TM-55-1510-219-10

TM-55-1510-219-10
TM 55-1510-219-10
TECHNICAL MANUAL
OPERATOR'S MANUAL
FOR
ARMY RC-12D AIRCRAFT
This copy is a reprint which includes current
pages from Change 1.
WARNING DATA
TABLE OF CONTENTS
INTRODUCTION
DESCRIPTION AND
OPERATON
AVIONICS
MISSION EQUIPMENT
OPERATING LIMITS AND
RESTRICTIONS
WEIGHT/BALANCE AND
LOADING
PERFORMANCE DATA
NORMAL PROCEDURES
EMERGENCY PROCEDURES
REFERENCES
ABBREVIATIONS AND TERMS
ALPHABETICAL INDEX
"Approved for public release; distribution is unlimited."
*This TM 55-1510-219-10 supersedes TM 55-1510-219-10
dated 25 May 1985 including all changes.
HEADQUARTERS, DEPARTMENT
OF THE ARMY
31 May 1991
URGENT
TM 55-1510-219-10
C3
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 12 May 1998
CHANGE
NO. 3
OPERATOR’S MANUAL
FOR
ARMY RC-12D AIRCRAFT
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited
TM 55-1510-219-10, 31 May 1991, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by vertical
bar in the margin. An illustration change is indicated by a miniature pointing hand.
2.
Remove pages
Insert pages
i and ii
58.1/(5-8.2 blank)
5-9 and 5-10
8-25 and 8-26
8-26.1/(8-26.2 blank)
------9-3 through 9-10
------INDEX-3 and INDEX-4
i and ii
5-8.1 and 5-8.2
5-9 and 5-10
8-25 and 8-26
8-26.1 and 8-26.2
(9-2.1 blank)/9-2.2
9-3 through 9-10
9-10.1/(9-10.2 blank)
INDEX-3 and INDEX-4
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
DENNIS J. REIMER
General, United States Army
Chief of Staff
Administrative Assistant to the
Secretary of the Army
DISTRIBUTION:
To be distributed in accordance with Initial Distribution Number (IDN) 310121, requirements for TM 551510-219-10.
URGENT
TM 55-1510-219-10
C2
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 12 December 1994
NO. 2
Operator's Manual
For
ARMY RC-12D AIRCRAFT
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.
TM 55-1510-219-10, 31 May 1991, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
5-9 and 5-10
Insert pages
5-9 and 5-10
2. Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
GORDON R. SULLIVAN
General, United States Army
Chief of Staff
MILTON H. HAMILTON
Administrative Assistant to the
Secretary of the Army
07780
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31-E, block no. 0121, requirements for TM 55-1510-219-10.
URGENT
URGENT
TM 55-1510-219-10
C1
CHANGE
NO. 1
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 7 August 1992
}
Operator's Manual
For
ARMY RC-12D AIRCRAFT
TM 55-1510-219-10, 31 May 1991, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar
in the margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
---5-9 and 5-10
8-15 through 8-18
8-25 and 8-26
----
5-8.1/5-8.2
5-9 and 5-10
8-15 through 8-18
8-25 and 8-26
8-26.1/8-26.2
2. Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
GORDON R. SULLIVAN
General, United States Army
Chief of Staff
Official:
MILTON H. HAMILTON
Administrative Assistant to the
Secretary of the Army
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31-E, block no. 0121, -10 & CL maintenance requirements for TM
55-1510-219-10.
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.
URGENT
TM 55-1510-219-10
WARNING PAGE
Personnel performing operations, procedures and practices which are included or implied in this technical manual
shall observe the following warnings. Disregard of these warnings and precautionary information can cause serious
injury or loss of life.
NOISE LEVELS
Sound pressure levels in this aircraft during some operating conditions exceed the Surgeon General's hearing
conservation criteria, as defined in TM MED 501. Hearing protection devices, such as the aviator helmet or ear plugs
shall be worn by all personnel in and around the aircraft during its operation.
STARTING ENGINES
Operating procedures or practices defined in this Technical Manual must be followed correctly. Failure to do so
may result in personal injury or loss of life.
Exposure to exhaust gases shall be avoided since exhaust gases are an irritant to eyes, skin and respiratory
system.
HIGH VOLTAGE
High voltage is a possible hazard around AC inverters, ignition exciter units, and strobe beacons.
USE OF FIRE EXTINGUISHERS IN CONFINED AREAS
Monobromotrifluoromethane (CF3Br) is very volatile, but is not easily detected by its odor. Although non toxic, it must be
considered to be about the same as other freons and carbon dioxide, causing danger to personnel primarily by reduction
of oxygen available for proper breathing. During operation of the fire extinguisher, ventilate personnel areas with fresh
air. The liquid shall not be allowed to come into contact with the skin, as it may cause frostbite or low temperature burns
because of its very low boiling point.
VERTIGO
The strobe/beacon lights should be turned off during flight through clouds to prevent sensations of vertigo, as a result of
reflections of the light on the clouds.
CARBON MONOXIDE
When smoke, suspected carbon monoxide fumes, or symptoms of lack of oxygen (hypoxia) exist, all personnel shall
immediately don oxygen masks, and activate the oxygen system.
FUEL AND OIL HANDLING
Turbine fuels and lubricating oils contain additives which are poisonous and readily absorbed through the skin. Do not
allow them to remain on skin.
SERVICING AIRCRAFT
When conditions permit, the aircraft shall be positioned so that the wind will carry fuel vapors away from all possible
sources of ignition. The fueling unit shall maintain a distance of 20 feet between unit and filler point. A minimum of 10
feet shall be maintained between fueling unit and aircraft.
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TM 55-1510-219-10
Prior to refueling, the hose nozzle static ground wire shall be attached to the grounding lugs that are located
adjacent to filler openings.
SERVICING BATTERY
Improper service of the nickel-cadmium battery is dangerous and may result in both bodily injury and equipment
damage. The battery shall be serviced in accordance with applicable manuals by qualified personnel only.
Wear rubber gloves, apron, and face shield when handling batteries. If corrosive battery electrolyte (potassium
hydroxide) is spilled on clothing or other material, wash immediately with clean water. If spilled on personnel,
immediately start flushing the affected area with clean water. Continue washing until medical assistance arrives.
JET BLAST
Occasionally, during starting, excess fuel accumulation in the combustion chamber causes flames to be blown
from the exhausts. This area shall be clear of personnel and flammable materials.
RADIOACTIVE MATERIAL
Instruments contained in this aircraft may contain radioactive material (TB 55-1500-314-25). These items present
no radiation hazard to personnel unless seal has been broken due to aging or has accidentally been broken. If seal is
suspected to have been broken, notify Radioactive Protective Officer.
RF BURNS
Do not stand near the antennas when they are transmitting.
OPERATION OF AIRCRAFT ON GROUND
At all times during a towing operation, be sure there is a man in the cockpit to operate the brakes.
Personnel should take every precaution against slipping or falling. Make sure guard rails are installed when using
maintenance stands.
Engines shall be started and operated only by authorized personnel. Reference AR 95-1.
Insure that landing gear control handle is in the DN position.
b
TM 55-1510-219-10
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 31 May 1991
TECHNICAL MANUAL
NO. 55-1510-219-10
Operator’s Manual
ARMY MODEL RC-12D AIRCRAFT
REPORTING OR ERRORS AND RECOMMENDING IMPROVEMENTS
You can help improve this manual. If you find any mistakes or if you know of any way to improve the procedures,
please let use know. Mail your letter, DA Form 2028 (Recommended Changes to Publications and Blank Forms), or
DA Form 2028-2 located in the back of this manual directly to: Commander, U.S. Army Aviation and Missile
Command, ATTN: AMSAM-MMC-LS-LP, Redstone Arsenal, AL 35898-5230. A reply will be furnished directly
to you. You may also send in your comments electronically to our E-mail address at <[email protected]>, or
by fax at (205) 842-6546 or DSN 788-6546. Instructions for sending an Electronic DA Form 2028 may be found at
the back of this manual immediately preceding the hard copy DA Forms 2028.
TABLE OF CONTENTS
INTRODUCTION
CHAPTER 1.
AIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATION
CHAPTER 2.
Section I. Aircraft
II. Emergency equipment
III. Engine and related systems
IV. Fuel systems
V. Flight controls
VI. Propellers
VII. Utility systems
VIII. Heating, ventilation, cooling, and environmental control system
IX. Electrical power supply and distribution system
X. Lighting
XI. Flight instruments
XII. Servicing, parking, and mooring
AVIONICS
CHAPTER 3.
Section I. General
II. Communications
III. Navigation
CHAPTER 4.
MISSION EQUIPMENT
Section I. Mission avionics
II. Aircraft survivability equipment
OPERATING LIMITS AND RESTRICTIONS
CHAPTER 5.
Section I. General
II. System limits
III. Power limits
IV. Loading limits
V. Airspeed limits, maximum and minimum
VI. Maneuvering limits
VII. Environmental restrictions
VIII. Other limitations
IX. Required equipment for various conditions of flight
PAGE
1-1
2-1
2-1
2-22
2-22
2-28
2-35
2-39
2-41
2-53
2-58
2-67
2-70
2-76
3-1
3-1
3-2
3-18
4-1
4-1
4-3
5-1
5-1
5-1
5-6
5-7
5-7
5-9
5-9
5-10
5-10
Change 3
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CHAPTER 6.
WEIGHT/BALANCE AND LOADING
Section I. General
II. Weight and balance
III. Fuel/oil
IV. Center of gravity
V. Cargo loading
CHAPTER 7.
PERFORMANCE
Section I. Introduction to performance
CHAPTER 8.
NORMAL PROCEDURES
Section I. Mission planning
II. Operating procedures and maneuvers
III. Instrument flight
IV. Flight characteristics
V. Adverse environmental conditions
VI. Crew duties
CHAPTER 9.
EMERGENCY PROCEDURES
Section I. Aircraft systems
APPENDIX A.
REFERENCES
APPENDIX B.
ABBREVIATIONS AND TERMS
INDEX
ii
6-1
6-1
6-3
6-8
6-8
6-8
7-1
7-1
8-1
8-1
8-1
8-20
8-21
8-23
8-26
9-1
9-1
A-1
B-1
INDEX-1
TM 55-1510-219-10
CHAPTER 1
INTRODUCTION
1-1. GENERAL.
1-5. APPENDIX B, ABBREVIATIONS AND TERMS.
These instructions are for use by the operator(s). They
apply to the RC-12D aircraft.
Appendix B is a listing of abbreviations and terms
used throughout the manual.
1-2. WARNINGS, CAUTIONS, AND NOTES.
1-6. INDEX.
Warnings, cautions, and notes are used to
emphasize important and critical instructions and are
used for the following conditions:
The index lists, in alphabetical order, every titled
paragraph, figure, and table contained in this manual.
Chapter 7, Performance Data, has an additional index
within the chapter.
WARNING
An operating procedure, practice, etc. ,
which if not correctly followed, could
result in personal injury or loss of life.
1-7. ARMY AVIATION SAFETY PROGRAM.
Reports necessary to comply with the safety
program are prescribed in AR 385-40.
CAUTION
An operating procedure, practice, etc.
, which, if not strictly observed, could
result in damage to or destruction of
equipment.
1-8.
DESTRUCTION OF ARMY MATERIEL TO
PREVENT ENEMY USE.
For information concerning destruction of Army
materiel to prevent enemy use, refer to TM 750-2441-5.
NOTE
An operating procedure, condition,
etc. , which is essential to highlight.
1-9. FORMS AND RECORDS.
Army aviators flight record and aircraft
maintenance records which are to be used by crew
members are prescribed in DA PAM 738-751 and TM
55-1500-342-23.
1-3. DESCRIPTION.
This manual contains the best operating
instructions and procedures for the RC-12D aircraft
under most circumstances.
The observance of
limitations, performance, and weight/balance data
provided is mandatory. The observance of procedures
is mandatory except when modification is required
because of multiple emergencies, adverse weather,
terrain, etc. Your flying experience is recognized, and
therefore basic flight principles are not included. THIS
MANUAL SHALL BE CARRIED IN THE AIRCRAFT AT
ALL TIMES.
1-10. EXPLANATION OF CHANGE SYMBOLS.
Changes, except as noted below, to the text and
tables, including new material on added pages, are
indicated by a vertical line in the outer margin extending
close to the entire area of the material affected;
exception: pages with emergency markings, which
consist of black diagonal lines around three edges, may
have the vertical line or change symbol placed along the
inner margins. Symbols show current changes only. A
miniature pointing hand symbol is used to denote a
change to an illustration. However, a vertical line in the
outer margin, rather than miniature pointing hands, is
utilized when there have been extensive changes made
to an illustration. Change symbols are not utilized to
indicate changes in the following:
1-4. APPENDIX A, REFERENCES.
Appendix A is a listing of official publications cited
within the manual applicable to and available for flight
crews.
1-1
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C - Basic mission and type symbol (cargo)
12 - Design number
D - Series symbol
a. Introductory material.
b. Indexes and tabular data where the change
cannot be identified.
1-12. USE OF WORDS SHALL, WILL, SHOULD, AND
MAY.
c. Blank space resulting from the deletion of text,
an illustration or a table.
Within this technical manual the word "shall" is
used to indicate a mandatory requirement. The word
"should" is used to indicate a nonmandatory but
preferred method of accomplishment. The word "may"
is used to indicate an acceptable method of
accomplishment. The word "will" is used to express a
declaration of purpose and may also be used where
simple futurity is required.
d. Correction of minor inaccuracies, such as
spelling, punctuation, relocation of material, etc.
, unless correction changes the meaning of
instructive information and procedures.
1-11. AIRCRAFT DESIGNATION SYSTEM.
The designation system prescribed by AR 70-50 is
used in aircraft designations as follows: EXAMPLE RC12D
1-13. PLACARD ITEMS.
All placard items (switches, controls, etc.
shown throughout this manual in capital letters.
R - Modified mission symbol (reconnaissance)
1-2
) are
TM 55-1510-219-10
CHAPTER 2
AIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATION
Section I. AIRCRAFT
2-1. INTRODUCTION.
2-6. EXHAUST DANGER AREA.
The purpose of this chapter is to describe the
aircraft and its systems and controls which contribute to
the physical act of operating the aircraft. It does not
contain descriptions of avionics or mission equipment,
covered elsewhere in this manual.
This chapter
contains descriptive information and does not describe
procedures for operation of the aircraft.
These
procedures are contained within appropriate chapters in
the manual. This chapter also contains the emergency
equipment installed. This chapter is not designed to
provide instructions on the complete mechanical and
electrical workings of the various systems; therefore,
each is described only in enough detail to make
comprehension of that system sufficiently complete to
allow for its safe and efficient operation.
Danger areas to be avoided by personnel while
aircraft engines are being operated on the ground are
depicted in Figure 2-5. Distance to be maintained with
engines operating at idle are shown. Temperature and
velocity of exhaust gases at varying locations aft of the
exhaust stacks are shown for maximum power. The
danger area extends to 40 feet aft of the exhaust stack
outlets. Propeller danger areas are also shown.
2-7. LANDING GEAR SYSTEM.
The landing gear is a retractable, tricycle type,
electrically operated by a single DC motor. This motor
drives the main landing gear actuators through a gear
box and torque tube arrangement, and also drives a
chain mechanism which controls the position of the nose
gear. Positive down-locks are installed to hold the drag
brace in the extended and locked position. The downlocks are actuated by overtravel of the linear jackscrews
and are held in position by a spring-loaded overcenter
mechanism. The jackscrew in each actuator holds all
three gears in the UP position, when the gear is
retracted. A friction clutch between the gearbox and the
torque shafts protects the motor from electrical overload
in the event of a mechanical malfunction. A 150ampere
current limiter, located on the DC distribution bus under
the center floorboard, protects against electrical
overload. Rotation of the dipole antenna is controlled
through the actuation of the right main gear up limit
switch. Gear doors are opened and closed through a
mechanical linkage connected to the landing gear. The
nose wheel steering mechanism is automatically
centered and the rudder pedals relieved of the steering
load when the landing gear is retracted. Air-oil type
shock struts, filled with compressed air and hydraulic
fluid, are incorporated with the landing gear. Gear
retraction or extension time is approximately six
seconds.
2-2. GENERAL.
The RC-12D is a pressurized, low wing, all metal aircraft
(figs. 2-1 and 2-2) powered by two PT6A41 turboprop
engines, and has all weather capability. Distinguishable
features of the aircraft are the slender, streamlined
engine nacelles, an aft rotating boom antenna, mission
antennas, antenna pods, wing tip pods, a T-tail and a
ventral fin below the empennage. The basic mission of
the aircraft is radio reconnaissance. Cabin entrance is
made through a stair-type door on the left side of the
fuselage.
2-3. DIMENSIONS.
Overall aircraft dimensions are shown in Figure 23.
2-4. GROUND TURNING RADIUS.
Minimum ground turning radius of the aircraft is
shown in Figure 2-4.
a. Landing Gear Control Switch. Landing gear
system operation is controlled by a manually actuated,
wheel-shaped switch placarded LDG GEAR CONTR,
UP and DN, on the left subpanel. The control switch
and associated relay circuits are protected by a 5ampere circuit breaker, placarded
2-5. MAXIMUM WEIGHTS.
Maximum takeoff gross weight is 14,200 pounds.
Maximum landing weight is 13,500 pounds. Maximum
ramp weight is 14,290 pounds.
2-1
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Figure 2-1. General Exterior Arrangement (Sheet 1 of 7)
2-2
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Figure 2-1. General Exterior Arrangement (Sheet 2 of 7)
2-3
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Figure 2-1. General Exterior Arrangement (Sheet 3 of 7)
2-4
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Figure 2-1. General Exterior Arrangement (Sheet 4 of 7)
2-5
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Figure 2-1. General Exterior Arrangement (Sheet 5 of 7)
2-6
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Figure 2-1. General Exterior Arrangement (Sheet 6 of 7)
2-7
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Figure 2-1. General Exterior Arrangement (Sheet 7 of 7)
2-8
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Figure 2-2. General Interior Arrangement
2-9
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Figure 2-3. Principal Dimensions
2-10
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Figure 2-4. Ground Turning Radius
2-11
TM 55-1510-219-10
Figure 2-5. Exhaust and Propeller Danger Areas
2-12
TM 55-1510-219-10
LANDING GEAR RELAY on the overhead circuit breaker panel (fig. 2-26).
b. Landing Gear Down Position Indicator Lights. Landing gear down position is indicated by three green lights on
the left subpanel, placarded GEAR DOWN. These lights may be checked by operating the ANNUNCIATOR TEST
switch. The circuit is protected by a 5-ampere circuit breaker, placarded LANDING GEAR IND, on the overhead circuit
breaker panel (fig. 2-26).
c. Landing Gear Position Warning Lights. Two red bulbs, wired in parallel and activated by microswitches
independent of the GEAR DOWN position indicator lights, are positioned inside the clear plastic grip on the landing gear
control switch. These lights illuminate whenever the landing gear switch is in either the UP or DN position and the gear is
in transit. Both bulbs will also illuminate should either or both power levers be retarded below approximately 79 to 81%
NI when the landing gear is not down and locked. To turn the switch lights OFF, during single-engine operation, the
power lever for the inoperative engine must be advanced to a position which is higher than the setting of the warning
horn microswitch. Extending the landing gear will also turn the lights off. Both red lights indicate the same warning
conditions, but two are provided for a fail-safe indication in the event one bulb
1 with the landing gear out of down and locked position
and the flaps extended beyond 40%, a warning horn
located in the overhead control panel will sound
intermittently. To prevent actuation of the warning horn
during long descents, a pressure differential "Q" switch
is connected into the copilot's static air line to prevent
the warning horn from sounding at airspeeds greater
than 140 KIAS. An altitude sensing switch is installed in
series with the "Q" switch to disable the warning horn at
altitudes above approximately 12,500 feet MSL. The
warning horn is enabled as the aircraft descends through
approximately 10,500 feet MSL. The warning horn
circuit is protected by a 5-ampere circuit breaker,
placarded
burns out. The circuit is protected by a 5-ampere circuit
breaker, placarded LANDING GEAR IND, on the
overhead circuit breaker panel (fig. 2-26).
d. Landing Gear Warning Light Test Button. A test
button, placarded HDL LT TEST, is located on the left
subpanel.
