylons
LIMITATIONS
(weight/tank)
Fuel in Tank
Airspeed
(KCAS/IMN)
Load
Factor
(g)
>500 lb
≤ 500 lb
550/1.6 above 15,000 ft MSL
520 KCAS at or below 15,000 ft MSL
600/1.3 without
FLIR Pod
570/1.3 with FLIR
Pod
Symmetrical
-3.0 to +3.0 at or below 475 KCAS
0 to +2.0 above 475 KCAS (1)
Rolling
-1.0 to +2.0 at or below 325 KCAS
0 to +2.0 above 325 KCAS (1)
LON
Lateral
Stick
Deflection
LON at or below 225 KCAS
Half between 225 and 325 KCAS
Smooth inputs of up to 3/4 inches above 325 KCAS
LON without FLIR
Pod
Smooth inputs of up
to 3/4 inches above
550 KCAS with
FLIR Pod
Rudder
Smooth inputs above 500 KCAS
Field Landing, FCLP,
T&G, Field Arrest
Any fuel state
Catapult
Less than or equal to 100 lb or greater than 2,700 lb
Carrier Arrestment,
Carrier T&G
Less than or equal to 800 lb
NOTE:
1. For load factor restrictions resulting in an aircraft envelope less than or equal to 2.5 g for symmetric and 2.0 g for rolling, no further
reduction in load factor is required as a result of aircraft weight being greater than basic flight design gross weight.
Simplified Limitations Schematic
> 500 lb Fuel
D Max Airspeed 520 KCAS/1.6 IMN
D Navigational Turns Only
(Ferry Jet or Tanker)
≤ 500 lb Fuel
D Max Airspeed 570 KCAS/1.3 IMN
D LON up to 550 KCAS
Note: The simplified limitations are presented as a rule-of-thumb for ease of memorization.
The actual NATOPS limits are shown in the upper portion of the table.
Figure 4-11. Station 4/8 480-Gal External Fuel Tank Carriage Limits (Rev A Pylons)
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ORIGINAL
A1-F18EA-NFM-000
LIMITATIONS
(weight/tank)
Station 6 ARS
Stowed and feathered
Altitude (ft MSL)
0 to 40,000
Airspeed (KCAS/IMN)
Load
Factor
(g)
Station 6 ARS, Station 3/4/8/9
480-Gal Carriage (5-wet)(1)
575/0.92
430/0.8 (2)
Symmetrical
-1.0 to +4.0 (3)
LON
Rolling
-1.0 to +3.0 (3,4)
Lateral Stick
Deflection
Full with any air to ground store
or
an EFT on any wing station
(roll limiting on);
otherwise Half (6)
Rudder
LON
Angle of Attack
LON
Field Landing, FCLP,
T&G, Field Arrest
Simultaneously large and abrupt
multi-axis inputs prohibited (4,5)
Flaps AUTO: -6° to + 15°
Flaps HALF or FULL: LON
Any fuel state
Catapult
≤ 100 lb or > 1,700 lb fuel
ARS: ≤ 100 lb or > 1,700 lb fuel
480-gal EFT: ≤ 100 lb or > 2,700 lb fuel
Carrier Arrestment,
Carrier T&G
≤ 500 lb fuel
ARS and/or Station 3/9 480-gal
EFT: ≤ 500 lb fuel
Station 4/8 480-gal EFT: ≤ 800 lb fuel
NOTES:
1. SUU-79 B/A (Rev B) Wing Pylons, AFC-315 to Lot 22 aircraft.
2. Abnormal wing tank fuel transfer sequence: If fuel is transferred (because of a failure condition) from
Station 4/8 tanks before the Station 3/9 tanks are empty, then airspeed limit is 300 KCAS/0.6 IMN,
whichever is less.
3. Nz is based on an aircraft GW of 66,000 lb. When aircraft gross weight is greater than 66,000 lb, reduce
symmetric load factor allowable to -0.5 to +3.5 and rolling load factor allowable to -0.5 to +2.5
4. Lateral stick input is restricted to half when the fuel in Station 3/9 is greater than 1,600 lb/tank. This
restriction is not applicable when in power approach configuration.
5. Large is defined as greater than half maximum input. Abrupt is defined as a control input rate of
maximum deflection in less than 1 second. Multi-axis is defined as simultaneous longitudinal, lateral, or
directional control inputs. A multi-axis input is allowed as long as it is not large and abrupt.
6. Half lateral stick displacement restriction does not apply when in power approach configuration.
Figure 4-12. Air Refueling Store Carriage Limits
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ORIGINAL
A1-F18EA-NFM-000
4.2.1 ARS Limitations
1. ARS operating limitations are listed in figure 4-13.
Limitation
(P/N 31-301-48310-2)
Condition
EFT loading
Station 4/8 only (3 wet)
Altitude
Limitation
(P/N 31-301-48310-3)
Station 4/8 (3 wet)
Station 3/4/8/9 (5 wet)
0 to 35,000 feet MSL
Power on
(RAT unfeathered)
180 to 300 KCAS
180 to 250 KCAS
(<5,000 feet MSL)
Hose extension
180 to 250 KCAS
180 to 275 KCAS
180 to 260 KCAS
(5,000 to 10,000 feet MSL)
180 to 275 KCAS
(>10,000 feet MSL)
Hose extended
180 to 300 KCAS/0.80 IMN max
Fuel transfer to receiver
180 to 300 KCAS/
0.80 IMN max
225 to 300 KCAS/0.80 IMN max
180 to 250 KCAS
(<5,000 feet MSL)
Hose retraction
180 to 200 KCAS
(<25,000 feet MSL)
180 to 210 KCAS
(≥25,000 feet MSL)
180 to 275 KCAS
180 to 260 KCAS
(5,000 to 10,000 feet MSL)
180 to 275 KCAS
(>10,000 feet MSL)
Figure 4-13. ARS Operating Limitations (Sheet 1 of 2)
2. Prohibited maneuvers with ARS installed:
a. Placing the LTD/R switch to ARM.
b. Afterburner operation with the hose extended.
3. ARS carriage limitations are listed in figure 4-12.
4. ARS operating limitations for P/N 31-301-48310-4 and -5 are as follows:
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ORIGINAL
A1-F18EA-NFM-000
Condition
EFT loading
Limitation (P/N 31-301-48310-4/5)
Station 4/8
(3 wet)
Station 3/4/8/9
(5 wet)
Altitude
0 to 35,000 feet MSL
Power on
(RAT unfeathered)
Hose extension/
retraction
220-300 KCAS
180 to 275 KCAS
(<5,000 feet MSL)
180 to 250 KCAS
(<5,000 feet MSL)
190 to 275 KCAS
(5,000 to 7,500 feet MSL)
190 to 260 KCAS
(5,000 to 7,500 feet MSL)
200 to 275 KCAS
(7,501 to 10,000 feet MSL)
200 to 260 KCAS
(7,501 to 10,000 feet MSL)
200-275 KCAS
10,001 to 15,000 feet MSL
220-275 KCAS
15,001 to 35,000 feet MSL
Hose extended
180 to 300 KCAS/0.80 IMN max
180 to 300 KCAS/0.80 IMN max
(<5,000 feet MSL)
190 to 300 KCAS/0.80 IMN max
(5,000 to 7,500 feet MSL)
Fuel transfer to
receiver
200 to 300 KCAS/0.80 IMN max
(7,501 to 15,000 feet MSL)
220 to 300 KCAS/0.80 IMN max
15,001 to 35,000 feet MSL
Figure 4-13. ARS Operating Limitations (Sheet 2)
5. Prohibited maneuvers with ARS installed:
a. Placing the LTD/R switch to ARM.
b. Afterburner operation with the hose extended.
6. ARS carriage limitations are listed under External Stores Limitations.
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A1-F18EA-NFM-000
4.2.2 ATFLIR Limitations
1. These limits are ATFLIR specific, apply to any ATFLIR configuration without inboard 480
gallon tanks, and do not pertain to asymmetric loads. ATFLIR limitations are listed in
figure 4-14.
Failure to use eye protection when firing the LASER may result in severe
eye damage and possible blindness.
ATFLIR Limitations
Altitude
Airspeed
(KCAS)
Nz
(g’s)
Between 10,000 and 18,000
feet MSL
< 10,000 feet MSL
> 18,000 feet MSL
< 550
Between
550 and 600
> 600
< 550
> 550
LON
Symmetric
LON
-3.0 to +7.0
-1.0 to 4.0
LON
-3.0 to 7.0
LON
Rolling
LON
-1.0 to +4.0
(1)
LON
NOTE:
1. Smooth inputs up to ¾ inch deflection for navigational turns or as required to maintain roll attitude.
Figure 4-14. ATFLIR Limitations
4.3 OPERATING LIMITATIONS (LOT 21)
4.3.1 Prohibited Maneuvers
Systems 1. Flight in RVSM controlled airspace.
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A1-F18EA-NFM-000
4.4 AESA RADAR LIMITATIONS
Flight testing of the AN/APG-79 AESA radar has shown that this system can interfere with, and
significantly degrade the performance of, various precision approach systems including PAR and
ACLS. Aircrew operating AN/APG-79 AESA equipped aircraft shall be aware that directing
AN/APG-79 AESA transmissions toward landing facilities (ship or shore) may cause strong electromagnetic interference (EMI) affecting the safety of aircraft on approach. To the maximum extent
possible, aircrew shall avoid directing AN/APG-79 AESA radar transmissions towards facilities with an
operating active PAR or ACLS (shipboard or shore-based). Unless specifically authorized for
operational necessity or required for safety of flight, aircrew should cease AN/APG-79 AESA radar
transmissions whenever operating within 10 nm of PAR and ACLS facilities during IMC/Case III
conditions.
• AN/APG-79 AESA radar transmissions specifically directed towards
an operating PAR or ACLS within 10 nm have caused strong electromagnetic interference (EMI) with these radar-based systems and
caused intermittent, unannunciated anomalies that resulted in erroneous azimuth/glideslope commands being sent to aircraft on precision
approach, increasing the risk of controlled flight into terrain (CFIT).
Proper flight instrument crosscheck using all available means
(TACAN, DME, and RADALT) is essential for all aircraft during PAR
and ACLS approaches to detect abnormal precision approach parameters.
• The AGR mode of the AN/APG-79 AESA radar was observed to cause
some of the highest levels of EMI on affected PAR and ACLS systems.
Ensure the AN/APG-79 radar is in SIL, STBY, or OFF prior to
selection of TDC to the HUD in order to enable the steering dot in the
center of the velocity vector (which commands AGR).
• AN/APG-79 AESA radar transmissions within 10 nm directed towards
civil airfields operating ASDE-X ground traffic management systems
have caused EMI. ASDE-X is an FAA-designed, radar-based system
designed to reduce runway incursions by tracking departing, arriving,
and taxiing aircraft plus ground vehicles at major civilian airfields.
AN/APG-79 AESA-generated EMI results in ASDE-X system shutdown, increasing the risk to taxiing aircraft and airport vehicles of
unintentional ground or runway incursion.
NOTE
AN/APG-79 AESA radar does not interfere with SPN-41 ICLS systems
or non-precision radar approach systems (ASR). AN/APG-79 has not
been tested against other civil systems (ILS).
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ORIGINAL W/IC 33
SEE IC # 33
A1-F18EA-NFM-000
NOTE
• Adverse EMI effects from AN/APG-79 AESA equipped aircraft acting
as overhead/recovery tankers (not on approach) have been observed
during SPN-42 and SPN-46 ACLS testing.
• The SPN-42 (shore-based ACLS) and SPN-46 (shipboard ACLS) have
shown degraded lock-on and command guidance anomalies due to
EMI interference from AN/APG-79 AESA radar transmisisons. For
the SPN-46 system, these anomalies are more severe and more
probable with the Block 1 Upgrade to the SPN-46 system.
• Through 2011, the FAA has stated the ASDE-X system will be fielded
at the following locations: KATL, KBDL, KBOS, KBWI, KCLT,
KDCA, KDEN, KDFW, KEWR, KFLL, KHIK, KHOU, KIAD,
KIAH, KJFK, KLAX, KLGA, KLSV, KMCO, KMDW, KMEM,
KMIA, KMKE, KMSP, KORD, KPHL, KPHX, KPVD, KSAL,
KSAN, KSDF, KSEA, KSNA, and KSTL.
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A1-F18EA-NFM-000
PART II
INDOCTRINATION
Chapter 5 - Indoctrination
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CHAPTER 5
Indoctrination
5.1 INITIAL QUALIFICATION
Initial NATOPS qualification in F/A-18E/F series aircraft shall be obtained by satisfactory
completion of the CNO approved course of instruction at an F/A-18E/F Fleet Replacement Squadron
(FRS). The minimum training requirements to be completed prior to initial NATOPS qualification for
pilots and weapon system operators (WSO) are set forth below.
5.1.1 Minimum Ground Training Requirements. The following minimum ground training requirements shall be successfully completed prior to first flight in an E/F-series aircraft:
1. FRS academic familiarization syllabus to include:
a. Aircraft systems and procedures.
b. Emergency procedures review.
c. Instrument flight training.
d. Immediate action, open, and closed book NATOPS exams.
e. Cockpit orientation.
f. Ejection, egress, and survival equipment checkout.
g. Preflight checkout.
2. FRS simulator familiarization syllabus in the Tactical Operational Flight Trainer (TOFT).
5.1.2 Minimum Flight Training Requirements. The following minimum flight training requirements shall be successfully completed prior to initial NATOPS qualification:
1. FRS familiarization flight phase to include a minimum of 10 hours first pilot time (FPT) in
E/F-series aircraft (5 hours if currently NATOPS qualified in F/A-18A-D), and a minimum of 10
hours special crew time (SCT) (5 hours if currently NATOPS qualified in F/A-18B/D) in F-series
aircraft for WSOs.
5.2 FOLLOW-ON TRAINING
Follow-on ground training for each activity may vary according to local conditions, field facilities,
requirements from higher authority, and the immediate unit Commanding Officer’s estimate of
squadron readiness.
Follow-on flight training should include aircraft and weapon systems instruction, normal and
emergency procedures, simulators (if available), and evaluation of aircrew performance. Local
command requirements, squadron mission, and other factors may influence the actual flight training
syllabus and the sequence in which it is completed.
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ORIGINAL
A1-F18EA-NFM-000
5.3 CURRENCY REQUIREMENTS
Minimum requirements to maintain currency after initial qualification shall be established by the
unit Commanding Officer but will involve no less than 5 hours of first pilot time and two takeoffs and
landings in E/F-series aircraft in the previous 90 days for pilots, and 5 hours of special crew time in
F-series aircraft in the previous 90 days for WSOs. Additionally, an annual NATOPS evaluation with
a grade of at least conditionally qualified is required for both pilots and WSOs.
5.3.1 Regaining Currency. Requalification of those crewmembers whose currency has lapsed shall
include a familiarization flight(s), to include enough flight time to enable crewmembers to attain the
requirement of 5 flight hours and 2 takeoff and landings in E/F series within 90 days. Specifics of
familiarization flight(s) at discretion of Commanding Officer of the unit having custody of the aircraft.
5.4 REQUIREMENTS FOR VARIOUS FLIGHT PHASES
The specific requirements for various flight phases conducted during initial training in E/F-series
aircraft are listed in figure 5-1.
Flight Phase
Day Solo/Crew Solo
Night Solo/Crew Solo
Solo/Crew Solo Cross
Country
Initial CQ
Pilot General
WSO
4 FAM flights *
Pilots with current
NATOPS in
Type/Model, and
F/A-18 CAT1 or CAT2
transition syllabus
complete
1 FAM flight *
Day Solo/Crew Solo qualified, 1 night FAM flight
*
F/A-18E/F instrument and NATOPS qualified
50 F/A-18E/F hours
(FPT)
NATOPS Qualified
15 F/A-18E/F hours
(FPT)
* CNO approved syllabus
Figure 5-1. Requirements for Various Flight Phases During Initial Training
5.4.1 Instrument Evaluation Flights. Instrument evaluation flight requirements are delineated in
OPNAVINST 3710.7 series and the NATOPS Instrument Flight Manual. Instrument evaluation
flights may be conducted by designated military aviators or Naval Flight Officers designated in writing
by their Commanding Officer.
5.4.2 Instrument Qualification. In accordance with OPNAVINST 3710.7 series, instrument ratings
shall be valid in all aircraft in which the pilot is instrument qualified regardless of the model in which
the check was flown. A pilot may be considered to be instrument qualified in an aircraft when he/she
has completed the evaluation as outlined in each respective NATOPS manual and has met the
requirements for an instrument rating as outlined in OPNAVINST 3710.7 series.
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ORIGINAL
A1-F18EA-NFM-000
5.4.2.1 Requirements for Instrument Qualification in E/F-Series Aircraft. Each aircrew is required
to be instrument qualified in F/A-18E/F series aircraft.
Pilots must have a current instrument rating prior to flight in actual instrument conditions. Figure
5-2 delineates the pilot requirements for instrument qualification along with ceiling and visibility
restrictions for instrument rated pilots. When a pilot becomes fully instrument qualified in F/A-18E/F
series aircraft, the ceiling and visibility requirements shall be the minimums authorized by
OPNAVINST 3710.7 series, namely field minimums not less than 200/½.
WSOs must attend instrument ground school and satisfactorily complete a written examination to
be instrument qualified.
5.4.3 Ceiling/Visibility Requirements. The ceiling and visibility requirements for takeoff and
landing for pilots who are not fully instrument qualified in E/F-series aircraft are delineated in figure
5-2.
F/A-18E/F Actual
Flight Experience
Pilot With Current
Instrument Rating,
F/A-18A-D NATOPS
Qualification and
>300 hours in model
Pilot With Current
Instrument Rating
and >500 Hours in
Tactical Aircraft
IP in rear of F(T)
Pilot With Current
Instrument Rating
300/1
None
1000/3 for takeoff/
landing, and training in
clear air mass
Complete 1st FAM
flight *
Circling mins for takeoff/
landing, and training in
clear air mass
Complete 2nd FAM
flight *
F/A-18E/F instrument
qualified
Remain VMC and VFR
1000/3 for takeoff/
landing, and training in
clear air mass
Remain VMC and VFR
Complete 2nd FAM
flight *, and IAC in rear
cockpit or complete all
FAM flights * ++
−
Circling mins for takeoff/
landing, and training in
clear air mass
1000/3 for takeoff/
landing, and training in
clear air mass
F/A-18E/F NATOPS
check and >10 hours
(FPT)
−
F/A-18E/F instrument
qualified
Circling mins for
takeoff/landing, and
training in clear air mass
F/A-18E/F NATOPS
check and >40 hours
(FPT)
−
−
F/A-18E/F instrument
qualified
* CNO approved syllabus
++ Excluding CQ introduction
Figure 5-2. Pilot Ceiling and Visibility Restrictions Prior to Instrument Qualification
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A1-F18EA-NFM-000
5.5 WAIVERS
Unit Commanding Officers are authorized to waive, in writing, minimum flight and/or training
requirements in accordance with OPNAVINST 3710.7 series.
5.6 PERSONAL FLYING EQUIPMENT
The minimum requirement for personal flying equipment is contained in OPNAVINST 3710.7
series. In addition, all F/A-18E/F aircrew shall use the latest available flight safety and survival
equipment authorized by the Aircrew Personal Protective Equipment Manual (NAVAIR 13-1-6).
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ORIGINAL
A1-F18EA-NFM-000
PART III
NORMAL PROCEDURES
Chapter 6 - Flight Preparation
Chapter 7 - Shore-Based Procedures
Chapter 8 - Carrier-Based Procedures
Chapter 9 - Special Procedures
Chapter 10 - Functional Checkflight Procedures
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A1-F18EA-NFM-000
CHAPTER 6
Flight Preparation
6.1 MISSION PLANNING
6.1.1 General. All aircrew shall be responsible for preflight planning and preparation of required
charts, route navigation computations including fuel planning, checking weather and NOTAMS, and
for filing required flight plans. Refer to Part XI, Performance Data or approved fuel planning software,
to determine fuel consumption and profile. Planned minimum on deck fuel should not be less than
1,800 lb. The aircrew shall refer to applicable tactical publications to plan specialized missions.
6.1.2 Flight Codes. The proper flight classification and flight purpose codes to be assigned to
individual flights are established by OPNAVINST 3710.7 (Series).
6.2 BRIEFING/DEBRIEFING
6.2.1 Briefing. The flight leader is responsible for the briefing of each aircrew in the flight on all
aspects of the mission to be flown. A standard briefing guide shall be used in conducting the briefing.
Briefs shall include applicable ADMIN, TAC ADMIN, and MISSION CONDUCT. Aircrew qualified
to assume the mission lead shall record all data necessary to complete the mission. The briefing guide
should include the following:
6.2.1.1 NATOPS Admin Briefing Guide
General
Time hack
Objectives
Mission (Primary, Secondary)
Training
Julian date, event number
Times
Walk
Start
Check In
Taxi
Takeoff
Land
Debrief
Line up
Callsigns
Aircraft assigned
Crew (Msn Commander, Alternate Lead)
A/A TACAN
Radar channels, search block
Loadout Gross Weight, Max Trap
Comm Plan
Frequencies
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A1-F18EA-NFM-000
Controlling Agencies
IFF Procedures
Alpha Check
Waypoint Plan
Weather, NOTAMs
Launch, Mission, Recovery, Divert
Sunrise, Sunset, Moonrise, Moonset
Water / Air Temperature
Joker / Bingo / Fuel Ladder
Preflight
Aircraft
Ordnance
Ground / On-deck
Line / Deck and Start Procedures
Final checks
Clearance
Arming
Marshal
Taxi
Takeoff / Launch
Duty Runway / Ships Heading
Ships Posit, PIM
Type Takeoff / Case Departure
Takeoff Data (NWLO, T/O, Abort speeds, Distance)
Catapult Endspeed, Trim (asymmetrical)
Takeoff Checks
Departure Procedures
En Route
Rendezvous (Location, Speed)
En Route Formation
Route of Flight
Op Area
Range Info, Altitudes, Restrictions
Target Time, Range Event Number
Controlling Agency
Entry / Exit procedures
RTB / Recovery
Rendezvous (Location, Speed)
Battle Damage Checks
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A1-F18EA-NFM-000
Formation
Controlling Agency
Route of Flight
Airfield Recovery Procedures
Ship Recovery Procedures
Case Recovery
Marshal
Recovery Time
Type entry
Overhead, Break Interval
Straight-In / GCA
Type Landing
Post Landing
Clearing Landing Area
Configuration Changes
Comm
Taxi
De-Arming
Parking (Line / Hotpits / Hotseat)
Contingencies
Allowable Slide Time
Go / No Go Criteria
Fallouts, Spares
Bent Radar / Sensor / Weapon
Hung / Unexpended Ordnance
Weather
Emergencies
Abort, Field Arrestment
Loss of Brakes, Emergency Cat Flyaway
Inflight Emergencies / System Failures
NORDO, Lost Comm / Lost Sight
Midair, Bird Strike
Divert / BINGO
Ejection / SAR
ORM
Training Rules
ACM
NVG
LAT
LATT
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ORIGINAL
A1-F18EA-NFM-000
Aircrew Coordination
Refer to Chapter 28.
6.2.1.2 NATOPS Tactical Admin Briefing Guide.
Environmentals
Sun / Moon
Winds
Conning Altitudes, Conn Check
Cloud cover
Decks (Hard / Soft)
Weapons Checks
G-warm, Inverted Check
Expendables Check
Fence Checks Complete
CVRS - Tapes
Knock It Off / Terminate Calls
Fuel & G Checks
6.2.2 Debriefing. Post-flight debriefing is an integral part of every flight. The flight leader shall
conduct a mission debrief to include ADMIN, TAC ADMIN, SAFETY OF FLIGHT, and MISSION
CONDUCT. Emphasis shall be placed on identifying and correcting errors and poor techniques.
Debrief shall include all available aircrew and be conducted in a timely manner.
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A1-F18EA-NFM-000
CHAPTER 7
Shore-Based Procedures
NOTE
• F/A-18F WSO responsibilities are italicized.
• For F/A-18F crew coordination specifics, including individual aircrew
responsibilities and ICS communications, refer to Chapter 29, Crew
Coordination Standards.
7.1 PREFLIGHT CHECKS
7.1.1 In Maintenance Control. The A-sheet must be checked for aircraft status, configuration,
armament loading, and servicing prior to manning the aircraft. Review the aircraft discrepancy book
(ADB) for (1) all outstanding discrepancies and (2) at least the last 10 flights worth of discrepancies
and corrective action. Weight and balance clearance is the responsibility of the Maintenance
Department.
7.1.2 Inspection of RCS Reduction Features. The most critical RCS reduction features/treatments
include (1) EMIS III radar bulkhead shields, (2) canopy and windshield coatings, (3) engine inlet
devices, and (4) outer moldline mismatch/gap control. To ensure that the survivability characteristics
of the aircraft are retained, attention should be focused on the following areas:
1. On missions where the full RCS reduction potential of the aircraft is desired (typically wartime
environment only), ensure the twelve missionized EMIS III radar bulkhead shields are installed.
Additionally, ensure all SUU-79 pylons are fitted with their LO hardware (CAD access covers and
four bolt fairings).
2. Typically, if minor damage to canopy or windshield coatings is visually acceptable for flight, the
RCS reduction potential of the coatings should be retained.
3. Typically, if minor damage to the inlet lip/duct RAM coatings or to the inlet device are acceptable
from a FOD standpoint, the RCS reduction potential of the coatings/device should be retained.
4. Mismatches and gaps in the outer moldline of the aircraft can substantially reduce RCS reduction
potential. Care should be taken to note and repair damage to RAM coatings and FIP seals,
particularly around frequently opened panels. Doors and panels should be flush with the surrounding structure and gaps should be filled. In general, a rule of thumb for an acceptable amount of
panel/structure mismatch is no greater than the thickness of a PCL cover. With mismatches greater
than that width, some RCS reduction potential is lost. The most critical gaps are those aligned
perpendicular to the longitudinal axis of the aircraft (e.g., vertical gaps between side panels and 3-9
line gaps between underside panels). Gaps and mismatches that run along the longitudinal axis of
the aircraft are less critical.
In general, at least 75% of the perimeter of every door should exhibit good FIP seal integrity (e.g.,
sealed and flush). RAM coating damage should not exceed 25% of the total RAM area in any particular
location (e.g., around flap hinges or main landing gear door edges). Multi-layer RAM patches forward
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ORIGINAL
A1-F18EA-NFM-000
Figure 7-1. Exterior Inspection
of the inlet should show no sign of disbonding or peeling. All conductive tape, the windshield aft arch
termination strip, NLG blade seals, canopy and wing conductive bulb seals, and TEF/rudder boots
should be fully bonded, with no loose or peeling corners or edges. It is normally acceptable to trim loose
materials that are noticed just prior to flight.
7.1.3 Exterior Inspection. The exterior inspection (figure 7-1) is divided into 20 areas, beginning at
the left forward fuselage and continuing clockwise around the aircraft. Check doors secure and be alert
for loose fasteners, cracks, dents, leaks, and other general discrepancies. Close inspection shall be given
to the left and right nose sections for moldline defects (bends, dents, dings, cracking or blistering of
exterior coatings, or other surface discrepancies) that will affect the pitot-static system and RVSM
capability.
1. Nose landing gear
a. Drag brace/fairing - CHECK CONDITION
b. Drag brace ground safety pin - REMOVED
c. Holdback fitting - CHECK CONDITION
d. Tires and wheels - CHECK CONDITION
e. Ensure key washer not in direct contact with wheel hub.
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ORIGINAL
A1-F18EA-NFM-000
f. Strut piston chrome exposed - 6 INCHES
g. Launch bar - CHECK CONDITION
h. Nosewheel steering assembly - CHECK CONDITION
i. Tiedown rings (2) - CHECK FOLDED AGAINST STRUT
j. Taxi and approach lights - CHECK CONDITION
k. Strut pressure gauges (2) - CHECK IN THE GREEN
l. Retract actuator - CHECK CONDITION
m. Strut - CHECK CONDITION
2. Nose wheelwell
a. Maintenance code switch - SELECT RESET MOMENTARILY (applies power to the SMS
processor).
b. Doors and linkages - CHECK CONDITION
3. Nose section (left side)
a. Fuselage surface - CHECK CONDITION (no bends, dents, dings, cracking or blistering of
exterior coatings, or other surface discrepancies)
b. Safety switches - CHECK
(1) Expendables - Yellow when out
(2) Gun electrical - Orange when out
(3) Gun holdback - Orange when out
c. Gun - PREFLIGHT
d. AOA probe - CHECK CONDITION
(1) Smooth, concentric rotation through the full range of travel to include while gently pulling
and pushing the AOA probe.
(2) No bends, dents, dings, cracking or blistering of exterior coatings, or other surface
discrepancies.
e. Pitot tube - CHECK CONDITION (no bends, dents, dings, cracking or blistering of exterior
coatings, or other surface discrepancies).
f. Pitot static drains (4) - CLOSED (underside)
g. Forward antennas - CHECK CONDITION
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ORIGINAL
A1-F18EA-NFM-000
(1) Blade antenna (Comm 1, DL, IFF)
(2) Hump antenna (ALR-67 Low band array)
(3) Flush chevron antenna (ICLS, ACLS)
h. Radome - CHECK SECURE (2 points)
4. Nose section (top)
a. Gun blast diffuser and gun port - CLEAR
5. Nose section (right side)
a. Fuselage surface - CHECK CONDITION (no bends, dents, dings, cracking or blistering of
exterior coatings, or other surface discrepancies)
b. Radome - CHECK SECURE (2 points)
c. AOA probe - CHECK CONDITION
(1) Smooth, concentric rotation through the full range of travel to include while gently pulling
and pushing the AOA probe.
(2) No bends, dents, dings, cracking or blistering of exterior coatings, or other surface
discrepancies.
d. Pitot tube - CHECK CONDITION (no bends, dents, dings, cracking or blistering of exterior
coatings, or other surface discrepancies).
e. Refuel cap - ON
f. Refuel door (8R) - CLOSED/SECURED
6. Forward fuselage (right side)
a. Aft blade antenna (Comm 2, TCN) - CHECK CONDITION
b. Flush LEX antenna (ALQ-165 low/high band transmitter) - CHECK CONDITION
c. SMS processor - CHECK WEAPON/FUZE CODES
d. DOOR 13R - CLOSED/SECURED
e. Right engine intake - CLEAR
f. Heat exchanger ram air inlet (top, inside intake) - CLEAR
g. Chaff/flare dispensers (2) - PREFLIGHT (ensure chaff/flare buckets or access covers
installed).
7. External fuel tank(s) - PREFLIGHT
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ORIGINAL
A1-F18EA-NFM-000
a. Refuel cap - DOWN, LOCKED, ARROW FORWARD
b. Precheck valve - DOWN, FLUSH
8. Station 7 missile, NFLR, or LDT (if installed) - PREFLIGHT
9. Right main landing gear - CHECK
a. Tire and wheel - CHECK CONDITION
b. Brake wear indicator - EXTENDED (not flush or below flush)
c. Planing link - CHECK CONDITION
d. Strut - CHECK CONDITION
e. Tiedown rings (2) - CHECK SPRING CONDITION
f. Landing gear pin - REMOVED
10. Right wing
a. Flush LEF antenna (ALR-67/ALQ-165 receivers) - CHECK CONDITION
b. LEF - CHECK CONDITION
c. Pylons and external stores - PREFLIGHT
d. Wingfold area - CHECK CONDITION AND VERIFY WINGFOLD PIN REMOVED
e. Position lights - CHECK CONDITION
f. LAU-7 - ENSURE NITROGEN BOTTLE INSTALLED AND DOORS SECURE
g. Wingtip AIM-9 (if installed) - PREFLIGHT
h. Aileron - CHECK CONDITION, FAIRED WITH WINGS FOLDED
If the wings are folded, note the position of the ailerons. If the aileron
locking pins do not restrain the ailerons in the faired position, ensure the
ailerons are moved to a faired or outboard position prior to engine start
to preclude damage to the ailerons and TEFs.
i. TEF - CHECK CONDITION
11. Right main wheelwell
a. Hydraulic filter indicators (delta-Ps) - NOT POPPED
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ORIGINAL
A1-F18EA-NFM-000
b. APU accumulator gauge - CHECK (3,000 psi nominal)
c. APU handpump handle - STOWED AND PINNED
d. Doors and linkages - CHECK CONDITION
e. Landing gear downlock and retract actuators - CHECK CONDITION
12. Aft fuselage (right side)
a. Vertical tail and rudder - CHECK CONDITION
b. Strobe light - CHECK CONDITION
c. Fuel vent outlet - CLEAR
d. Light/antenna radomes - CHECK CONDITION
(1) Tail light (top)
(2) ALQ-165 low/high band receiver (middle)
(3) ALR-67 receiver (bottom)
e. Dump outlet - CLEAR
f. Stabilator - CHECK CONDITION
g. Exhaust nozzle and afterburner section - CHECK CONDITION
13. Arresting hook area
a. Arresting hook - CHECK CONDITION (make sure cotter key installed in hook point attach
bolt)
b. Arresting hook pin - REMOVED
c. Towed decoy stand-off bar (when installed) - CHECK CONDITION
14. Aft fuselage (left side)
a. Exhaust nozzle and afterburner section - CHECK CONDITION
b. Stabilator - CHECK CONDITION
c. Vertical tail and rudder - CHECK CONDITION
d. Antenna radomes - CHECK CONDITION
(1) ALQ-165 high band transmitter (top)
(2) ALQ-165 low band transmitter (middle)
III-7-6
ORIGINAL
A1-F18EA-NFM-000
(3) ALR-67 receiver (bottom)
e. Dump outlet - CLEAR
f. Strobe light - CHECK CONDITION
g. Fuel vent outlet - CLEAR
15. Aft fuselage (underside)
a. APU intake (screened) and exhaust ducts - CLEAR
b. ATS exhaust ducts (screened) - CLEAR
c. ALE-50 dispenser - CHECK CONDITION (make sure live or protected magazine installed)
16. Left main wheelwell
a. Hydraulic filter indicators (delta-Ps) - NOT POPPED
b. Doors and linkages - CHECK CONDITION
c. Landing gear downlock and retract actuators - CHECK CONDITION
d. Landing gear pin - REMOVED
17. Left wing
a. TEF - CHECK CONDITION
b. Aileron - CHECK CONDITION, FAIRED WITH WINGS FOLDED
If the wings are folded, note the position of the ailerons. If the aileron
locking pins do not restrain the ailerons in the faired position, ensure the
ailerons are moved to a faired or outboard position prior to engine start
to preclude damage to the ailerons and TEFs.
c. Wingtip AIM-9 (if installed) - PREFLIGHT
d. LAU-7 - ENSURE NITROGEN BOTTLE INSTALLED AND DOORS SECURE
e. Position lights - CHECK CONDITION
f. Wingfold area - CHECK CONDITION AND VERIFY WINGFOLD PIN REMOVED
g. Pylons and external stores - PREFLIGHT
h. LEF - CHECK CONDITION
III-7-7
ORIGINAL
A1-F18EA-NFM-000
i. Flush LEF antenna (ALR-67/ALQ-165 receivers) - CHECK CONDITION
18. Left main landing gear
a. Tire and wheel - CHECK CONDITION
b. Brake wear indicator - EXTENDED (not flush or below flush)
c. Planing link - CHECK CONDITION
d. Strut - CHECK CONDITION
e. Tiedown rings (2) - CHECK SPRING CONDITION
19. Station 5 missile or TFLR (if installed) - PREFLIGHT
20. Forward fuselage (left side)
a. Chaff/flare dispensers (2) - PREFLIGHT (ensure chaff/flare buckets or access covers
installed).
b. Left engine intake - CLEAR
c. Heat exchanger ram air inlet (top, inside intake) - CLEAR
d. Fuel cavity drains (underside) - VERIFY NO LEAKS
e. Loose fasteners - CHECK
f. Flush LEX antenna (ALQ-165 low/high band transmitter) - CHECK CONDITION
7.1.4 Before Entering Cockpit.
1. Ensure all doors forward of the intakes are secured properly.
2. Boarding ladder - SECURE (2 points)
3. Fuselage (upper surface)
a. Spoilers - CHECK CONDITION
b. Upper blade antenna (Comm 1, DL, TCN) - CHECK CONDITION
c. ECS auxiliary duct doors - CHECK DOWN/CONDITION
d. Maintenance handle - CHECK STOWED
4. (LOTs 21-24) Over-the-shoulder cameras (2) - CHECK SECURE
5. CVRS tapes - INSTALL IN RECORDERS (if desired)
III-7-8
ORIGINAL
A1-F18EA-NFM-000
6. EMI shields (covers over bay behind ejection seat) - CLOSED/SECURED
Ensure the EMI shields are properly closed and secured prior to closing
the canopy to prevent damage to the shields and the canopy actuation
link.
7. Ejection seat SAFE/ARMED handle - SAFE
8. Ejection seat(s) - PREFLIGHT
a. Manual override handle - FULL DOWN and LOCKED
b. Right pitot - STOWED
c. Ballistic gas quick-disconnect - CONNECTED (indicator dowel flush or slightly protruding)
d. Top latch plunger locking indicator - FLUSH WITH THE END OF THE PLUNGER
If the top latch plunger locking indicator is not flush, the seat could come
loose on the mounting rails.
e. Catapult manifold valve - CHECK (hoses and manifold connected; retaining pin installed)
f. Parachute withdrawal line - CONNECTED/SECURED
g. Parachute container lid - SECURE
h. Left pitot - STOWED
i. Electronic sequencer - NOT ACTIVATED
(1) Indicator should be BLACK (not activated).
(2) White - CHECK THERMAL BATTERIES NOT ACTIVATED
j. Thermal batteries - NOT ACTIVATED
(1) Indicator should be WHITE or PINK (not activated).