Failure of the landing gear switch to
illuminate red, when this test button is pressed, indicates
two defective bulbs or a circuit fault. The circuit is
protected by a 5-ampere circuit breaker, placarded
LANDING GEAR RELAY CONTROL, on the overhead
circuit breaker panel (fig. 2-26).
e.
Landing Gear Warning Horn. When either
power lever is retarded below approximately 80 ± 1% N
Figure 2-6. Subpanels
2-13
TM 55-1510-219-10
anism. This is accomplished through use of an alternate
engage handle located adjacent to the landing gear
alternate extension handle. To disengage the landing
gear motor, pull the alternate extension handle up and
turn it clockwise. When this handle is pulled, the
landing gear motor is disconnected from the system and
the alternate drive system is locked to the gearbox and
motor. With the alternate drive locked in, the landing
gear may be manually extended by pumping the
alternate extension handle until the three GEAR DOWN
position indicator lamps are illuminated.
Refer to
Chapter 9 for additional information on emergency gear
extension procedures.
LANDING GEAR WARN, located on the overhead
circuit breaker panel (fig. 2-26).
f. Landing Gear Warning Horn Test Switch.
The landing gear warning horn may be tested by the test
switch on the right subpanel. The switch, placarded
STALL WARN TEST OFF LDG GEAR WARN TEST,
will sound the landing gear warning horn and illuminate
the landing gear position warning lights when moved to
the momentary LDG GEAR WARN TEST position. The
circuit is protected by a 5-ampere circuit breaker,
placarded LANDING GEAR WARN, on the overhead
circuit breaker panel (fig. 2-26).
i. Tires. The aircraft is equipped with dual 22 x 6.
75 x 10, 8 ply rated, tubeless, rim inflation tires on the
main gear. The nose gear is equipped with a single 22 x
6. 75 x 10, 8 ply rated, tubeless, rim inflated tire.
g. Landing Gear Safety Switches. A safety switch
on each main landing gear shock strut controls the
operation of various aircraft systems that function only
during flight or only during ground operation. These
switches are mechanically actuated whenever the main
landing gear shock struts are extended (normally after
takeoff), or compressed (normally after landing). The
safety switch on the right main landing gear strut
activates the landing gear control circuits, cabin
pressurization circuits and the flight hour meter when
the strut is extended. This switch also activates a downlock hook, preventing the landing gear from being raised
while the aircraft is on the ground. The hook, which
unlocks automatically after takeoff, can be manually
over-ridden by pressing down on the red button,
placarded DN LOCK REL located adjacent to the
landing gear switch. If the over-ride is used and the
landing gear control switch is raised, power will be
supplied to the warning horn circuit and the horn will
sound. The safety switch on the left main landing gear
strut activates the left and right engine ambient air
shutoff valves when the strut is extended.
j. Steerable Nose Wheel. The aircraft can be
maneuvered on the ground by the steerable nose wheel
system. Direct linkage from the rudder pedals (fig. 2-9)
to the nose wheel steering linkage allows the nose wheel
to be turned 12° to the left of center or 14° to the right.
When rudder pedal steering is augmented by the main
wheel braking action, the nose wheel can be deflected
up to 48°either side of center. Shock loads which would
normally be transmitted to the rudder pedals are
absorbed by a spring mechanism in the steering linkage.
Retraction of the landing gear automatically centers the
nose wheel and disengages the steering linkage from
the rudder pedals.
k. Wheel Brake System. The main wheels are
equipped with multiple-disc hydraulic brakes actuated by
master cylinders attached to the rudder pedals at the
pilot's and copilot's position. Braking is permitted from
either set of rudder pedals. Brake fluid is supplied to the
system from the reservoir in the nose compartment.
The toe brake sections of the rudder pedals are
connected to the master cylinders which actuate the
system for the or responding wheels. No emergency
brake system is provided.
CAUTION
Continued pumping of handle after
GEAR DOWN position indicator lights
(3) are illuminated could damage the
drive
mechanism,
and
prevent
subsequent gear retraction.
WARNING
Repeated and excessive application of
brakes, without allowing sufficient
cooling time between applications, will
cause loss of braking efficiency, and
may cause brake or wheel failure, tire
blowout, or destruction of wheel
assembly by fire.
h. Manual Landing Gear Extension System.
Manual landing gear extension is provided through a
manually powered system as a backup to the electrically
operated system. Before manually extending the gear,
make certain that the landing gear switch is in the down
position with the LANDING GEAR RELAY circuit
breaker pulled.
During a manual landing gear
extension, the landing gear motor must be disengaged
from the landing gear drive mech-
2-14
TM 55-1510-219-10
and the other outside the door. When either handle is
rotated, three rotating-cam-type latches on either side of
the door capture posts mounted on the cargo door. In
the closed position, the door becomes an integral part of
the cargo door. A button adjacent to the door handle
must be depressed before the handle can be rotated to
open the door. A bellows behind the button is inflated
when the aircraft is pressurized to prevent accidental
unlatching and/or opening of the door. A small round
window just above the second step permits observation
of the pressurization safety bellows. A placard adjacent
to the window instructs the operator that the safety lock
arm is in position around the bellows shaft which
indicates a properly locked door. Pushing the red button
adjacent to the window will illuminate the inside door
mechanism. A CABIN DOOR annunciator light in the
caution/advisory panel will illuminate if the door is not
closed and all latches fully locked.
2-8. PARKING BRAKE.
CAUTION
Parking brakes shall not be set during
flight.
Dual parking brake valves are installed below the
cockpit floor. Both valves can be closed simultaneously
by pressing both brake pedals to build up pressure, then
pulling out the handle placarded PARKING BRAKE, on
the left subpanel. Pulling the handle full out sets the
check valves in the system and any pressure being
applied by the toe brakes is maintained. Parking brakes
are released when the brake handle is pushed in. The
parking brake may be set from either cockpit position.
Parking brakes shall not be set during flight.
2-9. ENTRANCE AND EXIT PROVISIONS.
b. Cargo Door. A swing-up door (fig. 2-7), hinged
at the top, provides cabin access for loading cargo or
bulky items. After initial opening force is applied, gas
springs will completely open the cargo door
automatically. The door is counterbalanced and will
remain in the open position. A door support rod is used
to hold the door in the open position, and to aid in
overcoming the pressure of the gas spring assemblies
when closing the door. Once closed, the gas springs
apply a closing force to assist in latching the door. A
rubber seal around the door, seals the pressure vessel
while in flight. The door locking mechanism is operated
only from inside the aircraft, and is operated by two
handles, one in the bottom forward portion of the door
and the other in the upper aft portion of the door. When
the upper aft handle is operated per placard instructions,
two rotating cam-type latches on the forward side of the
door and two on the aft side rotate, capturing posts
mounted on the fuselage side of the door opening. The
bottom handle, when operated per placard instructions,
actuates four pin lug latches across the bottom of the
door. A button on the upper aft handle must be pressed
before the handle can be released to open the door. A
latching lever on the bottom handle must be lifted to
release the handle before the lower latches can be
opened. These act as additional aids in preventing
accidental opening or unlatching of the door. The cabin
and cargo doors are equipped with dual sensing circuits
to provide the crew remote indication of cabin/cargo
door security. An annunciator light placarded CABIN
DOOR will illuminate if the cabin or cargo door is open
and the BATT switch is ON. If the battery switch is
OFF, the annunciator will illuminate only if the cargo
door is not securely closed and latched. The cargo door
sensing circuit receives power from the hot battery bus.
NOTE
Two keys are provided in the loose
tools and equipment bag. Both keys
will fit the locks on the cabin door,
emergency hatch, tailcone access
door and the right and left nose
avionics doors.
a. Cabin Door.
CAUTION
Structural damage may be caused if
more than one person is on the cabin
door at a time. The door is weight
limited to less than 300 pounds.
A swing-down door (fig. 2-7), hinged at the
bottom, provides a stairway for normal and emergency
entry and exit. Two of the steps are movable and fold
flat against the door in the closed position. A step folds
down over the door sill when the door opens to provide
a platform (step) for door seal protection. A plastic
encased cable provides support for the door in the open
position, a handhold, and a convenience for closing the
door from inside. A hydraulic damper permits the door
to lower gradually during opening. A rubber seal around
the door seals the pressure vessel while the aircraft is in
flight. The door locking mechanism is operated by
either of the two mechanically interconnected handles,
one inside
2-15
TM 55-1510-219-10
Figure 2-7. Cabin and Cargo Doors
2-16
TM 55-1510-219-10
while exerting downward force to overcome the pressure
of gas spring assemblies, Then remove support rod
from door as gas spring assemblies pass over-center
position.
CAUTION
Insure the cabin door is closed and
locked.
Operating the cargo door
while the cabin door is open may
damage the door hinge and adjacent
structure.
2. Cargo door Pull closed, using finger hold
cavity in fixed cabin door step.
(1) Opening cargo door.
3. Handle access door (upper aft corner of
door) -Unfasten and open.
CAUTION
Avoid side loading of the gas springs
to prevent damage to the mechanism.
4. Handle Pull handle down until it latches in
closed position.
1. Handle access door (lower forward corner
of door) -Unfasten and open.
5. Handle access door Secure.
2. Handle Lift hook and move to OPEN
position.
6. Handle access door (lower forward corner
of door) -Unfasten and open.
3. Handle access door Secure.
7. Handle Move to full forward position.
4. Handle access door (upper aft corner of
door) -Unfasten and open.
8. Safety hook Check locked in position by
pulling aft on handle.
5. Handle Press button and lift to OPEN
position then latch in place.
9. Handle access door Secure.
c. Cabin Emergency Hatch. The cabin emergency
hatch, placarded EXIT-PULL, is located on the right
cabin sidewall just aft of the copilot's seat. The hatch
may be released, from the inside with a pull-down
handle. A flush mounted pull out handle allows the
hatch to be released from the outside. The hatch is of
the non-hinged, plug type which removes completely
from the frame when the latches are released. The
hatch can be key locked, from the inside, to prevent
opening from the outside. The inside handle will unlatch
the hatch whether or not it is locked, by overriding the
locking mechanism. The keylock should be unlocked
prior to flight to allow removal of the hatch from the
outside in the event of an emergency. The key remains
in the lock when the hatch is locked and can be
removed only when the hatch is unlocked. The key slot
is in the vertical position when the hatch is unlocked.
Removal of the key from the lock before flight assures
the pilot that the hatch can be removed from the outside
if necessary.
6. Handle access door Secure.
7. Door support rod Attach one end to cargo
door ball stud (on forward side of door).
8. Support rod detent pin Check in place.
9. Cabin door sill step Push out on and allow
cargo door to swing open. Gas springs
will automatically open the door.
10. Door support rod Attach free end to ball
stud on forward fuselage door frame.
(2) Closing cargo door.
CAUTION
Avoid side loading of the gas springs
to prevent damage to the mechanism.
2-10. CABIN DOOR CAUTION LIGHT.
1. Door support rod Detach from fuselage
door frame ball stud, then firmly grasp
free end of rod
As a safety precaution, two illuminated MASTER
CAUTION lights on the glare shield and a steadily
illuminated CABIN DOOR yellow caution light on the
annunciator panel indicates the cabin door is not closed
and locked. This circuit is protected by 5-ampere circuit
breakers placarded ANN PWR and ANN IND, located on
the overhead circuit breaker panel (fig. 2-26).
2-17
TM 55-1510-219-10
each seat. Pulling up on the handle, allows the seat to
move up and down.
Both seats have adjustable
headrests and armrests which will raise and lower for
access to the cockpit. Handholds on either side of the
overhead panels and a fold-away protective pedestal
step are provided for pilot and copilot entry into the
cockpit (fig. 2-9). For the storage of maps and the
operator's manual, pilot and copilot seats have an
inboard-slanted, expandable pocket affixed to the lower
portion of the seat back. Pocket openings are held
closed by shock cord tension.
2-11. WINDOWS.
a. Cockpit Windows. The pilot and copilot have
side windows, a windshield and storm windows, which
provide visibility from the cockpit. The storm windows
may be opened on the ground or during unpressurized
flight.
b.
Cabin Windows. The outer cabin windows, of twoply construction, are the pressure type and are integral
parts of the pressure vessel. All cabin windows are
painted over except for the two small aft windows
(movable curtains are provided to cover these windows
when it is desited to seal out light).
b. Pilot and Copilot Seat Belts and Shoulder
Harnesses. Each pilot and copilot seat is equipped with
a lap-type seat belt and shoulder harness connected to
an inertia reel. The shoulder harness belt is of the "Y"
configuration with the single strap being contained in an
inertia reel attached to the base of the seatback. The
two straps are worn with one strap over each shoulder
and fastened by metal loops into the seat belt buckle.
The spring loading at the inertia reel keeps the harness
snug but will allow normal movement required during
flight operations.
2-12. SEATS.
a. Pilot and Copilot Seats. The pilot and copilot
seats (fig. 2-8) are separated from the cabin by
movable curtains. The controls for vertical height
adjustment and fore and aft travel are located under
each seat. The fore and aft adjustment handle is
located beneath the bottom front inboard corner of each
seat. Pulling up on the handle allows the seat to move
fore or aft. The height adjustment handle is located
beneath the bottom front outboard corner of
The inertia reel is designed with a locking device that
will secure the harness in the event of sudden forward
movement or an impact action.
Figure 2-8. Pilot and Copilot Seats
2-18
TM 55-1510-219-10
Figure 2-9. Cockpit
2-19
TM 55-1510-219-10
Figure 2-10. PT6A-41 Engine (Sheet 1 of 2)
2-20
TM 55-1510-219-10
Figure 2-10. PT6A-41 Engine (Sheet 2 of 2)
2-21
TM 55-1510-219-10
Section II: EMERGENCY EQUIPMENT
romethane (CF3Br) or decomposition
products should be avoided.
The
liquid shall not be allowed to come
into contact with the skin, as it may
cause frost bite or low temperature
burns because of its very low boiling
point.
2-13. DESCRIPTION.
The equipment covered in this section includes all
emergency equipment, except that which forms part of a
complete system. For example, landing gear system,
etc. Chapter 9 describes the operation of emergency
exits and location of all emergency equipment.
2-14. FIRST AID KITS.
One hand-operated fire extinguisher is mounted
below the pilot's seat and a second extinguisher, located
on the left cabin sidewall, aft of the cabin door. They
are of the monobromotrifluoromethane (CF3Br) type.
The extinguisher is charged to a pressure of 150 to 170
PSI and emits a forceful stream. Use an extinguisher
with care within the limited area of the cabin to avoid
severe splashing.
Four first aid kits are included in the survival kit.
2-15. HAND-OPERATED FIRE EXTINGUISHER.
WARNING
Repeated or prolonged exposure to
high
concentrations
of
monobromotrifluo-
NOTE
Engine fire extinguisher systems are
described in Section III.
Section III. ENGINES AND RELATED SYSTEMS
fuel control, the oil pumps, the refrigerant compressor
(right engine), the starter/generator, and the turbine
tachometer transmitter. The reduction gearbox forward
of the power turbine provides gearing for the propeller
and drives the propeller tachometer transmitter, the
propeller overspeed governor, and the propeller
governor. A torque limiter is incorporated on the front
case of the propeller reduction gearbox to maintain
developed engine torque within design limits.
2-16. DESCRIPTION.
The aircraft is powered by two PT6A-41 turboprop
engines (fig. 2-10). The engine has a three stage axial,
single stage centrifugal compressor, driven by a single
stage reaction turbine. The power turbine, a two stage
reaction turbine, counter-rotating with the compressor
turbine, drives the output shaft. Both the compressor
turbine and the power turbine are located in the
approximate center of the engine with their shafts
extending in opposite directions. Being a reverse flow
engine, the ram air supply enters the lower portion of the
nacelle and is drawn in through the aft protective
screens. The air is then routed into the compressor.
After it is compressed, it is forced into the annular
combustion chamber, and mixed with fuel that is
sprayed in through 14 nozzles mounted around the gas
generator case. A capacitance discharge ignition unit
and two spark igniter plugs are used to start combustion.
After combustion, the exhaust passes through the
compressor turbine and two stages of power turbine and
is routed through two exhaust ports near the front of the
engine. A pneumatic fuel control system schedules fuel
flow to maintain the power set by the gas generator
power lever. The accessory drive at the aft end of the
engine provides power to drive the fuel pumps,
2-17. ENGINE COMPARTMENT COOLING.
The forward engine compartment including the
accessory section is cooled by air entering around the
exhaust stack cutouts, the gap between the propeller
spinner and forward cowling, and exhausting through
ducts in the upper and lower aft cowling.
2-18. AIR INDUCTION SYSTEMS GENERAL.
Each engine and oil cooler receives ram air
ducted from an air scoop located within the lower
section of the forward nacelle. Special components of
the engine induction system protect the power plant
from icing and foreign object damage.
2-22
TM 55-1510-219-10
Figure 2-11. Pedestal
2-23
TM 55-1510-219-10
2-19. FOREIGN OBJECT DAMAGE CONTROL.
(4) If for any reason the vane does not attain
the selected position within approximately 15 seconds, a
yellow #1 VANE FAIL or #2 VANE FAIL light illuminates
on the caution/advisory panel. In this event, the manual
backup system should be used. When the vane is
successfully positioned with the manual system, the
yellow annunciator lights will extinguish.
The engine has an integral air inlet screen
designed to obstruct objects large enough to damage the
compressor.
2-20. ENGINE ICE PROTECTION SYSTEMS.
b. Engine Air Inlet Deice System.
a. Inertial Separator.
CAUTION
After the ice vanes have been
manually extended, they may be
mechanically actuated only.
No
electrical extension or retraction
shall be attempted as damage to
the actuator may result. Linkage in
the nacelle area must be reset prior
to operation of the electric system.
(1) Description. Hot engine exhaust gas is
utilized for heating the air inlet lips to prevent the
formation of ice. Hot exhaust gas is picked up inside
each engine exhaust stack and carried by plumbing to the
inlet lip. The gas flows through the inside of the lip to the
bottom where it is allowed to escape.
(2) Engine air inlet deice system switches.
(Provisions only system not installed. ) Two switches
placarded ENG INLET LIP HEAT #1, #2 OFF/ ON,
located on the overhead control panel, operate solenoid
valves in the exhaust system of each engine. These
valves control the flow of hot exhaust gases to the inlet air
lip assemblies.
An inertial separation system is built into each
engine air inlet to prevent moisture particles from entering
the engine inlet plenum under icing conditions.
A
movable vane and a bypass door are lowered into the
airstream when operating in visible moisture at 5'C or
colder, by energizing electrical actuators with the
switches, placarded ICE VANE RETRACT EXTEND,
located on the overhead control panel. A mechanical
backup system is provided, and is actuated by pulling the
T-handles just below the pilot's subpanel placarded ICE
VANE #1 ENG #2 ENG. Decrease airspeed to 160 knots
or less to reduce forces for manual extension. Normal
airspeed may then be resumed.
(3) Fuel heater. A oil-to-fuel heat exchanger,
located on the engine accessory case, operates
continuously and automatically to heat the fuel sufficiently
to prevent ice from collecting in the fuel control unit.
Each pneumatic fuel control line is protected against ice.
Power is supplied to each fuel control air line jacket
heater by two switches actuated by moving the condition
levers in the pedestal out of the fuel cutoff range. Fuel
control heat is automatically turned on for all engine
operations.
(1) The vane deflects the ram airstream slightly
downward to introduce a sudden turn in the airstream to
the engine, causing the moisture particles to continue on
undeflected, because of their greater momentum, and to
be discharged overboard.
2-21. ENGINE FUEL CONTROL SYSTEM.
a. Description.
The basic engine fuel system
consists of an engine driven fuel pump, a fuel control unit,
a fuel flow divider, a dual fuel manifold and fourteen fuel
nozzles.
(2) While in the icing flight mode, the extended
position of the vane and bypass door is indicated by green
annunciator lights, #1 VANE EXT and #2 VANE EXT.
b. Fuel Control Unit.
One fuel control unit is
mounted on the accessory case of each engine. This unit
is a hydro-mechanical metering device which determines
the proper fuel schedule for the engine to produce the
amount of power requested by the relative position of its
power lever. The control of developed engine power is
accomplished by adjusting the engine compressor turbine
(N1) speed. N speed controlled by varying the amount of
fuel injected into the combustion chamber through the
fuel nozzles.
Engine shutdown is accomplished by
moving the appropriate condition lever to the full aft,
FUEL CUT-OFF position, which shuts off the fuel supply.
(3) In the non-ice protection mode, the vane
and bypass door are retracted out of the airstream by
placing the ice vane switches in the RETRACT position.
The green annunciator lights will extinguish. Retraction
should be accomplished at 15°C and above to assure
adequate oil cooling.
The vanes should be either
extended or retracted; there are no intermediate positions.
2-24
TM 55-1510-219-10
2-25. ENGINE FIRE DETECTION SYSTEM.
2-22. POWER LEVERS.
CAUTION
Moving the power levers into
reverse range without the engines
running may result in damage to
the reverse linkage mechanism.
a. Description. A flame surveillance system is
installed on each engine to detect external engine fire and
provide alarm to the pilot. Both nacelles are monitored,
each having a control amplifier and three detectors.
Electrical wiring connects all sensors and control
amplifiers to DC power and to the cockpit visual alarm
units. In each nacelle, one detector monitors the forward
nacelle, a second monitors the upper accessory area, and
a third the lower accessory area. Fire emits an infrared
radiation that will be sensed by the detector which
monitors the area of origin. Radiation exposure activates
the relay circuit of a control amplifier which causes signal
power to be sent to cockpit warning systems. An
activated surveillance system will return to the standby
state after the fire is out. The system includes a
functional test switch and has circuit protection through
the FIRE DETR circuit breaker. Warning of internal
nacelle fire is provided as follows: the red MASTER
WARNING lights on the glareshield illuminate
accompanied by the illumination of a red warning light in
the appropriate fire control T-handle placarded FIRE
PULL (fig. 2-28). Fire detector circuits are protected by a
single 5-ampere circuit breaker, placarded FIRE DETR,
located on the overhead circuit breaker panel (fig. 2-26).
Two power levers are located on the control
pedestal (fig. 2-11). These levers regulate power in the
reverse, idle, and forward range, and operate so that
forward movement increases engine power.
Power
control is accomplished through adjustment of the N1
speed governor in the fuel control unit.
Power is
increased when Ni RPM is increased. The power levers
also control propeller reverse pitch. Distinct movement
(pulling up and then aft on the power lever) by the pilot is
required for reverse thrust. Placarding beside the lever
travel slots reads POWER. Upper lever travel range is
designated INCR (increase), supplemented by an arrow
pointing forward. Lower travel range is marked IDLE,
LIFT and REVERSE. A placard below the lever slots
reads: CAUTION REVERSE ONLY WITH ENGINES
RUNNING.
2-23. CONDITION LEVERS.
b. Fire Detection System Test Switch. One
rotary switch placarded FIRE PROTECTION TEST on the
copilot's subpanel is provided to test the engine fire
detection system. Before checkout, battery power must
be on and the FIRE DETR circuit breaker must be closed.
Switch position DETR 1, checks the area forward of the
air intake of each nacelle, including circuits to the cockpit
alarm and indication devices. Switch position DETR 2,
checks the circuits for the upper accessory compartment
of each nacelle. Switch position DETR 3, checks the
circuits for the lower accessory compartment of each
nacelle. Each numbered switch position will initiate the
cockpit indications previously described.
Two condition levers
are located on the control
pedestal (fig. 2-11). Each lever starts and stops the fuel
supply, and controls the idle speed for its engine. The
levers have three placarded positions: FUEL CUTOFF,
LO IDLE, and HIGH IDLE. In the FUEL CUTOFF
position, the condition lever controls the cutoff function of
its engine-mounted fuel control unit. From LO IDLE to
HIGH IDLE, they control the governors of the fuel control
units to establish minimum fuel flow levels. LO IDLE
position sets the fuel flow rate to attain 52 to 55% (at sea
level) minimum NI and HIGH IDLE position sets the rate
to attain 70% minimum NI. The power lever for the
corresponding engine can select NI from the respective
idle setting to maximum power. An increase in low idle Ni
will be experienced at high field elevation.
c. Erroneous
Fire
Detection
System
Indications.
During ground test of the engine fire
detection system, an erroneous indication of system fault
my be encountered if an engine cowling is not closed
properly, or if the aircraft is headed toward a strong
external light source. In this circumstance, change the
aircraft heading to enable a valid system check.
2-24. FRICTION LOCK KNOBS.
Four friction lock knobs, placarded FRICTION
LOCK, are provided on the control pedestal (fig. 211) to
adjust friction drag against the engine power, propeller
RPM, and fuel condition levers. These knobs prevent the
levers from creeping. When rotated clockwise, each knob
increases the friction that opposes movement of the
affected lever. Counterclockwise movement of the knob
decreases friction.
2-26. ENGINE FIRE EXTINGUISHER SYSTEM.
a. Description. The fire extinguisher system
utilizes an explosive squib and valve which, when
2-25
TM 55-1510-219-10
opened, allows the distribution of the pressurized
extinguishing agent through a plumbing network of spray
nozzles strategically located in the fire zones of the
engines.
2-27. OIL SUPPLY SYSTEM.
a. The engine oil tank is integral with the air-inlet
casting located forward of the accessory gearbox. Oil for
propeller operation, lubrication of the reduction gearbox
and engine bearings is supplied by an external line from
the high pressure pump. Two scavenge lines return oil to
the tank from the propeller reduction gearbox. A noncongealing external oil cooler keeps the engine oil
temperature within the operating limits. The capacity of
each engine oil tank is 2. 3 U. S. gallons. The total
system capacity for each engine, which includes the oil
tank, oil cooler, lines, etc. , is 3. 5 U. S. gallons. The
oil level is indicated by a dipstick attached to the oil filler
cap. Oil grade, specification and servicing points, are
described in Section XII, Servicing.
b. Fire Pull Handles. The fire control handles,
which are used to arm the extinguisher system are
centrally located on the pilot's instrument panel (fig. 228), immediately below the glareshield. These controls
receive power from the hot battery bus. The fire detection
system will indicate an engine fire by illuminating the
master fault warning light on the pilot's and copilot's
glareshield and the respective #1 or #2 FIRE PULL lights
in the fire control T-handles. Pulling the fire control Thandle will electrically arm the extinguisher system and
close the fuel firewall shutoff valve for that particular
engine. This will cause the red light in the PUSH TO
EXTINGUISH switch and the respective red #1 and #2
FUEL PRESS light in the warning annunciator panel to
illuminate.
Pressing the lens of the PUSH TO
EXTINGUISH switch (after lifting one side of its springloaded clear plastic guard) will fire the squib, expelling all
the agent in the cylinder at one time. The respective
yellow caution light, #1 or #2 EXTGH DISCH on the
caution/advisory annunciator panel will illuminate and
remain illuminated until the squib is replaced.
b. The oil system of each engine is coupled to a
heat exchanger unit (radiator) of fin-and-tube design.
These exchanger units are the only airframe mounted part
of the oil system and are attached to the nacelles below
the engine air intake. Each heat exchanger incorporates
a thermal bypass which assists in maintaining oil at the
proper temperature range for engine operation.
2-28. TORQUE LIMITER
c. Fire Extinguisher System Test Switch. A rotary
test switch, placarded FIRE PROTECTION TEST, is
located on the copilot's subpanel. The test functions,
placarded EXTGH #1 #2, are arranged on the left side of
the switch and provide a test of the pyrotechnic cartridge
circuitry. During Before Exterior Check, the pilot should
rotate the test switch through the two positions and verify
the illumination of the green SQUIB OK light on the
PUSH TO EXTINGUISH switch and the corresponding
yellow #1 or #2 EXTGH DISCH light on the
caution/advisory annunciator panel.
A torque limiter is installed in the torquemeter
pressure transmitter boss on the front case of the
propeller reduction gearbox. The limiter incorporates a
sealed bellows, connected directly to torquemeter oil
pressure, which works against an externally adjustable
torque-limit spring. Torquemeter oil pressure is sensed at
the inside of the bellows assembly, and during normal
operating conditions the torquemeter oil pressure is not
high enough to compress the torque-limit spring. When
high torque pressure is sensed, the bellows will stretch
and compress the spring assembly. This will permit a
limited amount of governing air pressure to be bled from
the fuel control unit and will continue to do so until engine
speed is reduced, resulting in a proportional reduction in
engine torque pressure.
d. Fire Extinguishing System Supply Cylinder
Gages. A gage, calibrated in PSI, is mounted on each
supply cylinder for determining the level of charge and
should be checked during preflight (table 2-1).
2-26
TM 55-1510-219-10
2-29. ENGINE CHIP DETECTION SYSTEM.
NOTE
The system should be turned OFF
during extended ground operation
to prolong the life of the igniter
plugs.
a. Auto Ignition Switches. Two switches placarded
ENG AUTO IGN #1 or #2, control the auto ignition
systems. The ARM position initiates a readiness mode
for the auto ignition system of the corresponding engine.
The OFF position disarms the system. Each switch is
protected by a corresponding START CONTR #1 or #2 5ampere circuit breaker on the overhead circuit breaker
panel (fig. 2-26).
A magnetic chip detector is installed in the bottom
of each engine nose gearbox to warn the pilot of oil
contamination and possible engine failure. The sensor is
an electrically insulated gap immersed in the oil,
functioning as a normally-open switch. If a large metal
chip or a mass of small particles bridge the detector gap,
a circuit is completed, sending a signal to illuminate a red
annunciator panel indicator light placarded #1 or #2 CHIP
DETR and the MASTER WARNING lights. Chip detector
circuits are protected by two 5-ampere circuit breakers,
placarded CHIP DETR #1 and #2 on the overhead circuit
breaker panel (fig. 2-26).
b. Auto Ignition Lights. If an armed auto ignition
system changes from a ready condition to an operating
condition (energizing the igniter plugs in the engine) a
corresponding green annunciator panel light will
illuminate. The annunciator panel light is placarded #1 or
#2 IGN ON and indicates that the igniters are energized.
The auto ignition system is triggered from a ready
condition to an operating condition when engine torque
drops below approximately 20%. Therefore, when an
auto ignition system is armed, the igniters will be
energized continuously during the time when an engine is
operating at a level below approximately 20% torque.
The auto ignition lights are protected by 5-ampere
IGNITOR CONTR #1 or #2 circuit breakers, located on
the overhead circuit breaker panel (fig. 2-26).
2-30. ENGINE IGNITION SYSTEM.
a. The basic ignition system consists of a solid state
ignition exciter unit, two igniter plugs, two shielded ignition
cables, pilot controlled IGNITION AND ENGINE START
switches and the ENG AUTO IGN switch. Placing an
IGNITION AND ENGINE START switch to ON (forward)
will cause the respective igniter plugs to spark, igniting
the fuel/air mixture sprayed into the combustion chamber
by the fuel nozzles. The ignition system is activated for
ground and air starts, but is switched off after combustion
light up.
b. One three-position toggle switch for each engine,
located on the overhead control panel, will initiate starter
motoring and ignition in the ON/ ENG START position, or
will motor the engine in the STARTER ONLY (aft) position
(fig. 2-18). The switches are placarded #1 ENG START
or #2 ENG START to designate the appropriate engine.
The ON switch position completes the starter circuit for
engine rotation, energizes the igniter plugs for fuel
combustion, and activates the IGN ON light on the
annunciator panel. At center position the switch is OFF.
Two 5-ampere circuit breakers on the overhead circuit
breaker panel, placarded IGNITOR CONTR #1 and #2,
protect ignition circuits. Two 5ampere circuit breakers on
the overhead circuit breaker panel, placarded START
CONTR #1 and #2, protect starter control circuits (fig. 226).
2-32. ENGINE STARTER-GENERATORS.
One start-generator is mounted on each engine
accessory drive section. Each is able to function either as
a starter or as a generator. In the starter function, 28
volts DC is required to power rotation. In the generator
function, each unit is capable of 400 amperes DC output.
Each starter circuit is protected by a 5-ampere circuit
breaker placarded START CONTR, located on the
overhead circuit breaker panel. An automatic starter cutoff is installed in each starter-generator to provide
automatic termination of the start cycle at approximately
40% NI. This rotational speed is sensed by a magnetic
sensor internal to the starter-generator. For additional
information on the starter-generator system refer to
Section IX.
2-31. AUTO IGNITION SYSTEM.
If armed, the auto ignition system automatically
provides combustion re-ignition of either engine should
accidental flameout occur. The system is not essential to
normal engine operation, but is used to reduce the
possibility of power loss due to icing or other conditions.
Each engine has a separate auto ignition control switch
and a green indicator light placarded #1 or #2 IGN ON, on
the annunciator panel. Auto ignition is accomplished by
energizing the two igniter plugs in each engine.
2-33. ENGINE INSTRUMENTS.
The engine instruments are vertically mounted
near the center of the instrument panel (fig. 2-28).
2-27
TM 55-1510-219-10
d. Oil Pressure/Oil Temperature Indicators. Two
gages on the instrument panel (fig. 2-28) panel register
oil pressure in PSI and oil temperature in °C. Oil pressure
is taken from the delivery side of the main oil pressure
pump. Oil temperature is transmitted by a thermal sensor
unit which senses the temperature of the oil as it leaves
the delivery side of the oil pressure pump. Each gage is
connected to pressure transmitters installed on the
respective engine. Both instruments are protected by
5ampere circuit breakers, placarded OIL PRESS and OIL
TEMP #1 or #2, on the overhead circuit breaker panel
(fig. 2-26).
a. Turbine Gas Temperature Indicators. Two TGT
gages on the instrument panel (fig. 2-28) are calibrated in
degrees Celsius.
Each gage is connected to
thermocouple probes located in the hot gases between
the turbine wheels. The gages register the temperature
present between the compressor turbine and power
turbine for the corresponding engine.
b. Engine Torquemeters. Two torquemeters on the
instrument panel (fig. 2-28) indicate torque applied to the
propeller shafts of the respective engines. Each gage
shows torque in percent of maximum using 2 percent
graduations and is actuated by an electrical signal from a
pressure sensing system located in the respective
propeller reduction gear case.
Torquemeters are
protected by individual 0. 5ampere circuit breakers
placarded TORQUEMETER #1 or #2 on the overhead
circuit breaker panel (fig. 2-26).
e. Fuel Flow Indicators.
Two gages on the
instrument panel (fig. 2-28) register the rate of flow for
consumed fuel as measured by sensing units coupled into
the fuel supply lines of the respective engines. The fuel
flow indicators are calibrated in increments of hundreds of
pounds per hour. Both circuits are protected by 1/2
ampere circuit breakers placarded FUEL FLOW #1 or #2,
on the overhead circuit breaker panel (fig. 2-26).
c. Turbine Tachometers. Two tachometers on the
instrument panel (fig. 2-28) register compressor turbine
RPM (N1) for the respective engine. These indicators
register turbine RPM as a percentage of maximum gas
generator RPM.
Each instrument is slaved to a
tachometer generator attached to the respective engine.
Section IV. FUEL SYSTEM
WARNING lights. Refer to Chapter
9. All time in this category shall be
entered on DA Form 2408-13 for the
attention
of
maintenance
personnel.
2-34. FUEL SUPPLY SYSTEM.
The engine fuel supply system (fig. 2-12) consists
of two identical systems sharing a common fuel
management panel and fuel crossfeed plumbing. Each
fuel system consists of five interconnected wing tanks, a
nacelle tank, an auxiliary inboard fuel tank. A fuel
transfer pump is located within each auxiliary tank.
Additionally, the system has an engine-driven boost
pump, a standby fuel pump located within each nacelle
tank, a fuel heater (engine oilto-fuel heat exchanger unit),
a tank vent system, a tank vent heating system and
interconnecting wiring and plumbing. Refer to Section IV
for fuel grades and specifications. Fuel tank capacity is
shown in Table 2-2.
A gear driven boost pump, mounted on each
engine supplies fuel under pressure to the inlet of the
engine-driven primary high-pressure pump for engine
starting and all normal operations. Either the enginedriven boost pump or standby pump is capable of
supplying sufficient pressure to the enginedriven primary
high-pressure pump and thus maintain normal engine
operation.
b. Standby Fuel Pumps. A submerged, electricallyoperated standby fuel pump, located within each nacelle
tank, serves as a backup unit for the engine-driven boost
pump. The standby pumps are switched off during
normal system operations. A standby fuel pump will be
operated during crossfeed to pump fuel from one system
to the opposite engine. The correct pump is automatically
selected when the CROSSFEED switch is activated.
Each standby fuel pump has an inertia switch included in
the power supply circuit. When subjected to a 5 to 6 G
shock loading, as in a crash situation, the inertia switch
will remove electrical power from the standby fuel pumps.
The
standby
fuel
pumps
are
protected
a. Engine Driven Boost Pumps.
CAUTION
Engine operation using only the
enginedriven
primary
(high
pressure)
fuel
pump
without
standby pump or engine-driven
boost pump fuel pressure is limited
to 10 cumulative hours.