(2) Black or purple is UNSAT (activated).
k. Console oxygen/comm lines - CONNECTED/SECURED
l. Survival kit - CHECK
(1) Oxygen/comm lines - CONNECTED/SECURED
III-7-9
ORIGINAL
A1-F18EA-NFM-000
(2) Emergency oxygen gauge - IN THE BLACK
(3) Seat pan - CHECK SECURED TO SEAT (pull up on front end to test security)
m. Radio beacon lanyard - SECURED TO COCKPIT FLOOR (make sure lanyard and quick
release connector are positioned forward of the underseat rocket motor tubes)
n. Lap belts - SECURE (pull up strongly on each belt to make sure bolt fittings are engaged in
the seat)
o. Leg restraint lines - CHECK
Check that leg restraint lines are secured to seat and floor and are not twisted. Check that lines
are routed first through the thigh garter ring, then through the lower garter ring, and then
routed outboard of the thigh garter ring before the lock pins are inserted into the seat just
outboard of the snubber boxes.
Failure to route the restraint lines properly through the garters could
cause serious injury during ejection/emergency egress.
p. Ejection seat firing initiators - CHECK FIRING LINKAGE CONNECTED TO SEARS
q. Parachute risers - CHECK (ensure risers are routed down the forward face of the parachute
container and are routed behind the retaining strap; pull on risers to check ease of operation).
r. SEAWARS - CHECK FOR PROPER INSTALLATION
s. (SJU-17A (V)1/A, 2/A, and 9/A) Backpad adjustment handle - SET TO DESIRED POSITION
For solo flight in the F/A-18F 9. Rear cockpit - SECURE
a. Ejection seat SAFE/ARMED handle - SAFE
b. Ejection control handle pin - VERIFY REMOVED
c. (LOTs 21-24) EMERG BRK handle - STOWED
The position of the EMERG BRK handle is the only indication that
emergency braking is selected; no warning or caution is displayed. The
EMERG BRK handle(s) must be fully stowed in both cockpits to ensure
that normal braking with anti-skid is available.
d. CANOPY JETT handle - OUTBOARD AND DOWN/PIN REMOVED
e. L(R) DDI and MPCD knobs - OFF
III-7-10
ORIGINAL
A1-F18EA-NFM-000
f. Comm 1 and 2 knobs - OFF
g. EJECTION MODE handle - SOLO/COLLAR INSTALLED
h. SEAT CAUT MODE switch - SOLO/PIN INSTALLED
i. Leg restraints, lap belts, parachute risers, JHMCS QDC - SECURED/STOWED
j. Loose items - SECURED
In trainer configured aircraft k. Control stick - CHECK SECURE
l. UFCD adapter - VERIFY NOT INSTALLED
Forward stick throw is restricted if a rear cockpit control stick and UFCD
adapter are both installed.
m. Throttles - CHECK CONDITION
7.1.5 Interior Checks - Pilot.
Do not place any item on the glare shield, as scratching the windshield is
probable.
1. Leads, leg restraints, and harness - SECURE/ADJUST
Connect oxygen, g suit, QDC (if applicable) and communications leads. Check routing of JHMCS.
UHVI does not interfere with oxygen hose. Check QDC is securely connected or stowed if not in use.
Fasten and secure leg restraint garters and lines. Check leg garters buckled and properly adjusted
with hardware on inboard side of the legs. Connect and adjust lap belt straps. Attach parachute
Koch fittings to harness buckles. Check operation of shoulder harness locking mechanism.
The leg restraint lines must be buckled at all times during flight to ensure
that the legs will be pulled back upon ejection. This enhances seat
stability and prevents leg injury by keeping the legs from flailing
following ejection.
III-7-11
ORIGINAL
A1-F18EA-NFM-000
The JHMCS UHVI must be properly routed through the torso bundle
flue under the survival vest and the QDC secured in the QMB to ensure
that no entanglement exists with the oxygen hose. Misrouting of the
JHMCS UHVI may allow the QDC to rub against the oxygen hose
disconnect causing unintentional oxygen/communications disconnect in
flight.
2. Ejection control handle - CLEAR
3. Ejection control handle pin - VERIFY REMOVED
Left console 1. Circuit breakers - IN
2. Manual canopy handle - STOWED
3. MC and HYD ISOL switches - NORM
4. OBOGS control switch - OFF
5. OXY FLOW knob - OFF
6. OBOGS monitor pneumatic BIT plunger - VERIFY UNLOCKED AND FULLY EXTENDED
Inadvertent rotation of the OBOGS monitor pneumatic BIT plunger
while pressed can result in the locking of the plunger in a maintenance
position and may result in intermittent OBOGS DEGD cautions and lead
to hypoxia. Rotation of the BIT plunger disengages the locking slot
allowing the plunger to extend and move freely when pushed.
7. COMM 1/IFF ANT SEL switches - AUTO/BOTH
8. COMM panel - SET
a. RLY and GXMT switches - OFF
b. ILS CHANNEL/ILS switch - SET/UFCD
c. CRYPTO, MODE 4, (IFF) MASTER switches - NORM/OFF/NORM
9. VOL panel - SET AS DESIRED
10. FCS GAIN switch - NORM/GUARD DOWN
11. APU switch - OFF
III-7-12
ORIGINAL
A1-F18EA-NFM-000
12. PROBE switch - RETRACT
13. EXT TANKS switches - NORM
14. DUMP switch - OFF
15. INTR WING switch - NORM
16. GEN TIE CONTROL switch - NORM/GUARD DOWN
17. EXT LT panel - SET
a. EXT LT IDENT knob - NORM
b. FORM knob - AS REQUIRED
c. POS knob - AS REQUIRED
d. STRB switch - BRT/DIM/OFF (as required)
18. Throttles - OFF
19. External lights master switch - FORWARD
20. BRK PRESS switch - CHECK (2,600 psi min)
Instrument panel 1. PARK BRK handle - SET
2. LDG/TAXI LIGHT switch - OFF
3. ANTI SKID switch - ON
4. SELECT JETT knob - SAFE
5. FLAP switch - FULL
6. LAUNCH BAR switch - RETRACT
7. LDG GEAR handle - DN
8. Landing gear handle mechanical stop - FULLY ENGAGED
9. CANOPY JETT handle - FORWARD
10. MASTER ARM switch - SAFE
11. EMERG JETT button - NOT PRESSED IN
III-7-13
ORIGINAL
A1-F18EA-NFM-000
12. FIRE and APU FIRE warning lights - NOT PRESSED IN
NOTE
If a FIRE light is depressed, approximately 1/8 inch of yellow and
black stripes will be visible around the outer edges of the light.
13. L(R) DDI, HUD, and MPCD knobs - OFF
NOTE
Power to the UFCD is controlled by the MPCD knob, so the UFCD
knob does not need to be OFF.
14. COMM 1 and 2 knobs - OFF
15. CVRS mode switch - OFF
16. ALT switch - BARO or RDR
17. ATT switch - AUTO
18. Standby attitude reference indicator - CAGED
19. IR COOL switch - OFF
20. SPIN switch - NORM/GUARD DOWN
21. HOOK handle - UP
22. WINGFOLD switch - SAME AS WING POSITION
23. AV COOL switch - NORM
Pedestal Panel (LOTs 23 and up) 1. ECM JETT button - NOT PUSHED IN
2. JAMMER switch - OFF
3. RWR switch - OFF
4. DISPENSER switch - OFF
5. AUX REL switch - NORM
6. RUD PED ADJ lever a. SET PEDAL POSITION FULL FORWARD
D Cycle left and right pedal to check for binding.
III-7-14
ORIGINAL
A1-F18EA-NFM-000
b. SET PEDAL POSITION AS DESIRED FOR FLIGHT
D Locks securely when RUD PED ADJ lever released.
Restrain the rudder pedals during adjustment. Unrestrained release of
the rudder pedals may damage the rudder pedal mechanism.
Ensure the rudder pedals are locked in position after adjustment. Failure
to lock the rudder pedals may result in uncommanded forward rudder
pedal movement inflight.
Pedestal Panel (LOTs 21 - 22) 1. ECM JETT button - NOT PUSHED IN
2. DISPENSER switch - OFF
3. ECM knob - OFF
4. DECOY switch - OFF
5. AUX REL switch - NORM
6. RWR switch - OFF
7. Clock - CHECK AND SET
8. RUD PED ADJ lever a. SET PEDAL POSITION FULL FORWARD
D Cycle left and right pedal to check for binding.
b. SET PEDAL POSITION AS DESIRED FOR FLIGHT
D Locks securely when RUD PED ADJ lever released.
Right console 1. Circuit breakers - IN
2. GEN switches - NORM
3. BATT switch - OFF
4. ECS panel - SET
a. MODE switch - AUTO
III-7-15
ORIGINAL
A1-F18EA-NFM-000
Selection of MAN with the ECS mode switch is prohibited. Selecting
MAN while the aft cooling fan shutoff valve is open may cause the fan to
overspeed resulting in a catastrophic fan failure potentially leading to loss
of OBOGS.
b. CABIN TEMP knob - AS DESIRED
c. CABIN PRESS switch - NORM
d. BLEED AIR knob - OFF
e. ENG ANTI ICE switch - OFF
f. PITOT ANTI ICE switch - AUTO
5. DEFOG handle - MID RANGE
6. WINDSHIELD switch - OFF
7. INTR LT panel - SET
a. CONSOLES, INST PNL, and FLOOD knobs - AS DESIRED
b. CHART and WARN/CAUT knobs - AS DESIRED
c. MODE switch - DAY, NITE, or NVG (as required)
8. Sensor control panel - SET
a. FLIR, LTD/R, and LST/NFLR switches - OFF/SAFE/OFF
b. INS and RADAR knobs - OFF
9. NVG storage container - CHECK SECURE
7.1.6 Interior Checks - WSO.
1. Leads, leg restraints, and harness - SECURE/ADJUST
Connect oxygen, g suit, QDC (if applicable) and communications leads. Check routing of JHMCS
UHVI does not interfere with oxygen hose. Check QDC is securely connected or stowed if not in use.
Fasten and secure leg restraint garters and lines. Check leg garters buckled and properly adjusted
with hardware on inboard side of the legs. Connect and adjust lap belt straps. Attach parachute
Koch fittings to harness buckles. Check operation of shoulder harness locking mechanism.
III-7-16
ORIGINAL
A1-F18EA-NFM-000
• The leg restraint lines must be buckled at all times during flight to
ensure that the legs will be pulled back upon ejection. This enhances
seat stability and prevents leg injury by keeping the legs from flailing
following ejection.
• The JHMCS UHVI must be properly routed through the torso bundle
flue under the survival vest and the QDC secured in the QMB to
ensure that no entanglement exists with the oxygen hose. Misrouting
of the JHMCS UHVI may allow the QDC to rub against the oxygen
hose disconnect causing unintentional oxygen/communications disconnect in-flight.
2. Ejection control handle - CLEAR
3. Ejection control handle pin - VERIFY REMOVED
In trainer configured aircraft 1. Control stick - CHECK SECURE
2. UFCD adapter - VERIFY NOT INSTALLED
Forward stick throw is restricted if a rear cockpit control stick and UFCD
adapter are both installed.
3. Throttles - CHECK CONDITION
4. RUD PED ADJ lever - ADJUST PEDAL POSITION
Left console 1. OXY FLOW knob - OFF
2. Left hand controller - CHECK SECURE
3. CANOPY JETT handle - OUTBOARD AND DOWN
4. VOL panel - SET AS DESIRED
5. RECCE panel - SET
a. POD PWR knob - OFF
b. ATARS switch - OFF
III-7-17
ORIGINAL
A1-F18EA-NFM-000
Instrument panel 1. (LOTs 21-24) EMERG LDG GEAR handle - STOWED
2. (LOTs 21-24) EMERG BRK handle - STOWED
The position of the EMERG BRK handle is the only indication that
emergency braking is selected: no warning or caution is displayed. The
EMERG BRK handle(s) must be fully stowed in both cockpits to ensure
that normal braking with anti-skid is available.
Due to friction in the EMERG BRK handle mechanism, the handle may
not return to the fully stowed position unless positively pushed.
3. L(R) DDI and MPCD knobs - OFF
NOTE
Power to the UFCD is controlled by the MPCD knob, so the UFCD
knob does not need to be OFF.
4. EMERG JETT button − NOT PRESSED IN
5. COMM 1 and 2 knobs - OFF
6. Standby attitude reference indicator - CAGED
7. EJECTION MODE handle - NORM
Right console 1. Right hand controller - CHECK SECURE
2. INTR LT panel - SET
a. CONSOLES, INST PNL, and FLOOD knobs - AS DESIRED
b. CHART and WARN/CAUT knobs - AS DESIRED
3. NVG storage container - CHECK SECURE
7.2 ENGINE START
A self-contained (battery/APU) start is the primary method for starting the engines. The aircraft
also has provisions for starting on external power, external air, or opposite engine bleed air (crossbleed)
for circumstances when that may be appropriate (e.g., alert launch, low battery, maintenance, engine
restart after APU shutdown, etc.). As such, the steps for a ″normal″ battery/APU start are numbered
below, while steps for alternate starting sources are lettered.
III-7-18
ORIGINAL
A1-F18EA-NFM-000
With an external power start, all electrical systems are operative. With a battery start, power is
available to operate the APU and engine fire warning systems, the caution lights panel, the intercom
system between the aircrew and the ground crew, the cockpit utility light, and the EFD backup display.
The right engine is normally started first in order to provide normal hydraulics to the brakes. During
first engine battery start, the EFD RPM indication typically jumps from 0 to 5 or 10%, and light-off
is indicated by TEMP rising from a minimum reported value of approximately 190°C. When the
corresponding generator comes online (approximately 60%N2 rpm), the engine crank switch returns to
OFF. After both generators are online, the APU will run for 1 minute and then shut down
automatically.
• To prevent engine damage during start, if an engine was not idled
(75% N2 rpm or less) for 5 minutes prior to shutdown and a restart
must be made between 15 minutes and 4 hours after shutdown, the
engine must be motored for 1 minute at 29%N2 or greater before
restart.
• To prevent vibration and damage to compressor blades, do not allow
N2 rpm to dwell between 26 to 29%during engine motoring.
7.2.1 Intercockpit Communications (F/A-18F). The following Challenge/Response voice communications are mandatory:
Challenge
Response
Pilot - ICS check
WSO - Loud and clear
Pilot - Fire warning
WSO - Roger (Optional)
Pilot - Starting APU left/right
WSO - Roger (Optional)
WSO - Good waypoint zero
Pilot - Roger (Optional)
Pilot - Canopy
WSO - Clear/standby
7.2.2 Engine Start Checks. In the F/A-18F, the WSO must monitor pilot procedures, EFD indications, and plane captain signals to ensure maximum safety during engine start.
1. BATT switch - ON
III-7-19
ORIGINAL
A1-F18EA-NFM-000
2. Battery gauge - CHECK
NOTE
Nominal voltage for a ″good″ battery should be 23 to 24 vdc. Minimum
battery voltage is that which provides a successful engine start (i.e.,
APU remains online and the EFD remains powered to provide
indications of RPM and TEMP). EFD blanking and/or uncommanded
APU shutdown should be anticipated with a battery voltage at or
below approximately 18 vdc. If a weak battery results in an
unsuccessful engine start attempt, the battery should be charged or
replaced prior to takeoff, since the battery provides the last source of
electrical redundancy for the FCCs.
3. Caution Lights Panel - CHECK CABIN light on (if CPWS installed)
4. ICS - CHECK
With external electrical power a. EXT PWR switch - RESET
b. GND PWR switches 1, 2, 3, and 4 - B ON (hold for 3 seconds)
c. L(R) DDI, HUD, and MPCD knobs - ON (both cockpits)
d. COMM 1 and 2 knobs - ON/VOLUME AS DESIRED (both cockpits)
e. LT TEST switch - TEST (both cockpits)
f. MPCD/UFCD - ENTER DESIRED WAYPOINTS
All starts 5. FIRE warning test - PERFORM
a. FIRE test switch - TEST A (hold until all lights and aural warnings indicate test has been
successfully passed)
b. FIRE test switch - NORM (pause 7 seconds or cycle BATT switch for system reset)
c. FIRE test switch - TEST B (hold until all lights and aural warnings indicate test has been
successfully passed)
NOTE
During a successful FIRE warning test, ALL of the following lights
should illuminate in each TEST position: both FIRE lights (all 4
bulbs), the APU FIRE light (all 4 bulbs), and both L and R BLEED
warning lights. Additionally, the following voice aural warnings should
be heard in order: ″Engine fire left, engine fire right, APU fire, bleed
air left, bleed air right″ (each repeated twice).
III-7-20
ORIGINAL
A1-F18EA-NFM-000
NOTE
• A complete FIRE warning test is performed in each TEST position
because it is difficult to recognize a single unlit bulb in a FIRE light.
Since an aural warning does not annunciate if any of the FIRE or
BALD loops are bad, lack of an aural warning is the best cue to the
aircrew of a test failure.
• Failure to pause in NORM for at least 3 seconds between TEST A and
TEST B results in a false BALD failure MSP code.
6. Forward MPCD and UFCD knobs - ON
NOTE
Forward MPCD and UFCD need to be turned on to display the
backup HUD format and L/R ATS cautions.
If APU start 7. APU ACC caution light - VERIFY OFF
8. APU switch - ON (READY light within 30 seconds)
To prevent an APU running engagement and to prevent APU exhaust
torching, a minimum of 2 minutes must elapse between APU shutdown
and another APU start.
NOTE
If an APU fire or overheat condition is detected on the ground, the
APU fire extinguishing system will automatically shutdown the APU
and, after 10 seconds, will discharge the extinguisher bottle.
If external air start a. BLEED AIR knob - OFF
All starts -
Regardless of the engine start air source utilized, the corresponding GEN
switch should be ON, as the generator provides primary overspeed cutout
protection for the ATS.
9. ENG CRANK switch - R
III-7-21
ORIGINAL
A1-F18EA-NFM-000
10. Right throttle - IDLE (10% N2 minimum. Oil pressure should be a minimum of 10 psi within 30
seconds. Maximum transient EGT during start is 871°C).
NOTE
During ground starts only, the FADEC will automatically cut back fuel
flow to prevent EGT from exceeding 815°C. If required, fuel flow will
be reduced to the point of engine flameout. While this mechanization
is provided to prevent engine damage due to an overtemp, the aircrew
should not rely on it to prevent a hot start.
11. CFIT voice warnings - CHECK (OFP 18E: ″ROLL-LEFT...ROLL-LEFT″) (OFP 13E: ″ROLLOUT...ROLL-OUT″)
NOTE
MC1 does an ACI configuration check after the generator comes on
line during a cold start power-up by commanding the above voice
warning. If the ACI does not contain the appropriate software, the
″ROLL-LEFT...ROLL-LEFT″ voice warning is not heard and the MC1
will assume an incompatible ACI is installed in the aircraft, GPWS/
TAWS voice warnings are not available and only the recovery arrow
will be displayed during a CFIT condition.
12. Battery gauge - VERIFY 28 vdc
NOTE
If the battery gauge fails to reach approximately 28 vdc with one
generator online, a battery charger malfunction has occurred which
requires maintenance action prior to flight.
13. L(R) DDI, HUD, and MPCD knobs - ON (both cockpits)
During a battery start of the right engine, recognition of a R ATS caution
will be delayed until the displays are turned on. Once the displays are
powered, verify that the R ATS caution is not set.
14. HMD switch (if applicable) - ON
15. EFD - CHECK
III-7-22
ORIGINAL
A1-F18EA-NFM-000
Ground idle RPM
TEMP
FF
OIL
NOZ
61% minimum
250° to 590°C
600 to 900 pph
35 to 90 psi (warm oil)
77% to 83%
NOTE
Following the initial start of each engine, engine anti-ice airflow will
turn on automatically 45 seconds after the engine reaches idle power
and will remain on for 30 seconds, provided the throttle remains at
IDLE. The corresponding LHEAT or RHEAT advisory will be
displayed during this engine anti-ice functional test.
If external power start a. External electrical power - DISCONNECT
If APU or crossbleed start 16. BLEED AIR knob - NORM
NOTE
The bleed air shutoff valves close during the fire warning test, so the
BLEED AIR knob must be rotated from OFF to NORM with ac power
applied to reset the valves.
17. LT TEST switch - TEST (both cockpits)
For a crossbleed start ensure the APU switch is OFF. The operating engine should be advanced
to a minimum of 80% N2.
18. ENG CRANK switch - L
19. Left throttle - IDLE (10%N2 minimum)
20. ENG CRANK switch - CHECK OFF
21. EFD - CHECK
If external air start a. BLEED AIR knob - NORM
7.3 BEFORE TAXI CHECKS
1. WYPT 0 and MVAR - CHECK/SET
2. GPWS/TAWS - CHECK BOXED
3. INS knob - CV or GND (PARK BRK SET)
III-7-23
ORIGINAL
A1-F18EA-NFM-000
4. RADAR knob - OPR
5. FLIR and LST/FLR switches - AS DESIRED
6. UFCD avionics - TURN ON
a. RALT - ON/SET
b. TCN - ON, T/R, CH SET
c. IFF - ON/MODES UNBOXED
7. WINGFOLD switch - SPREAD
8. FCS RESET button - PUSH (verify RSET advisory displayed)
NOTE
Prior to takeoff (cycle to WoffW), a successful FCS RESET
automatically clears all BLIN codes.
If no reset (RSET advisory displayed) a. FCS exerciser mode - INITIATE (push the FCS RESET button while holding the FCS BIT
consent switch up)
In standard or warm conditions, do not initiate the FCS exerciser mode
multiple times in an attempt to get a successful FCS RESET. In such
conditions, multiple initiations may excessively elevate hydraulic system
temperatures, increasing actuator and hydraulic pump seal wear and
potentially decreasing component life.
b. FCS RESET button - PUSH (verify RSET advisory displayed)
After successful FCS reset 9. FLAP switch - AUTO
10. FCS IBIT - PERFORM
a. FCS BIT consent switch - HOLD UP THEN PRESS THE FCS OPTION
b. AOA warning tone - VERIFY ANNUNCIATION AT FCS IBIT COMPLETION
c. FCS A and FCS B BIT status - VERIFY GO
d. FCS display - VERIFY NO BLIN CODES
III-7-24
ORIGINAL
A1-F18EA-NFM-000
Flight with BLIN codes could result in a FCS failure and aircraft loss.
Pressing the FCS RESET button simultaneously with the paddle switch
does not correct BIT detected FCS failures; it simply clears the BLIN
codes from the FCS display. FCS IBIT must be re-run after clearing
BLIN codes to ensure that previously detected failures no longer exist. If
BLIN codes remain following IBIT, maintenance action is required to
identify and correct failures in the FCS.
NOTE
• With the wings folded, both ailerons are X’d out, but no aileron BLIN
codes should be displayed. Even with wings folded, there are aileron
functions tested that may reveal FCS failures via valid BLIN codes.
• For FCS IBIT to start, the FCS BIT consent switch must be held for
at least 2 seconds. If not held for the required time, FCS A and FCS B
will indicate RESTRT on the BIT status line. If RESTRT is displayed, select STOP on the FCS-MC sublevel display and then repeat
the initiation procedure.
• The FCS will not enter IBIT if the throttles are above 14° THA or
NWS is engaged.
• Do not operate any FCS related switches or move the stick or rudder
pedals while FCS IBIT is running, as this may produce false failure
indications.
• With the wings folded, a BIT status indication of GO will only be
displayed for approximately 2 seconds before reverting to a DEGD
indication. BIT status will return to GO when the wings are spread and
locked.
• If the FCS IBIT fails, FCS A and FCS B will indicate DEGD on the
BIT status line. Note surface X’s and/or BLIN codes and contact
maintenance personnel.
11. Trim - CHECK (check pitch, roll, and yaw trim for proper movement in all directions)
NOTE
It is not possible to trim the stabilators to negative values (TED) with
WonW.
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A1-F18EA-NFM-000
12. T/O TRIM button - PRESS UNTIL TRIM ADVISORY DISPLAYED (stabilators 4° NU)
NOTE
If the TRIM advisory does not appear, longitudinal trim is not set for
takeoff. The CHECK TRIM caution will be displayed when both
throttles are advanced beyond 27° THA if the stabilators are trimmed
less than 3.5° TEU with the launch bar up (field takeoff) or 6.5° TEU
with the launch bar down (carrier takeoff).
13. Controls - CHECK (tolerance ±1°)
a. Control stick - CYCLE
(1) Full aft
- CHECK 24° NU STABILATOR (check left and right stabilators
track symmetrically within ±1° of each other)
At certain ejection seat heights, the A/A weapon select switch may hook
the EJECTION handle at the full aft stick deflection. With an armed
seat, inadvertent ejection initiation may occur if the stick returns to a
neutral position.
(2) Full fwd - CHECK 20° ND STABILATOR (check left and right stabilators
track symmetrically within ± 1° of each other)
(3) Full L/R - CHECK 30° DIFFERENTIAL STABILATOR (21° with tanks or A/G stores on
any wing station)
- CHECK DIFFERENTIAL TEFs
b. FLAP switch - HALF
c. Rudder pedals - CYCLE RUDDERS 40° L/R
d. FLAP switch - FULL (carrier-based)
e. TRIM - SET FOR CATAPULT LAUNCH (carrier-based)
14. PROBE, speedbrake, LAUNCH BAR switches and HOOK handle - CYCLE (LAUNCH BAR
optional for shore based operations.)
Ensure HOOK handle is fully up to minimize the probability of inadvertent hook drops during
catapult launch.
15. Pitot and AOA heat check - PERFORM
a. PITOT ANTI ICE switch - ON
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A1-F18EA-NFM-000
b. Make sure ground crew verify proper operation
c. PITOT ANTI ICE switch - AUTO
Failure of both AOA probe heaters in icing conditions may cause a sharp
uncommanded nose down attitude, uncontrollable by normal stick forces
or paddle switch actuation.
16. AV COOL switch emergency check (if ground personnel present)
a. AV COOL switch - EMERG then release
b. Make sure ground crew verify proper operation and stows emergency scoop.
17. APU - VERIFY OFF
18. FLBIT option - SELECT
19. BINGO - CHECK/SET
20. CVRS - AS DESIRED (both cockpits)
21. Standby attitude reference indicator - UNCAGE AND ERECT (both cockpits)
22. Altimeter setting - SET (both cockpits)
a. Altimeter setting displayed on HUD.
b. HUD altitude displayed within ±30 feet of parking spot elevation.
c. Standby altimeter within ±60 feet of parking spot elevation.
NOTE
If the standby altimeter barometric pressure is adjusted during the
FCS IBIT, the altitude reading displayed in the HUD will not change
until the IBIT is complete.
23. INS - CHECK
a. Alignment status - VERIFY COMPLETE
b. GPS HERR/VERR - VERIFY WITHIN LIMITS
c. INS knob - NAV or IFA
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A1-F18EA-NFM-000
NOTE
• Prior to placing the INS switch to IFA for a GPS alignment or for
AINS position keeping, ensure valid GPS data is available. AINS
position keeping is normally available when GPS HERR and VERR
are each less than 230 feet. Double digit GPS HERR/VERR (less than
100 feet) should guarantee AINS position keeping is available.
• Selecting IFA without good GPS data and without a complete carrier
or ground alignment will cause the INS to attempt to perform a radar
IFA and will halt/prevent alignment. If this occurs, return the INS
knob to GND or CV (as appropriate).
d. Verify HUD airspeed indicates less than 50 kts.
24. MUMI/ID - SELECT/ENTER DATE and FLT
25. Stores page - VERIFY PROPER STORE INVENTORY AND STATION STATUS
26. ZTOD/LTOD - BOX TO ENABLE HUD DISPLAY (if desired)
NOTE
• The TIMEUFC option is removed from the HSI format if INS
alignment is being performed (GND, CV, or IFA GPS).
• At GPS power up (first GEN online), SDC time and date are
automatically sent to the GPS to aid the acquisition of satellites. After
satellite acquisition, the GPS backloads satellite time to the SDC thus
synchronizing the SDC with precise GPS time. This GPS time
backload is only performed once per flight after the initial MC1 power
up.
• Manually changing ZTOD, LTOD, or the DATE with WonW resets
SDC time and/or date and reinitializes the GPS (even if GPS had a
good satellite acquisition). GPS reinitialization will delay the availability of AINS position keeping. GPS time synchronization will not be
available until a subsequent MC1 power-up (cold start). ARC-210
radios will need to be re-synched to enable HAVE QUICK operations.
27. Weapons/sensors - ON/BIT CHECK (as required)
28. BIT page - NOTE DEGD/FAIL INDICATIONS
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A1-F18EA-NFM-000
29. HMD - ALIGN (both cockpits)
NOTE
Canopy must be down and locked to align HMD/AHMD.
(CVRS record HMD if desired)
a. SUPT/HMD/ALIGN page - SELECT
b. Superimpose the HMD alignment cross on the HUD/BRU alignment cross.
c. Cage/Uncage button - PRESS and HOLD until ALIGNING turns to ALIGN OK or ALIGN
FAIL
If ALIGN FAIL d. Repeat steps b and c.
If ALIGN OK and HMD alignment crosses are not coincident with HUD/BRU alignment cross d. Perform FINE ALIGN.
(1) With FA DXDY displayed, use TDC to align azimuth and elevation HMD alignment
crosses with the HUD/BRU alignment cross.
(2) Cage/Uncage button - PRESS and RELEASE
(3) With FA DROLL displayed, use TDC to align the roll axis HMD alignment crosses with the
HUD/BRU alignment cross.
(4) Cage/Uncage button - PRESS and RELEASE
If satisfied with alignment e. ALIGN - UNBOX
30. Standby attitude data - CHECK
a. ATT switch - STBY
b. Verify INS attitude data is replaced by standby attitude data on the HUD and check
agreement of standby and INS data.
c. ATT switch - AUTO
31. OBOGS system - CHECK
a. OBOGS control switch - ON
b. OXY FLOW knob - ON/MASK ON (both cockpits)
c. OBOGS flow - CHECK
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A1-F18EA-NFM-000
d. OBOGS monitor electronic BIT pushbutton - PRESS AND RELEASE
e. Verify OBOGS DEGD caution set and removed (within 15 seconds).
f. OXY FLOW knob - OFF/MASK OFF (both cockpits)
Continued operation and use of the OBOGS system with an OBOGS
DEGD caution may result in hypoxia.
7.4 TAXI CHECKS
1. Canopy - EITHER FULL UP OR FULL DOWN FOR TAXI
Taxiing with the canopy at an intermediate position can result in canopy
attach point damage and failure.
2. Normal brakes - CHECK
3. Nosewheel steering - CHECK IN HIGH MODE L/R
NOTE
When using brakes, apply firm, steady brake pedal pressure. Use
nosewheel steering whenever possible, minimizing differential braking.
Avoid dragging brakes or light brake applications except as necessary
for drying wet brakes. Wet brakes can degrade brake effectiveness by
as much as 50%. Hard momentary braking with wet brakes during taxi
can reduce drying time. At heavy gross weight, make all turns at
minimum speed and maximum practical radius.
7.5 TAKEOFF
7.5.1 Before Takeoff Checks.
For MAX power catapult launches 1. ABLIM option - BOX
2. ABLIM advisory - VERIFY DISPLAYED
For all takeoffs 3. CHKLST page (figure 7-2)
a. FUEL TYPE - VERIFY
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A1-F18EA-NFM-000
Figure 7-2. Checklist Display
b. T.O. checklist - COMPLETE
Ensure the WINGFOLD switch is lever-locked in the SPREAD position.
If the wings are commanded to unlock or fold during a catapult shot, the
wings will unlock, the ailerons will fair, the wings may fold partially, and
the aircraft will settle.
WSO must make sure the EJECTION MODE handle is in AFT INITIATE (NORM) and, in Lot 21-24, aircraft the EMERG LDG GEAR and
EMERG BRK handles are fully stowed.
NOTE
EJECT SEL is displayed in the F/A-18F only.
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A1-F18EA-NFM-000
NOTE
Rear cockpit command eject is enabled when the EJECTION MODE
handle is in the AFT INITIATE position. When a passenger unfamiliar
with the F/A-18F occupies the aft cockpit, the NORM position may be
utilized.
4. Canopy - CHECK CLEAR/CLOSED
Prior to operating the canopy switch, confirm aircrew are clear to reduce
the potential for injury.
5. OXY FLOW knob - ON/MASK ON (both cockpits)
It is possible to place the OXY FLOW knob in an intermediate position
between the ON and OFF detents, which may result in a reduced flow of
oxygen. The OXY FLOW knob should always be fully rotated to the ON
or OFF detent position.
6. IFF sublevel - BOX REQUIRED MODES
7. PARK BRK handle - FULLY STOWED
8. ENG page - CHECK ENGINES AT MIL (if desired)
N1 RPM
N2 RPM
EGT
FF
NOZ POS
OIL PRESS
86 to 98%
88 to 100%
720 to 932°C
11,000 pph max
0 to 45% open
80 to 150 psi
7.5.2 Normal Takeoff. Predictions for takeoff performance (nosewheel liftoff speed, takeoff speed,
takeoff distance, and abort speed) should be calculated in the preflight brief based on aircraft
configuration and expected ambient conditions. These predictions are based on the following
technique: both engines stabilized at 80%N2 rpm, simultaneous brake release and throttle advance to
MIL or MAX, ½-aft (2.5 inches) stick rotation at the predicted nosewheel liftoff speed. This technique
should be used when ambient conditions and performance predictions warrant minimizing takeoff roll.
Review these numbers prior to takeoff.
The takeoff checklist should be completed prior to taking the duty runway. For single-ship takeoffs,
taxi to runway centerline and allow the aircraft to roll forward slightly to center the nosewheel. Begin
the takeoff roll by releasing the brakes, advancing the throttles from IDLE to MIL, and checking EGT
and RPM. If an afterburner takeoff is desired, further advance the throttles to MAX (full forward).
Check for proper afterburner light-off as indicated by both nozzles opening. As the aircraft accelerates
during the takeoff roll, track runway centerline using small rudder pedal inputs (e.g., NWS
commands). NWS is the most effective means of directional control during takeoff. Differential
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ORIGINAL
A1-F18EA-NFM-000
braking is much less effective and should therefore be avoided. The NWS system (low gain)
incorporates a yaw rate feedback input from the FCCs, which is designed to suppress directional PIO
tendencies by increasing directional damping during takeoff.
At nominal takeoff CG, aft stick will be required to rotate the aircraft. Approaching the predicted
nosewheel liftoff speed, ease the stick back to approximately 1/3 to 1/2 aft stick (1-1/2 to 2-1/2 inches).
Hold this input until the velocity vector rises to approximately 3 to 5°. Capture and climb/accelerate
at the desired flight path angle.
When clear of the ground with a positive rate of climb, raise the LDG GEAR handle and place the
FLAP switch to AUTO. In a flat takeoff attitude with MAX power selected, the aircraft will accelerate
rapidly towards gear speed. If required, reduce power to MIL or below to ensure the landing gear is up
and locked (light in the LDG GEAR handle is out) before passing 250 KCAS.
Takeoff performance is greatly affected by gross weight, center of gravity,
power setting, stabilator position, and ambient conditions. Under adverse
conditions (e.g., hot, heavy, and forward CG), takeoff speeds may be
significantly higher than those routinely seen at nominal conditions.
Knowing the aircraft’s predicted takeoff performance should prevent a
high speed abort in what is a normally functioning aircraft.
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ORIGINAL
A1-F18EA-NFM-000
• Under the most extreme conditions (e.g., hot, heavy, and forward CG),
nosewheel liftoff speed may exceed the nose tire limitation (195 KGS).
The takeoff technique and/or the aircraft configuration may need to
be adjusted to remain within limitations.
• Large aft stick inputs, particularly with CG near the aft limit, can
result in significant over-rotation. With pitch attitude above 10°, the
trailing edge of the stabilators can impact the ground if a large forward
stick input is used to check the over-rotation. Above 14° pitch
attitude, the engine exhaust nozzles may contact the ground. Therefore, pitch attitude shall not exceed 10° on takeoff.
• Takeoff with significant standing water (greater than 1/4 inch) on the
runway may cause water ingestion, which in extreme cases can cause
engine stalls, flameouts, AB blowouts, and/or engine FOD.
7.5.3 Crosswind Takeoff. Crosswind takeoffs should be performed using the normal takeoff technique. However, the pilot should expect to make slightly larger and more frequent rudder pedal inputs
to track runway centerline. As the aircraft accelerates and the ailerons become effective, lateral stick
into the wind may be desired to maintain wings level throughout the remainder of the takeoff roll and
rotation. As the aircraft becomes light on the main wheels, the aircraft will tend to yaw into the wind.
Slight main tire scrubbing can be expected. Allow the aircraft to crab into the wind at takeoff, while
continuing to maintain runway centerline during the gear transition and early climbout.
When calculating crosswind component for takeoff or landing, use the full
value of any reported gusts in your calculations.
7.5.4 After Takeoff Checks.
When definitely airborne 1. LDG GEAR handle - UP
2. FLAP switch - AUTO
7.6 AIRBORNE CHECKS
7.6.1 Climb. For safe maneuverability of the aircraft, up to 350 KCAS may be required up to 10,000
feet. For optimum climb performance, refer to A1-F18EA-NFM-200.
7.6.2 10,000 Foot Checks.
1. Cabin altimeter - VERIFY 8,000 FEET
2. Fuel transfer - CHECK INTERNAL and EXTERNAL
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A1-F18EA-NFM-000
3. RALT - CHECK/SET to 5,000 FEET
7.6.3 Cruise. Maximum range and maximum endurance data can be found in the performance charts
contained in A1-F18EA-NFM-200. Maximum range cruise is approximated by establishing 3.0° AOA,
but no faster than Mach 0.85. Maximum endurance cruise is approximated by establishing 3.7° AOA.