This
condition
is
indicated
by
illumination of either #1 or #2 FUEL
PRESS lights and the simultaneous
illumination of both MASTER
2-28
TM 55-1510-219-10
Table 2-2. Fuel Quantity Data
breakers placarded AUXILIARY TRANSIEK #1 or #2,
located on the overhead circuit breaker panel (fig. 2-26).
by two 10-ampere circuit breakers placarded STANDBY
PUMP #1 or #2, located on the overhead circuit breaker
panel (fig. 2-26), and four 5ampere circuit breakers (2
each in parallel) on the hot battery bus.
NOTE
In turbulence or during maneuvers,
the NO FUEL XFR lights may
momentarily illuminate after the
auxiliary
fuel
has
completed
transfer.
c. Fuel Transfer Pumps. The auxiliary tank fuel
transfer system automatically transfers the fuel from the
auxiliary tank to the nacelle tank without pilot action.
Motive flow to a jet pump mounted in the auxiliary tank
sump is obtained from the engine fuel plumbing system
downstream from the engine driven boost pump and
routed through the transfer control motive flow valve.
The motive flow valve is energized to the open position
by the control system to transfer auxiliary fuel to the
nacelle tank to be consumed by the engine during the
initial portion of the flight. When an engine is started,
pressure at the engine driven boost pump closes a
pressure switch which, after a 30 to 50 second time delay
to avoid depletion of fuel pressure during starting,
energizes the motive flow valve. When auxiliary fuel is
depleted, a low level float switch de-energizes the motive
flow valve after a 30 to 60 second time delay provided to
prevent cycling of the motive flow valve due to sloshing
fuel. In the event of a failure of the motive flow valve or
the associated control circuitry, the loss of motive flow
pressure when there is still fuel remaining in the auxiliary
fuel tank is sensed by a pressure switch and float switch,
respectively, which illuminates a light placarded #1 or #2
NO FUEL XFR on the annunciator panel. During engine
start, the pilot should note that the NO FUEL XFR lights
extinguish 30 to 50 seconds after engine start. The NO
FUEL XFR lights will not illuminate if auxiliary tanks are
empty. A manual override is incorporated as a backup for
the automatic transfer system. This is initiated by placing
the AUX TRANSFER switch, located on the fuel
management panel (fig. 2-13) to the OVERRIDE position.
This will energize the transfer control motive flow valve.
The transfer systems are protected by 5-ampere circuit
d. Fuel Gaging System. The total fuel quantity in
the left or right main system or left or right auxiliary tank
is measured by a capacitance type fuel gaging system.
Two fuel gages, one for the left and one for the right fuel
system, read fuel quantity in pounds. Refer to Section XII
for fuel capacities and weights. A maximum of 3% error
may be encountered in each system. However, the
system is compensated for fuel density changes due to
temperature excursions. In addition to the fuel gages,
yellow No. 1 or No. 2 NAC LOW lights on the caution/
advisory annunciator panel illuminate when there is
approximately 20 minutes of fuel per engine remaining
(on standard day, at Sea Level, Maximum Cruise Power
consumption rate). The fuel gaging system is protected
by individual 5 ampere circuit breakers placarded QTY
IND and QTY WARN #1 or #2, located on the overhead
circuit breaker panel (fig. 226). A mechanical spiral float
gage (fig. 2-13) is installed in the auxiliary fuel tank to
provide an indication of fuel level when servicing the tank.
The gage is installed on the auxiliary fuel tank cover,
adjacent to the filler neck. A small sight window in the
upper wing skin permits observation of the gage.
e. Fuel Management Panel. The fuel management
panel (fig. 2-13) is located on the cockpit overhead
between the pilot and copilot. It contains the fuel gages,
standby fuel pump switches, the crossfeed valve switch
and a fuel gaging system control switch and transfer
control
switches
are
also
installed.
2-29
TM 55-1510-219-10
Figure 2-12. Fuel System Schematic
2-30
TM 55-1510-219-10
Figure 2-13. Fuel Management Panel and Auxiliary Tank Mechanical Gage
(3) Fuel transfer control switches.
Two
switches on the fuel management panel (fig. 2-13),
placarded AUX TRANSFER OVERRIDE AUTO control
operation of the fuel transfer pumps. During normal
operation both switches are in AUTO which allows the
system to be automatically actuated by fuel flow to the
engine. If either transfer system fails to operate, the fault
condition is indicated by two illuminated MASTER
CAUTION lights on the glareshield and a steadily
illumihated yellow #1 or #2 NO FUEL XFR light on the
caution annunciator panel.
(1) Fuel gaging system control switch.
A
switch on the fuel management panel (fig.
2-13)
placarded FUEL QUANTITY, MAIN AUXILIARY, controls
the fuel gaging system. When in the MAIN position the
fuel gages read the total fuel quantity in the left and right
wing fuel system. When in the AUXILIARY position the
fuel gages read the fuel quantity in the left and right
auxiliary tanks only.
(2) Standby fuel pump switches.
Two
switches, placarded STANDBY PUMP ON located on the
fuel management panel (fig. 2-13) control a submerged
fuel pump located in the corresponding nacelle tank.
During normal aircraft operation both switches are off so
long as the engine-driven boost pumps function and
during crossfeed operation. The loss of fuel pressure,
due to failure of an engine driven boost pump will
illuminate the MASTER WARNING lights on the
glareshield and will illuminate the #1 FUEL PRESS or #2
FUEL PRESS on the warning annunciator panel. Turning
ON the STANDBY PUMP will extinguish the FUEL
PRESS lights. The MASTER WARNING lights must be
manually cleared.
(4) Fuel crossfeed switch. The fuel crossfeed
valve is controlled by a 3-position switch (fig, 2-13),
located on the fuel management panel, placarded
CROSSFEED OFF. Under normal flight conditions the
switch is left in the OFF position. During emergency
single engine operation, it may become necessary to
supply fuel to the operative engine from the fuel system
on the opposite side. The crossfeed system (fig. 2-15) is
placarded for fuel selection with a simplified diagram on
the overhead fuel control panel. Place the standby fuel
pump switches in the off position when crossfeeding. A
lever lock switch, placarded CROSSFEED, is moved from
the center OFF position to the left or to the right,
depending on direction of fuel flow. This opens the
crossfeed valve and energizes the standby
NOTE
Both standby pump switches shall
be off during crossfeed operation.
2-31
TM 55-1510-219-10
Figure 2-14. Gravity Feed Fuel Flow
2-32
TM 55-1510-219-10
Figure 2-15. Crossfeed Fuel Flow
2-33
TM 55-1510-219-10
An additional drain for the extended range fuel system
line extends through the bottom of the wing center section
adjacent to the fuselage. Anytime the extended range
system is in use, a part of the preflight inspection would
consist of draining a small amount of fuel from this drain
to check for fuel contamination. Whenever the extended
range system is removed from the aircraft and the fuel
line is capped off in the fuselage, the remaining fuel in the
line shall be drained.
pump on the side from which crossfeed is desired. During
crossfeed operation with firewall fuel valve closed,
auxiliary tank fuel will not crossfeed. When the crossfeed
mode is energized, a green FUEL CROSSFEED light on
the caution/advisory panel will illuminate. Crossfeed
system operation is described in Chapter 9.
The
crossfeed valve is protected by a 5-ampere circuit breaker
placarded CROSSFEED VALVE located on the overhead
circuit breaker panel (fig. 2-26).
h. Fuel Drain Collector System. Each engine is
provided with a fuel drain collector system to return fuel
dumped from the engine during clearing and shutdown
operations back into its respective nacelle tank. The
system draws power from the #4 feeder bus. Fuel
transfer is completely automatic. Fuel from the engine
flow divider drains into a collector tank mounted below the
aft engine accessory section. An internal float switch
actuates an electric scavenger pump which delivers the
fuel to the fuel purge line just aft of the fuel purge shutoff
valve. A check valve in the line prevents the backflow of
fuel during engine purging. The circuit breaker for both
pumps is located in the fuel section of the overhead
circuit breaker panel; placarded SCAVENGER PUMP. A
vent line, plumbed from the top of the collector tank, is
routed through an inline flame arrestor and then
downward to a drain manifold on the underside of the
nacelle.
f. Firewall Shutoff Valves.
CAUTION
Do not use the fuel firewall shutoff
valve to shut down an engine,
except in an emergency.
The
engine-driven
highpressure fuel
pump obtains essential lubrication
from fuel flow. When an engine is
operating, this pump may be
severely damaged (while cavitating)
if the firewall valve is closed before
the condition lever is moved to the
FUEL CUTOFF position.
The fuel system incorporates a fuel line shutoff
valve mounted on each engine firewall. The firewall
shutoff valves close automatically when the fire
extinguisher T-handles on the instrument panel are pulled
out. The firewall shutoff valves receive electrical power
from the main buses and also from the hot battery bus
which is connected directly to the battery. The valves are
protected by circuit breakers placarded FIREWALL
VALVE #1 or #2 on the overhead circuit breaker panel
(fig. 2-26), and FIREWALL SHUTOFF #1 or #2 on the
hot battery bus circuit breaker board.
i. Fuel Vent System. Each fuel system is vented
through two ram vents located on the underside of the
wing adjacent to the nacelle. To prevent icing of the vent
system, one vent is recessed into the wing and the
backup vent protrudes out from the wing and contains a
heating element. The vent line at the nacelle contains an
inline flame arrestor.
j. Engine Oil-to-Fuel Heat Exchanger. An engine
oil-to-fuel heat exchanger, located on each engine
accessory case, operates continuously and automatically
to heat the fuel delivered to the engine sufficiently to
prevent the freezing of any water which it might contain.
The temperature of the delivered fuel is thermostatically
regulated to remain between 21°C and 32°C.
g. Fuel Sump Drains. A sump drain wrench is
provided in the aircraft loose tools to simplify draining a
small amount of fuel from the sump drain. There are five
sump drains and one filter drain in each wing (Table 2-3).
Table 2-3. Fuel Sump Drain Locations
2-34
TM 55-1510-219-10
will supply the required pressure. Operation under this
condition will result in an unbalanced fuel load as fuel
from one system will be supplied to both engines while all
fuel from the system with the failed engine driven and
standby boost pumps will remain unused. A triple failure,
which is highly unlikely, would result in the engine driven
primary pump operating without inlet head pressure.
Should this situation occur, the affected engine can
continue to operate from its own fuel supply on its enginedriven primary high-pressure fuel pump.
2-35. FUEL SYSTEM MANAGEMENT.
a. Fuel Transfer System. When the auxiliary tanks
are filled, they will be used first. During transfer of
auxiliary fuel, which is automatically controlled, the
nacelle tanks are maintained full. A check valve in the
gravity feed line from the outboard wing prevents reverse
fuel flow. Normal gravity transfer of the main wing fuel
into the nacelle tanks will begin when auxiliary fuel is
exhausted. The system will gravity feed fuel only to its
respective nacelle tank, i. e. left or right. Fuel will not
gravity feed through the crossfeed system.
2-36. FERRY FUEL SYSTEM.
b. Operation With Failed Engine-Driven Boost Pump
or Standby Pump. Two pumps in each fuel system
provide inlet head pressure to the engine driven primary
high-pressure fuel pump. If crossfeed is used, a third
pump, the standby fuel pump from the opposite system,
Provisions are installed for connection to long
range fuel cells.
Section V. FLIGHT CONTROLS
forward or aft position to provide adequate leg room for
the pilot and copilot. Adjustment is accomplished by
depressing the lever alongside the rudder pedal arm and
moving the pedal forward or aft until the locking pin
engages in the selected position.
2-37. DESCRIPTION.
The aircraft's primary flight control systems
consist of conventional rudder, elevator and aileron
control surfaces. These surfaces are manually operated
from the cockpit through mechanical linkage using a
control wheel for the ailerons and elevators, and
adjustable rudder/brake pedals for the rudder. Both the
pilot and copilot have flight controls. Trim control for the
rudder, elevator and ailerons is accomplished through a
manually actuated cable-drum system for each set of
control surfaces.
The autopilot has provisions for
controlling the position of the ailerons, elevators, and
rudder. Chapter 3 describes the operation of the autopilot
system.
b. Yaw Damp System. A yaw damp system is
provided to aid the pilot in maintaining directional stability
and increase ride comfort. The system may be used at
any altitude and is required for flight above 17,000 feet. It
must be deactivated for takeoff and landing. The yaw
damp system is a part of the autopilot. Operating
instructions for this system are contained in Chapter 3.
The system is controlled by a YAW DAMP switch
adjacent to the ELEV TRIM switch on the pedestal
extension.
c. Rudder Boost System. A rudder boost system is
provided to aid the pilot in maintaining directional stability
resulting from an engine failure or a large variation of
power between the engines. Incorporated in the rudder
cable system are two pneumatic rudder boosting servos
that actuates the cables to provide rudder pressure to
help compensate for asymmetrical thrust.
2-38. CONTROL WHEELS.
Elevator and aileron control surfaces are operated
by manually actuating either the pilot's or copilot's control
wheel (fig. 2-16). Switches are installed in the grips of
each control wheel for operation of pitch trim microphone,
autopilot disconnect, transponder identification, and chaff
dispenser. A manually wound 8-day clock is installed in
the center of the pilot's control wheel, and a digital
clock/timer is installed in the center of the copilot's control
wheel. A map light switch is mounted adjacent to each
clock.
For information on operation of the digital
clock/timer, refer to Section XI.
(1) During operation, a differential pressure
valve accepts bleed air pressure from each engine.
When the pressure varies between the bleed air systems,
the shuttle in the differential pressure valve moves toward
the low pressure side. As the pressure difference reaches
a preset tolerance, a switch closes on the low pressure
side which activates the rudder boost system. This
system is designed only to help compensate for
asymmetrical thrust.
Appropriate trimming is to be
accomplished
by
the
pilot.
Moving
2-39. RUDDER SYSTEM.
a. Rudder Pedals. Aircraft rudder control and nose
wheel steering is accomplished by actuation of the rudder
pedals from either pilot's or copilot's station (fig. 2-9).
The rudder pedals may be individually adjusted in either a
2-35
TM 55-1510-219-10
Figure 2-16. Control Wheels
2-36
TM 55-1510-219-10
rudder bellcrank with the guide hole in the floor. Remove
the locks in reverse order, i. e. , rudder pin, control
column pin, and power control clamp.
either or both of the bleed air valve switches on the
overhead control panel to PNEU & ENVIRO OFF position
will disengage the rudder boost system.
NOTE
Condition levers must be in LOW
IDLE position to perform rudder
boost check.
2-41. TRIM TABS.
Trim tabs are provided for all flight control
surfaces. These tabs are manually activated, and are
mechanically controlled by a cable-drum and jackscrew
actuator system, except the right aileron tab which is of
the fixed bendable type. Elevator and aileron trim tabs
incorporate neutral, anti-servo action, i. e. , as the
elevators or ailerons are displaced from the neutral
position, the trim tab maintains an "as adjusted" position.
The rudder trim tab incorporates anti-servo action, i. e. ,
as the rudder is displaced from the neutral position the
trim tab moves in the same direction as the control
surface. This action increases control pressure as rudder
is deflected from the neutral position.
(2) The system is controlled by a switch
located on the extended pedestal placarded RUDDER
BOOST ON OFF, and is to be turned on before flight. A
preflight check of the system can be performed during the
run-up by retarding the power on one engine to idle and
advancing power on the opposite engine until the power
difference between the engines is great enough to
activate the switch to turn on the rudder boost system.
Movement of the appropriate rudder pedal (left engine
idling, right rudder pedal moves forward) will be noted
when the switch closes, indicating the system is
functioning properly for low engine power on that side.
Repeat the check with opposite power settings to check
for movement of the opposite rudder pedal. The system
is protected by a 5-ampere circuit breaker placarded
RUDDER BOOST, located on the overhead circuit
breaker panel (fig. 2-26).
a. Elevator Trim Tab Control. The elevator trim tab
control wheel placarded ELEVATOR TAB DOWN, UP, is
on the left side of the control pedestal and controls a trim
tab on each elevator (fig. 210). The amount of elevator
tab deflection in degrees from a neutral setting, is
indicated by a position arrow.
b. Electric Elevator Trim. The electric elevator trim
system is controlled by an ELEV TRIM PUSH ON PUSH
OFF switch located on the pedestal, dual element thumb
switches on the control wheels, a trim disconnect switch
on each control wheel and a circuit breaker on the
overhead circuit breaker panel. The PUSH ON -PUSH
OFF switch must be in the ON position to operate the
system. The dual element thumb switch is moved
forward for trimming nose down, aft for nose up, and
when released returns to the center (off) position. Any
activation of the trim system through the copilot's trim
switch can be cancelled by activation of the pilot's switch.
Operating the pilot's and copilot's switches in opposing
directions simultaneously results in no trim action. A
preflight check of the switches should be accomplished
before flight by moving the switches individually on both
control wheels. No one switch alone should operate the
system; operation of elevator trim should occur only by
movement of pairs of switches.
The trim system
disconnect is a bi-level, push button, momentary type
switch, located on the outboard grip of each control wheel.
Depressing the switch to the first of two levels disconnects
the autopilot and yaw damp system, and the second level
disconnects the electric trim
NOTE
With brake deice on, rudder boost
may be inoperative.
2-40. FLIGHT CONTROLS LOCK.
CAUTION
Remove control locks before
towing the aircraft or starting
engines. Serious damage could
result in the steering linkage if
towed by a tug with the rudder lock
installed.
Positive locking of the rudder, elevator and
aileron control surfaces, and engine controls (power
levers, propeller levers, and condition levers) is provided
by a removable lock assembly (fig. 2-17) consisting of
two pins, and an elongated U-shaped strap interconnected
by a chain.
Installation of the controls lock is
accomplished by inserting the U-shaped strap around the
aligned control levers from the copilot's side; then the
aileron/elevator locking pin is inserted through a guide
hole in the top of the pilot's control column assembly, thus
locking the control wheel. The rudder is held in a neutral
position by an L-shaped pin which is installed through a
guide hole in the floor aft of the pilots rudder pedals. The
rudder pedals must be centered to align the hole in the
2-37
TM 55-1510-219-10
Figure 2-17. Control Locks
position indicator on the forward control pedestal. Full
flap extension and retraction time is approximately 11
seconds. The flap control switch is located on the control
pedestal. No emergency wing flap actuation system is
provided. With flaps extended beyond the APPROACH
position, the landing gear warning horn will sound, unless
the landing gear is down and locked. The circuit is
protected by a 20-ampere circuit breaker, placarded FLAP
MOTOR POWER, located on the overhead circuit breaker
panel (fig. 2-26).
system. The system can be reset by pressing the ON
OFF switch on the pedestal to ON again.
c. Aileron Trim Tab Control. The aileron trim tab
control, placarded AILERON TAB -LEFT, RIGHT, is on
the control pedestal and will adjust the left aileron trim tab
only (fig. 2-10). The amount of aileron tab deflection,
from a neutral setting, as indicted by a position arrow, is
relative only and is not in degrees. Full travel of the tab
control moves the trim tab 7-1/2 degrees up and down.
d. Rudder Trim Tab Control. The rudder trim tab
control knob, placarded RUDDER TAB LEFT, RIGHT, is
on the control pedestal, and controls adjustment of the
rudder trim tab (fig. 2-10). The amount of rudder tab
deflection, in degrees from a neutral setting, is indicated
by a position arrow.
a. Wing Flap Control Switch. Flap operation is
controlled by a three-position switch with a flapshaped
handle on the control pedestal (fig. 2-10). The handle of
this switch is placarded FLAP and switch positions are
placarded: FLAP UP, APPROACH, and DOWN. The
amount of downward extension of the flaps is established
by position of the flap switch, and is as follows: UP 0%,
APPROACH 40%, and DOWN 100%. Limit switches,
mounted on the right inboard flap, control flap travel. The
flap control switch, limit switch, and relay circuits are
protected by a 5-ampere circuit breaker, placarded FLAP
CONTR located on the overhead circuit breaker panel
(fig. 2-26). Flap positions between UP and APPROACH
cannot be selected. For intermediate flap positions
between AP-
2-42. WING FLAPS.
The all-metal slot-type wing flaps are electrically
operated and consist of two sections for each wing.
These sections extend from the inboard end of each
aileron to the junction of the wing and fuselage. During
extension, or retraction, the flaps are operated as a single
unit, each section being actuated by a separate jackscrew
actuator. The actuators are driven through flexible shafts
by a single, reversible electric motor.
Wing flap
movement is indicated in percent of travel by a flap
2-38
TM 55-1510-219-10
b.
Wing Flap Position Indicator.
Flap
position in percent of travel from "O" percent (UP) to
100 percent (DOWN), is shown on an indicator,
placarded FLAPS, located on the subpanel (fig. 2-6).
The approach and full down or extended flap position is
14 and 34 degrees, respectively. The flap position
indicator is protected by a 5-ampere circuit breaker,
placarded FLAP CONTR, located on the overhead
circuit breaker panel (fig. 2-26).
PROACH and DOWN, the APPROACH position acts as
an off position. To return the flaps to any position
between full DOWN and APPROACH, place the flap
switch to UP and when desired flap position is obtained,
return the switch to the APPROACH detent. In the
event that any two adjacent flap sections extend 3 to 5
degrees out of phase with the other, a safety
mechanism is provided to discontinue power to the flap
motor.
Section VI. PROPELLERS
green and are placarded #1 AUTOFEATHER (left engine)
and #2 AUTOFEATHER (right engine). The system will
remain inoperative as long as either power lever is
retarded below 90% Ni position, unless TEST position of
the AUTOFEATHER SWITCH is selected to disable the
power lever limit switches. The system is designed for
use only during takeoff and landing and should be turned
off when establishing cruise climb. During takeoff or
landing, should the torque for either engine drop to an
indication between 16 21%, the autofeather system for
the opposite engine will be disarmed. Disarming is
confirmed when the AUTOFEATHER light of the opposite
engine becomes extinguished. If torque drops further, to
a reading between 9 and 14%, oil is dumped from the
servo of the effected propeller allowing a feathering
spring and counter-weights to move the blades into
feathered position.
Feathering also causes the
AUTOFEATHER light of the feathered propeller to
extinguish.
At
this
time,
both
annunciator
AUTOFEATHER lights are extinguished, the propeller of
the defective engine has feathered, and the propeller of
the operative engine has been disarmed from the
autofeathering capability. Only manual feathering control
remains for the second propeller.
2-43. DESCRIPTION.
A three blade aluminum propeller is installed on
each engine. The propeller is of the full feathering,
constant speed, counter-weighted, reversible type,
controlled by engine oil pressure through single action,
engine driven propeller governors. The propeller is flange
mounted to the engine shaft. Centrifugal counterweights,
assisted by a feathering spring, move the blades toward
the low RPM (high pitch) position and into the feathered
position. Governor boosted engine oil pressure moves
the propeller to the high RPM (low pitch) hydraulic stop
and reverse position. The propellers have no low RPM
(high pitch) stops; this allows the blades to feather after
engine shutdown.
Low pitch propeller position is
determined by the low pitch stop which is a mechanically
actuated, hydraulic stop. Beta and reverse blade angles
are controlled by the power levers in the beta and reverse
range.
2-44. FEATHERING PROVISIONS.
Both manual and automatic propeller feathering
systems are provided. Manual feathering is accomplished
by pulling the corresponding propeller lever aft past a
friction detent. To unfeather, the propeller lever is pushed
forward into the governing range.
An automatic
feathering system, will sense loss of torque and will
feather an unpowered propeller. Feathering springs will
feather the propeller when it is not turning.
b. Propeller Autofeather Switch. Autofeathering is
controlled by an AUTOFEATHER switch on the overhead
control panel (fig. 2-18). The threeposition switch is
placarded ARM, OFF and TEST, and is spring-loaded
from TEST to OFF. The ARM position is used only during
takeoff and landing. The TEST position of the switch,
enables the pilot to check readiness of the autofeather
systems, below 88% to 92% N1, and is for ground
checkout purposes only.
a. Automatic Feathering. The automatic feathering
system provides a means of immediately dumping oil
from the propeller servo to enable the feathering spring
and counterweights to start feathering action of the blades
in the event of an engine failure. Although the system is
armed by a switch on the overhead control panel,
placarded AUTOFEATHER ARM OFF TEST, the
completion of the arming phase occurs when both power
levers are advanced above 90% Ni at which time both
indicator lightson the caution/advisory annunciator panel
indicate a fully armed system. The annunciator lights are
c. Autofeather Lights. Two green lights on the
caution/advisory
annunciator
panel,
placarded
AUTOFEATHER #1 and #2. When illuminated, the lights
indicate that the autofeather system is armed. Both lights
will be extinguished if either propeller has been
autofeathered or if the system is disarmed by retarding a
power lever. Autofeather circuits are protected by one 5ampere circuit breaker placarded
2-39
TM 55-1510-219-10
excessive RPM if the left propeller is feathered while the
synchrophaser is on, the synchrophaser has a limited
range of control from the manual governor setting.
Normal governor operation is unchanged but the
synchrophaser will continuously monitor propeller RPM
and reset the governor as required. A magnetic pickup
mounted in each propeller overspeed governor and
adjacent to each propeller deice brush block transmits
electric pulses to a transistorized control box installed
forward of the pedestal. The right propeller RPM and
phase will automatically be adjusted to correspond to the
left. To change RPM, adjust both propeller controls at the
same time. This will keep the right governor setting within
the limiting range of the left propeller.
If the
synchrophaser is on but is unable to adjust to the right
propeller to match the left, the actuator has reached the
end of its travel. To recenter, turn the switch off,
synchronize the propellers manually, and turn the switch
back on.
AUTO FEATHER, located on the overhead circuit breaker
panel (fig. 2-26).
2-45. PROPELLER GOVERNORS.
Two governors (a constant speed governor, and
an overspeed governor) control propeller RPM. The
constant speed governor, mounted on top of the propeller
reduction housing, controls the propeller through its entire
range. The propeller control lever operates the propeller
by means of this governor. If the constant speed
governor should malfunction and request more than 2000
RPM, the overspeed governor cuts in at 2080 RPM and
dumps oil from the propeller to limit RPM. A solenoid,
actuated by the PROP GOV TEST switch located on the
overhead control panel (fig. 2-18), is provided to reset
the overspeed governor to approximately 1830 to 1910
RPM for test purposes. If the propeller sticks or moves
too slowly to prevent an overspeed condition, the power
turbine governor, contained within the constant speed
governor housing, acts as a fuel topping governor. When
the propeller reaches 2120 RPM, the fuel topping
governor limits the fuel flow to the gas generator, thereby
reducing the power driving the propeller.
During
operation in the reverse range, the power turbine
governor is reset to approximately 95% propeller RPM
before the propeller reaches a negative pitch angle. This
insures that the engine power is limited to maintain a
propeller RPM of somewhat less than that of the constant
speed governor setting. The constant speed governor
will, therefore, always sense an underspeed condition and
direct oil pressure to the propeller servo piston to permit
propeller operation in beta and reverse range.
b. Control Box. The control box converts any pulse
rate differences into correction commands, which are
transmitted to a stepping type actuator motor mounted on
the right engine cowl forward support ring. The motor
then trims the right propeller governor through a flexible
shaft and trimmer assembly to exactly match the left
propeller. The trimmer, installed between the governor
control arm and the control cable, screws in or out to
adjust the governor while leaving the control lever setting
constant. A toggle switch installed adjacent to the
synchrophaser turns the system on. With the switch off,
the actuator automatically runs to the center of its range
of travel before stopping to assure normal function when
used again. To operate the system, synchronize the
propeller in the normal manner and turn the
synchrophaser on. The system is designed for in-flight
operations and is placarded to be off for take-off and
landing. Therefore, with the system on and the landing
gear extended, the master caution lights will illuminate
and a yellow light on the caution/advisory annunciator
panel, PROP SYNC ON, will illuminate.
2-46. PROPELLER TEST SWITCHES.
Two two-position switches on the overhead
control panel (fig. 2-18), are provided for operational
testing of the propeller systems. Placarding above the
switches reads PROP GOV TEST. Each switch controls
test circuits for the corresponding propeller. In the test
position, the switches are used to test the function of the
corresponding overspeed governor. Refer to Chapter 8,
for test procedure. Propeller test circuits are protected by
one 5-ampere circuit breaker placarded PROP GOV.
located on the overhead circuit breaker panel (fig. 2-26).
c. Synchroscope.
The propeller synchroscope,
provides an indication of synchronization of the
propellers. If the right propeller is operating at a higher
RPM than the left, the face of the synchroscope, a black
and white cross pattern, spins in a clockwise rotation.
Left, or counterclockwise, rotation indicates a higher RPM
of the left propeller. This instrument aids the pilot in
obtaining complete synchronization of propellers. The
system is protected by a 5-ampere circuit breaker
placarded PROP SYNC, located on the overhead circuit
breaker panel (fig. 2-26).
2-47. PROPELLER SYNCHROPHASER.
a. Operation.
The propeller synchrophaser
automatically matches the RPM of the right propeller
(slave propeller) to that of the left propeller (master
propeller) and maintains the blades of one propeller at a
predetermined relative position with the blades of the
other propeller. To prevent the right propeller from losing
2-40
TM 55-1510-219-10
in dusty conditions, to prevent
obscuring the operator's vision.
2-48. PROPELLER LEVERS.
Two propeller levers on the control pedestal (fig,
2-10), placarded PROP, are used to regulate propeller
speeds. Each lever controls a primary governor, which
acts to regulate propeller speeds within the normal
operation range. The full forward position of the levers is
placarded TAKEOFF, LANDING AND REVERSE and
also HIGH RPM. The full aft position of the levers is
placarded FEATHER. When a lever is placed at HIGH
RPM, the propeller may attain a static RPM of 2,000
depending upon power lever position. As a lever is
moved aft, passing through the propeller governing range,
but stopping at the feathering detent, propeller RPM will
correspondingly decrease to the lowest limit. Moving a
propeller lever aft past the detent into FEATHER will
feather the propeller.
CAUTION
To prevent an asymmetrical thrust
condition, propeller levers must be
in HIGH RPM position prior to
propeller reversing.
The propeller blade angle may be reversed to
shorten landing roll. To reverse, propeller levers must be
positioned at HIGH RPM (full forward), and the power
levers are lifted up to pass over the IDLE detent, then
pulled aft into REVERSE setting. One yellow caution
light, placarded REV NOT READY, on the
caution/advisory annunciator panel, alerts the pilot not to
reverse the propellers. This light illuminates only when
the landing gear handle is down, and if propeller levers
are not at HIGH RPM (full forward). This circuit is
protected by a 5ampere circuit breaker, placarded
LANDING GEAR RELAY, located on the overhead circuit
breaker panel (fig. 2-26).
2-49. PROPELLER REVERSING.
CAUTION
Do not move the power levers into
reverse range without the engine
running. Damage to the reverse
linkage mechanisms will occur.
2-50. PROPELLER TACHOMETERS.
Two tachometers on the instrument panel register
propeller speed in hundreds of RPM (fig. 2-28). Each
indicator is slaved to a tachometer generator unit
attached to the corresponding engine.
CAUTION
Propeller reversing on unimproved
surfaces should be accomplished
carefully to prevent propeller
erosion from reversed airflow and,
Section VII. UTILITY SYSTEMS
2-51. DEFROSTING/DEFOGGING SYSTEM.
4. Cabin air, copilot air, pilot air, and defrost
air controls As required.
a. Description. The defrosting/defogging system is
an integral part of the heating and ventilation system.
The system consists of two warm air outlets connected by
ducts to the heating system. One outlet is just below the
pilot's windshield and the other is below the copilot's
windshield. A push-pull control, placarded DEFROST
AIR, on the pilot's subpanel, manually controls airflow to
the windshield. When pulled out, defrosting air is ducted
to the windshield. As the control is pushed in, there is a
corresponding decrease in airflow.
b. Automatic Operation.
c. Maximum Windshield Defrosting.
1. Pilot air, copilot air IN.
2. Cabin air and defrost air controls Out
3. Cabin
temperature
switch MAN HEAT.
mode
selector
4. Cold air outlets As required.
1. Vent blower switches As required.
5. Manual temperature switch As required.
2. Cabin temperature mode selector switch
AUTO
d. Manual Operation. If the automatic temperature
control should fail to operate, the tempera-
3. Cabin temperature control As required.
2-41
TM 55-1510-219-10
to the OFF position when released. After the cycle, the
boots will remain in the vacuum hold down condition until
again actuated by the switch.
ture (of defrost air and cabin air) may be controlled
manually by manipulating the CABIN TEMP MODE
control switch between the OFF and MAN HEAT
positions.
(4) Either engine is capable of providing
sufficient bleed air for all requirements of the surface
deicer system. Check valves in the bleed air and vacuum
lines prevent backflow through the system during singleengine operation. Regulated pressure is indicated on a
gage, placarded PNEUMATIC PRESSURE, located on
the copilots subpanel.
2-52. SURFACE DEICER SYSTEM.
a. Description. Ice accumulation is removed from
each inboard and outboard wing leading edge, and both
horizontal stabilizers by the flexing of deicer boots which
are pneumatically actuated. Engine bleed air, from the
engine compressor, is used to supply air pressure to
inflate the deicer boots, and to supply vacuum, through
the ejector system, for boot hold down during flight. A
pressure regulator protects the system from over inflation.
When the system is not in operation, a distributor valve
applies vacuum to the boots for hold down.
2-53. ANTENNA DEICE SYSTEM.
a. Description. The antenna deice system removes
ice accumulation from the inboard wing dipole antennas
and the aft rotating boom dipole antenna. Pressure
regulated bleed air from the engines supplies pressure to
inflate the boots. To assure operation of the system in the
event of failure of one engine, a check valve is
incorporated in the bleed air line from each engine to
prevent loss of pressure through the compressor of the
inoperative engine. Inflation and deflation phases are
controlled by a distributor valve.
Deice boots are
intended to remove ice after is has formed rather than to
prevent it's formation. For the most effective deicing
operation, allow at least 1/8 to 1/4 inch of ice to form on
the boots before attempting ice removal. Very thin ice
may crack and cling to the boots instead of shedding.
CAUTION
Operation of the surface deice
system in ambient temperatures
below -40°C can cause permanent
damage to the deice boots.
b. Operation.
(1) Deice boots are intended to remove ice
after it has formed rather than prevent its formation. For
the most effective deicing operation, allow at least 1/2
inch of ice to form on the boots before attempting ice
removal. Very thin ice may crack and cling to the boots
instead of shedding.
NOTE
Never cycle the system rapidly.
This may cause the ice to
accumulate outside the contour of
the inflated boots and prevent ice
removal.
NOTE
Never cycle the system rapidly, this
may cause the ice to accumulate
outside the contour of the inflated
boots and prevent ice removal.
b. Antenna Deice System Switch. The antenna
deice system is controlled by a switch placarded ANT
DEICE, SINGLE-OFF-MANUAL located on the overhead
control panel (fig. 2-18). The switch is spring loaded to
return to the OFF position from the SINGLE or MANUAL
position. When the switch is set to the single position, the
system will run through one timed inflation-deflation cycle.
When the switch is held in the MANUAL position the
boots will inflate and remain inflated until the switch is
released.
(2) A three position switch on the overhead
control panel placarded DEICE MANUAL OFF SINGLE
CYCLE AUTO, controls the deicing operation. The switch
is spring loaded to return to the OFF position from
SINGLE CYCLE AUTO or MANUAL. When the SINGLE
CYCLE AUTO position is selected, the distributor valve
opens to inflate the wing boots. After an inflation period
of approximately 6 seconds, an electronic timer switches
the distributor to deflate the wing boots and a 4 second
inflation begins in the horizontal stabilizer boots. When
these boots have inflated and deflated, the cycle is
complete.
c. Forward Wide Band Data Link Antenna Radome
Anti-Ice. The forward wide band data link antenna
radome anti-ice system utilizes engine bleed air to
prevent the formation of ice on the radome.
(3) If the switch is held in the MANUAL
position, the boots will inflate simultaneously and remain
inflated until the switch is released. The switch will return
2-42
TM 55-1510-219-10
Power to the two heating elements on each blade, the
inner and outer element, is cycled by the timer in the
following sequence: right propeller outer element, right
propeller inner element, left propeller outer element, left
propeller inner element. Loss of one heating element
circuit on one side does not mean that the entire system
must be turned off. Proper operation can be checked by
noting the correct level of current usage on the ammeter.
An intermittent flicker of the needle approximately each
30 seconds indicates switching to the next group of
heating elements by the timer.
The system is controlled by a switch placarded RADOME
ANTI-ICE-ON located on the overhead control panel. A
temperature sensing element installed in the discharge
duct measures the hot air temperature as it leaves the
mixing chamber. When the hot air temperature reaches
130 degrees Fahrenheit, the green RADOME HEAT light
in the mission caution/ advisory panel will illuminate. If
the air temperature exceeds 200 degrees Fahrenheit, a
heat detection switch, located adjacent to the temperature
sensing element, will close the solenoid shutoff valves,
and the yellow RADOME HOT light located on the
mission caution/advisory panel will illuminate.
The
shutoff valves will automatically open after the air
temperature cools to approximately 130 degrees
Fahrenheit. The system is protected by a 7. 5 ampere
circuit breaker placarded RADOME ANTI-ICE located on
the overhead control panel.
c. Manual Operation. The manual propeller deice
system is provided as a backup to the automatic system.
A control switch located on the overhead control panel,
placarded PROP INNER OUTER, controls the manual
override relays. When the switch is in the OUTER
position, the automatic timer is overridden and power is
supplied to the outer heating elements of both propellers
simultaneously. The switch is of the momentary type and
must be held in position until the ice has been dislodged
from the propeller surface. After deicing with the outer
elements, the switch is to be held in the INNER position to
perform the same function for the inner elements of both
propellers. The loadmeters will indicate approximately a
5% increase of load per meter when manual prop deice is
operating. The prop deice ammeter will not indicate any
load in the manual mode of operation.
2-54.
PROPELLER ELECTROTHERMAL ANTI-ICE
SYSTEM.
a. Description. Electrothermal anti-ice boots are
cemented to each propeller blade to prevent ice formation
or to remove the ice from the propellers. Each thermal
boot consists of one outboard and one inboard heating
element, and receives electrical power from the deice
timer. This timer sends current to all propeller deice
boots and prevents the boots from overheating by limiting
the time each element is energized. Four intervals of
approximately 30 seconds each complete one cycle.