7.6.3.1 Cruise Check.
1. Cabin altimeter - MONITOR
Aircraft Altitude
Less than 8,000 feet
Cabin Altitude
Ambient
8,000 to 24,500 feet
8,000 feet
Greater than 24,500 feet
Alt x 0.4 (rule of thumb)
A slowly increasing cabin pressure altimeter may be the first or only
warning of a gradual loss of cabin pressurization.
7.6.3.2 RVSM Checks
When at assigned altitude on HUD 1. AOA Crosscheck (REQUIRED ONCE)
a. Compare L and R AOA values on FCS page.
b. If L and R AOA values differ by more than 2°, notify ATC that the aircraft is no longer RVSM
compliant.
2. Altitude Crosscheck (REQUIRED PERIODICALLY)
With MC OFP H5E AND UP a. Compare STBY CHK value (HSI/DATA/(A/C)) to standby altimeter. These are uncorrected
altitudes.
Otherwise a. Add standby altimeter error (Standby Altimeter Error table below) to standby altimeter and
compare to HUD altitude.
All aircraft b. If altitudes differ by more than 250 feet, notify ATC that the aircraft is no longer RVSM
compliant.
c. If an ″X″ appears to the right of the HUD baro altitude box, notify ATC that the aircraft is no
longer RVSM compliant.
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ORIGINAL
A1-F18EA-NFM-000
Standby Altimeter Error
Mach Number
Standby Altimeter Error (Feet)
0.50
0.60
0.65
0.70
0.75
0.80
0.85
0.90
0.92
120
150
175
200
220
280
260
330
380
7.7 LANDING CHECKS
7.7.1 Descent/Penetration. The windshield may fog rapidly under conditions of very high aircraft
descent rates and high humidity. In such conditions, consider preheating the windshield by placing the
DEFOG handle to HIGH and, if necessary, by placing the WINDSHIELD switch to either ANTI ICE
or RAIN. The maximum comfortable cockpit temperature should be maintained to aid in windshield
defog.
Normal instrument penetration is 250 KCAS with a 4,000 to 6,000 feet per minute descent rate. For
safe maneuverability of the aircraft, up to 350 KCAS may be required below 10,000 feet. Refer to
A1-F18EA-NFM-200, for optimum descent profiles. Before starting descent, perform the following:
7.7.1.1 Descent/Penetration Checks.
1. HOOK handle/HOOK BYPASS switch - AS REQUIRED/DESIRED
2. Exterior lights - SET FOR LANDING
3. Visual ID IDENT knob - NORM
4. ENG ANTI ICE switch - AS REQUIRED
5. PITOT ANTI ICE switch - AUTO
6. DEFOG handle - HIGH (if required)
7. WINDSHIELD switch - AS REQUIRED
8. Altimeter setting - CHECK
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ORIGINAL
A1-F18EA-NFM-000
9. RALT - CHECK/SET
10. NAV master mode - SELECT (compare HUD with standby flight instruments and standby
compass).
11. Navaids/MAG VAR - CROSSCHECK
12. ILS - ON/CHANNEL SET (if required)
13. IFF - AS DIRECTED
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ORIGINAL
A1-F18EA-NFM-000
Figure 7-3. Typical Field Landing Pattern
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SEE IC # 34
A1-F18EA-NFM-000
14. Weapons/sensors - OFF AS REQUIRED
7.7.2 VFR Landing Pattern Entry. See figure 7-3. Typically, the VFR landing pattern can be entered through
several methods: the break, downwind entry, VFR straight-in, or low approach/touch-and-go from a GCA.
Regardless of the entry method, enter the pattern at the altitudes and airspeeds prescribed by local course rules.
A normal break is performed by executing a level turn to downwind with the throttles reduced to IDLE and the
speedbrake function enabled (if required to reduce airspeed). The desired abeam distance is 1.0 to 1.3 nm. The
g-level required to achieve the desired abeam distance will be a fallout of break airspeed.
As airspeed decelerates below 250 KCAS, lower the LDG GEAR handle and place the FLAP switch to FULL.
If enabled, the speedbrake function will retract automatically when the FLAP switch is moved from the AUTO
position. Continue to decelerate to on-speed AOA (8.1 deg). Longitudinal trim inputs are required with the flaps
in HALF or FULL. The MI code for on-speed AOA is unit 14, address 15743, data 3300.
In-flight Memory Inspect (MI) of FCC (UNIT 14 or 15) addresses (ADDR)
greater than six digits long is prohibited since it may cause all four FCC channels
to shut down which will result in loss of aircraft control.
With MC OFP H3E AND UP, the pitch trim AOA value is displayed on the HUD while trimming and for two
seconds after trimming, and continuously on the FCS page with WoffW and flaps in HALF or FULL. The HUD
value is displayed with or without ATC engaged but will not be displayed with autopilot engaged. If the autopilot
is ″paddled off″ and AOA is greater than or equal to 6°, pitch trim is automatically set to on-speed. Trim the
aircraft hands-off and on-speed. Compare airspeed and AOA. Onspeed AOA is approximately 136 KCAS at
44,000 lb gross weight (max trap). Subtract (add) 1½ KCAS for each 1,000 lb decrease (increase) in gross weight.
Complete the landing checklist. When wings level on downwind, descend to pattern altitude (600 ft AGL for the
low pattern). Ensure the ground track pointer is on the exact reciprocal of runway heading.
7.7.2.1 Landing Checks.
1. Landing checklist - COMPLETE:
WHEELS
FLAPS
HOOK
ANTI SKID
HARNESS
DISPENSER
EJECT SEL
AOA
2. Report - AFT INITIATE, 3 DOWN AND LOCKED, FLAPS FULL (HALF), AOA CROSSCHECKED
7.7.3 VFR Landing Pattern and Approach. At the abeam position, pick a spot on the ground as a reference
point. (At the ship, TACAN will be used to adjust abeam distance). Remember this abeam position, as all abeam
distance corrections will use it as a reference. From the abeam position, time 20 seconds to arrive at a no-wind
180° position. To compensate for winds, subtract one second for each knot of final approach headwind
component. At the 180, roll into 27 - 30° AOB, add power, and adjust rate of descent to 300 to 400 fpm. Maintain
on-speed AOA. This should place the velocity vector about 1° below the horizon with its wingtip below the
horizon bar. If required, adjust rate of descent to arrive at the 90° position at 450 ft AGL. Develop an instrument
scan for the turn from the 180 to the 90, because an instrument scan will be required at the ship.
At the 90, glance at runway centerline and the lens and adjust AOB to arrive on extended centerline.
From the 90, rate of descent must be increased by reducing power and adjusting the velocity vector to
1½ to 2° below the horizon, on-speed. This will produce a rate of descent of 400 to 500 fpm to arrive
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ORIGINAL W/IC 34
A1-F18EA-NFM-000
at the 45° position at 320-370 feet AGL. From the 45, continue to increase rate of descent to
approximately 500-600 fpm with a power reduction to arrive at ″the start″ on centerline, at 220 to 250
feet AGL, with 650 to 750 fpm rate of descent, on-speed. The optimum rate of descent will vary with
glideslope angle, approach speed, and headwind component.
The approach turn from a pattern altitude greater than 600 ft AGL is slightly different. At the 180,
adjust rate of descent between 400 - 700 fpm to arrive at the 90 at approximately 500 ft AGL. This
requires a power reduction at the 180 rather than a power addition. Power will need to be added at the
90 to break the rate of descent to 400 to 500 fpm in order to arrive at the 45 at the same flight conditions
as the low pattern.
7.7.4 Pattern Adjustments. Deviations to the standard no-wind pattern will be required based on
headwind, crosswind, approach speed, and starts by adjusting abeam distance. Adjust the ground
reference point and fly exactly the same AOB as the previous pass. Correct for long-in-the-groove or
not-enough-straight-away starts by adjusting the timing from the abeam to 180° positions. Correct for
high or low starts by adding or subtracting 20 to 50 feet from the target altitudes at and inside of the
90. The purpose of pattern adjustments is to determine a repeatable pattern technique which will
produce consistent starts.
7.7.5 Final Approach. The desired final approach is flown by maintaining a centered ball to
touchdown on runway centerline and on-speed. Timely, well-controlled power corrections will be
required to capture and/or maintain the desired glideslope. A complete discussion of glideslope
geometry and glideslope corrections will be covered during the FRS training syllabus and/or by
squadron LSOs.
7.7.6 ATC Approaches. If an ATC approach is desired, engage ATC when wings level on downwind
at or near on-speed AOA. With ATC engaged, the aircraft must still be manually trimmed to on-speed
AOA. Unlike a manual throttles approach, nose position (i.e., velocity vector placement) now controls
power. Fly the same pattern as a manual approach. Coming off the 180, roll into 27 to 30° AOB and
lower the velocity vector approximately 1 to 2° below the horizon. ATC will add power as the aircraft
rolls into the turn. Reposition the velocity vector to maintain 300 to 400 fpm rate of descent. Passing
through the 90, lower the velocity vector slightly to pick up a 400 to 500 fpm rate of descent. Rolling
wings level in the groove, lower the velocity vector further to about 3°. Power corrections required to
adjust glideslope are made by repositioning the velocity vector with forward or aft stick inputs. For best
results, make small corrections in velocity vector placement and be smooth. Avoid large, rapid, cyclic
stick motion or ″stick pumping″ as these inputs can produce a PIO with the autothrottles.
Although ATC is capable of handling almost all glideslope corrections, the stick inputs required to
successfully correct large deviations can be difficult to make. In general, if the ball is more than 1 ball
from the center, consider disengaging ATC and executing a manual pass.
7.7.7 FPAH/ROLL - ATC Approaches. The FPAH/ROLL autopilot mode, when utilized with ATC,
provides an alternative method for landing the aircraft. The FPAH/ROLL mode is designed to reduce
pilot workload by maintaining flight path angle (FPA) and roll attitude. When the velocity vector is
positioned as desired and the stick is neutralized, the autopilot maintains the current FPA and roll
attitude, making corrections for wind gusts or disturbances as required. Repositioning the velocity
vector with longitudinal or lateral stick inputs changes the reference FPA and/or roll attitude that the
autopilot holds when the stick is released. In FPAH/ROLL, aircraft response to longitudinal stick
inputs is slightly sluggish compared to CAS while response to lateral stick inputs is essentially the
same.
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ORIGINAL
A1-F18EA-NFM-000
Once the velocity vector is placed in the desired position, the stick is neutralized, and the pilot
essentially monitors autopilot progress. Corrections should be small and applied only when required.
Learning to make appropriate corrections and to stay out-of-the-loop when corrections are not
required takes practice to achieve good results. With practice, smooth, consistent landings can be
achieved even in gusty wind conditions.
NOTE
Use of FPAH/ROLL without ATC may result in more difficult AOA
control and is not recommended.
7.7.7.1 FPAH/ROLL - ATC Approach Technique (field only). If an FPAH/ROLL - ATC approach is
desired, engage ATC when wings level on downwind and trim for on-speed AOA. Select FPAH/ROLL
from the A/P sublevel on the UFCD, and ensure both modes are boxed.
Fly the standard landing pattern utilizing the numbers and velocity vector positioning described in
the ATC Approaches paragraph. A push and roll is required to establish the approach turn. Once the
velocity vector is positioned, neutralize the stick and monitor autopilot progress. No back stick should
be required in the turn. Passing through the 90 and approaching the start, push forward stick to lower
the velocity vector and establish the desired rate of descent and then neutralize the stick. If on
glideslope, roll wings level in the groove using only lateral stick inputs. Longitudinal stick inputs should
not be required, as the autopilot compensates automatically to maintain FPA. Similarly, if on
glideslope, make lineup corrections solely with lateral stick.
If the ball is not centered, adjust the velocity vector (i.e., reference FPA) up or down accordingly and
allow the autopilot to fly the aircraft back to glideslope. Approaching a centered ball, adjust the
velocity vector to the desired flightpath and neutralize the stick. The autopilot should then maintain
FPA (ideally a centered ball) and compensate automatically for gusts. Make corrections with small,
discrete longitudinal stick inputs and evaluate the correction before applying another. If the ball is
centered and stable, the system works best if longitudinal inputs are minimized. There may be
noticeable pitch motion, similar to what is seen on a Mode-1 ACLS approach, as the airplane responds
to gusts, but FPA should be stable.
FPAH/ROLL is less capable at handling large deviations than CAS - ATC. In general, if the ball is
more than 1 ball from the center, consider disengaging FPAH/ROLL with the paddle switch and
executing an ATC or manual pass.
7.7.8 Full Stop Landings. Maintain approach rate of descent and power setting by flying a centered
ball to touchdown or by placing the velocity vector at least 500 feet past the runway threshold. After
touchdown, place the throttles to IDLE and track runway centerline using small rudder pedal inputs.
The engines will not select ground idle until the aircraft has decelerated below 80 KCAS. While the
rudders are effective above 100 KCAS, NWS is the most effective means of directionally controlling the
aircraft during landing rollout. Low gain NWS is activated automatically at touchdown with weight on
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the nose landing gear and at least one main landing gear. Differential braking to maintain directional
control is not as effective and should normally be avoided.
Use of NWS HI during landing rollout is not recommended, as it may
lead to directional PIO due to the increased sensitivity of the NWS
system to rudder pedal inputs.
Engaging NWS HI while maintaining a rudder pedal input will greatly
increase nosewheel deflection and may cause loss of directional control.
7.7.9 Braking Technique. Under normal circumstances, the best results are attained by applying
moderate to heavy braking with one smooth application of increasing braking pressure as airspeed
decelerates towards taxi speed. Anti-skid is effective down to approximately 40 KGS. Below 40 KGS,
heavy brake pedal pressure should be relaxed to prevent tire skid. Below 35 KGS, steady but firm
brake pedal pressure should be applied. Steady, light brake applications should be avoided, as they
increase brake heating, do not significantly contribute to deceleration, and ultimately reduce braking
effectiveness. If desired, selecting aft stick (up to full) below 100 KCAS will increase TEU stabilator
deflection and aid in deceleration. Full aft stick increases down force on the main landing gear, as well
as significantly increasing drag due to large stabilator size.
Recommended braking speeds are based on tests conducted at sea level.
Ground speed may be significantly higher than calibrated airspeed at
airfields above sea level. Aircrew should consider available runway length
and field elevation to evaluate wheel brake usage and landing rollout
distance to avoid excessive brake heat build up and subsequent tire
deflation or wheel assembly fire when landing at airfields above sea level.
Maximum braking performance is attained by applying full brake pedal pressure (approximately 125
lb) immediately after touchdown. Anti-skid must be on to attain maximum braking performance and
to reduce the risk of a blown tire. Longitudinal pulsing may be felt as the anti-skid cycles. Approaching
40 KCAS, full brake pedal pressure should be relaxed to prevent tire skid.
7.7.9.1 Aerobraking Technique. Aerobraking is not required under most circumstances. However,
aerobraking is an effective method to slow heavy gross weight aircraft with a reduced risk of hot brakes
and fire, or to slow aircraft on wet runways. Aerobraking is authorized under the following conditions:
a. Crosswind 5 knots or less
b. Pitch attitude 10° or less
c. Greater than 80 KCAS
d. GAIN ORIDE not selected
e. No FCS AIR DAT or FLAP SCHED cautions
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f. Flap position not changed during aerobraking
After main landing gear touchdown, smoothly apply aft stick to capture a positive pitch attitude
with the waterline, not to exceed 10°. Directional control can be maintained with rudder pedal inputs
and wings can be leveled with lateral stick. At approximately 100 KCAS, center rudder pedals and
smoothly relax aft stick to allow the nose of the aircraft to fall. Avoid abrupt forward stick inputs to
derotate. Once the nosewheel is on the ground, proceed with normal braking technique. Stopping
distance using aerobraking should be approximately that experienced during normal braking.
Large, abrupt aft stick inputs, particularly with CG near the aft limit, can
result in significant over−rotation. With pitch attitude over 10°, the
trailing edge of the stabilators can impact the ground if a full forward
stick input is used to check the over-rotation. Above 14° pitch attitude,
the raised hook point or engine exhaust nozzles may contact the ground.
Therefore, pitch attitude shall not exceed 10° during aerobraking and
abrupt forward stick inputs to derotate should be avoided.
NOTE
Landing distance data in Chapter XI and the PCL are calculated on
maximum braking performance technique listed above. The effect of
aerobraking is not accounted for in the braking distance performance
charts.
7.7.10 Heavy Gross Weight Landings. The aircraft’s 50,600 lb GW field landing limitation provides
the capability to land with a significant amount of fuel and/or stores (approximately 16,000 lb of
bringback). Landing at heavy gross weight, however, requires that the pilot pay particular attention to
braking technique and overall brake usage to avoid excessive brake and wheel assembly heating, melted
fuse plugs, and deflated tires. The wheel assembly fuse plugs are designed to melt and deflate the tires
at temperatures below those which would result in catastrophic tire blowouts. Wheel assembly
temperatures do not, however, reach their peak until approximately 20 minutes after landing, e.g., it
takes 20 minutes for the heat (energy) imparted to the brake assembly at landing to transfer into the
wheel assembly. Due to this slow transfer of heat, it is not uncommon for an aircraft to pass a post flight
hot brakes check yet still melt a fuse plug in the line.
In general, the aircraft’s braking system is designed for landing under the following circumstances
without melting a fuse plug: land at 50,600 lb GW, maximum anti-skid braking at 115 KCAS, three taxi
stops from 30 KGS, park for 15 minutes, three more taxi stops from 30 KGS. If overall brake usage
exceeds these criteria, melted fuse plugs and deflated tires may result. Below approximately 46,000 lb
GW, brake usage following a maximum anti-skid landing (at or below 90%of approach speed) should
be unlimited. Therefore, any landing above 46,000 lb GW should be considered a heavy gross weight
landing.
7.7.10.1 Heavy Gross Weight Braking Technique. Above 46,000 lb GW, delay the initial brake
application to 115 KCAS or lower, if possible. Utilize aerobraking if desired and runway length is not
a factor, otherwise normal braking technique or maximum anti-skid braking is acceptable. Release the
brakes when desired taxi speed is reached. When clear of the runway, make a conscious effort to limit
taxi speed and minimize brake applications, particularly if maximum anti-skid braking was utilized. If
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overall brake usage is extensive, consider chocking the wheels and leaving the parking brake off to aid
in brake cooling and to limit the amount of heat transferred to the wheel assembly.
Recommended braking speeds are based on tests conducted at sea level.
Ground speed may be significantly higher than calibrated airspeed at
airfields above sea level. Aircrew should consider field elevation when
determining the calibrated airspeed at which brakes will be applied to
avoid excessive brake heat build up and subsequent tire delflation or
wheel assembly fire.
7.7.11 Crosswind Landings. During flight test, three crosswind landing techniques were evaluated:
full-crab-to-touchdown, half-crab-kickout, and wing-down-top-rudder. In general, the half-crabkickout technique works best and is recommended for all crosswinds up to 30 knots; the full-crab-totouchdown technique is acceptable for moderate crosswinds only; and the wing-down-top-rudder
technique is not recommended.
When calculating crosswind component for takeoff or landing, use the full
value of any reported gusts in your calculations.
7.7.11.1 Half-Crab Kickout Technique. In crosswinds up to 30 knots, best crosswind landing results
are attained by performing a half-crab-kickout technique. This technique reduces lateral and
directional oscillations after touchdown and minimizes landing gear side loads.
Fly a full crab approach (wings level, neutral pedals) to approximately 50 feet AGL. Immediately
prior to touchdown, apply one smooth rudder pedal input to ″kick out″ half of the crab angle. Maintain
wings level. Allow the initial directional oscillations to subside, then utilize the normal braking
technique. Stabilator braking with up to full aft stick does not degrade directional control and may be
used to aid deceleration. Lateral stick into the wind will be required and is recommended to maintain
wings level during landing rollout.
Avoid removing half the crab angle too early or removing more than half of the crab angle. This may
cause the aircraft to drift downwind prior to touchdown and increases directional transients after
landing.
7.7.11.2 Full-Crab-to-Touchdown Technique. The landing gear is capable of absorbing the sideloads
imparted during a full-crab-to-touchdown landing in crosswinds up to 30 knots. However, in
crosswinds above approximately 15 knots, the aircraft response produced by this technique can be
uncomfortable. When the main gear contact the ground, the aircraft swerves downwind to align with
the runway and rolls away from the crosswind and into the runway. This roll excursion can be as much
as 8°. Two to three directional oscillations can be expected before the aircraft settles out and tracks
straight. While this motion is controllable, lateral stick inputs to level the wings must be timely, and
rudder pedal inputs must be judicious to control the directional transients. For this reason, a
full-crab-to-touchdown technique is not recommended in crosswinds over 15 knots.
In crosswinds below 15 knots, the roll into the runway and ensuing directional oscillations are small,
and the aircraft tends to track straight soon after touchdown. Fly a full-crab approach (wings level,
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neutral pedals) all the way to touchdown. Apply lateral stick to keep the wings level, allow the small,
initial directional oscillations to subside, and then utilize the normal braking technique.
7.7.11.3 Wing-Down-Top-Rudder Technique. Even in light to moderate crosswinds, a wing-downtop-rudder approach requires up to full rudder pedal displacement and an excessive bank angle (as
much as 10°) to balance the aircraft with no drift. Landing in this attitude is uncomfortable and should
be avoided. Additionally, any rudder pedal input applied at touchdown produces a large directional
excursion when NWS automatically engages. For these reasons, a wing-down-top-rudder technique is
not recommended.
7.7.12 Wet Runway Landings. Wet runway conditions can induce hydroplaning during landing
rollout. The minimum total hydroplaning speeds of the main landing gear tires (280 psi) and the nose
landing gear tires (150 psi) are 150 KGS and 110 KGS, respectively. Depending on runway conditions,
partial hydroplaning can occur at much lower speeds. If the nose tires are hydroplaning, the aircraft
may respond sluggishly to initial NWS commands. Under such circumstances, increasing rudder pedal
inputs may cause directional excursions when nose tire contact is established. If hydroplaning is
suspected, rudder pedal inputs should be kept as small as practicable.
For wet (standing water) runway conditions, reduce gross weight to the minimum practical. Land
on-speed or slightly slow with the power reduced to idle as soon as possible. Maintaining a constant
attitude and sink rate will help dissipate aircraft energy at touchdown. If directional control is
questionable, do not hesitate to add power, go around, and set up for an arrested landing. If directional
control is comfortable, use maximum anti-skid braking to minimize landing distance.
7.7.13 Asymmetric Stores Landings. The maximum lateral stores asymmetry for field landings is
29,000 ft-lb. For non-crosswind landings, the aircraft handles very much like a symmetrically loaded
aircraft. Trim the aircraft for wings level flight and fly a normal on-speed approach to touchdown.
During periods of moderate to heavy braking, expect the heavy wing to yaw forward. While easily
controlled with small rudder pedal inputs, this motion should be anticipated and countered quickly to
prevent a build up in yaw rate. Best results are attained by judiciously tracking runway centerline with
timely rudder pedal inputs.
For crosswind landings, use the half-crab kickout technique recommended for normal crosswind
landings. At touchdown, expect a slightly larger roll away from the crosswind and into the runway only
if the wind is into the light wing. Lateral stick into the wind will be required and is recommended to
maintain wings level during crosswind landing rollout, particularly when the wind is into the light wing.
Using this technique, asymmetric landings up to 29,000 ft-lb can be safely executed on a normal
3.25° glideslope up to 50,600 lb gross weight and in a 30 knot crosswind.
7.8 POST-FLIGHT CHECKS
7.8.1 After Landing. Do not taxi with the right engine shut down, as normal brakes and NWS are not
available.
7.8.1.1 After Landing Checks.
When clear of active runway 1. Ejection seat SAFE/ARMED handle(s) - SAFE (confirm status in both cockpits)
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2. EJECTION MODE handle - NORM
Make sure the ejection seat SAFE/ARMED handle is locked in the SAFE
position detent and that the word SAFE is completely visible on the
inboard side of the handle. If the handle will not lock in the detent or the
word SAFE is not completely visible, check to ensure that the ejection
control handle is fully stowed and attempt to resafe the seat. If unable to
properly safe the ejection seat, instruct line personnel to remain clear of
the cockpit until the seat is checked by qualified maintenance personnel.
3. Landing gear handle mechanical stop - CHECK FULLY ENGAGED
If the DOWNLOCK ORIDE button is pressed or the mechanical stop is
not fully engaged, the LDG GEAR handle can be raised on the ground,
and the main landing gear will retract.
4. FLAP switch - AUTO
5. T/O TRIM button - PRESS UNTIL TRIM ADVISORY DISPLAYED
6. Mask - OFF (confirm status both cockpits)
7. OBOGS system - SECURE
a. OXY FLOW knob - OFF (both cockpits)
b. OBOGS control switch - OFF
8. Canopy - EITHER FULL UP OR FULL DOWN FOR TAXI
• Taxiing with canopy at an intermediate position can result in canopy
attach point damage and failure.
• Prior to operating the canopy switch, confirm aircrew are clear and all
loose equipment is stowed to reduce the potential for injury and/or
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engine FOD.
NOTE
Once the ejection seat(s) are confirmed SAFE and the EJECTION
MODE handle is in the NORM position, it is safe to unstrap.
Adjusting seat height after the upper Koch fittings are removed may
damage the ejection seat trombone fittings.
7.8.2 Hot Refueling. When hot refueling for a subsequent flight, the RADAR switch may be left in
OPR or STBY. However, if feed tank fuel temperatures are approaching their 79°C limit, consider
turning off the radar to aid in RLCS/fuel cooling.
Hot refueling must be performed with the canopy closed. Expect the REFUEL DR caution to be
displayed when ground crew open door 8R to expose the single point refueling receptacle. If refueling
of external tanks is not desired, the appropriate EXT TANKS switches must be placed to STOP.
Otherwise, hot refueling through the single point receptacle will fill all internal and external tanks.
NOTE
When hot refueling in Lots 21 thru 25, the IFR probe must be
extended to refuel any external fuel tanks loaded on the inboard
stations (4 and 8) when external fuel tanks are loaded on the midboard
stations (3 and 9).
The EFD and/or FUEL display can be referenced to monitor refueling progress. Expect external
tanks to refuel slowly until the internal tanks are full.
If an internal tank refuel valve has failed or is leaking, that tank will overfill and direct fuel into the
aircraft vent system. If the aircraft vent tanks overflow, fuel will spill from the vertical tail vent outlets.
When hot refueling is complete, ensure that the fuel cap is properly installed and door 8R is closed:
the REFUEL DR caution should be out and the plane captain/final checker shall give the confirmation
signal. This signal is a cupped, open hand rotated counterclockwise then clockwise followed by a
thumbs up.
For a subsequent flight, expect final checks prior to taxi for takeoff. If placed to OFF prior to
refueling, the RADAR switch may be reselected to OPR when refueling is complete.
A failed or leaking refuel valve can cause rapid overfilling of the aircraft
vent system, fuel spillage from the vent outlet(s), and possible fire if fuel
spills on hot engine components. If this occurs, discontinue hot refueling
immediately.
7.8.3 Before Engine Shutdown Checks.
1. PARK BRK handle - SET
2. BIT display - RECORD DEGD/FAIL INDICATIONS
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3. Radar maintenance (BOA) codes - RECORD IF PRESENT
4. RADAR knob - OFF
5. FCS display - RECORD BLIN CODES
6. EFD - RECORD MSP CODES
7. INS - PERFORM POST FLIGHT UPDATE (if desired)
8. INS knob - OFF
9. Standby attitude reference indicator - CAGE (both cockpits)
10. HMD switch - OFF (both cockpits)
11. CRYPTO switch - AS REQUIRED
NOTE
Ensure the MIDS terminal is on, by ensuring L16 or TACAN is ON,
prior to any attempt to zeroize IFF Mode 4 Crypto Keys via the
CRYPTO switch.
12. Sensors, avionics, and CVRS - OFF
NOTE
The aircraft incorporates an avionics auto-shutdown feature which
powers down all UFCD controlled avionics when both throttles are
secured (ac power removed). Therefore, UFCD controlled avionics do
not need to be secured prior to shutdown.
13. EXT and INTR LT knobs - OFF (both cockpits)
14. Canopy - CHECK CLEAR/OPEN
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A high voltage (100,000 volt) static electrical charge may build up inflight
and be stored in the windscreen and canopy. If possible, ensure that
ground crew discharge the static electricity prior to egress. Otherwise,
avoid direct contact with the outside of the windscreen and canopy to
prevent electrical shock.
15. QDC - DISCONNECTED AND STOWED
Failure to disconnect QDC prior to pilot egress will damage the lower IRC
connection.
7.8.4 Engine Shutdown Checks.
1. Brake accumulator gauge - CONFIRM 3,000 PSI
2. Paddle switch - PRESS (disengage NWS)
3. Confirm 5 minute engine cool down.
NOTE
Before engine shutdown, both engines should be operated at ground
idle (75%N2 or less) for 5 minutes to allow engine temperatures to
stabilize and to prevent engine seizure and rotor damage.
4. BLEED AIR knob - OFF
NOTE
If an engine is shutdown before placing the BLEED AIR knob to OFF,
the corresponding primary bleed air shutoff valve may not fully close,
resulting in residual engine fumes in the cockpit on subsequent start of
that engine.
5. Throttle - OFF (alternate sides)
6. Verify proper switching valve operation.
After hydraulic pressure decays through 500 psi a. FLAP switch - FULL
b. If aileron, rudder, or LEF surfaces X and the Xs do not clear after one FCS reset attempt,
maintenance action is required.
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c. If one FCS reset attempt was required to reset surface Xs, cycle FLAP switch to AUTO then
back to FULL. If Xs reappear, maintenance action is required.
7. FCS page - Verify no channel is completely Xd out.
NOTE
If an FCS channel is completely Xd out with one engine shutdown,
that channel is not being powered by essential bus backup, and
maintenance action is required.
8. COMM 1 and 2 knobs - OFF (both cockpits)
9. L (R) DDI, HUD, and MPCD knobs - OFF (In the F/A-18F, confirm all COMM and display
knobs OFF in both cockpits).
10. Other throttle - OFF
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When amber FLAPS light illuminates 11. BATT switch - OFF
Due to FCS keep alive circuitry, uncommanded flight control movement
may occur for up to 10 seconds after the BATT switch is placed to OFF
if residual hydraulic pressure is still present.
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CHAPTER 8
Carrier-Based Procedures
8.1 GENERAL
The CV and LSO NATOPS Manuals are the governing publications for carrier-based operations and
procedures. All flight crewmembers shall be familiar with CV NATOPS procedures prior to carrier
operations.
8.2 DAY OPERATIONS
8.2.1 Preflight Checks.
1. Exterior Inspection - Perform IAW NATOPS
Conduct a normal preflight inspection with particular attention given to the landing gear, day ID
light, struts, tires, and arresting hook. Check the underside of the fuselage and stabilators for
possible arresting cable damage. Note the relationship of the APU exhaust port and the arresting
hook to the deck edge and, for example, catwalk fire extinguishers. If APU exhaust is a factor, the
aircraft may need to be respotted prior to start. Do not lower the hook during poststart checks
unless the hook point will drop onto the flight deck. A hook check may have to be delayed until
the aircraft is taxiied forward. Make sure sufficient clearance exists for cycling ALL control
surfaces.
The maximum wind allowed for canopy opening is 60 kt. Opening the
canopy in headwinds of more than 60 kt or in gusty or variable wind
conditions may result in damage to or loss of the canopy.
2. Interior Checks - Perform IAW NATOPS with two exceptions:
a. External lights master switch - OFF (Required for proper operation of the Day ID strobe light
on the nose landing gear)
b. ANTI SKID switch - OFF
Ensure the ANTI SKID switch is OFF for all carrier operations to ensure
that full brake authority is available (including locking a tire).
8.2.2 Hangar Deck Operation. Occasionally the aircraft may be manned on the hangar deck. Follow
the same procedures as those concerning flight deck operations.
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Tiedowns shall not be removed from the aircraft unless the emergency brake accumulator pressure
gauge indicates at least 2,600 psi. Emergency brakes shall be used for stopping the aircraft anytime it
is being moved while the engines are not running. The signal to stop an aircraft that is being towed is
either a hand signal or a whistle blast. The whistle signifies an immediate or emergency stop. Once in
the cockpit, leave the canopy open and helmet off to ensure hearing the whistle. Keep the taxi director
in sight at all times. If unable to see the taxi director, or if in doubt of safe aircraft movement, stop the
aircraft immediately.
If the aircraft is not already on the elevator, it will be towed or pushed (with the pilot in the cockpit)
into position to be raised to the flight deck. Ensure tiedowns are in place; set the parking brake; and
close the canopy. Ensure the parking brake is set anytime the aircraft is stopped on the elevator.
8.2.3 Engine Start. Do not start the engines until directed to do so by the tower/Air Boss, typically
30 minutes prior to the stated launch time. APU starts should be made whenever possible. Crossbleed
starts must be approved by the Air Boss due to the relatively high power setting required, and the
potential for injury from jet blast.
8.2.3.1 Before Taxi Checks.
1. Before Taxi Checks - Perform IAW NATOPS and ensure:
a. FLAP switch - FULL
b. TRIM - SET FOR CATAPULT LAUNCH
Ensure the T/O TRIM button is pressed until the TRIM advisory is displayed (stabilators 4°
TEU). Horizontal stabilator trim should be manually set for catapult launch IAW figure 8-1
Tables A thru G. Launches with less than 15 knot excess endspeed require additional trim to
compensate for the reduced launch speed. If the aircraft is loaded asymmetrically, lateral trim
(differential stabilator with WonW) should also be manually set IAW figure 8-1 Table G. Trim
laterally into the light wing (unloaded wing down). The trim settings are designed to keep roll
off less than 5° for 3 seconds after WoffW. Obviously, not all possible external store
configurations could be evaluated. Therefore, some external store configurations may exhibit
more or less roll off at the Table G trim setting. Launches above 15 knots excess would require
less lateral trim. Higher excess endspeeds, mis-set trim conditions were tested and the aircraft
is easily controlled with lateral stick. The key is to trim in the correct direction, which is
unloaded wing down.
Correct stabilator trim is critical to aircraft fly-away performance (hands-off). The stabilator
trim setting determines the aircraft’s initial pitch rate and sets the reference AOA that the FCS
attempts to hold after launch. Reference AOA is set to 12° when the stabilators are trimmed
to 6° TEU or higher. Between 4° and 6° TEU stabilator, reference AOA is steeply changed
from 4° to 12°. The recommended launch trim settings are designed to provide the aircraft
with a consistent 10° to 12°/sec pitch rate regardless of gross weight, CG, or catapult endspeed.
Trim settings above those recommended in tables D and E or launches with greater than 15
knot excess endspeed will maintain the 12° reference AOA but will be characterized by
increased pitch rates. Normal catapult launches are characterized by an initial rotation as high
as 13° AOA before AOA and pitch rate feedbacks reduce AOA to the reference value. For light
gross weight launches, peak pitch rates will be higher and peak AOA’s will be lower due to the
Vmc based launch speed. At heavier gross weights, a range of 10° thru 14° AOA can be
expected during launch and is the best compromise between minimizing sink-off-bow and
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ensuring controllability in the event of an engine failure. If stabilator trim is less than 6.5°, the
CK TRIM caution will be set when the throttles are advanced above 27° THA (FLAP switch
FULL).
c. External fuel tank quantities - CHECK
Do not catapult with partially full external fuel tank(s) (≤2,700 lbs). Fuel
sloshing may cause structural damage to the tanks, pylons, and/or
airframe.
8.2.4 Catapult Trim. See figure 8-1.
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CATAPULT TRIM CALCULATIONS
1. Enter with:
Example
Gross Weight
_______ (60K)
CG from Form-F
_______ (19%)
Lateral Weight Asymmetry
_______ (9,000 ft-lb)
2. Using Gross Weight and Table A, determine type power setting for launch (MIL or MAX)
Catapult Power Setting Requirements
Weight Board
(1,000 lb)
Power Setting
64 to 66 (66.8*)
MAX only
58 to 63
MAX (MIL optional if density
altitude is ≤ 3,000 ft)
46 to 57
MIL (MAX optional)
32 to 45
MIL only
Table A
* 5 Wet Tanker only
Example
Type Launch 60,000 lb with 3,500 ft DA
______ (MAX)
To reduce engine susceptibility to steam ingestion and compressor
stalls, transition from MIL to MAX during the catapult stroke shall
not be performed except in an emergency.
Figure 8-1. Launch Trim (Sheet 1 of 5)
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3. Using Gross Weight and Lateral Asymmetry, determine expected endspeed. Use Table B if
symmetric or the higher endspeed of Tables B and C if asymmetric.
Catapult Launch Endspeed
(Symmetrical Loading 0-2,500 ft-lb)
GW
(1,000)
Endspeed (MIN +15)
(KCAS)
MIL
MAX
66.8*
-
164*
66
-
161
65
-
64
-
63
165
62
163
61
161
51-60
160
≤50
153
160
153
Table B
* 5 Wet Tanker only
Catapult Launch Endspeed
(Station 2 -10 Asymmetric Loading)
Endspeed (Min +15)
(KCAS)
Weight Board
Designation
(xx,Xxx)
Table B
0
Asym Level 1 (2,501-9,000)
165
1
Asym Level 2 (9,001-17,000)
170
2
Asym Level 3 (17,001-29,000)
174
3
Station 2-10 Asymmetry
Level (ft-lb)
Sym Level 0 (0-2,500)
Table C
Example:
Expected Endspeed: 60 Klb, 9,000 ft-lb asymmetry, MAX Power_______ (165 KCAS)
Figure 8-1. Launch Trim (Sheet 2 of 5)
4. Determine required baseline longitudinal trim using Table D (MIL Power) and Table E (MAX
Power). Enter with launch endspeed from Table B or C and Form-F CG. Determine longitudinal trim
setting, interpolating between CG columns if required. The trim settings contained in Tables D and E
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are set up for 15 knot excess endspeed launches. Launches with greater than 15 knots excess will have
higher pitch rates but will maintain the same capture AOA target.