Current consumption is monitored by a propeller ammeter
on the copilot's subpanel.
Two 20-ampere circuit
breakers placarded PROP ANTI ICE LEFT and RIGHT
and 5-ampere propeller control circuit breaker placarded
PROP ANTI-ICE CONTR on the overhead circuit breaker
panel (fig. 2-26), protect the propeller electrothermal
deice system during manual operation. A 25 ampere
circuit breaker placarded PROP ANTI-ICE AUTO,
protects the system in automatic operation.
2-55. PITOT AND STALL WARNING HEAT SYSTEM.
CAUTION
Pitot heat should not be used for
more than 15 minutes while the
aircraft
is
on
the
ground.
Overheating may damage the
heating elements.
a. Pitot Heat. Heating elements are installed in the
pitot masts located on the nose. Each heating element is
controlled by an individual switch placarded PITOT ON
LEFT or RIGHT, located on the overhead control panel
(fig. 2-18). It is not advisable to operate the pitot heat
system on the ground except for testing or for short
intervals of time to remove ice or snow from the mast.
Circuit protection is provided by two 7. 5 amperes circuit
breakers, placarded PITOT HEAT, on the overhead circuit
breaker panel (fig. 2-26).
b. Automatic Operation. A control switch on the
overhead control panel placarded PROP OFF AUTO is
provided to activate the automatic system. A deice
ammeter on the center subpanel registers the amount of
current (14 to 18 amperes) passing through the system
being used. During AUTO operation, power to the timer
will be cut off if the current rises above 25 amperes.
Current flows from the timer to the brush assembly and
then to the slip rings installed on the spinner backing
plate. The slip rings carry the current to the deice boots
on the propeller blades. Heat from the boots reduces the
grip of the ice which is then thrown off by centrifugal
force, aided by the air blast over the propeller surfaces.
2-43
TM 55-1510-219-10
protected by a 5-ampere circuit breaker, placarded STALL
WARN, on the overhead circuit breaker panel.
CAUTION
The heating elements protect the
stall warning lift transducer vane
and face plate from ice, however, a
buildup of ice on the wing may
change or disrupt the airflow and
prevent the system from accurately
indicating an imminent stall.
2-57. BRAKE DEICE SYSTEM.
a. Description. A heated-air brake deice system
may be used on the ground or in flight with gear retracted
or extended. When activated, hot air is diffused by
means of a manifold assembly over the brake discs in
each wheel. Manual and automatic controls are provided.
There are two primary occasions which require brake
deicing. The first is when an aircraft has been parked in a
freezing atmosphere allowing the brake systems to
become contaminated by freezing rain, snow, or ice, and
the aircraft must be moved or taxied. The second
occasion is during flight through icing conditions with wet
brake assemblies presumed to be frozen, which must be
thawed prior to landing to avoid possible tire damage and
loss of directional control. Hot air for the brake deice
system comes from the compressor stage of both engines
obtained by means of a solenoid valve attached to the
bleed air system which serves both the surface deice
system and the pneumatic systems operation.
b. Stall Warning Heat.
The lift transducer is
equipped with anti-icing capability on both the mounting
plate and the vane. The heat is controlled by a switch
located on the overhead control panel placarded STALL
WARN. The level of heat is minimal for ground operation
but is automatically increased for flight operation through
the left landing gear saety switch. Circuit protection is
provided by a 15-ampere circuit breaker, placarded
STALL WARN, on the overhead circuit breaker panel (fig.
2-26).
2-56. PITOT AND STATIC SYSTEM.
a. Description. The pitot and static system supplies
static pressure to two airspeed indicators, two altimeters,
two vertical velocity indicators, and ram air to the
airspeed indicators. This system consists of two pitot
masts (one located on each side of the lower position of
the nose), static air pressure ports in the aircraft's exterior
skin on each side of the aft fuselage, and associated
system plumbing. The pitot head is protected from ice
formation by internal electric heating elements.
b. Operation. A switch on the overhead control
panel, placarded BRAKE DEICE, controls the solenoid
valve by routing power through a control module box
under the aisleway floorboards. When the switch is on,
power from a 5-ampere circuit breaker on the overhead
circuit breaker panel is applied to the control module. A
10-minute timer limits operation and avoids excessive
wheel well temperatures when the landing gear is
retracted. The control module also contains a circuit to
the green BRAKE DEICE ON annunciator light, and has a
resetting circuit interlocked with the gear uplock switch.
When the system is activated, the BRAKE DEICE ON
light should be monitored and the control switch selected
OFF after the light extinguishes otherwise, on the next
gear extension the system will restart without pilot action.
The control switch should also be selected OFF, if deice
operation fails to self-terminate after 10 minutes. If the
automatic timer has terminated brake deicer operation
after the last retraction of the landing gear, the landing
gear must be extended in order to obtain further operation
of the system.
b. Alternate Static Air Source. An alternate static air
line, which terminates just aft of the rear pressure
bulkhead, provides a source of static air for the pilot's
instruments in the event of source failure from the pilot's
static air line. A control on the pilot's subpanel placarded
PILOTS STATIC AIR SOURCE, may be actuated to
select either the NORMAL or ALTERNATE air source by
a two position selector valve. The valve is secured in the
NORMAL position by a spring clip. Refer to Chapter 7 for
airspeed indicator and altimeter calibration information
when using the alternate air source.
c. Stall Warning System. The stall warning system
consists of a transducer, a lift computer, a warning horn,
and a test switch.
Angle of attack is sensed by
aerodynamic pressure on the lift transducer vane located
on the left wing leading edge. When a stall is imminent,
the output of the transducer activates a stall warning horn.
The system has preflight test capability through the use of
a switch placarded STALL WARN TEST OFFLDG GEAR
WARN TEST on the right subpanel. Holding this switch in
the STALL WARN TEST position actuates the warning
horn by moving the transducer vane. The circuit is
(1) The BLEED AIR FAIL lights may
momentarily illuminate during simultaneous operation of
the surface deice and brake deice systems at low N1
speeds. If the lights immediately extinguish. this may be
disregarded.
2-44
TM 55-1510-219-10
b. Normal Operation.
For normal operation,
switches for the FUEL VENTS anti-ice circuits are turned
ON as required during the BEFORE TAKEOFF
procedures (Chapter 8).
(2) During certain ambient conditions, use of
the brake deice system may reduce available engine
power, and during flight will result in a TGT rise of
approximately 20°C. Appropriate performance charts
should be consulted before brake deice system use. If
specified power cannot be obtained without exceeding
limits, the brake deice system must be selected off until
after takeoff is completed. TGT limitations must also be
observed when setting climb and cruise power. The
brake deice system is not to be operated above 15°C
ambient temperature except to test the system. The
system is not to be operated for longer than 10 minutes
(one deicer cycle) with the landing gear retracted. If
operation does not automatically terminate after
approximately 10 minutes following gear retraction, the
system must be manually selected off. During periods of
simultaneous brake deice and surface deice operation,
maintain 85% Nl or higher. If inadequate pneumatic
pressure is developed for proper surface deicer boot
inflation, select the brake deice system off. Both sources
of pneumatic bleed air must be in operation during brake
deice system use. Select the brake deice system off
during single-engine operation.
Circuit protection is
provided by a 5-ampere circuit breaker, placarded
BRAKE DEICE, on the overhead circuit breaker panel
(fig. 2-26).
2-59.
WINDSHIELD ELECTROTHERMAL ANTIICE
SYSTEM.
a. Description. Both pilot and copilot windshields
are provided with an electrothermal anti-ice system. Each
windshield is part of an independent electrothermal antiice system. Each system is comprised of the windshield
assembly with heating wires sandwiched between glass
panels, a temperature sensor attached to the glass, an
electrothermal controller, two relays, a control switch, and
two circuit breakers. Two switches, placarded WSHLD
ANTIICE NORMAL -OFF HI PILOT, COPILOT, located
on the overhead control panel (fig. 2-18) control system
operation.
Each switch controls one electrothermal
windshield system. The circuits of each system are
protected by a 5-ampere circuit breaker and a 50-ampere
circuit breaker which are not accessible to the flight crew.
The 50-ampere circuit breakers are located in the power
distribution panel under the floor ahead of the main spar.
The 5-ampere circuit breakers are located on panels
forward of the instrument panel.
b. Normal Operation.
Two levels of heat are
provided through the three position switches placarded
NORMAL in the aft position, OFF in the center position,
and HI after lifting the switch over a detent and moving it
to the forward position. In the NORMAL position, heat is
provided for the major portion of each windshield. In the
HI position, heat is provided at a higher watt density to a
smaller portion of the windshield. The lever lock feature
prevents inadvertent switching to the HI position during
system shutdown.
2-58. FUEL SYSTEM ANTI-ICING.
a. Description.
An oil-to-fuel heat exchanger,
located on each engine accessory case, operates
continuously and automatically to heat the fuel sufficiently
to prevent freezing of any water in the fuel. No controls
are involved. Two external fuel vents are provided on
each wing. One is recessed to prevent ice formation; the
other is electrically heated and is controlled by two toggle
switches on the overhead control panel placarded FUEL
VENT ON, LEFT or RIGHT (fig. 2-18). They are
protected by two 5-ampere circuit breakers, placarded
FUEL VENT HEAT, RIGHT or LEFT, located on the
overhead circuit breaker panel (fig. 2-26). Each fuel
control unit's pneumatic line is protected against ice by an
electrically heated jacket protected by a 7. 5ampere
circuit breaker located on the overhead circuit breaker
panel placarded FUEL CONTR HEAT, LEFT or RIGHT
(fig. 2-26).
2-60. PRESSURIZATION SYSTEM.
a. Pressure Differential. The pressure vessel is
designed for a normal working pressure differential of 6.
0 PSI, which will provide a cabin pressure altitude of 3870
feet at an aircraft altitude of 20,000 feet, and a nominal
cabin altitude of 9840 feet at an aircraft altitude of 31,000
feet.
b. Description. A mixture of bleed air from the
engines, and ambient air, is available for pressurization to
the cabin at a rate of approximately 10 to 17 pounds per
minute.
Approximately 85% NI is required when
operating with one engine. The flow control unit of each
engine controls the bleed air from the engine to make it
usable for pressurization by mixing ambient air with the
bleed air depending
CAUTION
To prevent overheat damage to
electrically heated anti-ice jackets,
the FUEL VENT heat switches
should not be turned ON unless
cooling air will soon pass over the
jackets.
2-45
TM 55-1510-219-10
safety valve will be opened and the cabin will be
depressurized to the aircraft altitude. In the PRESS
position, cabin altitude is controlled by the CABIN
CONTROLLER control. In the TEST position, the landing
gear safety switch is bypassed to enable testing of the
pressurization system on the ground.
Operating
instructions are contained in Chapter 8.
upon aircraft altitude and ambient temperature. On
takeoff, excessive pressure bumps are prevented by
landing gear safety switch actuated solenoids
incorporated in the flow control units. These solenoids,
through a time delay, stage the input of ambient air flow
by allowing ambient air flow introduction through the left
flow control unit first, ten seconds later, air flow through
the right flow control unit. The Bleed Air Switches,
located on the overhead control panel (fig. 2-1 8) operate
an integral electric solenoid which controls the bleed air to
the firewall shutoff valves.
d. Cabin Rate-of-Climb Indicator.
An indicator,
placarded CABIN CLIMB, is installed on the copilot's side
of the instrument panel (fig. 2-28). The cabin rate-ofclimb controller is calibrated in thousands-of-feet perminute change in cabin altitude.
c. Cabin Altitude and Rate-of-Climb Controller. A
control panel is installed on the copilot's inboard subpanel
for operation of the system. A knob (fig. 2-6), placarded
INC RATE controls the rate of change of pressurization.
A control, placarded CABIN CONTROLLER is used to set
the desired cabin altitude.
For proper cabin
pressurization, the CABIN CONTROLLER should be set
500 feet above cruise altitude. For landing select 500
feet above field pressure altitude. The selected altitude is
displayed on a mechanically coupled dial above the
control, placarded CABIN ALT-FT. Mechanically coupled
to the cabin altitude dial, placarded ACFTX1000. This
dial indicates the maximum altitude the aircraft may be
flown at to maintain the desired cabin altitude without
exceeding the design pressure differential. A switch,
placarded CABIN PRESS DUMP-PRESS-TEST, is
provided to control pressurization. The switch is spring
loaded to the PRESS position. In the DUMP position, the
e. Cabin Altitude Indicator. An indicator, placarded
CABIN ALT, is installed in the instrument panel (fig. 2-28)
above the cabin rate-of-climb indicator. The longer
needle indicates aircraft altitude in thousands-of-feet on
the outside dial. The shorter needle indicates pressure
differential in PSI on the inner dial. Maximum differential
is 6. 1 PSI.
f Outflow Valve. A pneumatically operated outflow
valve, located on the aft pressure bulkhead, maintains the
selected cabin altitude and rate-of-climb commanded by
the cabin rate-of-climb and altitude controller on the
copilot's instrument panel.
Figure 2-78. Overhead Control Panel
2-46
TM 55-1510-219-10
As the aircraft climbs, the controller modulates the outflow
valve to maintain a selected cabin rate of climb and
increases the cabin differential pressure until the
maximum cabin pressure differential is reached. At a
cabin altitude of 12,500 feet a pressure switch mounted
on the back of the overhead control panel completes a
circuit to illuminate a red warning annunciator light, ALT
WARN, to warn of operation requiring oxygen. This light
is protected by a 5ampere breaker, placarded PRESS
CONTR.
(3) The bleed air firewall shutoff valve in the
control unit is a spring loaded, bellows operated valve that
is held in the open position by bleed air pressure. When
the electric solenoid_ is shut off, or when bleed air
diminishes on engine shutdown (in both cases the
pressure to the firewall shutoff valve is cut off), the
firewall valve closes.
2-61. OXYGEN SYSTEM.
a. Description. The oxygen system (fig. 2-19) is
provided primarily as an emergency system, however, the
system may be used to provide supplemental (first aid)
oxygen. Two 64 cubic foot capacity oxygen supply
cylinders charged with aviator's breathing oxygen are
installed in the unpressurized portion of the aircraft behind
the aft pressure bulkhead. The pilot and copilot positions
are equipped with diluter demand type regulators, which
mix the proper amount of oxygen for a given amount of
air at altitude. Also a first aid oxygen mask is provided in
the cabin. Oxygen system pressure is shown by two
gages placarded OXYGEN SUPPLY PRESSURE, located
aft of the pilot's oxygen regulator control panel. Two
pressure reducers, located in the unpressurized portion of
the aircraft behind the aft bulkhead, lower the pressure in
the system to 400 PSI, and route oxygen to the regulator
control panels. Both cylinders are interconnected, so
refilling can be accomplished through a single filler valve
located on the aft right side of the fuselage exterior. A
pressure gage is mounted in conjunction with the filler
valve, and each cylinder has a pressure gage. Table 2-4
shows oxygen duration capacities of the system.
g. Pressurization Safety Valve. Before takeoff, the
safety valve is open with equal pressure between the
cabin and the outside air. The safety valve closes on
liftoff if the CABIN PRESS CONTR switch on the copilots
subpanel is in the PRESS mode. The safety valve
adjacent to the outflow valve provides pressure relief in
the event of failure of the outflow valve. This valve is
also used as a dump valve and is opened by vacuum
which is controlled by a solenoid valve operated by the
cabin pressure dump switch adjacent to the controller. It
is also wired through the right landing gear safety switch.
If either of these switches is open, or the vacuum source
or electrical power is lost, the safety valve will close to
atmosphere except at maximum differential pressure of 6.
1 PSI. A negative pressure relief diaphragm is also
incorporated into the outflow and safety valves to prevent
outside atmospheric pressure from exceeding cabin
pressure during rapid descent.
h. Drain. A drain in the outflow valve static control
line is provided for removal of accumulated moisture.
The drain is located behind the lower sidewall upholstery
access panel in the baggage section of the aft
compartment.
b. Regulator Control Panels. Each regulator control
panel contains a blinker-type flow indicator, a 500 PSI
pressure gage, a red emergency pressure control lever
placarded EMERGENCY NORMAL TEST MASK, a white
diluter control lever placarded 100% OXYGEN -NORMAL
OXYGEN, and a green supply control lever placarded ON
OFF.
i. Flow Control Unit. A flow control unit forward of
the firewall in each nacelle controls bleed air flow and the
mixing of ambient air to make up the total air flow to the
cabin for pressurization, heating, and ventilation. The
bleed air switches located on the overhead control panel
(fig. 2-18) operates an integral electric solenoid which
controls the bleed air to the firewall shutoff valves. A
normally open solenoid operated by the landing gear
safety switch controls the introduction of ambient air flow
to the cabin on takeoff.
(1) Diluter control lever. The diluter control
lever selects either normal or 100% oxygen, but acts to
select only when the emergency pressure control lever is
in the NORMAL position.
(1) The unit receives bleed air from the engine
into an ejector which draws ambient air into the nozzle of
the venturi. The mixed air is then forced into the bleed air
line routed to the cabin.
CAUTION
When not in use, the diluter control
lever should be left in the 100%
OXYGEN position to prevent
regulator contamination.
(2) Bleed air flow is controlled automatically.
When the aircraft is on the ground, circuitry from the
landing gear safety switch prevents ambient air from
entering the flow control unit to provide maximum
heating.
(2) Emergency pressure control lever. The
emergency pressure control lever has three positions.
2-47
TM 55-1510-219-10
Figure 2-19. Oxygen System Schematic
2-48
TM 55-1510-219-10
Table 2-4. Oxygen Flow Planning Rates vs Altitude
(All Flows in LPM Per Mask at NTPD)
Two positions control oxygen consumption for the
individual using oxygen, and the remaining position
serves for testing hose and mask integrity. In the
EMERGENCY position, the control lever causes 100%
oxygen to be delivered at a safe, positive pressure. In the
NORMAL position, the lever allows delivery of normal or
100% oxygen, depending upon the selection of the diluter
control lever. In TEST MASK position, 100% oxygen at
positive pressure is delivered to check hose and mask
integrity.
(3) Supply control lever. The supply control
lever (green), placarded ON OFF, turns the oxygen
supply on or off at the regulator control panel.
2-49
TM 55-1510-219-10
OXYGEN and continue breathing
undiluted oxygen until the danger
is past.
(4) Oxygen supply pressure gage.
WARNING
The 500 PSI oxygen pressure gage
provided on the oxygen control
panels should never indicate over
400 PSI. If the pressure exceeds
400 PSI, a malfunction of the
pressure reducer is indicated.
c. Oxygen Masks. Oxygen masks for the pilot and
copilot are provided as personal equipment. To connect a
mask into the oxygen system, the individual connects the
line attached to the mask to the flexible hose which is
attached to the cockpit sidewall. The microphone in the
oxygen mask is provided with a cord for connecting with
the helmet microphone jack. To test mask and hose
integrity, the individual places the supply control lever on
the regulator control panel to the ON position, puts on and
adjusts his mask, selects TEST MASK position, and
checks for leaks.
Whenever oxygen is inhaled, a blinker-vane
slides into view within the flow indicator window, showing
that oxygen is being released. When oxygen is exhaled,
the blinker vane vanishes from view.
d. Normal Operation.