Longitudinal Trim - MIL Power
Form - F
CG (%MAC)
Endspeed
(KCAS)
18
19
20
21
22
153
20
18
15
12
10
160
16
13
11
8
161
15
12
10
163
14
11
8
164
13
10
8
165
12
9
170
8
≥23
7
7
7
7
7
174
7
Catapult Launch Trim MIL Power - Table D
Note: A 10 knot excess endspeed launch would require 4° additional nose up trim from the nominal
settings.
Longitudinal Trim - MAX Power
Form - F
CG (%MAC)
19
20
21
22
23
22
20
17
14
12
9
160
19
16
13
11
8
161
18
15
13
10
163
17
14
11
8
164
16
13
10
8
165
15
12
10
170
11
9
Endspeed
(KCAS)
18
153
≥24
7
7
7
7
7
174
8
7
Catapult Launch Trim MAX Power - Table E
Note: A 10 knot excess endspeed launch would require 4° additional nose up trim from the nominal
settings.
Example:
Baseline Longitudinal Trim: 165 KCAS, 19% CG_______ (12°)
Figure 8-1. Launch Trim (Sheet 3 of 5)
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5. Longitudinal trim MUST be adjusted for the aft CG shift that occurs during normal fuel burn.
The CG can shift as much as 3% MAC (F/A-18E) or 1% MAC (F/A-18F) when Tank 2 fuel drops to
approximately 2,200 lb and Tank 1 fuel drops to approximately 1,000 lb. This CG shift can affect
longitudinal trim by as much as 7° and must be accounted for to prevent catapult launch with a
significant over-trim. Once Tank 1 has dropped to approximately 1,000 lb, fuel scheduling maintains
the CG at an essentially neutral position. Table F is a rule-of-thumb for decreasing longitudinal trim
based solely on Tank 1 fuel quantity. Decrease baseline longitudinal trim by the ‘‘Trim Delta’’ value
down to but in no case less than 7° TEU stabilator.
Trim Adjustments for Normal Fuel Burn
Trim Delta - (°)
Tank 1 Fuel
Quantity (lb)
F/A-18E
F/A-18F
2,100
-3
-
1,500
-5
-
1,000
-7
-2
Table F
Example
Baseline Longitudinal Trim from Step 4: _______ (12°)
Adjusted Longitudinal trim: Tank 1 fuel 2,000 lb _______(9°)
Failure to make Tank 1 fuel quantity trim adjustment will result in an
over trimmed condition, which may aggravate aircraft controllability,
particularly following a single engine failure.
NOTE
If longitudinal trim must be adjusted after differential stabilator has
been input for a lateral weight asymmetry, push the T/O TRIM
button, adjust longitudinal trim and re-input differential stabilator.
Figure 8-1. Launch Trim (Sheet 4 of 5)
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6. If asymmetric, determine required differential stabilator (lateral trim) from Table G. Input
differential stabilator after longitudinal trim has been set, trimming into the light wing (unloaded
wing down).
CATAPULT LAUNCH LATERAL TRIM
Station 2-10 Lateral Weight
Asymmetry (ft-lb)
Differential Stabilator Unloaded Wing Down (°)
0 - 2,500
0
2,501 - 5,500
1
5,501 - 9,500
2
9,501 - 13,500
3
13,501 - 16,500
4
16,501 - 19,500
5
19,501 - 25,500
6
25,501 - 29,000
7
Table G
Example:
Lateral weight asymmetry: _______________(9,000 ft-lb)
Differential Stabilator (unloaded wing down):_____________(2°)
Therefore, if you set longitudinal trim of 9° nose up, a 2° differential stabilator trim would result in
an 8/10 or 10/8 nose up stabilator trim (depending on asymmetric loading) setting on the DDI FCS
page.
Failure to input differential stabilator trim for catapult launches with
asymmetric stores can aggravate aircraft controllability, particularly
following a single engine failure.
Figure 8-1. Launch Trim (Sheet 5 of 5)
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8.2.5 Taxi. The canopy should be down with oxygen mask on and the ejection seat armed prior to
aircraft breakdown and during taxi. Taxiing aboard ship is similar to confined area taxiing ashore.
However, be aware of jet exhaust from other aircraft and the relative position of own nozzles. Typically,
the wings are folded until the aircraft is positioned behind the jet blast deflector (JBD), so full-time
NWS HI should normally be available. NWS HI is recommended for carrier operations and should
provide excellent turning capability for directional control aboard ship. Taxi speed should be kept
under control at all times, especially on wet decks, in the landing area, and approaching the catapult.
Taxi signals from the flight deck directors (yellow shirts) are mandatory.
Be prepared to use the emergency brakes should normal braking fail. In the event of loss of brakes,
inform the tower and lower the tailhook immediately to indicate brake loss to deck personnel.
8.2.6 Takeoff Checks.
For MAX power catapult launches 1. ABLIM option - BOX
2. ABLIM advisory - VERIFY DISPLAYED
All catapult launches 3. T.O. checklist - COMPLETE (from bottom to top - EJECT SEL thru TRIM)
8.2.6.1 Catapult Hook-Up. The aircraft will be taxiied over the JBD and aligned with the catapult
track. Approach the catapult track slowly, lightly riding the brakes with NWS engaged. Use the
minimum power required to keep the aircraft rolling. Close attention to taxi director signals is required
to properly align the aircraft with the catapult track entry wye. If the taxi director is obscured by steam
from the catapult, stop the aircraft.
Prior to taxi past the shuttle 4. Weight board - ″Roger″ gross weight and asymmetry level (if in accordance with figure 8.1, Tables
B and C). The hundreds place on the weight board designates the asymmetry level (see figure 8-1
Table C) in order to set the proper catapult settings for launch. For example, if the aircraft’s gross
weight is 60,000 lb with 9,000 ft-lb of asymmetry, the 9,000 ft-lb falls within asymmetry level 1,
and the aircrew will ″Roger″ a weight board that reads 60,100.
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5. WINGFOLD switch - SPREAD and report: SPREAD and LOCKED, BEER CANS DOWN,
CAUTION OUT, SWITCH LEVER-LOCKED
Ensure the WINGFOLD switch is lever-locked in the SPREAD position.
If the wings are commanded to unlock or fold during a catapult shot, the
wings will unlock, the ailerons will fair, the wings may fold partially, and
the aircraft will settle.
6. Missile arming - COMPLETE (if required)
When directed 7. L BAR switch - EXTEND (green LBAR light on)
8. NWS button - PRESS and HOLD (if required to position launch bar)
Once the launch bar has been lowered, do not engage NWS unless directed to do so, since catapult
personnel may be in close proximity to the launch bar. Once the launch bar enters the catapult track,
do not use NWS. The catapult crew will install the holdback bar as the aircraft taxis forward. Taxi
forward slowly, following the signals of the taxi director or Catapult Officer. When the launch bar
drops over the shuttle spreader, the aircraft will be stopped by the holdback bar engaging the
catapult buffer.
8.2.7 Catapult Launch.
When ″Take Tension″ and ″Launch Bar Up″ signals received 9. Throttles - MIL
10. L BAR switch - RETRACT (green LBAR light out)
Due to the close proximity of the FLAP and LAUNCH BAR switches,
ensure that the FLAP switch is not inadvertently placed to AUTO.
Launching with the flaps in AUTO will result in an excessive settle.
Failure to place the LAUNCH BAR switch to RETRACT prior to
catapult launch may result in hydraulic seal failure and possible loss of
HYD 2A.
11. Controls - CYCLE and report FREE and CLEAR (Takeoff Checks complete)
Wait 5 seconds and ensure all warning and caution lights are out.
12. Engine instruments - CHECK
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When ″Select AB″ signal received (MAX power launches only) 13. Throttles - MAX
When ready for launch 14. Salute with right hand. Hold throttles firmly against the detent and place head against the
headrest.
Throttle friction may be used to help prevent inadvertent retraction of the throttles during the
catapult stroke. If required, it can be overridden if afterburner is needed due to aircraft/catapult
malfunction. Immediately after the end of the catapult stroke, the aircraft will rotate to capture the
12° reference AOA (hands-off). To avoid PIO with the FCS, do not restrain the stick during catapult
launch or make stick inputs immediately after catapult launch. The pilot should attempt to remain
out of the loop but should closely monitor the catapult sequence.
To reduce engine susceptibility to hot gas reingestion and compressor
stalls, transition from MIL to MAX during the catapult stroke shall not
be performed except in an emergency.
Once safely airborne 15. LDG GEAR handle - UP
16. Clearing turn - PERFORM (if required)
With positive rate of climb and clearing turn complete 17. FLAP switch - AUTO
NOTE
During catapult launches performed at heavy gross weight, the TEFs
may begin to retract prior to FLAP switch actuation (at approximately
190 KCAS) in order to follow the loads alleviation schedule.
8.2.7.1 Catapult Suspend. To stop the launch while in tension on the catapult, signal by shaking the
head negatively and transmitting “SUSPEND, SUSPEND” on land/launch frequency. Do not use a
thumbs down signal or any hand signal that might be mistaken for a salute. The Catapult Officer will
reply with a “SUSPEND” signal followed by an “UNTENSION AIRPLANE ON CATAPULT” signal.
The shuttle spreader will be moved aft and the launch bar will automatically raise clear of the shuttle
spreader. Maintain power at MIL or MAX until the Catapult Officer steps in front of the aircraft and
gives the “throttle-back”. The same signals will be used when a catapult malfunction exists.
8.2.7.2 Catapult Endspeed Requirements. Catapult endspeeds are established to provide safe
flyaway during normal launch conditions and to allow the pilot to maintain aircraft control in the event
of a single engine failure. The catapult endspeeds are not based on single engine rate of climb (SEROC)
capability, nor do they guarantee single engine flyaway performance. The minimum endspeed
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requirement is calculated to provide sufficient airspeed and altitude to maintain aircraft control while
executing emergency catapult flyaway procedures.
F/A-18E/F minimum catapult launch endspeeds are governed by three limiting factors: Flaps FULL
minimum single engine control speed (Vmc), maximum longitudinal acceleration capability, and
sink-off-bow. Vmc is the airspeed below which the aircraft is not controllable with a single engine
failure. The Vmc airspeed governs the endspeed for most of the gross weight range in both MIL and
MAX power (up to 60K MIL and 65K MAX, see figure 8-1, Table B). Vmc is also a function of lateral
weight asymmetry; therefore, endspeed must be increased for asymmetric loadings (see figure 8-1,
Table C). The catapult endspeed above 60K in MIL is governed by aircraft longitudinal acceleration
capability which limits maximum gross weight for MIL power launches (see figure 8-1, Table A).
Endspeeds above 65K in MAX are governed by the aircraft CG 10 foot sink-off-bow limit. Actual
catapult endspeeds in the Aircraft Launching Bulletins are computed to launch at the minimum
endspeed plus 15 knots (Vmin +15) (figure 8-1, Table B and C). FULL flap launches are required to
meet wind-over-deck requirements at heavy gross weights. HALF flap launches have not been tested,
and would increase launch wind-over-deck by approximately 10 knots.
8.2.7.3 Catapult Launch Flyaway Characteristics. Launches at light gross weights are characterized
by higher pitch rate and attitude, higher rate of climb, and lower peak AOA when compared to heavy
gross weight launches. Forward stick may be required following the rotation to control pitch attitude
as the aircraft accelerates.
There is a noticeable difference in aircraft flyaway characteristics from light to heavy weights due to
the transition from the Vmc based launch speeds to either the longitudinal acceleration or sink-off-bow
based airspeeds. Heavy weight launches will be characterized by reduced pitch rates and attitudes, and
higher peak AOA when compared to the light weight launches. Light buffet may be felt as the aircraft
rotates through 11° AOA during launch at heavier gross weights. The longitudinal trim settings will
provide the required 10-12°/sec pitch rate and capture a target AOA of 12°; however, peak AOA may
reach 15° momentarily. Maintaining hands off the stick during rotation is crucial to optimizing launch
performance and reduces the tendency for pilot induced oscillations during rotation and initial
flyaway. With normal endspeed and steady deck conditions, the aircraft CG settles up to 3 feet. The
pilot perceives the catapult launch to be level, as rotation keeps the pilot’s eye approximately level even
though the aircraft CG sinks. With less than 15 knots of excess endspeed, more settle will occur up to
a maximum of 10 feet of settle with zero excess endspeed. Launches anticipated with less than the
normal 15 knot excess endspeed require additional longitudinal trim to compensate for the reduced
launch speed. A 10 knot excess endspeed launch would require 4° additional nose up trim from the
nominal settings in figure 8-1, Tables D and E.
8.2.8 Landing Pattern. Refer to Chapter 4, for carrier operating limitations. While maneuvering to
enter the traffic pattern, attempt to determine the sea state. This information will be of value in
predicting problems that may be encountered during the approach and landing.
Enter the carrier landing pattern at 800 feet AGL (figure 8-2) with the hook down. Make a level
break from a course parallel to the Base Recovery Course (BRC), close aboard to the starboard side of
the ship. Below 250 KCAS lower the gear and flaps. The speedbrake function automatically retracts
when the FLAP switch is moved to HALF or FULL. Descend to 600 feet AGL when established
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Figure 8-2. Carrier Landing Pattern
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downwind and prior to the 180° position. Complete the landing checklist and crosscheck AOA and
airspeed (136 KCAS at 44,000 lb GW minus 1.5 KCAS for each 1,000 lb decrease in GW).
NOTE
Flaps HALF or FULL may be used for landing provided the minimum
wind-over-deck (WOD) requirements of the Aircraft Recovery Bulletin
(ARB) are met. As WOD increases above 30 kt, handling qualities in
flaps HALF are slightly improved over flaps FULL and are
recommended to avoid ″settle at the ramp″ situations.
To assist in achieving the desired abeam distance of 1.1 to 1.3 nm: select the 10 nm scale on the HSI
display, select ship’s TCN, and adjust the course line to the BRC. On downwind fly to place the wingtip
of the HSI airplane symbol on the course line. Ensure the ground track pointer is on the exact
reciprocal of the BRC. Select ILS if desired and available.
With 25-30 kt winds over deck begin the 180° turn to the final approach when approximately abeam
the LSO platform or when the ″white″ of the round down becomes visible. Use an instrument scan from
the 180 to the 90. Fly the pattern as described in the VFR Pattern and Approach section of Chapter
7. Adjust the 90 altitude up slightly to account for the height of the ship’s deck, usually 500 feet AGL
versus 450 feet AGL. Target 360 feet crossing the wake. The rate of descent required to maintain
glideslope may be slightly less than on FCLP approaches due to wind over deck. Expect slightly higher
throttle settings. When the meatball is acquired, transmit “SIDE NUMBER, RHINO, BALL, (fuel
state in thousands of pounds to the nearest 100 pound), AUTO” (if using ATC for approach) e.g. ″206,
RHINO, BALL, 7.5, Auto″. If unable to see any or all of the following: the meatball, datums, or
centerline, transmit (SIDE NUMBER, CLARA/CLARA datums/CLARA lineup.″ (e.g. ″206,
CLARA″). See figure 8-3 for a typical Carrier Controlled Approach.
8.2.8.1 ATC Approach Mode Technique. Refer to the ATC Approaches section of Chapter 7 for
basics on ATC operations. ATC stick-to-throttle gains are designed to allow correction of settles or
updrafts with small, rapid stick movements. Close-in corrections are very critical. If a large attitude
correction for a high-in-close situation develops, the recommended procedure is to stop ball motion,
making no attempt to recenter the ball. A low-in-close condition is difficult to correct with ATC and
may result in an over-the-top bolter. It may be necessary to downgrade from ATC and fly manually to
safely recover from a low-in-close condition. The force required to manually disengage ATC is
significant and may prevent salvaging the pass. Large deviations from glideslope may be difficult to
correct with ATC. Typically, ATC should be disengaged if more than one ball from center (or upon
LSO direction) and the approach continued manually.
8.2.8.2 Glideslope. The technique for maintaining glideslope is basically the same as FCLP except
that more power may be required. Maintaining centerline will most likely require more line-up
corrections due to the angled deck. With rough seas and a pitching deck, some erratic ball movement
may be encountered. If this is the case, listen to LSO calls and attempt to average out ball movement
to maintain a safe, controlled approach.
8.2.8.3 Waveoff. When the waveoff signal is received, select MIL (MAX if required) and maintain
on-speed AOA with the E-bracket until rate of descent is arrested and 10° pitch attitude is captured
for climb to pattern altitude. Best rate of climb occurs at on-speed AOA regardless of loading or
configuration. This requires slight back stick pressure as the aircraft accelerates. If ATC is engaged,
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A1-F18EA-NFM-000
immediately disengage ATC or apply enough force to override ATC while advancing the throttles to
MIL or MAX. Do not over-rotate.
An in-close or late waveoff, coupled with an over-rotation can lead to an
in-flight engagement, which can severely damage the aircraft and/or
arresting gear.
8.3 ACL MODE 1 AND 1A APPROACHES
A typical Mode 1 and 1A approach is shown in figure 8-4. The Mode 1/1A approach does not require
ATC, but ATC should normally be used. The following procedure is typical for a Mode 1 (1A) approach
from marshal to touchdown (or 0.5 mile).
1. Request a Mode 1 or Mode 1A approach from Marshal.
2. HSI format - SELECT (box) ACL
When the ACL option is boxed, the LINK 4 format automatically appears on the LDDI, and the
ACL mode automatically starts its self test. At this time, the ILS, data link, and radar beacon are
automatically turned on (if not previously on), and IBIT is run on the data link and radar beacon
systems. Also, the uplinked universal test message is monitored for valid receipt.
3. Onboard ACL Capability - CHECK
a. LINK 4 format - CHECK FOR ACL 1
Mode 1/1A capability is not available if ACL 1 is not displayed.
b. BIT page, NAV Sublevel - Verify AUG GO/PBIT GO
An augmentor degrade does not inhibit ACL coupling. A degraded
augmenter may lead to a significant lineup error, most often tending
right-of-centerline.
4. Report departing marshal - ″SIDE NUMBER, COMMENCING″
5. Normal CCA - PERFORM
Descend at 250 KCAS and 4,000 fpm to 5,000 feet, (platform) then reduce rate of descent to 2,000
fpm. When selected, ILS steering is automatically displayed on the HUD once valid signals are
received and must be manually deselected, if the symbology is not desired.
a. At 5,000 ft MSL, report - ″SIDE NUMBER, PLATFORM″
b. Continue descent to 1,200 ft MSL.
c. At 10 nm, report - ″SIDE NUMBER, 10 MILES″
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6. LDG GEAR handle - DN (at 10 nm but NLT 8 nm)
7. FLAP switch - FULL or HALF
NOTE
• Flaps may be switched between FULL and HALF while remaining
coupled outside of one nautical mile from touchdown.
• When coupled, changing flap position inside one nautical mile from
touchdown is prohibited.
8. Landing checklist - COMPLETE
a. Check the LDDI for ID LT indication.
9. Slow to approach speed at 6 nm.
10. ATC - ENGAGE
11. RALT hold mode - ENGAGE (if desired)
ACL acquisition occurs at approximately 3.5 to 8 nm and is indicated by ACL RDY on the LINK 4
format and the data link steering (TADPOLE) on the HUD. It is desired but not required, to have
ACL coupled at least 30 seconds before tipover. T/C is replaced by MODE 1 on the LINK 4 format.
After ACL Acquisition 12. Report needle position - e.g., ″UP AND ON″ or ″UP AND RIGHT″.
For Mode 1, when directed 13. CPL option - SELECT on UFCD
If T/C is engaged, press CPL once to uncouple T/C then press CPL again to couple ACL. When
the aircraft is not coupled, ACL RDY is displayed on the HUD. ACL couple is indicated by CMD
CNT and MODE 1 on the LINK 4 format and CPLD P/R on the UFCD and HUD. At this time,
the uplinked command displays of heading, airspeed, altitude, and rate of descent are removed
from the LINK 4 format and the HUD.
14. When coupled, report - ″COUPLED″
15. When aircraft responds to automatic commands, report - ″COMMAND CONTROL″
For Mode 1A Approach 16. Downgrade to Mode 2 at 0.5 mile by
a. Paddle switch - PRESS
b. ATC button - DISENGAGE (if desired)
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Figure 8-3. Carrier Controlled Approach
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17. Report - ″SIDE NUMBER, RHINO, BALL or CLARA, FUEL STATE, AUTO″ (if ATC
engaged).
For Mode 1 Approach 11. Report - ″SIDE NUMBER, RHINO, BALL or CLARA, FUEL STATE, COUPLED″.
12. At approximately 12.5 seconds before touchdown, the uplinked 10 SEC cue is displayed on the
LINK 4 format and the HUD.
13. After touchdown, ACL and ATC should be automatically disengaged.
NOTE
After Mode 1 or 1A downgrade or touch-and-go, actuate the paddle
switch to ensure complete autopilot disengagement.
8.4 ACL MODE 2 APPROACH
A typical ACL Mode 2 approach is shown in figure 8-5. For a Mode 2 approach, the HUD data link
steering is used to fly a manual approach.
1. HSI format - SELECT (box) ACL
When the ACL option is boxed, the LINK 4 format automatically appears on the LDDI, and the
ACL mode automatically starts its self test. At this time, the ILS, data link, and radar beacon are
automatically turned on (if not previously on), and IBIT is run on the data link and radar beacon
systems. Also, the uplinked universal test message is monitored for valid receipt.
2. LINK 4 format - CHECK FOR ACL 1 or ACL 2
Mode 2 capability is not available if ACL 1 or ACL 2 is not displayed.
3. Report departing marshal - ″SIDE NUMBER, COMMENCING″
4. Normal CCA - PERFORM
Descend at 250 KCAS and 4,000 fpm to 5,000 feet, (platform) then reduce rate of descent to 2,000
fpm. When selected, ILS steering is automatically displayed on the HUD once valid signals are
received and must be manually deselected, if the symbology is not desired.
a. At 5,000 ft MSL, report - ″SIDE NUMBER, PLATFORM″
b. Continue descent to 1,200 ft MSL.
c. At 10 nm, report - ″SIDE NUMBER, 10 MILES″
5. LDG GEAR handle - DN (at 10 nm but NLT 8 nm)
6. FLAP switch - FULL (HALF if required)
7. Landing checklist - COMPLETE
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Figure 8-4. ACL Mode 1 and 1A Approaches
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a. Check the LDDI for ID LT indication.
8. Slow to approach speed at 6 nm.
9. ATC - ENGAGE (if desired)
10. RALT hold mode - ENGAGE (if desired)
ACL acquisition occurs at approximately 3.5 to 8 nm and is indicated by ACL RDY on the LINK 4
format and the data link steering (TADPOLE) on the HUD.
After ACL Acquisition 11. Report needle position - e.g., ″UP AND ON″ or ″UP AND RIGHT″.
12. Report - ″SIDE NUMBER, RHINO, BALL or CLARA, FUEL STATE, AUTO″ (if ATC
engaged).
8.5 ARRESTED LANDING AND EXIT FROM THE LANDING AREA
1. Fly an on-speed, on centerline, centered-ball approach all the way to touchdown.
At touchdown 2. Throttles - MIL
To reduce aircraft and arresting gear loads and required recovery windover-deck, selection of MAX power at touchdown shall not be performed
except in an emergency.
When forward motion ceases 3. Throttles - IDLE and allow the aircraft to roll aft.
When directed 4. Brakes - APPLY
5. HOOK handle - UP
If the wire does not clear the hook, the taxi director will signal to lower the hook for aircraft
pullback.
6. FLAP switch - AUTO
7. WINGFOLD switch - HOLD or FOLD (HOLD if wingtip missile dearming required)
8. NWS button - ENGAGE NWS HI
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When the come ahead signal is received, add power, release brakes, and exit the landing area
cautiously and expeditiously. Taxi the aircraft as directed. Do not use excessive power. If one or both
brakes fail, utilize the emergency brakes; advise the tower; and drop the arresting hook.
Once spotted, keep the engines running until the taxi director signals engine shutdown and the
aircraft is properly chocked and chained.
8.6 SECTION CCA
A section CCA may be necessary when a failure occurs which affects navigation aids, communications equipment, or other aircraft systems. Normally, the aircraft experiencing the difficulty flies the
parade position on the starboard side during the approach. When the meatball is sighted, but no lower
than 300 feet AGL, the section leader breaks away from the wingman in a climbing left turn. The
section leader should climb to 1,200 feet AGL, or below an overcast, in the bolter configuration, and
position himself at the wingman’s 11:00 o’clock position. If the wingman bolters or waves-off, he should
rendezvous in the bolter configuration on the section leader. If a wave-off is required prior to flight
break-up, the flight leader executes a climbing right turn to 1,200 feet AGL and follows the directions
of CATCC. Necessary lighting signals between aircraft are contained in Chapter 26.
NOTE
A section penetration should not be made to the ship with less than
non-precision minimums.
8.7 NIGHT OPERATIONS
8.7.1 General. Night carrier operations have a much slower tempo than day operations and it is the
pilot’s responsibility to maintain this tempo. Standard daytime hand signals from deck crew to pilot
are executed with light wands. The procedures outlined here are different from, or in addition to,
normal day carrier operations.
8.7.2 Preflight. Conduct the exterior preflight using a white-lensed flashlight. Ensure that the
exterior lights are properly set for night launch and the external lights master switch is OFF before
engine start. Ensure that instrument and console light knobs are on. This will reduce the brilliance of
the warning and advisory lights when the generators come online.
8.7.3 Before Taxi. Adjust cockpit lighting as desired and perform Before Taxi Checks.
8.7.4 Taxi. Slow and careful handling by taxi directors and pilots is mandatory. If any doubt exists
as to taxi director signals, stop the aircraft. At night it is very difficult to determine speed and motion
over the deck, so the pilot must rely on the taxi director signals, following them closely.
8.7.5 Catapult Hook-Up. Maneuvering the aircraft for catapult hook-up at night is identical to that
used in day operations; however, it is difficult to determine speed or degree of motion over the deck.
If the taxi director is obscured by steam from the catapult, stop the aircraft.
8.7.6 Catapult Launch. At night, catapult procedures are the same as daytime, except signals are
provided by lights instead of hand signals. The exterior lights are utilized to signal that the pilot is
ready for launch. After the control wipeout, select the ADI for display on a DDI or the UFCD in case
the HUD should be lost during or immediately after launch. When ready for launch, place external
lights master switch to ON.
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Figure 8-5. ACL Mode 2 Approach
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A1-F18EA-NFM-000
All exterior lights (position, formation, and strobes) should be on. If instrument meteorological
conditions are expected shortly after launch, the strobes may be left off at the discretion of the pilot.
After launch, monitor rotation of the aircraft to 12° AOA, crosschecking all instruments to ensure
a positive rate of climb. When comfortably climbing, retract the landing gear and flaps and proceed on
the departure IAW CV NATOPS.
8.7.7 Catapult Suspend. To stop the launch while in tension on the catapult, do not turn on the
exterior lights and transmit “SUSPEND, SUSPEND”. Maintain MIL/MAX power until the catapult
officer walks in front of the wing and gives the throttle-back signal. If the external lights master switch
has been placed on prior to ascertaining that the aircraft is down, transmit “SUSPEND, SUSPEND”
and turn off the exterior lights and leave the throttles at MIL until signaled to reduce power.
8.7.8 Night Landings. Night and instrument recoveries will normally be made using case III
procedures IAW CV NATOPS. Prior to departing marshal, change the IDENT switch on the exterior
lights panel to the NORM position. Make sure the strobe lights are flashing a 3 flash, pause, repeat
pattern.
8.7.9 AN/APG-79 AESA Considerations. Aircrew operating AN/APG-79 AESA equipped aircraft
shall be aware that directing AN/APG-79 AESA transmissions toward the ship while on approach may
cause strong electromagnetic interference (EMI) affecting the safety of aircraft on approach. Unless
specifically authorized for operational necessity or required for safety of flight, aircrew shall cease
AN/APG-79 AESA radar transmissions whenever operating within 10 nm of the ship while on
approach during IMC/Case III conditions. Reference Chapter 4.4 for AN/APG-79 AESA radar
limitations.
8.7.10 Arrestment and Exit From the Landing Area. During the approach, all exterior lights should
be on with the exception of the landing/taxi light. Following arrestment, immediately turn the external
lights master switch off. Taxi clear of the landing area following taxi director signals. If brakes are lost,
signal by lowering the hook, turning on exterior lights, and transmitting on land/launch frequency.
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CHAPTER 9
Special Procedures
9.1 FORMATION FLIGHT
9.1.1 Formation Taxi/Takeoff. During section taxi, ensure adequate clearance between flight lead’s
stabilator and wingman’s wing/missile rail is maintained. The leader will take position on the
downwind side of the runway with other aircraft in tactical order, maintaining normal parade bearing.
See figure 9-1. For three aircraft formations, line up with the lead on the downwind side, number 2 on
the centerline, and number 3 on the upwind side. Wingtip/launch rail overlap should not be required
but is permitted if necessary. For four plane formations, line up with the lead’s section on the
downwind half of the runway and other section on the upwind half. When Before Takeoff checks are
completed and the flight is in position, each pilot looks over the next aircraft to ensure the speed brake
is retracted (spoilers down), the flaps are set for takeoff, all panels are closed, no fluids are leaking,
safety pins are removed, rudders are toed-in, nosewheel is straight, and the launch bar is up. Beginning
with the last aircraft in the flight, a “thumb up” is passed toward the lead to indicate “ready for
takeoff”.
9.1.1.1 Section Takeoff. For section takeoff, all aspects of the takeoff must be prebriefed by the
flight leader, to include flap settings; use of nosewheel steering; power changes; power settings; and
signals for actuation of landing gear, flaps, and afterburner. Engines are run up to approximately 80%,
instruments checked, and nosewheel steering low gain ensured. On signal from the leader, brakes are
released and throttles are advanced to military power minus 2% rpm. If afterburner is desired, the
leader may go into mid range burner immediately without stopping at military power. Normal takeoff
techniques should be used by the leader, with the wingman striving to match the lead aircraft attitude
as well as maintain a position in parade bearing with wingtip separation. The gear and flaps are
retracted on signal. Turns into the wingman shall not be made at altitudes less than 500 feet above
ground level.
9.1.2 Aborted Takeoff. In the event of an aborted takeoff, the aircraft aborting must immediately
notify the other aircraft. The aircraft not aborting should add max power and accelerate ahead and out
of the way of the aborting aircraft. This allows the aborting aircraft to steer to the center of the runway
and engage the arresting gear, if required.
9.1.3 Parade. The parade position is established by superimposing the front of the wingtip missile
rail over the pilot’s headbox. Superimposing the two establishes a bearing line and step down. Proper
wingtip clearance is set by reference to the exhaust nozzles. When the left and right nozzles are aligned
so that there is no detectable curve to the nozzles, the reference line is correct. The intersection of the
reference line with the bearing line is the proper parade position. See figure 9-2.
Parade turns are either standard (VFR) or instrument turns. During day VFR conditions, turns
away from the wingman are standard turns. To execute, when lead turns away, the wingmen roll the
aircraft about its own axis and increase power slightly to maintain rate of turn with the leader. Lateral
separation is maintained by increasing g. Proper step down is maintained by keeping the lead’s fuselage
on the horizon.
Turns into the wingmen and all IFR or night turns in a parade formation are instrument turns.
During instrument turns maintain a parade position relative to the lead throughout the turn.
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Figure 9-1. Formation Takeoff Runway Alignments
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Figure 9-2. Formations (Sheet 1 of 2)
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Figure 9-2. Formations (Sheet 2 of 2)
After initially joining up in echelon, three and four plane formations normally use balanced parade
formation. In balanced parade number 3 steps out until the exhaust nozzles on number 2 are flush. This
leaves enough space between number 3 and lead for number 2 to cross under into echelon.
When it is necessary to enter IFR conditions with a three or four plane formation, the lead directs
the flight to assume fingertip formation. In this formation number 3 moves up into close parade on the
lead. All turns are instrument turns.
9.1.4 Balanced Cruise Formation. The balanced cruise position is a looser formation which allows
the wingmen more time for visual lookout. Balanced cruise provides the wingmen with a cone of
maneuver behind the leader which allows the wingman to make turns by pulling inside the leader, and
requires little throttle change.
The balanced cruise position is defined by the wingman aligning his headbox with the front of lead’s
wingtip missile rail and headbox, and lead’s arresting hook fairing with the opposite wing formation
light. The wingmen are free to maneuver within the cone established by that bearing line on either
wing. In a division formation, number 3 should fly the bearing line but always leave adequate room for
number 2 and lead. Number 4 flies balanced cruise about number 3.
9.1.5 Section Approaches/Landing. The aircraft is comfortable to fly in formation, even at the low
airspeeds associated with an approach and landing. The rapid power response enhances position
keeping ability. The formation strip lighting provides a ready visual reference at night and the dual
radios generally ensure that intra-flight comm is available.
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During section approaches all turns are “instrument” turns about the leader. When a penetration is
commenced the leader retards power to 75% rpm and descends at 250 KIAS. If a greater descent rate
is required the speed brake may be used. Approximately 5 miles from the final approach fix or GCA
pickup the lead gives the signal for landing gear.
9.1.5.1 Section Landing. If a section landing is to be made, lead continues to maintain ON-SPEED
for the heaviest aircraft and flies a centered ball pass to touchdown on the center of one side of the
runway. Wingman flies the normal parade position, taking care not to be stepped up.
When “in-close”, wingman adds the runway to his scan and takes a small cut away from the lead to
land on the center of the opposite side of the runway while maintaining parade bearing. Use care to
ensure that drift away from the lead does not become excessive for the runway width. Remember,
flying a pure parade position allows 4 feet of wingtip clearance.
The wingman touches down first and decelerates on that half of the runway as an individual. Do not
attempt to brake in section. If lead must cross the wingman’s nose to clear the duty, the wingman calls
“clear” on comm 2 when at taxi speed and with at least 800 feet between aircraft. The lead stops after
clearing the runway and waits for the wingman to join for section taxi.
9.2 AIR REFUELING (RECEIVER)
Air refueling shall be conducted in accordance with NATO publication ATP-56, Air-to-Air Refueling
Procedures.
NOTE
The KC-10, KC-130, KC-135 tankers, F/A-18E/F and S-3 aircraft with
a 31-301 (A/A42R-1) buddy store are authorized tankers for air
refueling. Maximum refueling pressure is 55 psi.
9.2.1 Air Refueling Checklist. The air refueling checklist should be complete prior to plug-in.
1. RADAR - STBY/SILENT/EMCOM (″nose cold″)
2. MASTER ARM switch - SAFE (″switches safe″)
3. ALE-50 transmit power - OFF (if decoy deployed)
4. INTR WING switch - NORM (or as required)
5. EXT TANK switch(es) - AS DESIRED
If feed tank fuel level is critical, selecting STOP ensures the fastest transfer of fuel to the feed
tanks.
NOTE
For ARS configured aircraft: If fueling of the ARS is not desired
during aerial refueling, as the receiver, CTR ORIDE must be selected
since CTR STOP will not prevent fuel from entering the ARS.
Selecting CTR ORIDE will pressurize all external fuel tanks and
significantly reduce refueling rate.
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6. PROBE switch - EXTEND
7. Visor - DOWN (recommended)
For night air refueling 8. Exterior lights - SET FOR REFUELING
9. Tanker lights - REQUEST AS DESIRED
9.2.2 Refueling Technique. The following procedures, as applied to tanker operations, refer to single
drogue refueling from the F/A-18E/F and the aerial refueling store. All other tanking evolutions are
dependent on the type of tanker being utilized. Refer to Chapter 26, Visual Communications, for
proper hand signals during air refueling operations.
A sharp lookout doctrine must be maintained due to the precise flying imposed on both the tanker
and receiver pilots. Other aircraft in the formation may assist the tanker in maintaining a sharp
lookout. Refueling altitudes and airspeeds are dictated by receiver and/or tanker characteristics
balanced with operational needs. This typically covers a practical envelope from the surface to 40,000
feet and 180 to 300 KCAS (while engaged), depending on the buddy store part number. (See figure
4-13).
9.2.2.1 Approach. When cleared to commence an approach and the refueling checklist is complete,
assume a ready position 10 to 15 feet in trail of the drogue with the refueling probe in line both
horizontally and vertically. Once in a stabilized position, trim the aircraft and make sure the tanker
ready light (amber) is on. Referencing the probe and drogue for alignment, increase power to establish
a 3 to 5 knot closure rate.
• If the tanker ready light is not on, do not engage drogue until signaled
by tanker aircraft as hose-reel response may be inoperative, causing
damage to tanker and receiver aircraft.
• Avoid damage to the right AOA probe by contact with the basket as a
4 channel AOA failure may result.
• An excessive closure rate may cause a violent hose whip following
contact and/or increase the danger of structural damage to the aircraft
in the event of misalignment.
NOTE
An insufficient closure rate results in the pilot fencing with the drogue
as it oscillates in close proximity to the aircraft nose.
Make small corrections during the approach phase using the rudder pedals for lateral misalignment
and longitudinal stick for vertical misalignment. Avoid lateral stick inputs as they cause both vertical
and lateral probe displacement. During the final phase of the approach, the drogue has a tendency to
move up and to the right as it passes the nose of the receiver aircraft due to the aircraft-to-drogue air
stream interaction.
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9.2.2.2 Missed Approach. A missed approach is executed by reducing power and backing to the rear
with a 3 to 5 knot opening rate. Execute a missed approach if:
1. The receiver probe and the drogue basket cannot be properly aligned during the final phase of the
approach.
2. The receiver probe passes forward of the drogue basket.
3. The receiver probe impinges on the rim of the drogue basket.
4. Any unsafe condition develops.
Analyze alignment problems prior to commencing another approach.
9.2.2.3 Contact. When the receiver probe engages the basket, it seats itself into the reception
coupling and a slight ripple is evident in the refueling hose. The drogue and hose must be pushed
forward 5 feet by the receiver aircraft before fuel transfer can be started. This position is evident by
the tanker ready light (amber) going out and the (green) fuel transfer light (green) coming on. During
refueling, maintain a position directly behind and slightly below tanker aircraft.
NOTE
If streaming fuel is observed around the refueling probe, the probe is
not properly seated in the drogue. Disengage, stabilize in the ready
position, and then reengage the drogue.
9.2.2.4 Disengagement. The receiver aircraft disengages by reducing power in order to open from
the tanker at 3 to 5 knots. Back straight away and down, following the flight path of the tanker. The
receiver probe separates from the reception coupling when the hose reaches full extension. When clear
of the drogue, place the PROBE switch in the RETRACT position. Make sure that the PROBE UNLK
caution display is out before resuming normal flight operations.
• Disengagement must be made straight back, parallel to the tanker
flight path, and descending along the natural trail angle of the hose to
prevent damage to the tanker and/or refueling aircraft.
• When installed on the F/A-18 E/F tanker, the ARS hose/drogue/
coupling exhibits a strong tendency to re-center in its natural trail
position. Off-center disconnects may result in drogue contact and
damage to the aircraft. Avoid off-center disconnects and maintain a
constant separation rate until clear.
9.2.2.5 KC-10 Refueling Operations. The KC-10 tanker is equipped with a centerline hose reel
system and/or two Wing Aerial Refueling Pods (WARP). Maximum in-flight refueling airspeed and
altitude for the F/A-18E/F when refueling from the KC-10 is 275 KCAS and 25,000 feet with an
optimum airspeed of 220 KCAS. At airspeeds above 250 KCAS, tanker induced light turbulence causes
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random drogue movement of 2 to 3 feet while 1 foot of movement will be encountered at airspeeds less
than 250 KCAS. The recommended closure rate is 2 to 3 knots.
When joining a flight of receiver aircraft, do not close astern of the KC-10
within 1 to 3 miles from co-altitude to 500 feet below. Loss of aircraft
control can occur if wake turbulence is encountered.
Excessive closure rates may exceed the capabilities of the take-up reel. If
this should happen, a sine wave develops in the hose. Immediate
disengagement is required to prevent damage to the aircraft.
9.2.2.6 KC-135 Refueling Operations. The KC-135 may be configured with a Multi-Point Refueling
System (MPRS) and/or a Boom to Drogue Adapter (BDA) kit.
9.2.2.6.1 KC-135 BDA Refueling. The KC-135 hose has a fixed length of 9 feet attached by a
swiveling coupling to the end of a telescoping boom. The hose terminates in a hard, non-collapsible
drogue and has no reel retraction capability. The following refueling parameters are recommended:
• Airspeed of 200 to 275 KCAS or Mach 0.8 (whichever is less).
• Closure rate of 2 knots or less.
Aerial refueling from the KC-135 is fundamentally different from the standard Navy hose-drogue
systems. After assuming a standard ready position, add power to create a closure rate of 2 knots or less.
Due to the short length of hose and the weight of the drogue, the aircraft-to-drogue air stream
interaction is minimized.
Excessive closure rates (greater than 2 knots) may result in damage to the
aircraft or the refueling drogue.
Once contact has been made, the drogue must be pushed in approximately 4 feet and held in that
position within ±2 feet fore and aft for fuel to flow (the hose forms a U-shape when in the correct
position). If the F/A-18E/F is positioned too far aft with the hose near the trail position, slight aft or
radial movement results in disconnect. The potentially more hazardous situation occurs when the
drogue is pushed too far forward, such that the hose could be looped around the drogue on the probe.
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When disengaging, align the drogue with the boom and back straight away with reference to the
boom.
Off-center disconnects can result in damage to the refueling probe or
nozzle because of the excessive sideloads generated by the KC-135
boom-drogue adapter.
9.2.2.6.2 KC-135 MPRS Refueling. The KC-135 MPRS incorporates the use of wing tip mounted
aerial refueling pods to support receivers designed for hose/drogue refueling operations. The refueling
hose is slightly shorter than the KC-130 and located near the wing tips. The extreme outboard wing
location subjects the hose and drogue to wing tip flowfield disturbances at higher refueling speeds.
• Maximum recommended refueling speed 285 KCAS (up to 300 KCAS/0.86 IMN allowed)
• Optimum refueling airspeeds 260 - 285 KCAS.
While flying at the approach position (20 ft aft of the drogue), small lateral trim inputs may be
required to counter a tendency to roll toward the tanker. Deviations inboard and outboard may require
additional lateral stick inputs. Deviations of more than 10 feet high can result in a strong sideslip (on
right tanker wing, full left ball). Light buffet is a good indication to reposition down with respect to the
tanker.
Maintaining a slow controlled constant closure rate (less than 5 knots) will result in the best
engagement results. Tanking at 300 knots is demanding due to increased bow wave effects and high
drogue position.
9.3 AIR REFUELING (TANKER)
Air refueling shall be conducted in accordance with NATO publication ATP-56, Air-To-Air
Refueling Procedures.
9.3.1 Air Refueling Store (ARS). The ARS is a missionized component for in-flight refueling. The
ARS control panel (figure 9-3) provides power, fuel transfer operation, fuel status indicators, BIT
status, and normal hose extension/retraction, and hose jettison capability.
Figure 9-3. ARS Control Panel
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9.3.1.1 Power Switch (PWR). The PWR switch provides electrical and hydraulic power to the store.
NOTE
For the -4/-5 ARS stores, once powered on (220 KCAS min), airspeed
may be reduced to as slow as 180 KCAS to transfer fuel below 5,000 ft
MSL. Attempts to power on below 220 KCAS will be unsuccessful and
accelerating to 220 KCAS will not recover the RAT. Store power must
be turned off and airspeed increased to 220 KCAS or greater before
re-applying power.
ON
Electrical power is routed in the store and the ram air turbine (RAT) unfeathers, which
provides hydraulic power.
OFF
Feathers the RAT and removes electrical and hydraulic power.
DUMP
Disabled and safety guarded.
NOTE
If the hose is extended, the PWR switch is bypassed (cannot be turned
OFF).
9.3.1.2 STORE Switch. The STORE switch controls fuel transfer between the store and tanker (own
aircraft).
FROM
Pressurizes the store to transfer fuel from store to own aircraft.
OFF
Depressurizes store.
TO
Replenish ARS with own aircraft fuel.
9.3.1.3 REFUEL Display. The four digit display indicates fuel (in pounds) delivered or scheduled, a
three digit BIT code - when commanded, or a 0E when a serious malfunction occurs. The display
initially powers up with 2,500 lb scheduled.
9.3.1.4 Refuel RST Button. A momentary pushbutton that returns LBS scheduled or delivered to
original settings. Pounds scheduled returns to 2,500 pounds or previously scheduled quantity. Pounds
delivered returns to zero.
9.3.1.5 Refuel Data Display Switch. The refuel data display switch is a three position toggle switch
which is spring-loaded to the delivery (DEL) position. The switch can be toggled to the BIT CODE or
scheduled (SCH) positions.
BIT
CODE
Three digit number indicates a malfunction code.
DEL
Pounds of fuel being transferred (25 lb increments).
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ORIGINAL
A1-F18EA-NFM-000
SCH
Pounds of fuel scheduled to be transferred. When scheduled point is reached, automatic
transfer is terminated. 2,500 lb automatically scheduled at power up.
9.3.1.6 Refuel SLEW Switch. The SLEW switch is a three position toggle switch which is springloaded to the center position. If the refuel data display switch is in the SCH position, moving the switch
up increases pounds of fuel scheduled for transfer, and down decreases the fuel scheduled for transfer.
In SCH, the SLEW switch moves the selected indications as follows:
1. 0 to 100 lb, 75 lb/sec in 25 lb increments
2. 100 to 1,000 lb, 300 lb/sec in 100 lb increments
3. >1,000 lb, 600 lb/sec in 200 lb increments
9.3.1.7 BITE Flag Indicator. The BITE flag indicator indicates failures in the control panel assembly.
9.3.1.8 ALERT Indicators. There are two ALERT lights on the right side of the ARS panel.
LOW
RESV
A red warning light illuminates when the hydraulic reservoir level drops below ½ full
and begins flashing when the reservoir level drops below ¼ full.
DUMP
ARS DUMP is not functional.
9.3.1.9 HOSE CUT/SAFE Switch. The HOSE CUT/SAFE switch is a two position switch springloaded to the SAFE position and held in SAFE by a spring-loaded guard. The switch is wired through
the WonW interlock to prevent accidental activation on the deck.
SAFE
Normal position.
CUT
Refueling hose is cut and crimped. ARS is deactivated hydraulically and electrically. In
flight, the switch must be held for several seconds.
9.3.1.10 Dimming (BRT/DIM) Switch. The BRT/DIM switch controls the intensity of the receiver
pilot advisory lights on the ARS tailcone and control panel displays. BRT is used for day operations,
and DIM is used for night operations.
9.3.1.11 Status Lights. There are four status lights on the ARS control panel.
STOW
Green light comes on when hose is approximately one foot from complete stowage. If
hose/drogue is not fully stowed, the indication is through a MASTER CAUTION and
aural tone, and the ARS DROGUE caution appears on the DDI.
RDY
Green light illuminates when the hose is fully deployed and automatic hose response is
established.
PRESS Green light illuminates when hydraulic pressure drops below 1,700 psi. Light goes out
when pressure exceeds 2,000 psi.
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XFR
Green light illuminates when a minimum of 20 gallons/min of fuel is being transferred
to the receiver.
9.3.1.12 TRANS Switch. The three position transfer switch is lever locked in the AUTO position and
controls fuel transfer to the receiver aircraft.
OVRD
The ARS fuel pump is turned on regardless of hose position, fuel schedule, or fuel
remaining in the store. The OVRD position should only be used during emergency refueling situations.
AUTO
When connected to receiver aircraft, fuel flows when the following conditions are met:
• The hose is within the refueling range (approximately 5 to 20 feet of full trail),
• ARS (fuel is above 175 pounds) is not low on fuel; and,
• The scheduled amount of fuel is not exceeded.
OFF
Turns the pump off regardless of hose position.
9.3.1.13 HOSE Switch. The three position switch controls the hose reel. The HOSE switch lever locks
in the RETR and EXT positions and is spring-loaded to EXT from the RESET position.
RETR
The hose retracts or remains retracted.
EXT
The hose extends or remains extended.
RESET Reset is used to establish a new reference pressure (after release to EXT) when airspeed varies more than 10 KCAS from the last reference airspeed (either at extension
or at the time of the last RESET).
Do not select RESET when a receiver aircraft is plugged in. RESET
causes momentary loss of auto response and may damage receiver
aircraft.
Failure to reset hose reference pressure may result in hose auto retraction
if decelerating, or non-responsiveness if accelerating.
9.3.1.14 ARS DROGUE Caution. An ARS DROGUE caution on the DDI indicates that the ARS
PWR switch is OFF and the refueling drogue has not fully retracted.
9.3.2 ARS (Tanker) Procedures.
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9.3.2.1 Tanker Safety Precautions.
1. Do not extend the drogue when an ARS hydraulic leak is observed.
2. Do not actuate the speed brake function during any part of the refueling operation.
3. Single point failure of either the fuel/no air valve or ARS high level float sensor may result in fuel
discharge from the ARS vent during refueling as a receiver (ground or inflight). Inability to
transfer fuel from ARS to internal tanks, or failure to successfully complete pre-checks on ARS
prior to ground refueling (hot pit or truck), are indications of single point failures. If failures are
indicated, hot pit refueling and/or inflight refueling is prohibited.
NOTE
• When tanks 1 and 4 are empty, no more fuel can be transferred to the
ARS. With normal fuel transfer, this occurs at a normal aircraft fuel
state of 4,200 to 4,900 pounds.
• Anytime four external fuel tanks are loaded on wing stations (3, 4, 8,
and 9), selecting ORIDE on LI/RI external transfer switch will
improve dump performance and external transfer rate by commanding
simultaneous transfer of all external tanks vs. normal transfer
sequence (tanks on Stations 3/9 must be empty prior to tanks on
Stations 4/8 transferring). Performing this function imposes airspeed
limitations defined in Figure 4-12.
4. Avoid use of barometric altitude hold in turbulent conditions or whenever receiver is having
difficulty achieving basket contact.
9.3.2.2 ARS Interior Checks 1. PWR switch - OFF
2. STORE switch - OFF
3. HOSE switch - RETR
4. PWR switch - OFF
5. Fuel TRANS switch - OFF
6. Light switch - BRT (day), DIM (night)
7. REFUEL data display switch - DEL
8. HOSE CUT switch - SAFE/GUARD DOWN
NOTE
If the post guillotine shutdown relay in the tail section of the store has
not been reset following guillotine maintenance or actuation, the
system will not operate.
9. EXT LT IDENT knob - Select appropriate strobe pattern
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9.3.2.3 System Check - Prestart. Have the ground crew rotate the Ram Air Turbine (RAT)
counterclockwise, while facing aft, until the RAT blades are at full feather.
1. External electrical power - APPLY
a. EXT PWR switch - RESET
b. GND PWR switches 1, 2, 3, and 4 - B ON (hold for 3 seconds)
2. RAT unfeather test - PERFORM
a. ARS PWR switch - ON
b. STOW light and PRES light - CHECK ON
c. Make sure ground crew verify proper operation of RAT.
Prior to placing the ARS PWR switch to ON, make sure ground crew are
clear of the ARS as the RAT blades will move to the unfeather position
very rapidly.
3. BIT codes - CHECK
a. REFUEL DATA DISPLAY switch - BIT CODE
b. Refuel data display - CHECK (No codes should be present. If any codes are present, have
ground crew reset display).
4. BITE test - PERFORM
a. BITE TEST button - PRESS
b. ARS control panel lights - CHECK (All should illuminate for 10 seconds, then go out.)
c. Refuel data display - CHECK (Display counts from 00 to 500 in 25 pound increments, and then
returns to 00. If it indicates 0E, a serious malfunction has occurred and the ARS is down.)
d. Make sure ground crew verify proper operation of all tail cone lights.
5. ARS PWR switch - OFF (RAT remains unfeathered and control panel lights extinguish.)
9.3.2.4 System Check - Airborne.
1. HOSE switch - RETR
2. ARS PWR switch - ON (white STOW light illuminates)
3. BITE test - PERFORM
a. BITE TEST button - PRESS
b. ARS control panel lights - CHECK (All should illuminate for 10 seconds, then go out.)
c. Refuel data display - CHECK (Display counts from 00 to 500 in 25 pound increments, and then
returns to 00 if no faults detected. If 0E is displayed, it indicates a serious malfunction has
occurred and the ARS is down).
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NOTE
• BITE check can only be performed with the hose stowed.
• Self-test of electronic and hydraulic systems takes approximately 90
seconds.
• If BITE is actuated with less than 300 gallons, or 200 pounds of fuel in
the store, invalid fault indications may result.
9.3.2.5 Drogue Extension.
1. LTDR switch - CHECK SAFE
• Permanent eye damage to the receiving aircrew can occur if laser is
armed and firing during aerial refueling, even when using ″eye safe″
mode.
• Do not attempt to extend the hose unless system self-test has been
successfully completed.
2. HOSE switch - RETR
3. PWR switch - ON (white STOW light illuminates)
NOTE
For the -4/-5 ARS stores, once powered on (220 KCAS min), airspeed
may be reduced to as slow as 180 KCAS to transfer fuel below 5,000 ft
MSL. Attempts to power on below 220 KCAS will be unsuccessful and
accelerating to 220 KCAS will not recover the RAT. Store power must
be turned off and airspeed increased to 220 KCAS or greater before
re-applying power.
4. Airspeed - Refer to ARS operating limitations, figure 4-13.
5. HOSE switch - EXT (White STOW light extinguishes; amber RDY light illuminates when drogue
reaches full trail. HOSE advisory appears on the DDI).
6. STORE switch - TO
7. Fuel TRANS switch - AUTO or OFF
9.3.2.6 Fuel Transfer. The amount of fuel to be transferred is automatically set at 2,500 pounds. To
change the scheduled amount, place the REFUEL switch to SCH, and the digital display shows the
amount scheduled. To increase or decrease the amount scheduled, hold the SLEW switch up or down,
respectively. When desired amount is shown on the digital display, release SLEW switch and return
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A1-F18EA-NFM-000
REFUEL switch to DEL. When fuel is transferred with the TRANS switch in AUTO, the digital
display reads upward until the fuel transfer is automatically stopped at the scheduled amount.
Pressing the REFUEL RST button resets the digital display to zero and reschedules 2,500 pounds, or
the previously scheduled amount when the switch is in the DEL position.
The TRANS switch should always be in the AUTO position for normal
fuel transfer. The OVRD position is an emergency condition that
overrides normal system operation and provides fuel flow regardless of
hose position.
9.3.2.7 Receiver Hook-up and Refueling. Ater the receiver aircraft engages and moves forward into
the refueling range, the store amber RDY light extinguishes. Fuel transfer commences if the FUEL
TRANS switch is in the AUTO position. The green XFR light illuminates when the transfer rate is 20
gallons per minute or higher.
9.3.2.8 Stopping Fuel Transfer. The receiver aircraft receives fuel until one of the following occurs:
1. The IFR probe disengages.
2. The receiver aircraft moves out past the fueling range outer limit, approximately 5 feet from full
trail (the amber ready light comes on).
3. The receiver aircraft moves in past the refueling range inner limit, approximately 25 feet from full
trail (the amber ready light flashes).
• Refueling cannot be stopped by placing the PWR switch to OFF with
the hose extended.
• Lack of hydraulic pressure causes a loss of hose response resulting in
hose instability and a potential hose whip incident. If either the red
PRESS or LOW RESV light comes on, aerial refueling should be
terminated and the hose retracted.
Transfer of fuel to the aerial refueling store continues until one of the following occurs:
1. ARS TRANS switch is placed to OFF.
2. The BINGO caution comes on.
3. Tanks 1 and 4 are empty (with normal fuel transfer, this occurs at a fuel state of approximately
5,000 pounds of fuel). The FQTY advisory comes on.
9.3.2.9 Emergency Fuel Transfer. If problems are encountered in obtaining fuel transfer from the
store to the receiver aircraft, a system is provided that bypasses some normal switch functions to
provide fuel transfer. After the receiver is engaged in the coupling, turn the TRANS switch to OVRD.
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This switch provides power to open the shutoff valve and energize the priority valve to allow hydraulic
flow to the fuel pump motor, and bypasses the low level switch and the fuel flow range switches.
• The receiver aircraft should not attempt engagement when the switch
is in OVRD since there will be fuel pressure in the coupling that
increases the force required to make an engagement. The force will be
above that which provides adequate hose response and a damaging
hose whip may result.
• The TRANS switch should be moved from OVRD to OFF prior to the
receiver disconnecting. Fuel pressure in the coupling increases the
disconnect forces by 200 pounds above normal. Momentary fuel spray
may also occur.
NOTE
The TRANS switch should be taken out of the OVRD position if the
green XFR light on the ARS control panel extinguishes.
9.3.2.10 Drogue Retraction.
1. TRANS switch - OFF
2. Airspeed - Refer to ARS operating limitations, figure 4-13.
Field arrestment or carrier landing with the drogue extended is not
recommended. Guillotine (HOSE - CUT) drogue in clear or uninhabited
areas.
3. HOSE switch - RETR
4. When the white STOW light illuminates, and the HOSE advisory clears, PWR switch - OFF
9.3.2.11 Transfer Fuel from ARS to Tanker (Own Ship). If it is desired to transfer fuel from the
refueling store, do the following:
1. STORE switch - FROM
9.3.2.12 Before Landing.
1. STORE switch - OFF
2. HOSE switch - RETR
3. BIT codes - CHECK
a. REFUEL DATA DISPLAY switch - BIT CODE
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b. Refuel data display - RECORD CODES
4. PWR switch - OFF
5. TRANS switch - OFF
9.3.3 ARS Jettison.
The ARS may be jettisoned in the same manner as other external stores.
9.3.4 ARS Limitations.
Refer to Chapter 4, Operating Limitations.
9.4 NIGHT VISION DEVICE (NVD) OPERATIONS
9.4.1 Effects on Vision. Flight techniques and visual cues used during unaided night flying also apply
to flying with night vision devices (NVD). The advantage of NVD is improved ground reference
provided through image intensifier systems (NVG/NAVFLIR). Dark adoption is unnecessary for the
effective viewing through night vision goggles (NVG). In fact, viewing through the NVG for a short
period of time shortens the normal dark adaptation period. After using NVG, an average individual
requires 1 to 3 minutes to reach the 30 minute dark adaptation level. Color discrimination is absent
when viewing the NVG image. The image is seen in a monochromatic green hue and is less distinct than
normal vision. Prolonged usage may result in visual illusions upon removal of the NVG. These illusions
include complement or green after-images when viewing contrasting objects. Illusions from NVG are
temporary and normal physiological phenomena and the length of time the effects last vary with the
individual.
• Aircrew are strongly cautioned against maneuvering above 3g with the
AN/AVS-9 in the up-locked (not in use but on helmet) position
because the NVD bracket cannot retain AN/AVS-9 under elevated
loads.
• Ejection wearing Night Vision Goggles is not recommended. Severe
neck injury may result.
9.4.2 Effects of Light. Any non-NVG compatible light source in the cockpit degrades the ability to
see with NVG. Filters are used to prevent stray or scattered light from reaching the NVG intensifiers,
which would cause the automatic gain control to reduce the NVG image intensification. Head down
displays (DDI, MPCD) are filtered to allow non-electrical-optical viewing of the display. Viewing areas
illuminated by artificial light sources with NVG (runway/landing lights, flares, or aircraft position
lights) limit the ability to see objects outside of the area.
NOTE
Bright ground lights may cause loss of ground references during
landing. Avoid looking directly at bright light sources to prevent
degrading NVG vision.
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The NAVFLIR is not affected by light sources and complements NVG use.
9.4.3 Weather Conditions. NAVFLIR and NVG provide a limited capability to see through visibility
restrictions such as fog, rain, haze, and certain types of smoke. As the density of the visibility
restrictions increases, a gradual reduction in light occurs. Use of an offset scanning technique will help
in alerting the pilot to severe weather conditions.
NOTE
Visibility restrictions produce a ‘‘halo’’ effect around artificial lights.
9.4.4 Object/Target Detection. Detection ranges are largely a function of atmospheric and environmental conditions. Moving targets with contrasting backgrounds or targets with a reflected or
generated light or heat sources can be identified at greater ranges when using NVD.
9.4.5 Flight Preparation. Flights with NVD require unique planning considerations that include
weather, moon phase/angle, illumination, ground terrain and shadowing effects. Tactical consideration
and procedures can be found in the Night Attack operational tactics guides (OTG).
9.5 SHORT AIRFIELD FOR TACTICAL SUPPORT (SATS) PROCEDURES
9.5.1 Landing Pattern. Approach the break point either individually or in echelon, parade formation,
at 250 KIAS. A 17 to 20 second break interval provides a 35 to 40 second touchdown interval. The
landing checklist should be completed and the aircraft should be at on-speed AOA/approach speed by
the 180° position.
9.5.2 Approach. Plan for and execute an on-speed approach. Pay particular attention to maintaining
the proper airspeed and correct lineup.
9.5.3 Waveoff. To execute a waveoff, immediately add full power and maintain optimum attitude.
Make all waveoffs straight ahead until clear of the landing area.
9.5.4 Arrested Landing. The aircraft should be on runway centerline at touchdown. Aircraft
alignment should be straight down the runway, with no drift. Upon touchdown, maintain the throttle
at the approach position. When arrestment is assured, retard the throttle to idle. Allow the aircraft to
roll back to permit the hook to disengage from the pendant. When directed by the taxi director, apply
both brakes to stop the rollback and raise the hook. If further rollback is directed, release brakes and
allow the aircraft to be pulled back until a brake signal is given. Apply brakes judiciously to prevent
the aircraft from tipping or rocking back.
Use extreme caution when taxiing on a wet SATS runway.
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9.5.5 Bolter. Bolters are easily accomplished. Simultaneously apply full power and retract the
arresting gear hook. Smoothly rotate the aircraft to a lift-off attitude and fly away.
• Bolters in GAIN ORIDE or with AOA failed require positive aft stick
during rotation, 1/2 aft stick is recommended. Deflections of less than
1/2 aft stick will result in excessive settle during bolters.
• If landing on a runway with a SATS catapult installed, care must be
taken to prevent engagement of the dolly arrester ropes with the
aircraft tailhook. Structural damage to the aircraft and catapult will
result.
9.6 HOT SEAT PROCEDURE
1. PARK BRK handle - SET
2. Paddle switch - PRESS (disengage NWS)
3. Left throttle - OFF
4. Throttle friction - MAX
5. Avionics - AS DESIRED
9.7 ALERT SCRAMBLE LAUNCH PROCEDURES
9.7.1 Setting the Alert. The alert/scramble aircraft shall be preflighted in accordance with NATOPS
normal procedures every 4 hours or as local directives dictate. The pre-alert turn shall consist of full
Plane Captain checks and full systems checks. Minimum requirements are:
1. Radar BIT status - GO
2. AIM-7 - TUNED (if loaded)
3. INS alignment status - OK
4. COMM 1 and 2 - SET TO LAUNCH FREQUENCY
5. Launch trim - SET ( IAW Catapult Trim Calculations, Chapter 8)
Before engine shutdown 6. INS known - OFF (10 seconds before engine shutdown)
NOTE
Do not switch INS to NAV during pre-alert turn so that STD HDG
option will be available for next alignment.
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7. CRYPTO switch - HOLD THEN NORM
8. Sensors and weapon systems - ON
9. COMM 1 and 2 knobs - ON
10. EMCON - AS DESIRED
11. Exterior and interior lights - SET
12. DDIs, MPCD, and HUD - ON
13. OBOGS control switch and OXY FLOW knob - OFF
14. Landing gear pins - REMOVED and STOWED
After engine shutdown 15. External electrical power -CONNECT (if applicable)
16. EXT PWR switch - RESET THEN NORM
17. GND PWR switches 1, 2, 3, and 4 - OFF
18. BATT switch - OFF
19. SINS cable - CONNECT (if required)
9.7.2 Alert Five Launch.
If on external power 1. GND PWR switches 1, 2, 3, and 4 - B ON (hold 3 seconds)
2. INS known - CV/GND
3. INS - STD HDG (if available)
4. BATT switch - ON
5. APU switch - ON (READY light within 30 seconds)
6. R engine - START
7. L engine - START
8. FCS RESET button - PUSH (verify RSET advisory displayed)
9. OBOGS control switch and OXY FLOW knob - ON
10. External electrical power - DISCONNECT (if applicable)
11. SINS cable - DISCONNECT (if applicable)
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12. INS knob - NAV, GYRO or IFA
13. T.O. checklist - COMPLETE
9.8 AIRBORNE HMD ACCURACY CHECKS
The procedures below shall be performed to verify JHMCS accuracy at any time system accuracy is
in question, including verifying the accuracy of the cockpit magnetic map. These procedures require an
airborne target.
If performing these procedures to determine if cockpit re-mapping is needed following maintenance,
only 9.8.2 Airborne HMD Accuracy Check with Radar is required. Cockpit re-mapping is not required
if 9.8.2 Airborne HMD Accuracy Check with Radar is successful.
9.8.3 Airborne HMD Accuracy Check with CATM/AIM-9X can be performed at the aircrew’s
discretion to verify accuracy in the high off-boresight field of regard.
NOTE
If preflight HMD Alignment occurred less than 15 minutes after
system powered on, repeat 9.8.1 HMD Alignment prior to any airborne
checks.
9.8.1 HMD Alignment.
(CVRS record HMD if desired)
1. SUPT/HMD/ALIGN page - SELECT
2. Superimpose the HMD alignment cross on the HUD/BRU alignment cross.
3. Cage/Uncage button - PRESS and HOLD until ALIGNING turns to ALIGN OK or ALIGN FAIL
If ALIGN FAIL 4. Repeat steps 2 and 3.
If ALIGN OK and HMD alignment crosses are not coincident with HUD/BRU alignment cross 4. Perform FINE ALIGN a. With FA DXDY displayed, use TDC to align azimuth and elevation HMD alignment crosses
with the HUD/BRU alignment cross.
b. Cage/Uncage button - PRESS and RELEASE
c. With FA DROLL displayed, use TDC to align the roll axis HMD alignment crosses with the
HUD/BRU alignment cross.
d. Cage/Uncage button - PRESS and RELEASE
If satisfied with alignment 5. ALIGN - UNBOX
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9.8.2 Airborne HMD Accuracy Check with Radar.
1. Select STT while in trail of an airborne target.
2. Compare HMD TD Box to HUD TD Box and target’s actual position (when in HUD FOV) and
compare HMD TD Box and target’s actual position (when NOT in HUD FOV) at various
azimuth/elevation angles (up to 45° laterally left and right and 45° in elevation).
If HMD and HUD TD Boxes are not nearly coincident or portion of target is not located within
HMD and HUD TD Boxes 3. Perform 9.8.1 HMD Alignment procedures.
4. Repeat steps 1 and 2.
If HMD Alignment does not correct 5. Consider re-mapping the cockpit.
9.8.3 Airborne HMD Accuracy Check with CATM/AIM-9X.
1. No L&S track selected.
2. Select AIM-9X (manual mode).
• Verify AIM-9X symbology on HMD and AUTO not displayed below 9X at bottom of display.
• Verify AIM-9X slaved to HMD.
3. Perform the following steps at various azimuth/elevation angles throughout the AIM-9X field of
regard and outside the radar field of regard until aircrew are confident that the HMD and AIM-9X
are pointing properly:
a. Place aiming cross on the target.
b. Cage/Uncage button - PRESS to command AIM-9X to enter track
• Verify AIM-9X enters track on the target.
• Verify in-track seeker circle is within one 9X circle size of touching the target.
c. Cage/Uncage button - PRESS to slave AIM-9X to HMD
d. Set up next azimuth/elevation angle and repeat steps a through c.
If the in-track seeker circle is not within one 9X circle size of touching the target 4. Perform 9.8.1 HMD Alignment procedures.
5. Repeat steps 1 through 3.
If HMD Alignment does not correct 6. Consider re-mapping the cockpit.
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CHAPTER 10
Functional Checkflight Procedures
10.1 GENERAL
The intent of functional checks is to determine whether the airframe, power plant, accessories, and
equipment are functioning per predetermined standards. The unique electronic built-in test (BIT),
fault detection, and fault isolation capabilities of the F/A-18E/F allow functional checks that have
historically been performed inflight to be performed on the ground. In general, engine control and
flight control system faults are reliably detected, annunciated, and, in most cases, functionally
bypassed by the aircraft control systems.
In most cases, functional checks for the F/A-18E/F will be performed on the ground by maintenance
personnel based on the requirements set forth in the maintenance work package for the component
being removed, replaced, and/or installed and not by a pilot on a dedicated FCF. Required
maintenance ground checks take advantage of the aircraft’s BIT and fault detection capability and
ensure the health of the component and the integrity of the installation.
10.1.1 Engine Functional Checks. Based on the engine component replaced, ground functional test
requirements for the engine may include any or all of the following: idle speed test (low power turn);
air, oil, fuel leak test (leak check); anti-ice test; MIL power test; MIN AB test; MAX power test (high
power turn); transient test; and/or shutdown test. For instance, a single engine removal/reinstallation,
a single engine replacement, or a dual engine removal/reinstallation requires a low power turn and a
leak check. A dual engine replacement requires a low power turn, leak check, and high power turn. A
FADEC replacement requires ALL functional checks. Additionally, a crossbleed start is required on all
engine reinstallations and replacements. Given the FADEC’s fault detection capability, successfully
completing these functional checks ensures that the engine is properly installed and is functioning
normally. All engine functionality that would be checked inflight is checked during the required
ground checks. Dedicated FCFs are, therefore, not required following engine related maintenance
actions.
10.1.2 Flight Control System Functional Checks. Functional test requirements for the aircraft FCS
include electronic rigging, an FCS maintenance BIT, and a test group (TG) for the specific actuator or
surface which was reinstalled or replaced. The FCS maintenance BIT requires operator intervention
and is the most comprehensive test of the FCS. The FCS maintenance BIT also performs unique tests
to verify the proper installation of a system component. Successfully completing these functional
checks ensures that all surfaces and actuators are properly installed and the FCS is functioning
normally.
Generally, dedicated FCFs are not required following actuator/surface related maintenance actions.
An exception involves the replacement of a LEF hydraulic drive unit (HDU). It is possible for a weak
LEF HDU to pass ground checks yet fail to drive the LEF to the proper position when the surface is
subjected to air loads. A weak HDU may manifest itself by a LEF split, a FLAP SCHED caution,
and/or a possible roll off. Therefore, following the replacement of a LEF HDU, a series of inflight
functional checks are required to test the new component at flight conditions that safely detect weak
HDUs.
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10.1.3 Landing Gear Functional Checks. Ground functional test requirements for the landing gear
system include the following: aircraft jack, LDG GEAR handle mechanical stop and DOWNLOCK
ORIDE button test, landing gear warning light and warning tone test, normal landing gear extension
and retraction, planing link failure test, and emergency landing gear extension (front and rear cockpits
in the F/A-18F). Successfully completing these functional checks ensures that the normal and
emergency landing gear systems are functioning normally. While a dedicated FCF is not required
following landing gear related maintenance, an airborne functional check of the emergency landing
gear system may nonetheless be desired. An airborne functional check, coded E, has been included to
perform this test, at the discretion of the Maintenance Officer, on a ″pro and go″ (FCF combined with
but before operational flight) basis.
10.2 FCF REQUIREMENTS
Figure 10-1 lists the FCF requirements for the F/A-18E/F. Where appropriate, functional checks are
grouped by system and are coded with a letter, A thru E, to identify the type of FCF profile to be flown.
These letter codes appear next to each required item or groups of items in the FCF checklist.
COMNAVAIRFORINST 4790.2 Series allows an FCF to be flown in combination with operational
flights at the discretion of the Commanding Officer, provided the operational portion is not conducted
until the FCF requirements have been completed and entered on the FCF checklist. Generally, a
profile ‘‘A’’ FCF is flown as a dedicated flight due to the number of required checks. However, due to
the limited number of required checks, profile ‘‘C’’ and ‘‘E’’ FCFs, as well as profile ‘‘D’’ FCFs required
solely by the reconfiguration of the rear cockpit, can be flown and are recommended to be flown as ‘‘pro
and go’s.’’
10.3 FCF QUALIFICATIONS
Aircrew who perform FCFs shall be qualified per OPNAVINST 3710.7 and must be designated in
writing by the Squadron Commanding Officer. For a profile ‘‘A’’ FCF, the complete FCF checklist shall
be utilized. For a profile ‘‘C’’ or ‘‘E’’ FCF, a special, abbreviated FCF checklist has been created which
incorporates only those checks required for a ‘‘C’’ and ‘‘E’’ profile. Prior to flight, FCF aircrew must
familiarize themselves with the FCF checklists and the specific functional checks required for the given
profile.
Historically, FCF checklists have only included FCF checks. To reduce confusion and provide a more
coherent checklist, the FCF checks presented in this chapter have been interleaved into the normal
NATOPS checklist. Specific FCF requirements are, therefore, highlighted in italics in this chapter and
in the FCF checklist which is utilized inflight. Additionally, check-off blocks, provided on the FCF
checklist, appear next to those items required by the FCF and not next to non-FCF, normal procedure,
items.
The FCF checklist shall be properly completed and promptly returned to Maintenance Control at
the completion of the FCF.
10.4 FCF PROCEDURES
FCFs shall be conducted with the minimum crew necessary to ensure proper operation of all
required equipment. FCF aircrew shall be given a thorough preflight briefing, coordinated by
Maintenance Control and given by appropriate QA and work center personnel. The briefing shall
describe maintenance performed, the requirements for that particular flight, and the expected results.
FCFs shall be performed using the applicable FCF checklist. The procedures contained in the FCF
checklist are presented in a recommended order. While the order of these functional checks may be
altered as required, the sequence of steps listed for any procedure is mandatory.
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If an FCF profile cannot be completed on a single flight due to time, fuel, operating area restrictions,
or other limiting factors, it is permissible to complete the remaining checks on a subsequent flight. This
subsequent flight may be flown by a different pilot, provided there is a thorough passdown, either
verbal or written, between the pilots.
Profile
A
Type of Checks/Requirements
Complete FCF profile
• Completion of SDLM, to be conducted by the rework facility.
• Acceptance of a newly assigned aircraft or upon receipt of an aircraft returned
from SDLM.
• Return to flight status of an aircraft that has not flown in 30 or more days.
• At the discretion of the Maintenance Officer (e.g., return to flight status of an
aircraft that has been excessively cannibalized).
NOTE
In an F/A-18F (missionized configuration), an FCF qualified
rear cockpit crewmember is required unless the Maintenance
Officer determines that the maintenance actions performed
do not require one.
B
Engine/FADEC/fuel control
• Not required.
C
LEF Checks
• Removal/reinstallation or replacement of a LEF HDU.
D
Rear cockpit checks of an F/A-18F (trainer configuration only)
• Acceptance of a newly assigned aircraft or upon receipt of an aircraft returned
from SDLM.
• Reconfiguration from missionized to trainer configuration.
NOTE
Aft crewmember is required.
E
Emergency landing gear extension
• At the discretion of the Maintenance Officer (e.g., following extensive maintenance on the landing gear system).