Oxygen pressure is
maintained at all times to the regulator control panels if
the cylinder shut-off valves are on and if there is pressure
in the cylinders. Each individual places the supply lever
(green) on his regulator control panel to the ON position,
and the diluter lever (white) to the NORMAL OXYGEN
position.
NOTE
Check to insure that the OXYGEN
SUPPLY PRESSURE gage registers
adequate pressure before each
flight. When oxygen is in use, a
check of the supply pressure
should be made at intervals during
flight to note the quantity available
and to approximate the supply
duration. The outside temperature
is reduced as an aircraft ascends to
higher altitudes. Oxygen cylinders
thus cooled by temperature change
will show a pressure drop. This
type of drop in pressure will raise
again upon return to a lower or
warmer altitude. A valid cause for
alarm would be the rapid loss of
oxygen pressure when the aircraft
is in level flight or descending;
should
this
condition
arise,
descend as rapidly as possible to
altitude which does not require the
use of oxygen.
e. Emergency Operation. For emergency operation,
the affected crew member selects the EMERGENCY
position of the emergency pressure control lever (red) on
his regulator control panel. This selection provides 100%
oxygen at a positive pressure, regardless of the position
of the diluter control lever on his panel.
f First Aid Operation. A first aid oxygen mask is
installed in the aft cabin area as a supplemental or
emergency source of oxygen. The mask is stowed behind
an overhead cover placarded FIRST AID OXYGEN PULL.
Removing the cover allows the mask to drop out of the
container, exposing a manual control valve, which
releases oxygen to the mask when placed in the ON
position. After using the mask, the manual valve in the
container must be turned OFF before stowing the mask
and replacing the cover.
g. Oxygen Duration Example Problem.
WANTED
Duration in minutes of oxygen
at 100% capacity.
KNOWN
Two man crew plus one passenger, cabin pressure altitude =
15,000 feet, crew masks, normal,
100% capacity.
METHOD
Find "two man crew plus one
pass" line, move right then down
to 15,000 - "normal" read "232.
1" minutes.
WANTED
Duration of oxygen for previous
example data at 84% of capacity.
KNOWN
232.1 minutes duration at 100%,
84% capacity, total aircraft flow
= 13.9 LPM.
WARNING
Pure
oxygen
will
support
combustion. Do not smoke while
oxygen is in use.
WARNING
If any symptoms occur suggestive
of
the
onset
of
hypoxia,
immediately set the emergency
pressure control lever to the
EMERGENCY position and descend
below 10,000 feet.
Whenever
carbon monoxide or other noxious
gas is present or suspected, set the
dilutor control lever to 100%
2-50
TM 55-1510-219-10
Table 2-5. Oxygen Duration in Minutes 128 Cubic Foot System
METHOD
Multiply 232.1 X 0.84 = 194.9
minutes. or Multiply 3,226 X
0.84 = 2709.8, divide by 13.9
LPM = 194.9 minutes.
WANTED
Duration of oxygen for complement at other cabin pressure altitude, at less than 100% capacity.
KNOWN
Cylinder at 84% capacity, 100%
capacity = 3,226 L, cabin pressure altitude = 21,000 feet.
I crew mask = 7.2 LPM (100%),
I passenger mask = 3.7 LPM
METHOD
Multiply 3,226 L X 0.84 = 2.
709.8 L, multiply 2 crew X 7.2
LPM = 14.4 LPM, multiply 1
passenger X 3.7 LPM, add 14.4
LPM crew plus 3.7 LPM passenger = 18.1 LPM.Divide 3,226 L
by 18.1 LPM = 178.2 minutes.
h. Oxygen Cylinder Capacity Example Problem.
WANTED
a. Percent of capacity at known pressure and
temperature.
c. Enter 1600 PSIG move up to 20° C line, move
right to 84%.
d. Move left on 84% line to -30° C line, move down
to 1250 PSIG.
WANTED
KNOWN
METHOD
2-62. WINDSHIELD WIPERS.
a. Description. Two electrically operated windshield
wipers, are provided for use at takeoff, cruise and landing
speed.
A rotary switch (fig.
2-18) placarded
WINDSHIELD WIPER, located on the overhead control
panel, selects mode of windshield wiper operation. An
information placard above the switch states: DO NOT
OPERATE ON DRY GLASS. Function positions on the
switch, as read clockwise, are placarded: PARK OFF
SLOW FAST. When the switch is held in the springloaded PARK setting, the blades will return to their normal
inoperative position on the glass, then, when released, the
switch will return to OFF position terminating windshield
wiper operation. The FAST and SLOW switch positions
are separate operating speed
b. Pressure when temperature decreases.
KNOWN
100% capacity pressure at
known temperature.
Temperature -- -30°C.
Move left along 100% line to
-30°C line and move down to
1420 PSIG.
Pressure = 1,600 PSIG stabilized cylinder temperature is estimated at 20°C decreased stabilized cylinder temperature is estimated at -30°C.
METHOD
2-51
TM 55-1510-219-10
Figure 2-20. Cylinder Capacity vs Pressure and Temperature
settings for wiper operation. The windshield wiper circuit
is protected by one 10-ampere circuit breaker, placarded
WSHLD WIPER, located on the overhead circuit breaker
panel (fig. 2-26).
The pilot and copilot have individual cigarette
lighters and ash trays mounted in escutcheons outboard
of their seats. The cigarette lighters are protected by a 5ampere circuit breaker, placarded CIGAR LIGHTER, on
the overhead circuit breaker panel (fig. 2-26).
CAUTION
Do not operate windshield wipers
on dry glass. Such action can
damage the linkage as well as
scratch the windshield glass.
2-64. ELECTRIC TOILET.
a. Description. An electric toilet is installed in the aft
cabin area. The circuit is protected by a 10ampere circuit
breaker located in the power distribution panel under the
floor ahead of the main spar.
b. Normal Operation. To start, turn WINDSHIELD
WIPER switch to FAST or SLOW speed, as desired. To
stop, turn the switch to the PARK position and release.
The blades will return to their normal. inoperative position
and stop. Turning the switch only to the OFF position will
stop the windshield wipers, without returning them to the
normal inactive position.
b. Operation.
A switch, placarded PRESS TO
FLUSH, is mounted on the seat assembly for operation of
the toilet. Pressing the switch applies DC power to the
motor which drives the pump. The pump applies flushing
fluid through a nozzle in the upper rim and washes the
inner surface of the bowl. Waste is carried to the waste
tank mounted below the bowl.
When desired, the
removable waste tank may be removed from the toilet for
servicing.
2-63. CIGARETTE LIGHTERS AND ASH TRAYS.
2-52
TM 55-1510-219-10
2-65. SUN VISORS.
able. When not needed as a sun shield, each visor may
be manually rotated to a position flush with the top of
the cockpit so that it does not obstruct view through the
windows.
CAUTION
When adjusting the sun visors,
grasp only by the top metal
attachment to avoid damage to
the plastic shield.
2-66. RELIEF TUBES.
One relief tube is provided, located immediately aft of
the cabin door, on the left side of the fuselage.
A sun visor is provided for the pilot and copilot
respectively (fig. 2-9). Each visor is manually adjust
Section VIII. HEATING, VENTILATION, COOLING, AND ENVIRONMENTAL CONTROL SYSTEM
2-67. HEATING SYSTEM.
environmental flow control unit shutoff valve and the
pneumatic shutoff valve are open; when set to the
ENVIRO OFF position, the environmental flow control
unit shutoff valve is closed, and the pneumatic bleed air
valve is open; in the PNEU & ENVIRO OFF position,
both are closed. For maximum cooling on the ground,
turn the bleed air valve switches to the ENVIRO OFF
position.
a. Description. Warm air for heating the cockpit
and mission avionics compartments and warm
windshield defrosting air is provided by bleed air from
both engines (fig. 2-21). Engine bleed air is combined
with ambient air in the heating and pressurization flow
control unit in each nacelle. If the mixed bleed air is too
warm for cockpit comfort, it is cooled by being routed
through an air-to-air heat exchanger located in the
forward portion of each inboard wing. If the mixed bleed
air is not too warm, the air-to-air heat exchangers are
bypassed. The mixed bleed air is then ducted to a
mixing plenum, where it is mixed with cabin recirculated
air. The warm air is then ducted to the cockpit outlets,
windshield defroster outlets, and to the floor outlets in
the mission avionics compartment.
(4) Cabin temperature mode selector switch.
A switch placarded CABIN TEMP MODE MAN COOL
MAN HEAT OFF AUTO A/C COLD OPN, located on the
overhead control panel, controls cockpit and mission
avionics compartment heating and air conditioning.
When the cabin temperature mode selector switch is set
to the AUTO position, the heating and air conditioning
systems are automatically controlled. Control signals
from the temperature control box are transmitted to the
bleed air heat exchanger bypass valves. Here the
temperature of the air flowing to the cabin is regulated
by the bypass valves controlling the amount of air
bypassing the heat exchangers.
When the cabin
temperature mode selector switch is set to the AUTO
position, the heating and air conditioning systems are
automatically controlled. When the temperature of the
cabin has reached the temperature setting of the cabin
temperature control rheostat, the automatic temperature
control allows hot air to bypass the air-to-air exchangers.
When both bypass valves are in the fully closed
position, allowing no air to bypass the heat exchangers,
the air conditioner begins to operate, providing
additional cooling. When the cabin temperature mode
selector switch is set to the A/C COLD OPN position, the
air conditioning system is in continuous operation. The
cabin temperature control rheostat, in conjunction with
the cabin temperature control sensor, provides
regulation of cockpit and mission equipment
compartment temperature. Bleed air heat is added as
re
(1) Bleed airflow control unit. A bleed air flow
control unit, located forward of the firewall in each
engine nacelle controls the flow of bleed air and the
mixing of ambient air to make up the total airflow to the
cabin for heating, windshield defrosting, pressurization
and ventilation. The unit is fully pneumatic except for
an integral electric solenoid firewall shutoff valve,
controlled by the bleed air switches located on the
overhead control panel (fig. 2-18) and a normally open
solenoid valve operated by the left landing gear safety
switch.
(2) Pneumatic bleed air shutoff valve. A
pneumatic shutoff valve is provided in each nacelle to
control the flow of bleed air to the surface, antenna and
brake device systems. These valves are controlled by
the bleed air valve switches located on the overhead
control panel (fig. 2-18).
(3) Bleed air valve switches. The bleed air
flow control unit shutoff valve and pneumatic bleed air
shutoff valves are controlled by two switches placarded
BLEED AIR VALVE OPEN ENVIRO OFF -PNEU &
ENVIRO OFF, located on the overhead control panel
(fig, 2-18). When set to the open position, both the
2-53
TM 55-1510-219-10
Figure 2-21. Environmental Systems
2-54
TM 55-1510-219-10
quired to maintain the temperature selected by the cabin
temperature control rheostat.
4. Cabin, cockpit and defrost air knobs
As required
c. Cabin Heating Mode.
1. Bleed air valve switches OPEN,
LEFT and RIGHT.
(5) Cabin temperature control rheostat. A
control knob placarded CABIN TEMP INCR, located on
the overhead control panel (fig.
2-18), provides
regulation of cabin temperature when the cabin
temperature mode selector switch is set to the AUTO
position, or the A/C COLD OPN position.
A e
temperature sensing unit in the cabin, in conjunction
with the setting of the cabin temperature control
rheostat, initiates a heat or cool command to the
temperature controller for the desired cockpit or mission
avionics compartment environment.
2. Cabin temperature mode selector
switch MAN HEAT.
3. Vent blower switches As required.
4. Manual temperature
required.
(6) Manual temperature control switch. A
switch placarded MANUAL TEMP INCR DECR, located
on the overhead control panel (fig. 2-18), controls
cockpit and mission avionics compartment temperature
with the cabin temperature mode selector switch set to
the MAN HEAT positions. The manual temperature
control switch controls cockpit and mission avionics
temperature by providing a means of manually changing
the amount that the bleed air bypass valves are opened
or closed. To increase cabin temperature the switch is
held to the INCR position.
To decrease cabin
temperature, the switch is held to the DECR position.
Approximately 30 seconds per valve is required to drive
the bypass valves to the fully open or fully closed
position. Only one valve moves at a time.
switch
As
5. Cabin, cockpit and defrost air knobs
As required.
2-68. AIR CONDITIONING SYSTEM.
a. Description. Cabin air conditioning is provided
by a refrigerant gas vapor cycle refrigeration system
consisting of a belt driven, engine mounted compressor,
installed on the #2 engine accessory pad, refrigerant
plumbing, N1 speed switch, high and low pressure
protection switches, condenser coil, condenser blower,
forward and aft evaporator, receiver dryer, expansion
valve and a bypass valve. The plumbing from the
compressor is routed through the right inboard wing
leading edge to the fuselage and then forward to the
condenser coil, receiver-dryer, expansion valve, bypass
valve, and forward evaporator, which are located in the
nose of the aircraft. A circuit breaker placarded AIR
COND CONTR, located on the overhead circuit breaker
panel (fig. 226), protects the compressor clutch circuit.
(7) Forward vent blower switch. The forward
vent blower is controlled by a switch placarded VENT
BLOWER AUTO LO HI, located on the overhead control
panel (fig. 2-18). In the auto position the fan will run at
low speed except when the cabin temperature mode
selector switch is set to the OFF position, in this case
the blower will not operate.
(1) Forward evaporator.
The forward
evaporator and blower supplies the cockpit, forward
ceiling outlets, and forward floor outlets. The forward
evaporator blower has a high speed which can be
selected by setting the VENT BLOWER switch, located
on the overhead control panel (fig. 2-18), to the HI
position. The forward vent blower is protected by a
circuit breaker located on the DC power distribution
panel, located in the forward equipment bay.
(8) Aft vent blower switch. The aft vent
blower is controlled by a switch placarded AFT VENT
BLOWER OFF AUTO ON, located on the overhead
control panel (fig. 2-18). The single speed blower
operates automatically through the cabin temperature
mode selector switch when the aft vent blower switch is
placed in the AUTO position.
The blower runs
continuously when the switch is placed in the ON
position, In the OFF position, the blower will not operate.
(2) Aft evaporator. The aft evaporator and
blower are located in the fuselage center aisle
equipment bay aft of the rear spar. Environmental air is
circulated through the evaporator in either manual or
automatic control mode. The rear evaporator supplies
the aft ceiling outlets, rear floor outlets, and toilet
compartment. Rear evaporator blower is protected by a
circuit breaker located on the DC power distribution
panel in the lower equipment bay.
b. Automatic Heating Mode.
1. Bleed air valve switches OPEN,
LEFT and RIGHT.
2. Cabin temperature mode selector
switch AUTO.
(3) High and low pressure limit switches.
High and low pressure limit switches are provided to
prevent compressor operation beyond operational
3. Cabin temperature control rheostat
As required.
2-55
TM 55-1510-219-10
limits. When the low or high pressure switches are
activated, a reset switch/light assembly located in the
nosewheel well is activated to prevent further
compressor operation.
2. Cabin temperature mode selector
switch MAN COOL.
(3) Air conditioning cold start mode. (Used if
ambient temperature is between 10°C and -25°C).
(4) Thermal sense switch. A 33° thermal
sense switch is installed on the forward evaporator.
This sense switch actuates a hot gas bypass valve
which bleeds off a portion of the refrigerant from the
forward evaporator, thereby preventing icing of the
evaporator.
NOTE
Setting the cabin temperature
mode selector switch to the A/C
COLD OPN position at ambient
temperatures below -25°C may
cause the air conditioner system
to exceed the compressor low
pressure limit switch setting,
illuminating the reset switch in
the nosewheel well, thereby
rendering the system inoperative
for the remainder of the flight.
(5) Condenser blower. A vane-axial blower
draws air through the condenser on the ground as well
as in flight. The current limiter for this blower is located
on the DC distribution panel in the lower equipment bay.
When the cabin temperature mode selector switch is set
to the A/C COLD OPN position, the condenser blower
will be off, and will remain off until the added high
pressure switch senses a compressor discharge
pressure equal to the pressure it is set to. The
condenser blower will then remain in operation until the
low pressure switch senses that the system pressure has
dropped to the pressure it is set to.
1. Bleed air valve switches OPEN,
LEFT and RIGHT.
2. Cabin temperature mode selector
switch A/C COLD OPN.
(6) Air conditioning cold start bypass valve.
Low ambient temperature operation of the air
conditioning system is accomplished by means of a
bypass valve located in the nose wheel well, an
additional high pressure switch, and an additional low
pressure switch, both of which are located in the right
inboard wing leading edge. The cold start bypass valve
opens when the cabin temperature mode selector switch
is set to the A/C COLD OPN position to enhance
refrigerant flow at low temperatures and improve air
conditioner performance.
3. Cabin temperature control rheostat
As required.
4. Cabin, cockpit and defrost air knobs
As required.
2-69. UNPRESSURIZED VENTILATION.
Ventilation is provided by two sources. One
source is through the bleed air heating system in both
the pressurized and unpressurized mode. The second
source of ventilation is obtained from ram air through
the condenser section in the nose through a check valve
in the vent blower plenum. Ventilation from this source
is in the unpressurized mode only with the CABIN
PRESS switch in the DUMP position. The check valve
closes during pressurized operation. Ram air ventilation
is distributed through the main ducting system to all
outlets. Ventilation air, ducted to each individual eyeball
cold air outlet, can be directionally controlled by moving
the ball in the socket. Volume is regulated by twisting
the outlet to open or close the valve.
b. Normal Operation.
(1) Automatic cooling mode.
1. Bleed air valve switches OPEN,
LEFT and RIGHT.
2. Cabin temperature mode selector
switch AUTO.
3. Cabin temperature control rheostat
As required.
2-70. ENVIRONMENTAL CONTROLS.
4. Cabin, cockpit and defrost air knobs
As required.
An environmental control section on the overhead
control panel (fig. 2-18) provides for automatic or
manual control of the system. This section contains all
the major controls of the environmental function
including bleed air valve switches, a vent blower control
switch, an aft vent blower switch, a manual temperature
switch for control of the heat exchanger valves, a cabin
temperature level control,
(2) Manual cooling mode.
1. Bleed air valve switches OPEN,
LEFT and RIGHT.
NOTE
For maximum cooling on the ground,
set the bleed air valve switches to the
ENVIRO OFF position.
2-56
TM 55-1510-219-10
and the cabin temp mode selector switch for selecting
automatic heating or cooling or manual heating or
cooling. Four additional manual controls on the main
instrument subpanels may be utilized for partial
regulation of cockpit comfort when the cockpit curtain is
closed and the cabin comfort level is satisfactory.
a. Heating Mode.
section. The warm bleed air is mixed with the cooled
air. The rear evaporator picks up recirculated cabin air
only.
(1) When the automatic control drives the
environmental system from a heat mode to a cooling
mode, the bypass valves close. When the left bypass
valve reaches a fully closed position, the refrigeration
system provided the right engine N1 speed is above
62°. When the bypass valve is opened to a position
approximately 30° from full closed, the refrigeration
system will turn off.
(1) If the cockpit is too cold:
1. Pilot and copilot air knobs As
required.