Figure 10-1. Functional Checkflight Requirements
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10.5 FCF CHECKLIST - PROFILE A
10.5.1 Plane Captain Brief.
1. Connect external power.
2. FCS ram air scoop check (manually restow)
3. REFUEL DR check
4. Normal engine starts
5. Alternate engine shutdowns
a. Fuel/air heat exchanger leak check
b. Switching valve checks
c. Crossbleed restarts
6. ECS ram air scoop check
7. Engine runups to check cautions
8. 4 down but only 3 up (launch bar down)
9. Probe light check
10. Tail light check
10.5.2 Preflight Checks.
1. Exterior Inspection - Perform IAW NATOPS
a. No loose or improperly installed panels.
b. External canopy switch - CHECK
• Canopy opens and closes smoothly.
• Returns to center (hold) position when released.
c. Boarding ladder operation - CHECK
• Ladder electrically deploys.
• Ladder extends, locks, unlocks, and stows correctly.
2. Interior Checks - Perform IAW NATOPS
a. No loose or improperly installed components (both cockpits).
b. Brake accumulator pressure gauge reads 2,600 psi minimum.
c. Canopy and windscreen: No distortion, blemishes, or cracks (both cockpits).
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10.5.3 Pre-Start Checks.
1. BATT switch - ON
2. Battery gauge - CHECK
• Nominal: 23 to 24 vdc
• FCF minimum: 18 vdc
3. Caution Lights Panel - CHECK CABIN light on (if CPWS installed)
4. ICS - CHECK (F/A-18F)
Apply external electrical power 5. EXT PWR switch - RESET
6. GND PWR switches 1, 2, 3, and 4 - B ON (hold for 3 seconds)
• Audibly verify avionics cooling fans are on.
7. COMM 1 and 2 knobs - ON/VOLUME AS DESIRED (both cockpits)
8. L(R) DDI, HUD, and MPCD knobs - ON (both cockpits)
a. Display IBIT - Select ALL
• Approximately 3 minutes required before TEST patterns displayed.
• No stuck pushtile indications (small circles).
• Push STOP when complete.
• All displays operative.
• Note DEGD indications if present.
b. All mode (day/night), brightness, and contrast controls for all cockpit displays - CHECK/SET
(both cockpits)
c. Display surfaces - CHECK (both cockpits)
• No burned phosphor spots on HUD or DDIs.
• No lineouts or burned liquid crystals on MPCD, UFCD, or EFD.
d. HUD symbology reject - CHECK
(1) Select REJ2
• Heading scale, command heading, heading caret, nav range (if displayed), bank angle, g,
and airspeed and altitude boxes are removed.
(2) Select NORM
e. HUD displayed radar altitude - CHECK
(1) UFCD/RALT - ON
(2) ALT switch - RDR
• HUD displays radar altitude and ″R″.
(3) ALT switch - BARO
• HUD displays barometric altitude.
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f. (AMCD) HUDBU advisory - NOT DISPLAYED
9. LT TEST switch - TEST (both cockpits)
• All warning and caution lights properly illuminate.
• Landing gear warning tone annunciates (front cockpit switch only).
a. AOA indexer brightness - CHECK AND SET
10. Seat adjustment - CHECK (both cockpits)
• Smooth through full range of travel.
• Do not hold switch against stops (no limit switches).
11. Rudder pedal adjustment - CHECK (both cockpits)
• Adjustment smooth through full range of travel.
a. SET PEDAL POSITION FULL FORWARD
D Cycle left and right pedal to check for binding.
b. SET PEDAL POSITION AS DESIRED FOR FLIGHT
• Locks securely when RUD PED ADJ lever released.
12. EXT and INTR lights - Check for proper operation to extent possible for ambient conditions (both
cockpits)
• Signal: Point 2 fingers at eyes.
13. FIRE warning test
a. FIRE test switch - TEST A (hold until all lights and aural warnings indicate test has been
successfully passed)
b. FIRE test switch - NORM (pause 7 seconds or cycle BATT switch for system reset)
c. FIRE test switch - TEST B (hold until all lights and aural warnings indicate test has been
successfully passed)
10.5.4 Engine Start Checks.
APU start 1. APU ACC caution light - VERIFY OFF
2. APU switch - ON (READY light within 30 seconds)
3. ENG CRANK switch - R
4. Right throttle - IDLE
• RPM
10% minimum
• TEMP
871°C maximum transient
• OIL
10 psi within 30 seconds
5. Battery gauge - VERIFY 28 vdc
• Battery charger failed if ≤ 24 vdc.
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6. EFD - CHECK
Ground idle • RPM
61% minimum
• TEMP
250° to 590°C
• FF
600 to 900 pph
• OIL
35 to 90 psi (warm oil)
• NOZ
77% to 83%
7. External electrical power - DISCONNECT
8. BLEED AIR knob - NORM
9. ENG CRANK switch - L
10. Left throttle - IDLE
• RPM
10% minimum
• TEMP
871°C maximum transient
• OIL
10 psi within 30 seconds
11. ENG CRANK switch - CHECK OFF
12. EFD - CHECK
10.5.5 Post-Start Checks.
1. APU automatic shutdown - CHECK
• APU shutdown 1 minute after second generator online.
2. WINDSHIELD ANTI ICE/RAIN removal - CHECK
a. WINDSHIELD switch - ANTI ICE
• Verify airflow along the canopy bow.
b. WINDSHIELD switch - RAIN
• Verify reduced airflow along the canopy bow.
c. WINDSHIELD switch - OFF
• Verify airflow is secured.
3. Canopy operation (front cockpit) - CHECK
a. CANOPY switch - CLOSE (half way)
• Canopy stops when switch is released.
b. CANOPY switch - OPEN then release
• Switch returns to HOLD position.
• Canopy moves to full open position.
c. (LOTs 26 and up) Repeat steps a-b for the aft canopy switch.
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d. (LOTs 26 and up) Front CANOPY switch - CLOSE, aft CANOPY switch - OPEN • Canopy should go up.
e. (LOTs 26 and up) Aft CANOPY switch - CLOSE, front CANOPY switch - OPEN • Canopy should go up. Position canopy as desired.
4. WINGFOLD switch - SPREAD
5. FCS RESET button - PUSH (verify RSET advisory displayed)
• No flight control surface Xs.
• No BLIN codes.
• No complete FCC channel failures.
After successful FCS reset 6. GAIN ORIDE - CHECK
With flaps FULL a. GAIN switch - ORIDE
• LAND advisory displayed.
• Amber FLAPS light on.
b. FLAP switch - AUTO
• CRUIS advisory displayed.
• Amber FLAPS light on.
c. GAIN switch - NORM/GUARD DOWN
• Amber FLAPS light out.
7. FCC keep-alive circuitry - CHECK
a. FCS CH circuit breakers - PULL IN SEQUENCE 1, 2, 3, AND 4
b. Immediately reset in sequence 1, 2, 3, 4.
• Complete within 7 seconds for valid test.
• No FCC channel completely Xd out.
• No FCS surface Xs and no BLIN codes.
Steps 8 thru 12 are to be performed on both engines (RIGHT then LEFT).
8. Engine FIRE light shutdown - PERFORM
a. Throttle affected engine - IDLE
b. FIRE light affected engine - PUSH
• FIRE EXTGH READY light comes on.
When BOOST LO caution appears, but no longer than 5 seconds c. Throttle affected engine - IMMEDIATELY OFF
• Master caution light comes on, and tone sounds when BOOST LO caution appears.
III-10-8
ORIGINAL
A1-F18EA-NFM-000
d. BIT/STATUS MONITOR/FXFR page - SELECT ON RDDI
• X COOL line reads CL (closed).
• Affected ENG SO line reads CL.
• CROSS FD line reads CL.
• Affected REC first value reads 0 (i.e., 0,0).
• FADEC HOT caution may appear and is not a failure indication.
e. FIRE light affected engine - RESET
• X COOL line reads O (open).
• Affected ENG SO line reads O.
• CROSS FD line reads O.
• Affected engine REC first value reads non-zero (e.g., 13,0).
Discontinue FCF upon failure of any item listed under b, c, d, and e above.
With affected engine below 10% N2 rpm 9. Fuel/air heat exchanger leak check - PERFORM
a. RBYP or LBYP option (affected side) - PUSH TO READ HX
• Signal: Pull tip of nose with thumb and index finger.
• No fuel leaking from the heat exchanger drains (forward lower inboard side of the inlet on the
affected side) (thumbs up from PC).
b. FXFR/RESET option - PUSH
• RBYP or LBYP option (affected side) reads BP.
10. Verify proper switching valve operation.
a. Note hydraulic pressure decay through 500 psi on the affected side.
b. Cycle lateral stick and rudder pedals and verify aileron and rudder surface movement.
c. If aileron, rudder, or LEF surfaces X, maintenance action is required.
11. GEN/electrical system checks - PERFORM
With affected GEN inoperative • Opposite GEN picks up all three busses.
• GEN TIE caution light out.
• All displays operative.
• No FCS Xs or channel failures.
a. BATT switch - OFF (opens bus tie)
• Busses are isolated on affected side.
• BATT SW and GEN TIE caution lights on.
With R GEN off • HUD and RDDI inoperative.
• LDDI, MPCD, and UFCD operative.
III-10-9
ORIGINAL
A1-F18EA-NFM-000
With L GEN off • HUD and RDDI operative.
• LDDI, MPCD, and UFCD inoperative.
b. BATT switch - ON
• BATT SW and GEN TIE caution lights out.
• All displays operative.
c. GEN switch opposite side - OFF (for at least 30 seconds)
• PMGs pickup essential bus.
• Battery gauge reads >24 vdc (26.5 vdc nominal).
• BATT SW caution light out.
d. GEN switch opposite side - NORM
• No complete FCC channel failures on FCS page.
• ENGINE LEFT/RIGHT voice alert and BLIN code 260 can be expected and are normal in the
conduct of this check.
12. Inoperative engine - CROSSBLEED START
Advance operating engine to a minimum of 80% rpm.
13. Repeat steps 8 thru 12 for the left engine.
• Restart left engine within 15 minutes, else motor for 1 minute at 29% rpm or greater before
restart (to preclude engine damage).
14. GEN TIE operation - CHECK
a. GEN TIE switch - RESET
• GEN TIE caution light on.
b. GEN TIE switch - NORM/GUARD DOWN
• GEN TIE caution light out.
15. WYPT 0 and MVAR - CHECK/SET
16. GPWS/TAWS - CHECK BOXED
17. INS knob - CV OR GND (PARK BRK SET)
18. RADAR knob - OPR
19. FLIR and LST/FLR switches - AS DESIRED
20. UFCD avionics - TURN ON
a. RALT - ON/SET
b. TCN - ON, T/R, CH SET
c. IFF - ON/MODES UNBOXED
21. MPCD/UFCD - ENTER DESIRED WAYPOINTS
III-10-10
ORIGINAL
A1-F18EA-NFM-000
22. Fuel system checks - PERFORM
On the RDDI a. SDC/sensor operation - CHECK
(1) SUPT/BIT/STATUS MONITOR page - SELECT
(2) SDC BIT option - SELECT
• SDC BIT status indicates GO.
(3) FXFR page - SELECT
• Do not take off with flashing parameters.
(4) FQTY page - SELECT
• Do not take off with flashing parameters.
On the LDDI b. Fuel quantity/cautions/advisories - CHECK
(1) SUPT/FUEL page - SELECT
• No fuel cautions or advisories displayed.
• No CG DEGD, EST, INV, INVALID, or INVALID TIMER.
• BINGO, TOTAL, and INTERNAL fuel quantities agree with EFD.
(2) FLBIT option - SELECT
• On RDDI, TK2FL indicates GO within 2 seconds.
• On RDDI, TK3FL indicates GO within 13 seconds.
• FUEL LO caution and voice alert activated within 13 seconds.
• FUEL LO caution removed 60 seconds after displayed.
(3) SDC RESET option - SELECT
• CAUT DEGD caution displayed for 3 seconds.
On the EFD c. BINGO caution - CHECK
(1) BINGO - SET 200 lb above INTERNAL fuel
• BINGO caution and voice alert activated.
(2) BINGO - SET 200 lb below INTERNAL fuel
• BINGO caution removed.
(3) BINGO - SET AS DESIRED
23. Hydraulic pressure gauge - CHECK (2,600 to 3,300 psi)
24. ECS system checks - PERFORM
a. DEFOG - CHECK
III-10-11
ORIGINAL
A1-F18EA-NFM-000
(1) DEFOG handle - LOW
• Minimum defog airflow and maximum cabin airflow.
(2) DEFOG handle - HIGH
• Progressively decreasing cabin airflow and increasing defog flow.
b. ECS modes - CHECK
• Signal: Punch open palm with fist.
(1) ECS MODE switch - OFF/RAM
• Cabin airflow stops.
• Cabin ram air scoop opens (thumbs up from PC).
If CPWS installed • CK ECS caution light on.
• MASTER CAUTION light on and tone sounds.
(2) ECS MODE switch - AUTO
• Cabin airflow resumes.
• Cabin ram air scoop closes (thumbs up from PC).
If CPWS installed • CK ECS caution light out.
c. CABIN TEMP knob - ROTATE BETWEEN COLD AND HOT
• Air temperature changes to agree with setting.
d. Cabin Pressurization - CHECK
(1) CABIN PRESS switch - DUMP
• Cabin depressurizes.
• Cabin airflow remains.
If CPWS installed • CK ECS caution light on.
• MASTER CAUTION light on and tone sounds.
(2) CABIN PRESS switch - RAM/DUMP
• Cabin remains depressurized.
• Cabin airflow stops.
• Cabin ram air scoop opens (thumbs up from PC).
If CPWS installed • CK ECS caution light on.
(3) CABIN PRESS switch - NORM
• Cabin pressurizes.
• Cabin airflow resumes.
• Cabin ram air scoop closes (thumbs up from PC).
If CPWS installed • CK ECS caution light out.
25. ENG ANTI ICE system - CHECK
a. ENG ANTI ICE switch - ON
• LHEAT and RHEAT advisories displayed.
III-10-12
ORIGINAL
A1-F18EA-NFM-000
b. ENG ANTI ICE switch - TEST
• INLET ICE caution displayed when switch held.
26. BLEED AIR system - CHECK
a. Throttles - IDLE
b. BLEED AIR knob - CHECK EACH POSITION INDIVIDUALLY
(1) R OFF
• R BLD OFF caution displayed.
• MASTER CAUTION light on and tone sounds.
• Left engine TEMP increases 5° to 90°C.
(2) MASTER CAUTION light - RESET
(3) Pause 5 seconds to allow Master Caution tone to reset.
(4) OFF
• L and R BLD OFF cautions displayed.
• MASTER CAUTION light on and tone sounds.
• Cabin airflow stops.
• ECS auxiliary duct doors close (thumbs up from PC).
If CPWS installed • CK ECS caution light on.
(5) L OFF
• R BLD OFF caution removed.
• Right engine TEMP increases 5° to 90°C.
• Cabin airflow resumes.
• ECS auxiliary duct doors open (thumbs up from PC).
If CPWS installed • CK ECS caution light out.
c. BLEED AIR knob - NORM
• L BLD OFF caution removed.
• MASTER CAUTION light out.
d. FIRE test switch - TEST A (for at least 2 seconds)
• L and R BLEED warning lights on while switch held.
• Voice alert sequence initiated.
• L and R BLD OFF cautions displayed.
• Cabin airflow stops.
e. BLEED AIR knob - CYCLE THRU OFF TO NORM
• L and R BLD OFF cautions removed.
• Cabin airflow resumes.
f. Repeat steps d and e for the TEST B position.
27. Mission computer operation - CHECK
III-10-13
ORIGINAL
A1-F18EA-NFM-000
LOTs 21-24:
a. SUPT MENU - SELECT ON LDDI
b. MC switch - 1 OFF
• MC1 and NO RATS cautions displayed.
• BIT, CHKLST, ENG, and ADI options removed from SUPT MENU.
• ACL option appears on HSI.
c. MC switch - NORM
• MC1 and NO RATS cautions removed.
• SUPT MENU options return.
d. TAC MENU - SELECT ON LDDI
e. MC switch - 2 OFF
• MC2 caution displayed.
• STORES option removed from TAC MENU.
f. MC switch - NORM
• MC2 caution removed.
• STORES option returns.
LOTs 25 and up:
a. SUPT MENU - SELECT ON RDDI
b. MC switch - 1 OFF
• MC1 caution displayed on MPCD.
• BIT and CHKLST options removed from SUPT MENU.
• (A/A Master mode) STORES option removed from TAC MENU.
• LDDI displays green square.
c. MC switch - NORM
• MC1 caution removed.
• LDDI display returns.
• SUPT MENU option returns.
d. TAC MENU - SELECT ON LDDI
e. MC switch - 2 OFF
•
•
•
•
MC2 caution displayed on MPDC.
RDDI displays green square.
BIT, FCS, and CHKLST options removed from the SUPT MENU.
(A/A Master mode) STORES option removed from TAC MENU.
f. MC switch - NORM
• MC2 caution removed.
• RDDI display returns.
• STORES option returns.
III-10-14
ORIGINAL
A1-F18EA-NFM-000
28. (LOT 25 and up) HUD backup operation - CHECK
a. MC switch - 1 OFF FOR 3 SECONDS
b. MC switch - 2 OFF
• Both DDIs display a green square followed by a flashing STANDBY.
• Backup HUD provided on MPCD and UFCD.
c. MC switch - NORM
10.5.6 Before Taxi Checks.
1. WINGFOLD system - CHECK
With wings spread and locked a. WINGFOLD switch - HOLD
• Ailerons fair and beer cans pop up.
• WING UNLK cautions displayed.
b. NWS button - PUSH (twice if required)
• Full-time NWS HI available.
c. WINGFOLD switch - FOLD
• Wingfold system and locking pins operate properly.
• Both ailerons Xd out.
d. WINGFOLD switch - SPREAD THEN HOLD
• Wings stop at intermediate position.
With NWS HI selected e. WINGFOLD switch - SPREAD
• Wings spread fully.
• Beer cans go down.
• WING UNLK caution removed.
• NWS HI reverts to NWS (low).
f. NWS - PADDLE OFF
2. Throttle position related cautions - CHECK
a. PARK BRK handle - SET
b. FLAP switch - AUTO
c. Stabilator trim - SET LESS THAN 3° NU
d. Ejection seat SAFE/ARMED handle(s) - SAFE (both cockpits)
e. Throttles - ADVANCE TO MIL MOMENTARILY (Do not allow engine RPM to exceed 80%.)
• CK FLAPS and PARK BRK cautions displayed momentarily.
III-10-15
ORIGINAL
A1-F18EA-NFM-000
• CHECK TRIM and CHECK SEAT cautions displayed.
• CHECK SEAT caution does not clear until seat(s) armed for takeoff.
f. T/O TRIM button - PRESS UNTIL TRIM ADVISORY DISPLAYED
• CHECK TRIM caution removed.
To the maximum extent possible, make sure wings are spread and locked prior to FCS
IBIT to make sure all aileron related tests are performed.
3. FCS RESET button - PUSH (if required)
• RSET advisory displayed.
• If wings are folded, both ailerons Xd out.
4. FCS IBIT - PERFORM
a. FCS BIT consent switch - HOLD UP THEN PRESS THE FCS OPTION
b. AOA warning tone - VERIFY ANNUNCIATION AT FCS IBIT COMPLETION
c. FCS A and FCS B BIT status - VERIFY GO (if wings not folded)
d. FCS display - VERIFY NO BLIN CODES
5. Trim - CHECK
a. Trim - FULL LEFT and UP (ailerons, rudders, and stabs)
• Control surfaces respond correctly.
b. T/O TRIM button - PRESS UNTIL TRIM ADVISORY DISPLAYED
c. Trim - FULL RIGHT and DOWN (ailerons, rudders, and stabs)
• Control surfaces respond correctly.
6. T/O TRIM button - PRESS UNTIL TRIM ADVISORY DISPLAYED (stabilators 4° NU)
7. Controls - CHECK (tolerance ±1°)
a. Control stick - CYCLE
(1) Full aft - CHECK 24° NU STABILATOR (check left and right stabilators track
symmetrically within ±1° of each other)
(2) Full fwd - CHECK 20° ND STABILATOR (check left and right stabilators track
symmetrically within ±1° of each other)
(3) Full L/R - CHECK 30° DIFFERENTIAL STABILATOR (21° with tanks or A/G stores on
any wing station)
- CHECK DIFFERENTIAL TEFs
b. FLAP switch - HALF
c. Rudder pedals - CYCLE RUDDERS 40° L/R
III-10-16
ORIGINAL
A1-F18EA-NFM-000
d. FLAP switch - FULL (carrier-based)
e. TRIM - SET FOR CATAPULT LAUNCH (carrier-based)
8. Five Down Checks
a. PROBE, speedbrake, LAUNCH BAR switches and HOOK handle - CYCLE
• Spoilers extend to 60° ±3° and retract in 3 seconds.
• SPDBRK light on when spoilers not fully retracted.
• Hook extends within 2 seconds and retracts within 4 seconds.
• Probe extends and retracts within 6 seconds.
• Probe light is on with the probe extended (thumbs up from PC).
• Launch bar extends (leave extended).
• Green LBAR advisory light on.
b. AV COOL switch emergency cooling check
• AV COOL switch - EMERG (FCS ram air scoop deploys)
9. Pitot and AOA heat check - PERFORM
a. PITOT ANTI ICE switch - ON
b. Make sure ground crew verify proper operation.
c. PITOT ANTI ICE switch - AUTO
10. CHECK TRIM caution - CHECK
With launch bar extended a. Stabilator trim - SET LESS THAN 6° NU
b. Throttles - Advance to MIL momentarily. Do not allow engine rpm to exceed 80%.
• CHECK TRIM caution displayed.
c. Stabilator trim - SET ABOVE 7° NU
• CHECK TRIM caution removed.
d. T/O TRIM button - PRESS UNTIL TRIM ADVISORY DISPLAYED
e. LAUNCH BAR switch - RETRACT
11. CVRS - AS DESIRED (both cockpits)
12. Standby attitude reference indicator - UNCAGE AND ERECT (both cockpits)
13. Altimeter setting - SET (both cockpits)
• Altimeter setting displayed on HUD.
• HUD altitude displayed within ±30 feet of parking spot elevation.
• Standby altimeter within ±60 feet of parking spot elevation.
14. INS - CHECK
III-10-17
ORIGINAL
A1-F18EA-NFM-000
a. PARK BRK handle - CYCLE
• INS alignment time flashes when PARK BRK released.
• Stops flashing after PARK BRK reset.
• On ANAV equipped aircraft the alignment time does not flash when PARK BRK is released
unless the aircraft moves.
b. Alignment status - VERIFY COMPLETE
• QUAL ″OK″ displayed within 6 minutes.
c. GPS HERR/VERR - VERIFY WITHIN LIMITS
When clear of overhead obstructions for 6 to 12 minutes • HERR and VERR less than 100 feet (with keyed MAGR).
d. INS knob - NAV (to check unaided drift)
e. Verify HUD airspeed indicates less than 50 kts.
15. MUMI/ID - SELECT/ENTER DATE and FLT
16. Stores page - Verify proper store inventory and station status.
17. ZTOD/LTOD - BOX TO ENABLE HUD DISPLAY (if desired)
18. Weapons/sensors - ON/BIT CHECK (as required)
19. BIT page - NOTE DEGD/FAIL INDICATIONS
20. Standby attitude data - CHECK
a. ATT switch - STBY
• Velocity vector disappears.
• Pitch ladder referenced to the W .
• INS ATT caution displayed.
b. Standby attitude reference indicator - ERECT
• HUD pitch ladder moves/coincides with the standby attitude reference indicator.
c. ATT switch - AUTO
21. OBOGS system - CHECK
a. OBOGS control switch - ON
b. OXY FLOW knob - ON/MASK(S) (both cockpits)
• System provides oxygen on demand.
• No excessive backpressure.
c. OBOGS monitor pneumatic BIT plunger - PRESS AND HOLD (do not rotate)
• OBOGS DEGD caution displayed within 65 seconds.
III-10-18
ORIGINAL
A1-F18EA-NFM-000
• Release plunger.
• Caution removed within 30 seconds.
Inadvertent rotation of the OBOGS monitor pneumatic BIT plunger
while pressed can result in the locking of the plunger in a maintenance
position and may result in intermittent OBOGS DEGD cautions and lead
to hypoxia. Rotation of the BIT plunger disengages the locking slot
allowing the plunger to extend and move freely when pushed.
d. OBOGS electronic BIT button - PRESS AND RELEASE
• OBOGS DEGD caution displayed and removed within 15 seconds.
e. OXY FLOW knob(s) - OFF (both cockpits)
• OBOGS flow stops.
22. Engine status/FADEC channel transfer - CHECK
a. ENG format - SELECT ON LDDI
• LEFT and RIGHT engine STATUS is NORM.
b. FADEC channel transfers - A TO B AND B TO A ON EACH ENGINE
• FADEC channels change with selection.
• No channel line-outs.
23. REFUEL DR caution - CHECK
• Signal: (refuel cap) twist hand with curled fingers.
• REFUEL DR caution displayed when PC opens door 8R.
• Caution removed when PC closes door 8R.
• PC manually restows scoop.
• Thumbs up from PC checks complete good.
10.5.7 Taxi Checks.
1. Canopy - EITHER FULL UP OR FULL DOWN FOR TAXI
2. Braking system - CHECK
a. Normal brakes - CHECK
• Nominal braking performance at taxi speed.
b. ANTI SKID switch - OFF
• SKID advisory displayed.
• Nominal braking performance at taxi speed.
c. ANTI SKID switch - ON
• SKID advisory clears.
d. EMERG BRK handle - PULL TO DETENT (both cockpits - separately for LOTs 21-25, front
cockpit only for LOTs 26 and up)
III-10-19
ORIGINAL
A1-F18EA-NFM-000
• Handle latches securely in detent.
• Nominal braking performance at taxi speed.
e. EMERG BRK handle - NORM
3. Nosewheel steering - CHECK IN HIGH MODE L/R
• NWS responds appropriately in NWS and NWS HI.
• NWS disengages when paddle switch pressed.
10.5.8 Shipboard Taxi/Takeoff Checks.
1. Canopy - CHECK CLEAR/CLOSED (canopy caution removed)
2. OXY FLOW knob(s) - ON/MASK(S) ON prior to tiedown removal
3. Checklist page
a. FUEL TYPE - VERIFY
b. ABLIM OPTION - BOX
4. ABLIM advisory - VERIFY DISPLAYED on appropriate DDI
5. PARK BRK handle - FULLY STOWED
6. T.O. checklist - COMPLETE from Bottom to Top
7. IFF - SQUAWK MODES /CODES as appropriate
8. Heading checks
• NOTE HSI heading matches BRC within ±3° .
• STBY magnetic compass within limits of compass card.
At catapult tension signal 9. Engine run-ups - PERFORM (together)
a. Throttles both engines - IDLE to MIL
b. ENG page - CHECK ENGINES AT MIL
• N1 rpm
86 to 98%
• N2 rpm
88 to 100%
• EGT
720 to 932°C
• FF
11,000 pph maximum
• NOZ POS
0 to 45% open
• OIL PRESS
80 to 150 psi
• THRUST
100% minimum on CAT officer/deck lighting signal
10. Afterburners - SELECT
• Both nozzles open correctly.
• Feed tanks remain full during takeoff, climb, and immediately following climb.
III-10-20
ORIGINAL
A1-F18EA-NFM-000
10.5.9 Shorebased Takeoff Checks.
1. Canopy - CHECK CLEAR/CLOSED, canopy caution removed
2. OXY FLOW knob(s) - ON/MASK(S) ON
3. Checklist page
a. FUEL TYPE - VERIFY
b. T.O. checklist - COMPLETE
4. PARK BRK handle - FULLY STOWED in position and hold
5. IFF - Squawk appropriate modes/codes
6. Heading sources - CHECK after runway lineup • HSI heading within ±3° of known runway heading.
• STBY magnetic compass within limits of compass card.
7. Engine run-ups - PERFORM (individually)
a. Throttles affected engines - IDLE to MIL
b. ENG page - CHECK ENGINE AT MIL
• N1 rpm
86 to 98%
• N2 rpm
88 to 100%
• EGT
720 to 932°C
• FF
11,000 pph maximum
• NOZ POS
0 to 45% open
• OIL PRESS
80 to 150 psi
• THRUST
100% minimum
c. Throttles affected engines - MIL to IDLE, pause 1 second, IDLE to MIL
• Engine responds with normal acceleration characteristics.
• No stall or stagnation.
d. Throttles affected engines - IDLE
e. Repeat steps a thru d for opposite engine.
When cleared for takeoff 8. Afterburner Takeoff - Perform IAW Chapter 7.
• Both nozzles open correctly.
• Feed tanks remain full during takeoff, climb, and immediately following climb.
10.5.10 After Takeoff Checks.
When definitely airborne 1. LDG GEAR handle - UP
• Gear retracts within 7 seconds.
III-10-21
ORIGINAL
A1-F18EA-NFM-000
10.5.11 Medium Altitude Checks (above 10,000 feet).
Altitude blocks are suggested ONLY to provide a logical sequence for the FCF procedures. Deviations
from these block altitudes are acceptable unless specified.
1. Cabin pressurization - CHECK (both cockpits)
Aircraft Altitude
Cabin Altitude
• <8,000 feet
Aircraft altitude (+0, -3,000 feet)
• 8,000 to 24,500 feet
8,000 feet (±500 feet)
2. Fuel transfer - CHECK INTERNAL and EXTERNAL
3. RALT - CHECK SET to 5,000 FEET
4. COMM - CHECK (both cockpits)
• Comm switches function normally.
• Both radios operative in transmit and receive.
• Preset and manual frequency selection operative.
5. Flight control damping - CHECK
a. Airspeed - Maintain 300 to 350 KCAS
b. Make small, abrupt pitch, roll, and yaw inputs.
• Aircraft response is appropriate.
• No oscillation tendencies noted.
6. Air refueling probe - CHECK
• Airspeed - Maintain below 300 KCAS.
a. PROBE switch - EXTEND
• Probe extends normally within 6 seconds.
b. PROBE switch - RETRACT
• Probe retracts normally within 6 seconds.
• No PROBE UNLK caution when retracted.
c. PROBE switch - EMERG EXTD
• Probe extends normally within 6 seconds.
d. PROBE switch - RETRACT
• Probe retracts normally within 6 seconds.
• No PROBE UNLK caution when retracted.
7. HOOK - CHECK
a. HOOK handle - DOWN
• HOOK light on while hook in transit.
• HOOK light out when hook fully extended.
b. HOOK handle - UP
III-10-22
ORIGINAL
A1-F18EA-NFM-000
8. Fuel dump - CHECK
a. BINGO - CHECK/SET just below internal fuel level
b. DUMP switch - ON
• Fuel dumps from both vertical tails.
c. BINGO - Run above internal fuel level
• DUMP switch returns to OFF automatically.
• Fuel dump stops.
9. ATC cruise mode - CHECK
• ATC advisory in HUD when selected.
• Throttles respond correctly.
• ATC holds calibrated airspeed when straight-and-level and during turns, climbs, and descents.
10. HUD symbology - CHECK (both cockpits)
In NAV master mode with WYPT or TCN boxed • The following indications are present - heading, airspeed, altitude, AOA, Mach number,
aircraft g, bank angle scale, velocity vector, flight path/pitch ladder, steering arrow (TCN),
and distance to WYPT or TCN.
• WYPT # displayed to right of distance when WYPT boxed.
• Three-letter identifier displayed to right of distance when TCN boxed.
• HUD and MPCD distance agree.
• HUD format available on DDI, MPCD, or UFCD (both cockpits).
11. HSI symbology - CHECK on MPCD
a. HDG/TK set switch - SLEW (both cockpits)
• Heading bug moves in correct direction.
• Digital display and bug setting agree.
b. CRS set switch - SLEW (both cockpits)
• Steering arrow rotates in correct direction.
• Digital display and steering arrow agree.
c. TCN bearing and range - CHECK
• Symbol displayed at appropriate position when compared to a waypoint or known landmark.
• Digital display and symbol agree.
12. Standby flight instruments - CHECK (both cockpits)
a. Standby rate of climb indicator
• Indicates ±100 fpm or less during level 1g flight.
• Pointer movement smooth during climbs/descents.
b. Standby attitude reference indicator
(1) Perform a 360° roll right and left.
• No gyro tumble.
III-10-23
ORIGINAL
A1-F18EA-NFM-000
(2) Perform a loop.
• Gyro indications are smooth thru bullseye.
c. Standby airspeed indicator
• Agrees with HUD.
• Pointer movement is smooth during airspeed changes.
d. Standby altimeter
• Agrees with HUD (accept -100 to 400 feet of error since value is uncorrected by FCC air data
function).
• Pointer and drum movement is smooth and does not hang up during thousand-foot changes.
13. INS/GPS operation - CHECK
a. TCN update - PERFORM
• Proper update mechanization.
• Reject update.
b. DSG update - PERFORM
• Proper update mechanization.
• Reject update.
c. GPS HERR/VERR - CHECK
• Less than 100 feet inflight (with keyed MAGR).
14. ADF receivers (if installed) - CHECK
a. ADF - BOX on COMM 1 and COMM 2 sub-levels (individually)
• Bearing within ± 5° off the nose.
• Bearing within ± 20° off the wing tip.
15. IFF operation - CHECK
• ATC reports valid mode 3 and C.
If/when possible a. IFF MASTER switch - EMERG
• ATC reports valid emergency squawk (7700).
16. Radar/HOTAS functionality - CHECK (both cockpits)
a. A/A master mode - CHECK
b. A/G master mode - CHECK
10.5.12 10,000 Feet Checks.
With an asymmetric FLIR pod on station 5 or 7 • Expect a gradually increasing amount of pod-induced roll-off during the accel to 500 KCAS on the FCS
RIG check or to Mach 0.9 to Mach 0.93 on the LEF/HDU stall check.
• In this configuration, FCS rigging is considered acceptable if the 300 and 400 KCAS points are passed.
• If encountered on the LEF/HDU stall check, this ″gradual″ roll-off is not indicative of an HDU stall.
III-10-24
ORIGINAL
A1-F18EA-NFM-000
1. FCS RIG check - 10,000 feet
Only perform if • Aircraft symmetrically loaded. (Asymmetric FLIR pod on station 5 or 7 acceptable for 300 and
400 KCAS points only.)
• External/internal wing tank fuel asymmetry less than 300 pounds.
a. Autopilot mode - Disengage in 1g flight
b. T/O TRIM button - PUSH (4 seconds minimum)
• Do not re-trim laterally or directionally for duration of check.
c. Stabilize at each incremental airspeed. Release controls from wings-level and record the
direction of roll-off and angle-of-bank (AOB) at the end of 10 seconds.
d. 300 KCAS
e. 400 KCAS
f. 500 KCAS
g. Mach 0.92
• Perform the Mach 0.92 check only if the AOB was > 30° at 300, 400, or 500 KCAS.
NOTE
RIG Check fails if any AOB > 30° (> 3° per second); however, for
diagnostics purposes, complete all applicable roll-off checks.
2. LEF/HDU stall check - 10,000 feet
a. G-warm - PERFORM
• 4g for 90°.
• 6g for 90°.
• -1g pushover to check for cockpit foreign objects.
b. FCS page - CHECK
• G-LIM value not Xd out.
• No G LIM 7.5G caution.
c. Speed - Accelerate to Mach 0.9 to Mach 0.93
d. Roll to 90° AOB, retard the throttles to IDLE, and smoothly pull to g-limiter.
At 25° AOA e. Terminate the maneuver.
• No ″abrupt″ rolling tendency.
• No FLAP SCHED caution.
• No BLIN code 256 (channel identifies weak HDU).
3. HYD system check - 10,000 feet
• May be accomplished in conjunction with the FCS RIG accel and LEF/HDU decel.
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A1-F18EA-NFM-000
a. Stabilize at 350 to 375 KCAS (less than Mach 0.65).
• HYD pressure is 3,000 psi (+300/-400).
b. Accelerate toward 450 KCAS.
• HYD pressure increases to 5,000 psi (+400/-500) by 420 KCAS.
c. Decelerate towards 300 KCAS.
• HYD pressure returns to 3,000 psi (+300/-400) by 330 KCAS.
4. Emergency landing gear extension - PERFORM (front cockpit)
a. FLAP switch - HALF
b. Slow below 170 KCAS.
c. LG circuit breaker - PULL
• Rear cockpit landing gear UNSAFE light on.
d. LDG GEAR handle - DN
e. LDG GEAR handle - ROTATE 90° CLOCKWISE then PULL TO DETENT
• LDG GEAR handle stays in detent.
• Gear extends within 30 seconds.
• APU ACCUM caution displayed.
f. HYD ISOL switch - ORIDE (until APU ACCUM caution removed - approximately 20 seconds)
With the LDG GEAR handle outboard (DN position) g. LDG GEAR handle - PUSH IN then ROTATE 90° CCW
Pause 5 seconds h. LG circuit breaker - RESET
5. (LOTs 21-25) Emergency landing gear extension - PERFORM (rear cockpit)
In front cockpit a. LDG GEAR handle - UP
b. FLAP switch - HALF
c. Slow below 170 KCAS.
In rear cockpit d. EMERG LDG GEAR handle - PULL TO DETENT
• Gear extends within 30 seconds.
• APU ACCUM caution displayed.
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In front cockpit e. LDG GEAR handle - DN
f. HYD ISOL switch - ORIDE (until APU ACCUM caution removed - approximately 20 seconds)
In rear cockpit g. EMERG LDG GEAR handle - ROTATE 45° CLOCKWISE and PUSH FULL IN
6. AOA warning tone - CHECK
With gear down and flaps HALF a. Increase AOA toward 15°.
• AOA warning tone comes on at 14 ± 0.5°.
7. PA throttle transients - 10,000 feet (INDIVIDUALLY)
With gear down, flaps HALF, and at onspeed AOA a. Throttle affected engine - IDLE to MAX
• Afterburner lights within 8 seconds.
b. Throttle affected engine - MAX to IDLE, pause 3 seconds, IDLE to MAX
• Afterburner lights within 8 seconds.
• Engine responds smoothly with no stall, stagnation, or flameout.
c. Repeat steps a and b for opposite engine.
8. LDG GEAR handle - UP
9. Wheels warning - CHECK
a. Descend below 7,500 feet MSL.
b. Reduce airspeed below 175 KCAS.
c. Establish rate of descent greater than 250 fpm.
• Landing gear warning light flashes.
• Landing gear warning tone sounds.
10. FLAP switch - AUTO
10.5.13 High Altitude (above 30,000 feet).
1. Cabin pressurization - MONITOR
Above 24,500 feet MSL, cabin pressurization shall remain within 5 psi differential of actual
altitude. A rule of thumb is altitude x 0.4.
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A1-F18EA-NFM-000
Aircraft Altitude
• Less than 30,000 feet
• 40,000 feet
Cabin Altitude
10,000 to 12,000 feet
15,000 to 17,000 feet
2. Throttle transients - 35,000 ± 2,000 feet (INDIVIDUALLY)
a. ENG ANTI ICE switch - CHECK OFF
b. Airspeed - Maintain 200 to 220 KCAS
c. Throttle affected engine - IDLE to MAX
• Afterburner lights within 12 seconds.
d. Throttle affected engine - MAX to IDLE, pause 3 seconds, IDLE to MAX
• Afterburner lights within 12 seconds.
• Engine responds smoothly with no stall, stagnation, or flameout.
e. Repeat steps a thru d for opposite engine.
10.5.14 10,000 Feet to Landing.
1. Fuel transfer - CHECK (throughout flight)
With external fuel available • External fuel transfers normally.
• Internal tanks fill/stay near full.
With external tanks empty • Tank 1 depletes to approximately 1,000 pounds prior to wing tanks depleting.
When wing tanks are empty • Tanks 1 and 4 fall in approximately ¼ ratio.
• No FUEL XFER caution.
With fuel in tanks 1 and 4 • Feed tanks stay at or near full (2,100 to 2,450 pounds).
2. RALT operation - CHECK
During descent through 5,000 feet AGL • Low altitude warning correctly comes on.
• Radar altitude tracks correctly during descent.
• Verify flashing B changes to a solid R when passing through 5,000 feet AGL.
3. TCN or WYPT course intercept - PERFORM
• Course deviation indicator in HUD corresponds with steering arrow on MPCD.
4. ILS/ACLS operation - CHECK (if available)
• Proper ILS and/or ACLS indications.
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10.5.15 Landing Checks.
1. Landing checklist - COMPLETE
2. ATC approach mode - CHECK
• ATC advisories in HUD when selected.
• Throttles respond correctly.
• Holds onspeed AOA during turns and on approach.
10.5.16 After Landing Checks.
1. Anti-skid system - CHECK
Above 75 KGS on landing a. Brake pedals - Apply full brake pressure
• Anti-skid cycles smoothly.
• No left or right pulling tendencies.
When clear of active runway 2. Ejection seat SAFE/ARMED handle(s) - SAFE (both cockpits)
3. EJECTION MODE handle - NORM (rear cockpit)
4. Landing gear handle mechanical stop - FULLY ENGAGED
5. FLAP switch - AUTO
6. T/O TRIM button - PRESS UNTIL TRIM ADVISORY DISPLAYED
7. Mask(s) - OFF (both cockpits)
8. OXY FLOW knob(s) - OFF (both cockpits)
9. OBOGS control switch - OFF
10. Canopy - EITHER FULL UP OR FULL DOWN FOR TAXI
10.5.17 Before Engine Shutdown Checks.
1. PARK BRK handle - SET
2. BIT display - RECORD DEGD/FAIL INDICATIONS
3. BIT/HYDRO-MECH page - VERIFY absence of FADEC fault codes
4. Radar maintenance (BOA) codes - RECORD IF PRESENT
5. RADAR knob - OFF
6. FCS display - RECORD BLIN CODES
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7. EFD - RECORD MSP CODES
8. INS - PERFORM POST FLIGHT UPDATE
• Maximum error is 1.5 nm per hour of operating time.
9. INS knob - OFF
10. Standby attitude reference indicator - CAGE (both cockpits)
11. HMD switch - OFF
12. Sensors, avionics, and CVRS - OFF
13. EXT and INTR lights knobs - OFF (both cockpits)
14. Canopy - CHECK CLEAR/OPEN
15. QDC - DISCONNECTED AND STOWED
10.5.18 Engine Shutdown Checks.
1. Brake accumulator gauge - CONFIRM 3,000 PSI
2. Paddle switch - PRESS (disengage NWS)
3. Confirm 5 minute engine cool down.
4. OBOGS control switch - OFF
5. BLEED AIR knob - OFF
6. Throttle - OFF (alternate sides)
7. Verify proper switching valve operation.
After hydraulic pressure decays through 500 psi a. FLAP switch - FULL
b. If aileron, rudder, or LEF surfaces X and the Xs do not clear after one FCS reset attempt,
maintenance action is required.
c. If one FCS reset attempt was required to reset surfaces Xs, cycle FLAP switch to AUTO then
back to FULL. If Xs reappear, maintenance action is required.
8. FCS page - Verify no channel is completely Xd out.
9. COMM 1 and 2 knobs - OFF (both cockpits)
10. L (R) DDI, HUD, and MPCD knobs - OFF (both cockpits)
11. Other throttle - OFF
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A1-F18EA-NFM-000
When amber FLAPS light illuminates 12. BATT switch - LEAVE ON
• Battery gauge reads 23 to 24 vdc (nominal).
• Automatic battery cutoff operates at 2 minutes (1 minute on LOT 21 aircraft).
13. FCF Profile A - COMPLETE
10.6 FCF CHECKLIST - PROFILE C
1. Perform engine start, taxi, and takeoff IAW NATOPS.
10.6.1 10,000 Feet Checks.
With an asymmetric FLIR pod on station 5 or 7 • Expect a gradually increasing amount of pod-induced roll-off during the accel to 500 KCAS on the FCS
RIG check or to Mach 0.9 to Mach 0.93 on the LEF/HDU stall check.
• In this configuration, FCS rigging is considered acceptable if the 300 and 400 KCAS points are passed.
• If encountered on the LEF/HDU stall check, this ″gradual″ roll-off is not indicative of an HDU stall.
1. FCS RIG check - 10,000 feet
Only perform if • Aircraft symmetrically loaded. (Asymmetric FLIR pod on station 5 or 7 acceptable for 300 and
400 KCAS points only.)
• External/internal wing tank fuel asymmetry less than 300 pounds.
a. Autopilot mode - Disengage in 1g flight
b. T/O TRIM button - PUSH (4 seconds minimum)
• Do not re-trim laterally or directionally for duration of check.
c. Stabilize at each incremental airspeed. Release controls from wings-level and record the
direction of roll-off and angle-of-bank (AOB) at the end of 10 seconds.
d. 300 KCAS
e. 400 KCAS
f. 500 KCAS
g. Mach 0.92
• Perform the Mach 0.92 check only if the AOB was > 30° at 300, 400, or 500 KCAS.
NOTE
RIG check fails if any AOB > 30° (> 3° per second); however, for
diagnostics purposes, complete all applicable roll-off checks.
2. LEF/HDU stall check - 10,000 feet
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A1-F18EA-NFM-000
a. G-warm - PERFORM
• 4g for 90°.
• 6g for 90°.
• -1g pushover to check for cockpit foreign objects.
b. FCS page - CHECK
• G-LIM value not Xd out.
• No G-LIM 7.5G caution.
c. Speed - Accelerate to Mach 0.9 to Mach 0.93
d. Roll to 90° AOB, retard throttles to IDLE, and smoothly pull to g-limiter.
At 25° AOA e. Terminate the maneuver.
• No ″abrupt″ rolling tendency.
• No FLAP SCHED caution.
• No BLIN code 256 (channel identifies weak HDU).
3. FCF Profile C - COMPLETE
10.7 FCF CHECKLIST - PROFILE D (REAR COCKPIT)
1. When a profile D is required solely by the reconfiguration of the rear cockpit, perform engine
start, taxi, and takeoff IAW NATOPS.
10.7.1 Preflight Checks.
1. UFCD adapter - VERIFY NOT INSTALLED
10.7.2 Before Taxi Checks.
1. Rudder pedal adjustment - CHECK
• Adjustment smooth through full range of travel.
a. SET PEDAL POSITION FULL FORWARD
D Cycle left and right pedal to check for binding.
b. SET PEDAL POSITION AS DESIRED FOR FLIGHT
• Locks securely when RUD PED ADJ lever released.
2. Stick and rudder pedals - CYCLE
• No binding through full travel.
3. Throttles - Advance to MIL momentarily. Do not allow engine rpm to exceed 80%.
• No binding or sticking through range of travel.
• No engine shutdowns when pulled to IDLE.
10.7.3 Taxi Checks.
1. Braking system - CHECK
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a. Normal brakes - CHECK
• Nominal braking performance at taxi speed.
b. EMERG BRK handle - PULL TO DETENT
• Handle latches securely in detent.
• Nominal braking performance at taxi speed.
c. EMERG BRK handle - NORM
2. Nosewheel steering - CHECK IN HIGH MODE L/R
• NWS responds appropriately in NWS and NWS HI.
• NWS disengages when paddle switch pressed.
10.7.4 Medium Altitude Checks (above 10,000 feet).
1. Flight control damping - CHECK
a. Airspeed - Maintain 300 to 350 KCAS
b. Make small, abrupt pitch, roll, and yaw inputs.
• Aircraft response is appropriate.
• No oscillation tendencies noted.
2. Throttles - CYCLE INTO AB
• Nominal engine response to throttle position.
• Afterburners light off normally and cancel when MIL selected.
3. COMM switch - CHECK
• Comm switch functions normally.
• Both radios operative in transmit and receive.
4. Speedbrakes - CHECK
a. Speedbrake switch - HOLD AFT
• Speedbrake surfaces extend normally.
• SPD BRK light on when surfaces not fully retracted.
b. Speedbrake switch - RELEASE
• Speedbrake surfaces retract fully.
In front cockpit c. Speedbrake switch - HOLD AFT
• Speedbrake surfaces extend normally.
In rear cockpit d. Speedbrake switch - HOLD FWD
• Speedbrake surfaces retract (rear cockpit override).
e. Speedbrake switches - RELEASE
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5. Radar/HOTAS functionality - CHECK
a. A/A master mode - CHECK
b. A/G master mode - CHECK
6. FCF Profile D COMPLETE
10.8 FCF CHECKLIST - PROFILE E
1. Perform engine start, taxi, and takeoff IAW NATOPS.
10.8.1 10,000 Feet Checks.
1. Emergency landing gear extension - PERFORM (front cockpit)
a. FLAP switch - HALF
b. Slow below 170 KCAS
c. LG circuit breaker - PULL
• Rear cockpit landing gear UNSAFE light on.
d. LDG GEAR handle - DN
e. LDG GEAR handle - ROTATE 90° CLOCKWISE then PULL TO DETENT
• LDG GEAR handle stays in detent.
• Landing gear extends within 30 seconds.
• APU ACCUM caution displayed.
f. HYD ISOL switch - ORIDE (until APU ACCUM caution removed - approximately 20 seconds)
With the LDG GEAR handle outboard (DN position) g. LDG GEAR handle - PUSH IN then ROTATE 90° CCW
Pause 5 seconds h. LG circuit breaker - RESET
2. (LOTs 21-25) Emergency landing gear extension - PERFORM (rear cockpit)
In front cockpit) a. LDG GEAR handle - UP
b. FLAP switch - HALF
c. Slow below 170 KCAS
In rear cockpit d. EMERG LDG GEAR handle - PULL TO DETENT
• Gear extends within 30 seconds.
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A1-F18EA-NFM-000
• APU ACCUM caution displayed.
In front cockpit e. LDG GEAR handle - DN
f. HYD ISOL switch - ORIDE (until APU ACCUM caution removed - approximately 20 seconds)
In rear cockpit g. EMERG LDG GEAR handle - ROTATE 45° CCW and PUSH FULL IN
In front cockpit 3. LDG GEAR handle - UP
4. FCF Profile E - COMPLETE
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PART IV
FLIGHT CHARACTERISTICS
Chapter 11 - Flight Characteristics
65 (Reverse Blank)
ORIGINAL
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CHAPTER 11
Flight Characteristics
11.1 HANDLING QUALITIES.
The F/A-18E/F flight control system (FCS) is designed to present handling qualities that provide
virtually carefree maneuvering of the aircraft throughout most of the flight envelope. As such, there are
some flight characteristics which are somewhat unique to the F/A-18E/F airplane. A thorough
understanding of these flight characteristics along with the details of the flight control system
described in Chapter 2 and the operating limitations detailed in Chapter 4, allows the pilot to safely
and effectively exploit the full capabilities of the airplane.
11.1.1 Flight Control Mode Effects on Handling Qualities. Handling qualities are dependent on
which mode the flight control system is operating. FCS mode is determined primarily by the FLAP
switch position: power approach (PA) mode with the FLAP switch in HALF or FULL or up/auto (UA)
mode with the FLAP switch in AUTO. However, if airspeed is above approximately 240 KCAS, the
flight controls switches to, or remains in, UA mode regardless of FLAP switch position. FCS control
laws are also designed to minimize transients when switching flight control modes.
11.1.2 Handling Qualities with Flaps HALF or FULL. The FCS employs full-time AOA and pitch rate
feedback with flaps HALF or FULL. Therefore, longitudinal trim is required to maintain constant
AOA and/or airspeed. Once trimmed to an AOA, the aircraft tends to remain at that AOA until
changed by longitudinal stick or trim. The stick force gradient with AOA is constant up to 12° AOA
and does not vary with aircraft gross weight or center of gravity. Above 12° AOA, increased AOA
feedback increases stick forces as an artificial stall warning cue. Handling qualities are excellent up to
the 14° AOA limit. Maximum AOA at full aft stick with flaps HALF or FULL is approximately 25°
AOA. However, due to degraded handling qualities and reduced departure resistance above 15° AOA,
particularly with abrupt inputs, flight at greater than 14° AOA with flaps HALF or FULL is
prohibited.
The FCS provides good lateral directional control of the aircraft. The rolling surface to rudder
interconnect (RSRI) function along with sideslip and sideslip rate feedback are used to coordinate
lateral inputs, reducing pilot workload by allowing feet-on-floor maneuvering for most situations.
11.1.2.1 Stalls with Flaps HALF or FULL. The aircraft does not exhibit a classic stall break with flaps
HALF or FULL and both configurations are very departure resistant up to the 14° AOA limit for
normal two-engine operation, even with symmetric and asymmetric store loadings (see Single Engine
Operation). Roll and yaw control remain positive up to the 14° AOA limit in either flap setting but is
better above 10° AOA with flaps HALF. With flaps FULL, a distinct longitudinal buffet is felt at or
above 11 to 12° AOA which serves as a stall warning cue. This buffet tends to be more pronounced at
heavier gross weights and with wing tank loadings but does not adversely affect climb performance or
handling qualities. Above the AOA limit, uncontrollable roll-offs are possible in either flap setting,
particularly with high lateral weight asymmetry store loadings. An intermittent warning tone will
sound beginning at 14° AOA with an increasing beep frequency as AOA increases up to full aft stick.
11.1.2.2 Takeoff and Landing. Low gain nosewheel steering (NWS) incorporates yaw rate feedback
to stabilize directional control during the takeoff and landing roll. Maintaining runway position
without NWS using differential braking alone may be difficult. Crosswinds have minimal effect on
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ORIGINAL
A1-F18EA-NFM-000
takeoff characteristics and only a small amount of lateral stick into the wind is required to keep the
wings level during the takeoff roll. Nosewheel lift-off speeds are dependent on CG location and aircraft
gross weight. At nominal and forward CG locations, the airplane requires aft stick to effect rotation.
Premature aft stick application during the takeoff roll can result in early nosewheel lift-off and
potential over-rotation, particularly with aft CG.
• Pitch attitudes in excess of 10° during takeoff rotation may result in
ground contact between engine exhaust nozzles and/or stabilators.
• With combinations of heavy gross weight, forward CG, high density
altitudes and late takeoff rotation, ground speed can exceed the
maximum nose gear tire speed of 195 knots ground speed (see
NATOPS performance charts).
Additionally, landing gear speed limits can be easily exceeded during shallow climbs after takeoff
with MAX power.
With large lateral weight asymmetries, there is a slight tendency to yaw into the heavy wing during
the initial ground roll and again during the takeoff rotation. Otherwise, takeoff characteristics are very
similar to symmetric store loadings. Directional trim may be required after takeoff for balanced flight
with store asymmetries. A small lateral-directional transient may occur during configuration changes
from flaps HALF to AUTO or from flaps AUTO to HALF. The lateral transient occurs since TEFs are
deflected differentially for lateral control with flaps AUTO and the additional lateral control results in
an associated directional transient due to the rolling-surface-to-rudder interconnect.
Normal approach and landing characteristics are excellent; with good speed stability and solid
lateral-directional handling qualities. With crosswinds, a wings-level crabbed approach with removal of
half the crab angle just prior to touchdown minimizes deviations from runway heading and landing
gear side loads during landings. Touchdown in a full crab angle results in an uncomfortable roll
opposite the crab angle and upwind drift, requiring large rudder pedal inputs to align the aircraft with
the runway. Likewise, removing the crab angle entirely results in downwind drift and directional
transients after touchdown. A wing down, top rudder approach results in excessive bank angle and is
not recommended.
With flaps HALF or FULL, handling qualities with large lateral weight asymmetries are virtually
identical to those with symmetric loadings; however, landing with crosswinds from the heavy wing side
results in less roll away from the wind at touchdown. With large lateral asymmetries, the aircraft will
fly with the heavy wing forward. Upon landing the aircraft will yaw into the heavy wing as the aircraft
straightens to its ground track. Once airborne on a touch and go or bolter, the aircraft will yaw away
from the heavy wing as the aircraft trims to a heavy wing forward position. Regardless of wind
conditions, the aircraft tends to yaw away from the heavy wing during periods of heavy braking but the
yaw is easily countered with a small rudder pedal input.
11.1.3 Flaps AUTO Handling Qualities. The FCS control laws create handling qualities that are
slightly different from aircraft with conventional flight control systems. The most apparent characteristics are the neutral speed stability at low AOA and the excellent maneuverability at high AOA.
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Neutral speed stability occurs since the FCS automatically attempts to keep the aircraft in 1g, zero
pitch rate flight. This has the effect of eliminating the need for frequent longitudinal trim adjustments,
lowering pilot workload for most tasks; however, some tasks are made slightly more difficult. For
example, during large airspeed changes, the aircraft may initially appear to be slightly out of trim for
a few seconds until FCS re-establishes 1g flight. Since pitch trim biases the FCS away from 1g flight,
any pitch trim used during large airspeed changes must be removed within a few seconds of
establishing the new airspeed and only adds workload. Additionally, during climbs or dives, a small but
constant forward stick force is required to maintain a constant pitch attitude and load factor. Again,
if pitch trim is used to eliminate these stick forces, additional short trim inputs will be required to
re-establish 1g flight, further increasing pilot workload. Another task with a slightly increased
workload is the instrument penetration/approach where neutral speed stability may cause difficulty in
maintaining a desired airspeed.
The longitudinal handling qualities are excellent with good pitch rate and damping that combine to
allow very aggressive maneuvering. FCS control laws modify aircraft response to stick inputs, creating
the effect of changing stick forces to provide pilot cueing in maneuvering flight. Actual stick forces for
a given stick displacement do not change with flight condition. Full forward and aft stick requires a 20pound push and 37-pound pull, respectively. At high airspeeds, the FCS is a g-command system
requiring 3.5 pounds of stick force per g. At medium airspeeds, the FCS acts as a hybrid pitch rate and
g-command system. Pitch rate feedback is used to increase apparent stick force per g as a cue of
decaying airspeed and available load factor. At low airspeed, the FCS is primarily an AOA command
system using AOA feedback above 22° to provide increasing stick force with increasing AOA. The
maximum commanded AOA is approximately 45° to 50° at full aft stick. Combined with the capability
to command high AOA is the ability to generate high nose-down pitch rates with large forward stick to
rapidly reduce AOA, particularly below approximately 200 KCAS. This nose-down pitch rate
capability is further enhanced as airspeed decreases to 150 KCAS. When airspeed is below 150 KCAS
and longitudinal stick is pushed far forward (greater than 1.7 inches), up to full stabilator, maximum
rudder flare-out, and LEX spoiler are commanded to rapidly get the aircraft nose moving down. This
FCS feature was added to enable pilots to rapidly reduce AOA when at low airspeed and high AOA for
quick nose repositioning. To maintain departure resistance, the enhanced nose-down pitch rate
capability is reduced when lateral stick is deflected more than one inch.
The g-limiter function in the FCS limits commanded load factor under most flight conditions to the
symmetric load limit (NzREF) based on gross weight below 57,400 pounds gross weight. Above 57,400
pounds, NzREF is held constant at 5.5g even though the allowable load factor may be below NzREF
(refer to G-Limiter). Above 57,400 pounds gross weight, the g-limiter does not provide adequate over-g
protection and pilot action may be required to prevent an aircraft overstress.
Very abrupt full aft stick commands with aft CG conditions can beat the g-limiter and cause a
positive over-g (811 MSP code). Likewise, very abrupt pushes can result in a negative over-g (925 MSP
code). Care should be taken during all abrupt maneuvers. During rolling maneuvers, the g-limiter also
reduces commanded load factor to 80% of NzREF. This feature can also be defeated with abrupt
lateral stick inputs at elevated-g. Abrupt full lateral stick inputs at NzREF may result in an aircraft
overstress without setting an 811 or 925 over-g MSP code (see figure 4-7, Acceleration Limits - Basic
Aircraft, Note 2).
Another flight characteristic related to g-limiter performance occurs during very high bleed rate
turns where even with full aft stick, load factor may be slightly less than NzREF. This can happen when
the aircraft decelerates much faster than the FCS can position the stabilators to maintain NzREF.
Additionally, during elevated-g maneuvering at transonic flight conditions, the g-limiter unloads the
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aircraft (and NzREF) by as much as 1.0 to 1.7g. This feature helps prevent an aircraft overstress that
could result from the classic aerodynamic phenomenon known as ‘‘transonic pitch-up’’ experienced
during elevated-g decelerations at transonic flight conditions.
At low angles of attack, the aircraft is extremely smooth with little sensation of changing airspeed
or Mach. However, at transonic flight conditions, the aircraft may exhibit a mild buffet, which is more
pronounced with empty wing pylons or interdiction loadings but is almost nonexistent with clean
wings. Buffet begins at approximately 0.88M, subsides by 0.95M, and presents a sensation much like
riding on a ″gravel road″. Airframe buffet is also noticeable in the cockpit while maneuvering at tactical
speeds between approximately 6° and 11° AOA. At low altitudes, this AOA range begins at
approximately 6g but at high altitude begins at 2g to 3g. Above this AOA range, the buffet sensation
at the cockpit subsides slightly but is still apparent in the airframe. Although this buffet at elevated-g
is present in all configurations, it is most apparent with empty wing pylons at transonic flight
conditions. Additionally, persistent but bounded wing rock or roll-off may occur at some flight
conditions if the maneuvers linger in the 8 to 13° AOA range. Gun tracking is still good in the presence
of buffet but workload is slightly increased. Formation flight in the buffet AOA region also exhibits a
slightly higher workload.
The speedbrake function provides very good deceleration capability at subsonic flight conditions.
Deploying the speedbrake function results in a small nose-up transient; a small nose-down transient
during retraction. These transients still allow the speedbrake function to be used comfortably during
formation flight. With speedbrake function fully deployed, the aircraft may feel sloppy in the yaw axis
during large rudder pedal inputs due to one rudder stalling. With lateral weight asymmetries, a small
sideforce may also be apparent when deploying the speedbrake function. At most supersonic flight
conditions up to Mach 1.5, the spoilers are the only active speedbrake surface due to limited
effectiveness of the other surfaces. Deceleration capability is still adequate with throttles at IDLE; with
one exception. When less than MIL thrust is selected above Mach 1.23, the engine fan speed lockup
feature (to prevent engine inlet instability) maintains MIL thrust levels, which has the side effect of
limiting deceleration capability until fan speed lockup deactivates at Mach 1.18.
Lateral-directional handling qualities are also excellent, particularly at high AOA. Roll rates and roll
damping combine to provide very agile roll control. The FCS attempts to maintain consistent roll
response throughout the 1g flight envelope. Additionally, rolling surface to rudder interconnects
coordinate lateral inputs, reducing pilot workload by allowing feet on the floor maneuvering under
most circumstances. Maximum roll rates are in the 200 to 225°/second range with clean wings and
approximately 130 to 150°/second with wing tanks and/or air-to-ground stores. Flight tests with wing
tank loadings demonstrated a very localized drop in maximum roll rate of approximately 50°/second
at Mach 0.92 to Mach 0.93, most notably at 20,000 feet. The FCS reduces maximum roll rate by 40 to
60°/second at high subsonic airspeeds and low altitudes (approximately Mach 0.90 below 10,000 ft),
due to structural load concerns. For additional structural loads concerns during negative-g rolls,
maximum roll rate capability is reduced to approximately 60 to 80°/second above approximately 550
KCAS.
The most obvious lateral-directional characteristic is the excellent maneuverability at high AOA as
a direct result of specific FCS high AOA control laws. At 25° AOA and above, rudder pedal deflections
no longer provide yaw control inputs but instead act entirely as a roll control (identical to lateral stick
input) by commanding aileron and differential stabilator with the RSRI commanding the required
rudder deflection for roll coordination. Rudder pedal inputs are summed with lateral stick inputs and
this combined input is limited to a value equal to a full lateral stick input. Therefore, applying pedal
opposite to lateral stick cancels lateral stick inputs proportional to the pedal input, i.e., full opposite
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A1-F18EA-NFM-000
pedal cancels a full lateral stick command resulting in zero roll rate. Between 13° and 25° AOA, rudder
pedal deflections gradually change from pure yaw controllers to pure roll controllers. This method of
control provides enhanced departure resistance at high AOA.
Some traditional yaw control with rudder pedal is returned at low airspeed and high AOA only when
the pilot applies lateral stick and rudder in the same direction. This feature is effective only at
airspeeds below 225 KCAS and between 25° and 40° AOA. During flight tests, the most effective
pirouette initiation was found at approximately 200 KCAS and 35° AOA (18E-006 and subsequent).
Enabling this feature outside of these conditions would compromise departure resistance. When this
feature is enabled, the sum of lateral stick and rudder pedal command is no longer limited to a value
equal to a full lateral stick input. The excess roll command is fed to the directional axis to command
sideslip. For example, adding full rudder pedal with a full lateral stick input provides a maximum roll
and yaw command. Alternatively, adding lateral stick to an existing full rudder pedal input has the
same effect. The resulting aircraft motion is a highly controllable nose-high to nose-low reversal.
Small lateral trim variances may occur without significant changes in airspeed, AOA, or Mach
number. These variances result from small changes in internal or external wing tank fuel asymmetry
and may require more frequent lateral trim inputs. Lateral trim changes may also be required as flight
conditions change with asymmetric store loadings or if one or more flight control surfaces are slightly
out of rig. Additionally, small sideslip excursions (1 to 3°) are common during steep climbs and
descents, even with symmetric store loadings. These excursions are non-oscillatory in nature and are
controllable with minimal rudder pedal inputs.
In general, flying qualities are also very good with large lateral weight asymmetries. The aircraft
tends to roll toward the heavy wing at elevated g such as during a pull off target during an air-to-ground
attack; away from the heavy wing at negative g. In each case, the roll is easily countered with lateral
stick. Additionally, roll coordination may be slightly degraded with large lateral stick inputs and may
require rudder pedal to maintain balanced flight. At high AOA, the aircraft tends to yaw away from the
heavy wing. Yaw-off should be expected above 25° AOA. Opposite rudder pedal may be required to
maintain controlled flight.
11.1.4 FLIR Carriage Handling Qualities. Flight characteristics with a single ATFLIR or TFLIR pod
produce roll-off in the direction of the pod of up to 12°/second at transonic (Mach 0.90 to Mach 1.05)
Mach numbers. Above Mach 1.05, the roll-off becomes less and eventually reverses direction above
Mach 1.15. Above Mach 0.90, sizable lateral and directional trim settings are required when changing
Mach number, when greater than Mach 0.90, to reduce roll-off and large side forces. Additionally,
lateral inputs are required under elevated load factor in order to maintain the same roll attitude. This
roll-off phenomenon is dominated by the aerodynamic asymmetry, not the lateral weight asymmetry.
Carriage of the following reduces or eliminates the roll-off tendency. During flight test, empty
inboard SUU-79 pylons reduced the aerodynamic asymmetry by 50%. Carriage of symmetric pods
(ATFLIR or TFLIR on station 5 and NAVFLIR on station 7) effectively canceled the aerodynamic
asymmetry. A 480-gallon external fuel tank on the centerline also reduced the aerodynamic asymmetry
to nearly zero. The influence of an ARS was not flight tested but is expected to reduce the roll-off
tendency. Empty SUU-79 pylons on the midboard stations displayed a slight improvement. An
AIM-120 on the opposite fuselage station was ineffective in reducing the roll-off.
The magnitude of the roll-off at peak conditions (Mach 0.95) can be trimmed out, however, slight
variations in Mach require large variations in both lateral and directional trim settings to maintain
balanced flight. This additional pilot workload should be considered during low altitude flight where
mission crosscheck time is critical.
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A1-F18EA-NFM-000
During flight simulation, level bomb deliveries (using the ATFLIR for target identification and
refinement) were flown in a simulated night environment with no outside visual reference. Uncommanded roll-off appeared as a rotating FLIR image similar to what is displayed during over flight of
the designated target. It is possible that uncommanded aircraft bank angle changes, seen as a rotating
FLIR image through the sensor, may be confused with the rotating image that results from target over
flight.
Uncommanded roll-off during heads down sensor operation may result in
an unusual aircraft attitude, disorientation, altitude loss, and possible
CFIT.
11.1.5 5-Wet (4-EFT and ARS) Loading Handling Qualities. The 5-Wet tanker loading handling
characteristics are unique given the very high aircraft gross weight and considerable aerodynamic
influence (in particular, drag) of the external stores.
Ground handling qualities are excellent although higher throttle settings can be expected to initiate
aircraft movement. Additionally, heavy gross weights require noticeable long braking distances even
during routine taxi evolutions.
Throughout most of the 5-Wet envelope, aircraft handling characteristics are excellent. The heavy
gross weight and aerodynamic influence of the external fuel tanks result in a more sluggish aircraft
response to control inputs in all axes. This characteristic is the most noticeable during in-flight
refueling as a receiver. Control inputs during receiver tanking should be small and deliberate. High
frequency, ″last ditch″ inputs should be avoided. Larger throttle inputs are required to initiate aircraft
closure; however, the heavier aircraft weight requires considerable power reductions to arrest closure
rates. In-flight tanking engagements should target 2-3 KCAS closure (see NATOPS Flight Manual
Performance Data, A1-F18EA-NFM-200, Chapter 7, In-flight Refueling). As fuel is received, the
change in gross weight may be dramatic. The 5-wet loading has the largest fuel fraction (ratio of fuel
weight to total weight) of all loadings. As the total aircraft weight increases, the power required for level
flight will increase. When operating at high gross weights (>64,000 lb) and slow 1 g flight conditions
(<200 KCAS), AOA approaches the 15° limit. In this flight regime, the aircraft response may be more
sluggish, but still safe. Taking into account the human factors associated with the refueling task and
ARS operations (e.g. increased head-in-cockpit demands), difficulty of flying near the AOA limit is
compounded. Pay particular attention to the task in this flight regime. Small momentary excursions
above the AOA limit (15°) will not cause departure from controlled flight.
11.1.5.1 5-Wet Loading with 5-Wet FCC Gains not Set. Flying the aircraft with fuel tanks on the
midboard stations causes the mission computer to request that the flight control system use 5-Wet
control law gains to optimize flying qualities. The logic for this determination is absent from
H1E−031U/032U MC OFP and therefore causes the aircraft to always fly without the 5-Wet control
laws even if there are tanks on the midboard stations. For this case, where midboard tanks are installed
and H1E−031U/032U is installed, there may be some looseness in the flaps AUTO longitudinal flying
qualities at certain localized flight conditions, but this is not considered to be a safety of flight issue and
will not interfere with in−flight refueling operations. There are no NATOPs restrictions for this
loading/gain set ″mismatch″ condition. The degradation, if any, may become apparent for CGs aft of
approximately 30% MAC. Flaps AUTO lateral/directional and all flaps HALF or FULL flying
qualities are unaffected by this issue.
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The H2E−MC OFPs have incorporated an FCCGN advisory that is annunciated when the 5-Wet
loading information is declared invalid. The advisory is designed to notify the pilot that 5-Wet FCC
control law gains are no longer being applied. The consequence of the gains changing while still in the
5-Wet loading is identical to the H1E software build discussion above. That is, there may be some
longitudinal looseness while in AUTO flaps, but this will only be noticeable for aft CGs. Within the
NATOPS approved 5-Wet loading airspeed envelope (430 KCAS/0.8M), the degradation in flying
qualities stemming from the mismatched gain set is not considered to be a safety of flight issue.
11.1.6 Dual Midboard with Outboard Stores Handling Qualities.
NOTE
Loadings with dual carriage stores on stations 3 or 9 and an adjacent
store on stations 2 or 10 may experience uncommanded pitch
transients of ±1/2g to ±1g, roll transients of up to 20°/second, and/or
slight PIO at Mach 0.88−0.95 and less than 4° AOA. Roll-offs are more
likely during negative-g maneuvers. Transient severity may worsen at
lower altitude. Transients may increase pilot workload.
11.2 AIR COMBAT MANEUVERING.
11.2.1 Air-to-Air Gun Tracking. The flight characteristics described above combine to provide a very
agile fighter throughout the flight envelope. Excellent pitch pointing capabilities and lift vector
placement allow rapid acquisition and excellent fine tracking of air-to-air targets with little fear of
departure from controlled flight, especially at high AOA. However, the high airspeed bleed rate
associated with continued maneuvering at high AOA requires additional pilot attention to energy
management. Also, the presence of airframe buffet at elevated-g and high subsonic Mach numbers may
induce small pitch and roll transients and require increased workload to maintain precise pipper
placement.
11.2.2 Over-the-Top Maneuvering. The aircraft exhibits excellent slow speed over-the-top maneuverability. Aft stick is required near the top of looping maneuvers to keep the nose tracking until the
nose is below the horizon and airspeed is increasing. If aft stick is not maintained, AOA feedback
results in nose-down stabilator which eventually reduces AOA below 22°. Once below 22° AOA, neutral
longitudinal stick results in an inverted, nose-high attitude with only a small amount of pitch rate as
the FCS attempts to maintain 1g. If airspeed is allowed to decay in this attitude, or is insufficient to
complete the maneuver, a tailslide may result (see Departure Characteristics).
11.2.3 Slow Speed Maneuvering. The excellent controllability and maneuverability at high AOA
provided by FCS control laws result in very precise nose pointing and gun tracking at extremely low
airspeeds. When large and abrupt heading reversals are required during offensive or defensive
maneuvering at high AOA and low airspeed, two features discussed earlier allow the pilot to accelerate
aircraft motion without compromising departure resistance. The first is the enhanced nose-down pitch
rate capability below 200 KCAS which allows very rapid nose-down pitch pointing to acquire the target
at the end of a flat scissors engagement or to rapidly reduce AOA to maximize energy addition. The
second is the ″pirouette″ turning capability at high AOA and low airspeed which allows very rapid and
controllable nose-high to nose-low heading reversals. These two features combine to significantly
enhance maneuverability at high AOA, allowing the pilot to quickly bring the nose to bear on air-to-air
opponents.
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A1-F18EA-NFM-000
11.3 OUT-OF-CONTROL FLIGHT (OCF)
11.3.1 Departure Resistance. A departure is defined by aircraft motion that is contrary to flight
control inputs. Flight test has shown the F/A-18E/F is very resistant to departure from controlled
flight with symmetric loadings. No departure tendencies were found for single-axis control inputs and
for the majority of multi-axis inputs. A few departure tendencies exist with multi-axis inputs, but these
were usually found to occur beyond the 360° bank angle change limitation (two exceptions are
described in 11.3.2.1). Nose high, slow speed maneuvers that result in insufficient maneuvering
airspeed or a tailslide will cause a departure. Overall, the aircraft is very departure resistant when flown
within NATOPS limits. Additionally, clean and multiple store loadings (including aft CG), have shown
no self-sustaining falling leaf mode in the F/A-18E/F. For all known departure modes, following
NATOPS out-of-control (OCF) recovery procedures result in rapid recovery.
11.3.2 Departure Characteristics. The typical F/A-18E/F departure occurs as a yaw divergence
(nose-slice) followed by an uncommanded roll in the same direction. Usually, a departure is preceded
by a buildup in sideforce. This sideforce is often accompanied by ″vortex rumble″ generated from
excessive sideslip. ″Vortex rumble″ may not be noticeable during aggressive maneuvering; therefore,
excessive sideforce provides the most reliable departure warning cue. The initial phase of the departure
is not particularly violent or disorienting unless it occurs at high airspeed or Mach number. The yaw
rate warning tone may not provide sufficient departure warning. Post-departure gyrations self-recover
with controls released. Application of controls during post-departure gyrations may delay recovery.
11.3.2.1 Maneuvering within NATOPS Limits. There are three typical departure cases found for
flight within NATOPS limits.
1. Forward corner inputs below 300 KCAS.
Lateral stick and/or pedal combined with forward inputs at low airspeed may cause a departure prior
to reaching 360° bank angle change limit if AOA transitions from positive to zero or negative during
the roll. The departure is characterized by a dwell at 0 g, followed by a sideslip build-up and
subsequent moderate yaw rate spike (40 to 50 °/sec) and AOA increase. Subtle differences in roll rate
and nose-down pitch rates being generated by such inputs make it difficult to predict whether a
departure or favorable (faster than normal) roll rates will occur.
Lateral stick and/or pedal combined with forward inputs at low airspeed
may cause a departure.
2. Lateral stick with pedal and forward stick at high altitude and supersonic speeds.
This input is predicted to cause a departure that could result in aircraft damage. This departure is
possible at supersonic airspeed above 40,000 ft MSL and within 360° of bank angle change.
3. Tailslides and over-the-top maneuvering with insufficient airspeed.
NATOPS prohibits zero airspeed tailslides and intentional departures or spins. The obvious
consequence of over-the-top maneuvering with insufficient airspeed is departure from controlled
flight. In general, tailslides that are purely vertical induce departures that are benign and quick to
recover. Tailslides with large yaw angle (nose vertical but 3 to 9 line off the horizon) or over-the-top
maneuvering with sizable bank angle and insufficient airspeed usually result in a ″sideslide″ type
motion. The resulting sideslide departure motion is similar to vertical but typically is accompanied
by an abrupt roll snap before the aircraft settles nose low. If the sideslide motion yields excessive
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ORIGINAL
A1-F18EA-NFM-000
gravity-induced sideslip at relatively slow airspeed (<90 KCAS), the resulting yawing motion from
the growing aerodynamic forces can quickly develop into a spin. If the aircraft falls inverted, then the
AOA will simultaneously build negative and may result in an inverted spin. Inverted spins from
sideslides are slower to recover than upright sideslide recoveries.
11.3.2.2 Maneuvering Outside of NATOPS Limits.
NOTE
The following describe the known departure characteristics of the
aircraft if flown outside of the NATOPS limits.
11.3.2.2.1 Exceeding 360° Roll Limit. Certain airspeed and control input combinations held for
greater than the NATOPS bank angle change limit of 360° can lead to departures. Lateral stick with
pedal and forward stick from high g near 300 KCAS may result in a severe departure, with yaw rates
reaching above 100°/sec a possibility. Another severe departure is possible when slowly pulling aft stick
(less than 1 inch/sec) while rolling with a full lateral stick input below 210 KCAS, if initiated near 1g.
This departure can generate yaw rates briefly in excess of 120°/sec and negative g spikes from -2 to -3
g. Each of these departures was found only to occur when controls were held beyond the 360° bank
angle change limit.
11.3.2.2.2 Exceeding Asymmetric Loading AOA Limits. Exceeding NATOPS limits for asymmetric
store loadings can also lead to departures. Aggressive longitudinal maneuvers that result in AOA
beyond NATOPS limits can lead to a benign departure that begins as a slow roll toward the heavy wing
and yaw away from the heavy wing that cannot be controlled with lateral stick or rudder pedal.
Recovery is immediate as soon as AOA is reduced to within limits. Large sideslips create a greater risk
of a more violent departure. At higher speeds, aggressive maneuvering at elevated-g above AOA limits
can result in sudden departures with little or no warning. If limits are exceeded and a departure does
occur, post departure gyrations rapidly transition to an upright spin away from the heavy wing.
Recovery from this type of departure has been demonstrated with up to a 24,000 ft-lb lateral weight
asymmetry following NATOPS OCF recovery procedures (see Spin Characteristics).
11.3.2.3 Maneuvering with Flight Control System Failures. Continued maneuvering with flight
control system failures such as surfaces failed off (X’s in all channels of that actuator), air data, or other
sensor failures can also lead to departure. The flight control system is designed to provide adequate
flying qualities with actuator and other FCS failures as long as AOA and load factor are maintained
within NATOPS limits. In the event of FCS failures, reducing AOA and load factor to wings level 1g
flight as soon as possible minimizes the possibility of a departure. If a departure does occur, following
OCF procedures results in the most rapid recovery.
11.3.3 Spin Characteristics. Entry into a spin is rare for symmetrically loaded F/A-18E/F. On a few
occasions, moderate yaw rate spins have developed from departures experiencing large gravity-induced
sideslip excursions (e.g. sideslips or nose-high slow airspeed flight while banked near 90°). If AOA is
simultaneously negative, the spin will be inverted.
For high lateral weight asymmetry loadings, the aircraft is extremely departure resistant within
NATOPS AOA limits. Exceeding AOA limits for high lateral weight asymmetries will most likely result
in an upright spin away from the heavy wing. Spin recoveries for less common upright spins into the
heavy wing, and inverted spins, can be delayed due to the oscillatory nature of the spin. Spin recovery
has been extensively proven for lateral asymmetries up to 14,000 ft-lb for both spins into and away
from the heavy wing with positive recoveries demonstrated in all cases.
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ORIGINAL
A1-F18EA-NFM-000
Spin characteristics and recovery with asymmetries greater than 14,000 ft-lb have not been fully
tested. A departure with 24,000 ft-lb of asymmetry resulted in a high yaw rate spin (greater than
100°/sec) within 3 to 5 seconds. Recovery occurred within two to three turns and about 10,000 feet of
altitude loss after applying NATOPS OCF recovery procedures. High yaw rate spins typically result in
longitudinal accelerations at the pilot seat as high as -3.5g (eyeballs out). Consequently, accomplishing
spin recovery procedures can be difficult with an unlocked seat harness.
Spin recovery is straightforward and reliable if the OCF procedures are followed and sufficient
altitude remains. If a spin is encountered (Spin recovery display on the DDI), recovery occurs very
shortly after initiating NATOPS OCF recovery procedures, particularly from inverted spins. Recovery
from spins with high lateral weight asymmetry may require an additional turn or two.
Selection of manual spin recovery mode (SPIN switch in RCVY) seriously degrades controllability, prevents recovery from any departure or
spin, and is prohibited.
NOTE
During highly oscillatory spins or spins that transform from upright to
inverted or from inverted to upright, the spin recovery display may
disappear momentarily.
11.4 DEGRADED MODE HANDLING QUALITIES.
The reliability of the FCS is very high and when failures do occur, usually occur singly. No single
electrical failure affects flying qualities and multiple FCS failures are required to degrade flying
qualities. Depending on which combination of failures has occurred, flying qualities may be considerably degraded. Degraded flying qualities associated with some of the more serious or more common
FCS failures are described here. Appropriate corrective action is presented in the Warning/Caution/
Advisory Displays, figure 12-1.
11.4.1 Single Engine Operation.
11.4.1.1 Flaps AUTO. Engine failure or shutdown with flaps AUTO results in no degradation in
handling qualities under most circumstances at low AOA. A small amount of yaw trim may be required
to counter asymmetric thrust effects. At high AOA, engine failure results in a yaw toward the failed
engine that is controllable by quickly reducing AOA and countering the yaw with rudder. During hard
maneuvering, a slight degradation in handling qualities may be noticeable at less than Mach 1.0
between approximately 400 to 500 KCAS where the hydraulic system normally operates at 5,000 psi.
At these conditions 5,000 psi operation is inhibited by the FCS to maintain a windmill air-start
capability. When 5,000 psi operation is inhibited, flying qualities may be degraded during aggressive
maneuvers since there may not be enough hydraulic power to fully deflect numerous flight control
surfaces. A reduction in departure resistance can also be expected anytime normal 5,000 psi hydraulic
system operation is inhibited.
11.4.1.2 Flaps HALF or FULL. Single engine minimum control speed (Vmc) is defined as the
minimum airspeed required to maintain controlled flight with one engine operating. Vmc airspeeds
were determined at 14° AOA for catapult launches, and 12° AOA for all other circumstances. Vmc
airspeed varies depending on AOA, lateral asymmetry, altitude, and day temperature. For an engine
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ORIGINAL
A1-F18EA-NFM-000
failure off the catapult, 14° AOA provides the best compromise between arresting rate of descent off
the bow and controllability (increased AOA helps arrest sink but also reduces lateral-directional
controllability). In all cases, once control is established (and sink is stopped), allow the aircraft to
accelerate to on-speed to provide the best flyaway handling qualities. When an engine fails in flaps
HALF or FULL, the first perceptible aircraft motion is a yaw toward the failed engine. The rudders are
the primary flight control surface used to counter the yaw caused by the operating engine. Using too
much rudder pedal is not harmful, but using too little rudder pedal may not counter the yaw and cause
controllability problems. In addition to yawing into the failed engine, the aircraft also tends to roll into
the failed engine. The natural pilot reaction is to oppose the roll with lateral stick, but the resulting
differential aileron deflection generates adverse yaw and increases the demand on the rudders to
maintain directional control. As AOA increases above 10° AOA, the aircraft becomes less directionally
stable and rudder control effectiveness deteriorates. In this instance, the rudders may become
saturated (surfaces against the stops). When saturated, the rudders cannot counter any additional
adverse yaw, resulting in an increase in sideslip and the potential for an adverse yaw departure. If
airspeed is too slow, the rudders cannot generate enough control power to oppose the yaw toward the
failed engine.
For some situations, controllability alone will not guarantee flyaway (e.g., excessive rate of descent),
but may only ensure controllability for a long enough period of time to complete the requisite
immediate action procedures and make a timely ejection decision, if warranted.
When single engine with the operating engine at MAX, the possibility of
an adverse yaw departure increases as AOA exceeds on-speed.
NOTE
• In straight and level flight, a small amount of lateral and/or directional
trim is required to maintain balanced flight.
• Loss of either HYD 1 or HYD 2 due to engine failure or hydraulic
pump failure does not effect flight control with flaps AUTO; however,
failure of either HYD 1 or HYD 2 with flaps in HALF or FULL may
cause uncommanded but controllable yaw and roll transients as the
switching valves cycle. These yaw and roll transients may last 3 to 6
seconds.
• To prevent repeated switching valve cycling, avoid stabilized flight
where engine windmill rpm results in hydraulic pressure fluctuations
between 800 and 2,000 psi.
11.4.1.3 Single Engine Waveoff. Refer to Chapter 16, paragraphs 16.1 and 16.3.
11.4.2 Leading Edge Flap Asymmetry. Leading edge flap asymmetries can occur when one of the
LEF hydraulic drive units (HDU) stalls/fails or the mechanical interconnect between the inboard and
outboard LEF surfaces fails. The most common LEF asymmetry results from a weak LEF HDU that
stalls (stops moving due to aerodynamic loading) during abrupt longitudinal maneuvers at high
airspeed and low altitude. When this happens, a roll-off away from the failing HDU as AOA or g is
increased followed by an abrupt roll-off in the opposite direction is typical. Failure detection logic in
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ORIGINAL
A1-F18EA-NFM-000
the FCS software is designed to provide advanced warning of a LEF asymmetry; however, during
extremely abrupt maneuvers, an HDU stall may not be detected in time to allow the pilot to abandon
the maneuver and avoid a large roll transient. The FCS software uses two specific monitors to detect
weak LEF HDUs. The first monitor results in a FLAP SCHED caution accompanied with BLIN 256.
It sets and holds the FLAP SCHED caution for 6 seconds following the HDU stall. The second monitor
results in a FLAP SCHED caution accompanied by a BLIN 537. When accompanied by a BLIN 537,
the FLAP SCHED caution is displayed only as long as the failing HDU condition exists (e.g., when
AOA >12°).
Avoid high-g maneuvers at low altitude if an HDU stall has been detected during the flight (BLIN
256 or 537 on the FCS status page) even if the FLAP SCHED caution has cleared. BLINs 256 and 537
indicate a weak/failing HDU that may result in very large roll transients and/or over-g during high-g
maneuvers.
11.4.2.1 LEF Failure Landing Handling Qualities (Symmetric or Asymmetric). With a LEF failure,
the LEF symmetric commands are frozen when the FLAP switch is set to AUTO; however, differential
LEF and TEF continue to be commanded. With the FLAP switch set to HALF or FULL, the LEF
commands are frozen while the TEF and aileron droop commands follow the normal schedules for PA.
Shore and ship based flight tests were conducted with LEF frozen at both symmetric and asymmetric
(up to 34°/5° LEF split) deflections. Straight in, on-speed approaches in flaps HALF are recommended
for all frozen symmetric or asymmetric LEF configurations. General flying qualities, as well as waveoff
and T&G performance are acceptable for all LEF configurations.
Light to moderate buffet is present in most of the LEF configurations, especially for AOA greater
than on-speed. The buffet is more pronounced with LEF deflections significantly less than the normal
scheduled positions. Selecting flaps FULL with small LEF deflections results in higher buffet levels at
lower AOAs than with the FLAP switch set to HALF. Noticeable buffet is normal near on-speed
conditions with either HALF or FULL flaps for the degraded LEF condition, but may be uncomfortable during maneuvering. Maintaining on-speed or slightly fast approach AOA (to minimize buffet)
results in the best flying qualities for any off-schedule symmetric or asymmetric (left/right) LEF
configuration. Where practical, flying slightly fast (6-7° AOA) minimizes exposure to the buffet. For
carrier landings, the tendency to fly the approach in a ″fast″ condition should be avoided because the
recovery WOD requirements are based on the on-speed approach airspeed. Flying a ″fast″ approach
may result in aircraft and/or arresting gear overstress upon arrestment.
Small (1-3°) roll-off due to buffet can be expected but is easily controllable with small lateral inputs.
Roll and line up control are not significantly different than normal scheduled positions. The aircraft
rolls faster into the lesser-deflected LEF and slower when rolling away from the lesser-deflected LEF.
Roll-off may also be experienced with AOA/pitch attitude changes. As AOA decreases, the aircraft rolls
away from the lesser-deflected LEF and as AOA increases, the aircraft rolls into the lesser-deflected
LEF. Lateral stick and/or trim easily counters this roll-off tendency. A small roll off will occur with
airspeed changes during a waveoff or bolter but are easily controlled with small lateral stick inputs.
Glideslope control degradations are more pronounced with LEF deflections significantly less than
the normal scheduled positions. The addition of wing stores will further degrade glideslope control.
The aircraft response to power corrections is sluggish and smaller in magnitude than normal flap
configurations. Therefore, compared to normal flap configurations, larger, longer and more anticipatory power corrections are required to effect a glideslope change. Power corrections translate to an
airspeed change first before a rate of descent change is noticed. The delay in aircraft response coupled
with the larger throttle inputs leads to the tendency to over−control the glideslope. The technique of
applying power and waiting for a glideslope change will lead to larger glideslope deviations.
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ORIGINAL
A1-F18EA-NFM-000
Anticipatory throttle inputs are key to controlling glideslope. The sluggish power response also
degrades waveoff performance slightly. When operating shipboard, the waveoff window for all LEF
failure conditions should be moved farther out and the LSOs should be made aware of the degraded
glideslope performance. Upon bolter/waveoff, climb−out attitude will appear flatter than normal. The
LSOs and PRIFLY personnel should monitor rate of climb as the primary indication of bolter/waveoff
performance. It is recommended aircrew perform a familiarization check of aircraft response to control
and throttle inputs, including a waveoff maneuver, prior to attempting a shipboard landing.
11.4.3 Trailing Edge Flap Failure. TEF failures may be caused by an actuator failure (mechanical or
three-four channel failure) or by a dual HYD 1A/2B circuit failure. TEF actuators continue to operate
following two channel failures. If a TEF actuator is shutdown, the surface is hydraulically or
aerodynamically driven to 5° TED and locked. If the left to right TEF asymmetry exceeds 6°, the
opposite TEF fails off and is also driven to 5° TED and locked.
11.4.3.1 TEF Failure Landing Handling Qualities. Shore and ship based flight tests have been
conducted with the TEF in the failed position of 5° TED. Straight in, 10° AOA approaches in flaps
HALF or FULL are recommended for TEF failures. Pitch, roll, and line up control are similar to those
for normal approaches. Approach speeds will be high and every attempt should be made to reduce gross
weight. If shore based, consideration should be given to making an arrested landing taking into account
maximum arresting gear engagement speed and nose tire limit.
With this failure, approach drag is reduced and approach power settings are less than normal. This
results in slower engine response to throttle changes compared to normal HALF or FULL flap
approaches. Flying a slightly slow approach (10° AOA) reduces approach speed and increases approach
power setting slightly. Flying qualities at 10° AOA are similar or improved over those observed at
on-speed AOA. The aircraft easily trims to and maintains 10°AOA. Glideslope maintenance will
dominate the approach task and the tendency is to relax AOA maintenance. The 10° AOA approach
reduces WOD requirements by approximately 12 to 16 KCAS and improves approach flying qualities.
When operating at the ship, the recovery WOD should be kept as close as possible to the Aircraft
Recovery Bulletin recommendations. Due to the large difference between WOD requirements at
on−speed and 10° AOA, it is imperative that AOA be maintained at 10°. The sight picture behind the
ship is altered, but the field of view over the nose is not degraded.
Slower engine response to throttle changes may result in excessive sink
rates under high WOD conditions. The recovery WOD should be kept as
close as possible to the Aircraft Recovery Bulletin recommendations.
The aircraft response to power corrections is sluggish and smaller in magnitude than for normal TEF
configurations. TEF failures require larger, longer, and more anticipatory power corrections to effect
a glideslope change. There is a tendency to over control the power due to the low approach power
setting and the longer time to effect a change. Power corrections translate to an airspeed change first
before a rate of descent change is noticed. Aggressive, well−timed, and anticipatory throttle inputs are
required for glideslope control. The technique of applying power and waiting for a glideslope change
will lead to larger glideslope deviations. Waveoff performance is degraded for TEF failure approaches.
The waveoff technique is the same as normal flap configurations; apply power, maintain AOA until
positive rate of climb is achieved, and capture 10° pitch attitude for the climb−out. Time to achieve
positive rate of climb is slower than for normal TEF configurations. At the ship, the waveoff window
for a TEF failure condition should be moved farther out. The LSOs should be aware of degraded
IV-11-13
ORIGINAL
A1-F18EA-NFM-000
waveoff and glideslope performance. Upon bolter/waveoff, climb−out attitude will appear flatter than
normal. The LSOs and PRIFLY personnel should monitor rate of climb as the primary indication of
bolter/waveoff performance. It is recommended aircrew perform a familiarization check of aircraft
response to control and throttle inputs, including a waveoff maneuver, prior to attempting a shipboard
landing.
11.4.4 Stabilator Failure. Since there is no mechanical reversion mode of the flight controls, the FCS
control laws automatically reconfigure in each axis to allow continued flight in the event of a single
stabilator failure. This failure mode, known as stabilator reconfiguration, or STAB RECON, is
designed to compensate for the loss of the contribution of the failed stabilator to pitch and roll control.
This is accomplished by disabling differential stabilator commands and using the other rolling surfaces
to counter the roll and yaw moments produced when the remaining stabilator responds to pitch axis
commands. Flight tests demonstrated excellent handling qualities with a stabilator failed off during
maneuvering flight and aerial refueling. The pitch axis is slightly sluggish, maximum roll rate is
noticeably lower, and roll coordination is slightly degraded.
In UA, with a failed stabilator, maximum roll rate is extremely low in the
transonic region below 20,000 feet, especially when rolling away from the
failed side. Significant roll and yaw coupling may occur with forward stick
inputs at Mach >1.4 and altitude >30,000 feet.
The degradation in roll coordination is characterized as a small amount of sideforce during lateral
inputs with flaps AUTO and noticeable sideslip with flaps HALF. Configuration changes exhibit a
slight roll toward the failed stabilator when transitioning from flaps AUTO to flaps HALF; away from
the failed stabilator when transitioning from flaps HALF to flaps AUTO. Flight tests demonstrated
that flight with a stabilator failed in flaps HALF was nearly transparent to the pilot from a handling
qualities perspective during normal approaches, waveoffs, and bolters or flared landings. A small but
controllable yaw away from the failed stabilator is apparent at touchdown. Additionally, pitch stick
inputs during the landing roll complicate directional control and should be avoided.
Extending the speedbrake may produce uncommanded roll into the failed stabilator. The uncommanded roll is more pronounced at low speed flight conditions due to the reduced aileron effectiveness
at the large trailing edge up aileron deflections commanded by the speedbrake function. The roll can
be balanced with lateral stick deflection.
In UA, with a failed stabilator, nose down pitching moment capability is degraded at low to moderate
airspeeds, especially for aft center of gravity and heavy wing store loadings. A safe level of nose down
pitch capability is available for flight below 10° angle of attack.
In UA, with a failed stabilator, do not exceed 10° AOA due to reduced
nose down pitch authority.
Carrier based flight tests (STAB RECON/flaps HALF) demonstrated that carrier approaches were
easily controlled but were characterized by slightly degraded flying qualities which required increased
pilot attention to the landing task. Slight rolling and/or yawing motions were apparent when making
longitudinal inputs. Multiple small lateral inputs were required to maintain a centered approach.
IV-11-14
ORIGINAL
A1-F18EA-NFM-000
During bolters, the aircraft yawed into the good stabilator when the flight control system deflected the
good stabilator trailing edge up (TEU) in preparation for aircraft nose down rotation. The yaw was
sudden and pronounced, but easily controlled with rudder to counter the yawing motion.
Flight with flaps FULL has not been demonstrated due to limited nose-up control authority from
the remaining good stabilator.
• Do not select flaps FULL for landing with a failed stabilator, because
longitudinal control authority may be insufficient for landing.
• With a failed stabilator, do not exceed 10° AOA in AUTO flaps with
wing stores or wing tanks.
11.4.5 GAIN ORIDE. While not a failure mode, GAIN ORIDE is prescribed for certain AOA or
pitot-static sensor failures to provide better or more predictable handling qualities (see Warning/
Caution/Advisory Displays, figure 12-1). With flaps AUTO, selecting GAIN ORIDE results in fixed
gains that correspond to Mach 0.80, 39,000 feet, and 250 KCAS. At flight conditions that deviate from
the fixed gains, a slight degradation in handling qualities should be expected. The aircraft is less
sensitive to longitudinal inputs, as less pitch rate is generated per given stick input. Lateral stick inputs
provide similar responses as in GAIN NORM. In GAIN ORIDE, the aircraft is more sensitive to
directional inputs. Regardless, handling qualities remain very good within the 10° AOA and 350 KCAS
NATOPS limits for GAIN ORIDE operation. If the airspeed limit is exceeded, self-sustained pitch
oscillations will start to occur above 375 KCAS, and the aircraft will become uncontrollable above 450
KCAS due to the fixed air data values in the flight control system gains. If the AOA limit is exceeded,
departures are likely since the fixed values of the air data and AOA severely reduce departure
resistance. Additionally, the aircraft will stall at a higher than normal airspeed due to the fixed position
of the LEFs.
With flaps HALF or FULL, GAIN ORIDE results in fixed gains that correspond to 8.1° AOA and
500 feet; handling qualities are best at these conditions and degrade slightly away from on-speed AOA.
At higher airspeeds in 1g flight, the aircraft will stabilize at lower than normal AOA, due to the TEF
position frozen at higher than normal deflections. While not dangerous, this characteristic is
uncomfortable. Also, at higher airspeeds, higher than normal aft stick force is required to maintain
flight path while in a turn. Flight is prohibited above 190 KCAS (flaps FULL) or 200 KCAS (flaps
HALF) due to these characteristics. Flight is also prohibited above 10° AOA due to the reduced stall
margin available with fixed LEF deflections.
Transition to or from landing configuration should be done in level flight at 180 KCAS. Transition
should not be made while in a bank due to the higher than normal aft stick forces required to maintain
flight path angle. Sideslip excursions may also occur if flap transition is made in a turn.
Carrier based flight tests (GAIN ORIDE/flaps HALF) demonstrated satisfactory approach handling
qualities. The aircraft remained easily controllable, though increased pilot attention to the landing task
IV-11-15
ORIGINAL
A1-F18EA-NFM-000
was required. During descent, small longitudinal inputs were required to maintain proper AOA and
pitch attitude. Approaches flown at conditions other than on-speed resulted in sluggish longitudinal
handling qualities.
Bolters in GAIN ORIDE or with AOA failed require positive aft stick
during rotation, 1/2 aft stick is recommended. Deflections of less than 1/2
aft stick will result in excess settle during bolters.
In GAIN ORIDE, AOA will tend to readily increase above 14° when
decelerating from a trimmed on-speed condition. Timely longitudinal
stick inputs will be required to prevent excessive sink rates and correct a
deceleration as power alone will not change the AOA or pitch attitude
sufficiently in GAIN ORIDE. Alpha tone is disabled in GAIN ORIDE
with FLAPS HALF or FULL.
11.4.6 AHRS Failure Flying Qualities. An AHRS channel failure is defined as the loss of both rate
and acceleration data (Xs in CAS P, CAS R, CAS Y, N ACC, and L ACC). Single or dual channel AHRS
failures should have no adverse effect on flying qualities. If a third channel failure is detected, all four
channels will be Xd out because the FCCs will be unable to confirm which channel is providing valid
data. If the third failure can be isolated to a particular channel, all four channels will be Xd out but
the flying qualities will be unaffected with the exception of a small degradation to the g-limiter. If a
third failure occurs but is not detected (two columns of Xs), or is detected but not isolated to a
particular channel (four columns of Xs), flying qualities will be somewhat degraded. When flying
qualities are degraded due to AHRS channel failures, poor roll coordination for large lateral inputs,
pitch coupling, and/or sluggish pitch response can be expected. Due to the higher reliability of AHRS
over previous rate and acceleration sensors, a four channel failure is highly unlikely. However,
simulator evaluation has shown that a complete four channel AHRS failure (no rate and acceleration
inputs to FCCs) is controllable for most of the flight envelope with the flaps in AUTO. Refer to figure
11-1 for AHRS channel failure indications and effects.
AHRS failure modes have not been flight tested. With a four channel
AHRS failure, the aircraft is not controllable with the flaps in HALF or
FULL. At altitudes above 25,000 feet, loss of control occurs below Mach
0.92. For loss of AHRS above 25,000 feet, maintain airspeed above Mach
0.92 while descending.
With a four channel AHRS failure at altitudes below 20,000 feet, flying qualities are optimum
between 190 to 210 KCAS, and are acceptable above 370 KCAS. When decelerating below 370 KCAS,
flying qualities are degraded until below 270 KCAS. The worst flying qualities are between 270 to 370
KCAS. Flying qualities do improve above 370 KCAS, with a further improvement at supersonic speeds.
Execute a straight-in on-speed approach with flaps in AUTO, and limit angle of bank to 20°. Pitch and
directional damping will be very low and roll coordination weak. If not positioned for landing by in the
IV-11-16
ORIGINAL
A1-F18EA-NFM-000
middle to in close, a wave-off and go-around should be executed. A clean, lightly loaded aircraft
exhibits the best flying qualities. Flying qualities are not affected by CG locations, and the aircraft can
be landed to the aft CG limit. An arrested landing should be made if airspeed permits. Avoid stabilator
braking.
Level of AHRS
Channel Failure
Indications
Effects
One Channel Failure
• Xs in one column (CAS P, CAS
R, CAS Y, N ACC, and L ACC)
• DEGD X
Two Channel (incremental failure)
• Xs in two columns (CAS P, CAS
R, CAS Y, N ACC, and L ACC)
• DEGD X
Two channel (simultaneous failure caused by loss of communication between the two AHRS
channels and the FCCs)
• Xs in two columns (CAS P, CAS
R, CAS Y, N ACC, and L ACC)
• DEGD X
Three channel (detected and isolated)
• Xs in four columns (CAS P,
CAS R, CAS Y, N ACC, and L
ACC)
• DEGD X
• PCAS, RCAS, and YCAS cautions
Small degradation to the g-limiter
Three channel (detected but not
isolated)
• Xs in four columns (CAS P,
CAS R, CAS Y, N ACC, and L
ACC)
• DEGD X
• PCAS, RCAS, and YCAS cautions
• Poor roll coordination for large
lateral inputs
• Sluggish pitch response
• Pitch coupling
Three channel (not detected)
• Xs in two columns (CAS P, CAS
R, CAS Y, N ACC, and L ACC)
• DEGD X
Four channel
• Xs in four columns (CAS P,
CAS R, CAS Y, N ACC, and L
ACC)
• DEGD X
• PCAS, RCAS, and YCAS cautions
No effects on flying qualities.
• Uncontrollable in flaps HALF
or FULL
• Poor roll coordination for large
lateral inputs
• Sluggish pitch response
• Pitch coupling
Figure 11-1. AHRS Channel Failure Indication and Effects
IV-11-17 (Reverse Blank)
ORIGINAL
A1-F18EA-NFM-000
PART V
EMERGENCY PROCEDURES
Chapter 12 - General Emergencies
Chapter 13 - Ground Emergencies
Chapter 14 - Takeoff Emergencies
Chapter 15 - Inflight Emergencies
Chapter 16 - Landing Emergencies
Chapter 17 - Ejection
Chapter 18 - Immediate Action
67 (Reverse Blank)
ORIGINAL
A1-F18EA-NFM-000
EMERGENCY INDEX
Conference X-ray telephone number (Inflight emergencies only)
314-232-9999 and 866-543-5444
CHAPTERS 12 THRU 17
Page
No.
A
ABORT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-14-2
AFTERBURNER FAILURE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-1
AILERON HINGE FAILURE - SUSPECTED, INBOARD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-47
ANGLE OF ATTACK (AOA) FAILURE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-44
AOA PROBE DAMAGE OR BINDING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-45
AOA PROBE SELECTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-46
ARRESTMENT - FIELD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-13
Arresting Gear Types. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-13
Arrestment Decision. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-14
Arrestment - Short Field. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-14
Arrestment - Long Field. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-14
ARS MALFUNCTIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-25
ARS Hose Fails to Retract. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-25
ARS Refueling Hose Jettison.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-25
ARS Hydraulic Pressure Light. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-25
No RDY Light. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-26
B
BARRICADE ARRESTMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-14
BOLTER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-3
BRAKE FAILURE/EMERGENCY BRAKES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-13-4
C
COCKPIT SMOKE, FUMES, OR FIRE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-16
COCKPIT TEMPERATURE HIGH. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-15
CONTROLLABILITY CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-22
CV RECOVERY MATRIX . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-18
D
DISPLAY MALFUNCTION - AMCD AIRCRAFT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-19
DITCHING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-17-4
DOUBLE TRANSFORMER-RECTIFIER FAILURE. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-10
DUAL AOA FAILURE ON TAKEOFF . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-47
DUAL MISSION COMPUTER (MC) FAILURE - AMCD AIRCRAFT . . . . . . . . . . . . . . . . . . . . . V-15-19
Emergency Index-1
ORIGINAL
A1-F18EA-NFM-000
Page
No.
E
EJECTION. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-17-1
Ejection Procedures. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-17-3
Ejection Seat Restrictions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-17-1
Airspeed during Ejection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-17-2
Injury Risks - Nude Weight Greater than 213 lb. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-17-1
Injury Risks - Nude Weight Less than 136 lb.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-17-1
High Altitude Ejection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-17-3
Low Altitude Ejection.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-17-2
EMERGENCY CATAPULT FLYAWAY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-14-1
EMERGENCY TANKER DISENGAGEMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-26
EMERGENCY EGRESS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-13-3
ENGINE FAILS TO START/HUNG START . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-13-1
EXTERNAL STORES JETTISON. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-23
F
FCS FAILURE INDICATIONS AND EFFECTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-27
FORCED LANDING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-4
FUSELAGE FUEL LEAK. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-4
G
GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-12-1
Immediate Action Items.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-12-1
Warnings, Cautions, and Advisories. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-12-1
GO AROUND . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-14-3
GROUND FIRE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-13-2
H
HOT BRAKES/BRAKE FIRE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-13-2
HOT START. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-13-1
HUNG START . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-13-1
HYDRAULIC FAILURES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-5
HYPOXIA/LOW MASK FLOW/NO MASK FLOW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-17
L
LANDING GEAR EMERGENCY EXTENSION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-5
LANDING GEAR FAILS TO RETRACT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-14-6
LANDING GEAR UNSAFE/FAILS TO EXTEND . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-4
LOSS OF CABIN PRESSURIZATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-18
LOSS OF DC ESSENTIAL BUS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-13-1
LOSS OF DIRECTIONAL CONTROL DURING TAKEOFF OR LANDING
(BLOWN TIRE, NWS FAILURE ) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-14-4
Emergency Index-2
ORIGINAL
A1-F18EA-NFM-000
Page
No.
O
OUT-OF-CONTROL FLIGHT (OCF) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-19
Departure from Controlled Flight. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-19
Spin. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-20
OCF Recovery Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-20
Post Departure Dive Recovery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-21
P
PLANING LINK FAILURE. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-12
R
RESTART . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-15-1
S
SEAWATER ENTRY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-17-4
SINGLE ENGINE APPROACH AND LANDING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-1
SINGLE ENGINE FAILURE IN LANDING CONFIGURATION . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-1
SINGLE ENGINE WAVEOFF/BOLTER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V-16-3
Emergency Index-3
ORIGINAL
A1-F18EA-NFM-000
WARN/CAUT/FCS/HYD/ADVIS
Warn/Caut/FCS/Hyd/Advis
Page No
Warn/Caut/FCS/Hyd/Advis
Page No
A/A ..................................................
A/G ..................................................
ABLIM ..............................................
ACI ...................................................
AHMD ...............................................
AIR DATA .........................................
ALGN Xd ..........................................
AM DL ..............................................
AMAD & PR L/R ...............................
ANTI ICE L/R ....................................
ANTISKID .........................................
AOA ..................................................
AOA TONE ........................................
APU ACCUM .....................................
APU FIRE ..........................................
ARS DROGUE ....................................
ATC FAIL ...........................................
ATS L/R ...........................................
AUTO PILOT .....................................
AV AIR HOT ......................................
BALT ................................................
BATT SW ..........................................
BAY DISCH ........................................
BAY FIRE ...........................................
BINGO ...............................................
BIT ...................................................
BLD OFF L/R ....................................
BLEED DUAL ....................................
BLEED Single ...................................
BOOST LO L/R .................................
BRK ACCUM .....................................
CABIN ..............................................
CANOPY ...........................................
CAS P/R/Y........................................
CAUT DEGD ......................................
CFIT Xd ............................................
CHECK SEAT ....................................
CHECK TRIM ....................................
CK ECS .............................................
CK FLAPS .........................................
CNI ...................................................
COM1H/2H ......................................
COM1L/2L .......................................
COM1S/2S .......................................
CDATA ..............................................
CONFG .............................................
CONSNT ...........................................
CPLD ................................................
CRUIS ...............................................
CW ....................................................
D BAD ..............................................
D LOW ..............................................
DC FAIL L/R......................................
DCSCS ..............................................
DECM ...............................................
DCOY ON ..........................................
DEPLOY .............................................
DEVC BLD L/R ..................................
DFIR OVRHT ......................................
DFIRS GONE......................................
DISCH................................................
DPLY ................................................
DUMP OPEN ....................................
EBC ...................................................
ECSDR...............................................
ECS ICING ........................................
V-12-71
V-12-71
V-12-64
V-12-64
V-12-64
V-12-7
V-12-64
V-12-64
V-12-7
V-12-7
V-12-7
V-12-38
V-12-8
V-12-8
V-12-3
V-12-8
V-12-39
V-12-9
V-12-9
V-12-10
V-12-64
V-12-10
V-12-10
V-12-11
V-12-11
V-12-64
V-12-12
V-12-3
V-12-4
V-12-13
V-12-13
V-12-13
V-12-14
V-12-40
V-12-14
V-12-64
V-12-14
V-12-14
V-12-14
V-12-14
V-12-15
V-12-64
V-12-64
V-12-64
V-12-64
V-12-65
V-12-71
V-12-65
V-12-65
V-12-71
V-12-65
V-12-65
V-12-15
V-12-65
V-12-65
V-12-71
V-12-15
V-12-15
V-12-15
V-12-15
V-12-71
V-12-71
V-12-16
V-12-65
V-12-65
V-12-16
EGT HIGH L/R ..................................
ENG L/R ...........................................
ENG VIB L/R ....................................
ERASE FAIL ......................................
EXT TANK ........................................
EXT XFER .........................................
FADEC ...............................................
FADEC HOT .......................................
FC AIR DAT ......................................
FCCGN .............................................
FCES ................................................
V-12-17
V-12-17
V-12-17
V-12-17
V-12-18
V-12-18
V-12-66
V-12-19
V-12-41
V-12-66
V-12-42
-V-12-47
V-12-42
V-12-35
FCS HOT ..........................................
FCS Initial ........................................
FCS AHRS 1/2
Ch Fail.......................................
FCS AHRS 4 Ch Fail ......................
FCS Ail Fail....................................
FCS AOA Fail .................................
FCS AOA 4 Ch Fail ........................
FCS Singl Ch Fail ..........................
FCS Rud Fail .................................
FCS Stab Fail.................................
FIRE .................................................
FLAMEOUT L/R ................................
FLAP SCHED ....................................
FLAPS (Amber) .................................
FLAPS OFF LEF ................................
FLAPS OFF TEF ................................
FLIR OVRHT ......................................
FPAH ................................................
FPAS ................................................
F-QTY ...............................................
FUEL HOT L/R .................................
FUEL INLT L/R..................................
FUEL LO ...........................................
FUEL XFER .......................................
FULL .................................................
GEAR HANDLE .................................
GEN L/R ..........................................
GEN TIE ...........................................
G-LIM 7.5G .......................................
G-LIM OVRD .....................................
GPS ...................................................
GPS DEGD .........................................
GPSMP ..............................................
GSEL ................................................
GTRK ................................................
GUN GAS ..........................................
HALF ................................................
HAND CNTRL ....................................
HDG ..................................................
HEAT L/R .........................................
HEAT Xd............................................
HEAT FAIL .........................................
HIAOA ...............................................
HMD ..................................................
HO