2. Defrost air knob As required.
(2) The CABIN TEMP INCR control provides
regulation of the temperature level in the automatic
mode. A temperature sensing unit in the cabin, in
conjunction with the control setting, initiates a heat or
cool command to the temperature controller for desired
cockpit and cabin environment.
3. Cabin air knob Pull out in small
increments. Allow 3 5 minutes after
each adjustment for system to
stabilize.
(2) If the cockpit is too hot:
1. Cabin air knob As required.
d. Manual Mode Control.
With the cabin
temperature mode selector in the MAN HEAT or MAN
COOL position, regulation of the cabin temperature is
accomplished manually with the MANUAL TEMP switch.
2. Pilot and copilot air knobs In as
required.
3. Defrost air knob In as required.
(1) In the MAN HEAT mode, the automatic
system is overridden and the system is controlled by
opening and closing the bypass valves (two) with the
MANUAL TEMP INCR DECR switch. To increase cabin
temperature, hold the switch at the INCR position, to
decrease cabin temperature, hold the switch in the
DECR position. Allow approximately 30 seconds per
valve to drive the bypass valves to the fully open or fully
closed position. Only one valve moves at a time.
b. Cooling Mode.
(1) If the cockpit is too cold:
1. Pilot and copilot air knob In as
required.
2. Defrost air knob In as required.
3. Overhead cockpit outlets As
required.
(2) If the cockpit is too hot:
1. Pilot and copilot air knobs Out as
required.
(2) With the cabin temperature selector
switch in the MAN COOL position, the automatic
temperature control system is bypassed. In the manual
cooling mode, the refrigeration system is on, providing
the right engine right engine N1speed is above 62%,
however, the bypass valves may be manually positioned
for the desired temperature. Hold the MANUAL TEMP
switch in the DECR position approximately one minute
to fully close air-to-air heat exchanger bypass valves.
2. Cabin air knob. Close in small
increments. Allow 3 5 minutes after
each adjustment for system to
stabilize. If CABIN AIR knob is
completely closed before obtaining
satisfactory cockpit comfort, it may
be necessary to place the aft vent
blower switch in the ON position to
activate the aft evaporator to
recirculate cabin air.
e. Bleed Air and Vent Control.
(1) Bleed air entering the cabin is controlled
by bleed air valve switches placarded BLEED AIR
VALVE OPEN ENVIRO OFF PNUE & ENVIRO OFF.
When the switch is in the OPEN position, the
environmental flow control unit and the pneumatic valve
are open. When the switch is in the ENVIRO OFF
position, the environmental flow control unit is closed
and the pneumatic bleed air valve is open; in the PNEU
& ENVIRO OFF position, both are closed.
For
maximum cooling on the
c. Automatic Mode Control. When the AUTO
mode is selected on the cabin temperature mode
selector switch, the heating and air conditioning systems
are automatically controlled. When the temperature of
the cabin has reached the selected setting, the
automatic temperature control allows heated air to
bypass the air-to-air exchangers in the wing center TM
2-57
TM 55-1510-219-10
low speed except when the CABIN TEMP mode selector
switch is placed in the OFF position. In the OFF
position, the blower will not operate.
ground, turn the bleed air valve switches to the ENVIRO
OFF position.
(2) The forward vent blower is controlled by a
switch placarded VENT BLOWER AUTO LOW HI. The
HI and LOW positions regulate the blower to two speeds
of operation. IN the AUTO position, the fan will run at
(3) The aft vent blower is controlled by a
switch placarded AFT VENT BLOWER OFF AUTO ON.
In the OFF position, the blower will not operate.
Section IX. ELECTRICAL POWER SUPPLY AND DISTRIBUTION SYSTEM
indicators associated with the DC supply system are
located on the overhead control panel (fig. 2-18) and
consists of a single battery switch (BATT), two generator
switches (#1 GEN and #2 GEN), and two voltloadmeters.
2-71. DESCRIPTION.
The aircraft employs both direct current (DC) and
alternating current (AC) electrical power. The DC
electrical supply is the basic power system energizing
most aircraft circuits. Electrical power is used to start
the engines, to power the landing gear and flap motors,
and to operate the standby fuel pumps, ventilation
blower, lights and electronic equipment. AC power is
obtained from DC power through inverters. The three
sources of DC power consist of one 20 cell 34ampere/hour battery and two 400ampere startergenerators. DC power may be applied to the aircraft
through an external power receptacle on the right
nacelle.
The starter-generators are controlled by
generator control units. The output of each generator
passes through a cable to the respective generator bus
(fig. 2-22). Other buses distribute por to aircraft DC
loads, and derive power from the generator buses. The
generators are paralleled to balance the DC loads
between the two units. When one of the generators is
not on line, aircraft DC power requirements continue to
be supplied from one of the other generators. Most DC
distribution buses are connected to both generator
buses but have isolation diodes to prevent power
crossfeed between the generating systems, when
connection between the generator buses is lost. Thus,
when either generator is lost because of a ground fault,
the operating generator will supply power for all aircraft
DC loads except those receiving power from the
inoperative generator's bus which cannot be crossfed.
When a generator is not operating, reverse current and
over-voltage protection is automatically provided. Two
inverters operating from DC power produce the required
single-phase AC power (figure 2-23). Three phase AC
electrical power for inertial navigation system and
mission avionics is supplied by two DC powered
inverters (figures 2-24 and 2-25).
a. Battery Switch. A switch, placarded BATT is
located on the overhead control panel (fig. 2-18) under
the MASTER SWITCH. The BATT switch controls DC
power to the aircraft bus system through the battery
relay, and must be ON to allow external power to enter
aircraft circuits. When the MASTER SWITCH is moved
aft, the BATT switch is forced OFF.
NOTE
With battery or external power
removed
from
the
aircraft
electrical system, due to fault,
power cannot be restored to the
system until the BATT switch is
moved to OFF/RESET, then ON.
b. Generator Switches. Two switches (fig. 218),
placarded #1 GEN and #2 GEN are located on the
overhead control panel under the MASTER SWITCH.
The toggle switches control electrical power from the
designated generator to paralleling circuits and the bus
distribution system. Switch positions are placarded
RESET, ON and OFF. RESET is forward (springloaded back to ON), ON is center, and OFF is aft.
When a generator is removed from the aircraft electrical
system, due either to fault or from placing the GEN
switch in the OFF position, the affected unit cannot have
its output restored to aircraft use until the GEN switch is
moved to RESET, then ON.
2-72. DC POWER SUPPLY.
c. Master Switch. All electrical current may be
shut off using the MASTER SWITCH gangbar (fig.
One nickel-cadmium battery furnishes DC power
when the engines are not operating. This 24-volt, 34ampere/hour battery, located in the right wing center
section, is accessible through a panel on the 2-58 top of
the wing. DC power is produced by two engine-driven
28 volt, 400-ampere starter-generators. Controls and
2-18) which extends below the battery and generator
switches. The MASTER SWITCH gangbar is moved
forward when a battery or generator switch is turned on.
When moved aft, the bar forces each switch to the OFF
position.
2-58
TM 55-1510-219-10
Figure 2-22. DC Electrical System (Sheet 1 of 3)
2-59
TM 55-1510-219-10
Figure 2-22. DC Electrical System (Sheet 2 of 3)
2-60
TM 55-1510-219-10
Figure 2-22. DC Electrical System (Sheet 3 of 3)
2-61
TM 55-1510-219-10
Figure 2-23. Single Phase AC Electrical System
2-62
TM 55-1510-219-10
Figure 2-24. Three Phase AC Electrical System
2-63
TM 55-1510-219-10
Figure 2-25. Mission Equipment Power System
2-64
TM 55-1510-219-10
Figure 2-26. Overhead Circuit Breaker Panel
extinguish within two to five minutes, indicating that the
battery is approaching a full charge. The time interval
will increase if the battery has a low state of charge, the
battery temperature is very low, or if the battery has
previously been discharged at a very low rate (i.e.,
battery operation of radios or lights for prolonged
periods). The caution light may also illuminate for short
intervals after landing gear and/or flap operation. If the
caution light should illuminate during normal steadystate cruise, it indicates that conditions exist that may
cause a battery thermal runaway. If this occurs, the
battery switch shall be turned OFF and may be turned
back ON only for gear and flap extension and approach
to landing. Battery may be used after a 15 to 20 minute
cool down period.
d. Volt-Loadmeters. Two meters (fig. 2-18), on
the overhead control panel display voltage readings and
show the rate of current usage from left and right
generating systems. Each meter is equipped with a
spring-loaded push-button switch which when manually
pressed will cause the meter to indicate main bus
voltage. Each meter normally shows output amperage
reading from the respective generator, unless the pushbutton switch is pressed to obtain bus voltage reading.
Current consumption is indicated as a percentage of
total output amperage capacity for the generating
system monitored.
e. Battery Monitor.
Nickel-cadmium battery
overheating will cause the battery charge current to
increase if thermal runaway is imminent. The aircraft
has a charge-current sensor which will detect a charge
current. The charge current system senses battery
current through a shunt in the negative lead of the
battery. Any time the battery charging current exceeds
approximately 7-amperes for 6 seconds or longer, the
yellow BATTERY CHARGE annunciator light and the
master fault caution light will illuminate. Following a
battery engine start, the caution light will illuminate
approximately six seconds after the generator switch is
placed in the ON position. The light will normally
f. Generator Out Warning Lights.
Two
caution/advisory annunciator panel lights inform the pilot
when either generator is not delivering current to the
aircraft DC bus system. These lights are placarded #1
DC GEN and #2 DC GEN. Two MASTER CAUTION
lights and illumination of either fault light indicates that
either the identified generator has failed or voltage is not
sufficient to keep it connected to the power distribution
system.
2-65
TM 55-1510-219-10
c.
Instrument AC Light. A red warning light,
located on the warning annunciator panel, placarded
INST AC, will illuminate if all instrument AC busses
should fail.
CAUTION
The GPU shall be adjusted to
regulate at 28 volts maximum to
prevent damage to the aircraft.
d. Inverter Control Switches. Two switches (fig.
2-18), placarded INVERTER #1 and #2 on the overhead
control panel give the pilot control of inverters singlephase AC power.
g. DC External Power Source. External DC power
can be applied to the aircraft through an external power
receptacle on the underside of the right wing leading
edge just outboard of the engine nacelle.
The
receptacle is installed inside of the wing structure and is
accessible through a hinged access panel. DC power is
supplied through the DC external plug and applied
directly to the battery bus after passing through the
external power relay. Turn off all external power while
connecting the power cable to, or removing it from, the
external power supply receptacle. The holding coil
circuit of the relay is energized by the external power
source when the BATT switch is in the ON position. The
GPU shall be adjusted to regulate at 28 volts maximum
to prevent damage to the aircraft battery.
e. Volt-Frequency Meters.
Two volt-frequency
meters (fig. 2-18) are mounted in the overhead control
panel to provide monitoring capability for both 115 VAC
buses. Normal display on the meter is shown in
frequency (Hz). To read voltage, press the button
located in the lower left corner of the meter. Normal
output of the inverters will be indicated by 115 VAC and
400 Hz on the meters.
f Three Phase AC Power Supply. Three phase
AC electrical power for operation of the inertial
navigation system and mission avionics is supplied by
either of two DC powered 3000 volt-ampere solid state
three phase inverters (fig. 2-24).
h. Security Keylock Switch. The aircraft has a
security keylock switch (fig. 2-18) installed on the
overhead control panel, placarded OFF ON. The switch
is connected to the battery relay circuit and must be ON
when energizing the battery master power switch. The
key cannot be removed from the lock when in the ON
position.
(1) Three phase inverter control switch. A
three position switch placarded #1 MISSION INV, OFF,
#2 MISSION INV, located on the mission control panel
controls three phase inverter operation.
i. Circuit Breakers. The overhead circuit breaker
panel (fig. 2-26) contains circuit breakers for most
aircraft systems. The circuit breakers on the panel are
grouped into areas which are placarded as to the
general function they protect. A DC power distribution
panel is mounted beneath the floor forward of the main
spar. This panel contains higher current rated circuit
breakers and is not accessible to the flight crew under
normal conditions.
(2) Three phase volt-frequency meter. A
three phase volt/frequency meter, mounted on the
mission control panel, monitors output of the selected
three phase inverter.
Frequency (Hz) is normally
displayed on the meter. To read voltage press button
located in the lower left corner of the meter.
(3) Three phase loadmeter. A three phase
loadmeter, mounted on the mission control panel,
monitors inverter output level.
2-73. AC POWER SUPPLY.
a. Single Phase AC Power Supply. AC power for
the aircraft is supplied by inverter units, numbered #1
and #2 (fig. 2-23) which obtain operating current from
the DC power system. Both inverters are rated at 750
volt-amperes and provide single phase output only.
Each inverter provides 115 volt and 26 volt 400 Hz AC
output. The inverters are protected by circuit breakers
mounted on the DC power distribution panel mounted
beneath the floor. Controls and indicators of the AC
power system are located on the overhead control panel
and on the caution/advisory annunciator panel.
(4) Three phase AC off annunciator light. An
indicator light placarded 3( AC OFF, located on the
mission annunciator panel indicates that three phase AC
power is not being supplied.
(5) Three phase AC external power. External
three phase AC power for operation of the inertial
navigation system or mission equipment, can be applied
to the aircraft through an external power receptacle
located on the underside of the left wing leading edge
just outboard of the engine nacelle. The receptacle is
installed inside the wing structure and is accessible
through a hinged access panel. The AC electrical
system is automatically isolated from the external power
source if the external power is over or under voltage,
over or under frequency, or has an improper phase
sequence.
b. AC Power WARNING/CAUTION Lights. Two
MASTER/CAUTION lights and the illumination of an
annunciator caution light #1 INVERTER or #2
INVERTER indicates an inverter failure.
2-66
TM 55-1510-219-10
(b) External AC power control switch.
A switch placarded AC EXT POWER, located on the
mission control panel controls application of three phase
AC power to the aircraft.
(a) External AC power annunciator
light. An annunciator light placarded EXT AC PWR ON,
located on the mission annunciator panel indicates that
an AC GPU plug is mated to the AC external power
receptacle and the External AC Power control switch is
On.
Section X. LIGHTING
2-74. EXTERIOR LIGHTING.
the lights is provided by a switch placarded ICE on the
overhead control panel (fig. 2-18). Prolonged use
during ground operation may generate enough heat to
damage the lens.
a. Description.
Exterior lighting (fig.
2-27)
consists of a navigation light on top of the aft section of
the vertical stabilizer, one navigation light on top and
bottom of each wing tip pod, two strobe beacons, one on
top of the vertical stabilizer and one on the underside of
the fuselage center section, dual landing lights, one taxi
light mounted on the nose gear assembly, a recognition
light located in each wing tip, and two ice lights, one
light flush mounted in each nacelle, positioned to
illuminate along the leading edge of each outboard wing.
f. Recognition Lights. A RECOG switch, located
in the pilot's subpanel LIGHTS section, controls a white
recognition light in each wing tip. When requested, this
steady, bright light is used for identification in the traffic
pattern. The recognition lights circuit is protected by a 7
1/2 ampere RECOG circuit breaker located on the
overhead circuit breaker panel (fig. 2-26).
b. Navigation Lights. The navigation lights are
protected by a 5-ampere circuit breaker placarded NAV
on the overhead circuit breaker panel (fig. 2-26).
Control of the lights is provided by a switch placarded
NAV-ON on the overhead control panel (fig. 2-18).
2-75. INTERIOR LIGHTING.
Lighting systems are installed for use by the pilot
and copilot and by the passengers in the cabin area.
The lighting systems in the cockpit are provided with
intensity controls on the overhead control panel. A
switch placarded MASTER PANEL LIGHTS on the
overhead control panel (fig. 2-18) provides overall onoff control for all engine instrument lights, pilot and
copilot instrument lights, overhead panel lights, console
and subpanel lights and the free air temperature light.
c. Strobe Beacons. The strobe beacons are dual
intensity units. They are protected by a 15ampere
circuit breaker placarded BCN on the overhead circuit
breaker panel (fig. 2-26). Control of the lights is
provided by a switch placarded BEACONDAY-NIGHT
(fig. 2-18). Placing the switch in the DAY position will
activate the high intensity white section of the strobe
lights for greater visibility during daytime operation.
Placing the switch in the NIGHT position activates the
lower intensity red section of the strobe lights.
a. Cockpit Lighting.
(1) Flight instrument lights. Each individual
flight instrument contains internal lamps for illumination.
The circuit is protected by a 7 1/2-ampere circuit
breaker placarded FLT INST on the overhead circuit
overhead circuit breaker panel (fig. 226). Control is
provided by two rheostat switches placarded PILOT or
COPILOT INST LIGHTS-OFF-BRT on the overhead
control panel (fig. 2-18). Turning the control clockwise
from OFF turns the lights on and increases their
brilliance.
d. Landing/Taxi Lights. Dual landing lights and a
single taxi light are mounted on the nose gear assembly.
The lights are controlled by switches, placarded
LANDING and TAXI, located in the LIGHTS section of
the pilot's subpanel.
In addition, these light are
extinguished whenever the landing gear is retracted.
The landing/taxi lights circuits are protected by 5ampere circuit breakers placarded LANDING and TAXI
respectively, located on the overhead circuit breaker
panel (fig.
2-26).
Additional circuit protection is
provided by 35-ampere and 15ampere circuit breakers
on the DC power distribution panel located beneath the
cockpit floor.
(2) Instrument indirect lights. Three lights are
mounted in the glareshield overhang along the top edge
of the instrument panel and provide overall instrument
panel illumination. The circuit is protected by a 5ampere circuit breaker placarded INST INDIRECT on
the overhead circuit breaker panel (fig. 2-26). Control
is provided by a rheostat switch placarded INST
INDIRECT LIGHTS-OFF-BRT on the overhead control
panel (fig. 2-18). Turning the
e. Ice Lights. The ice lights circuit is protected by
a 5-ampere circuit breaker placarded ICE on the
overhead circuit breaker panel (fig. 2-26). Control of
2-67
TM 55-1510-219-10
Figure 2-27. Exterior Lighting
2-68
TM 55-1510-219-10
by a 10-ampere circuit breaker placarded CABIN on the
overhead circuit breaker panel (fig. 2-26). Control is
provided by a switch placarded CABIN LIGHTSBRIGHT-DIM-OFF on the subpanel (fig. 2-6).
control clockwise from OFF turns the lights on and
increases their brilliance.
(3) Engine instrument lights. Each individual
engine instrument contains internal lamps for
illumination. The circuit is protected by a 7 1/2ampere
circuit breaker placarded FLT INST on the overhead
circuit breaker panel (fig. 2-26). Control is provided by
a rheostat switch placarded ENGINE INST LIGHTS OFF
BRT on the overhead control panel (fig. 2-18). Turning
the control clockwise from OFF turns the lights on and
increases their brilliance.
(2) Threshold and spar cover lights. A spar
cover light is installed on the left side of the sunken aisle
immediately aft of the main spar cover. Both circuits
are protected by a 5-ampere circuit breaker mounted
beneath the battery and connected to the emergency
battery bus. Both lights are controlled by the switch
mounted forward of the airstair door. If the lights are
illuminated, closing the cabin door will automatically
extinguish them.
(4) Flood light. A single overhead flood light
is installed. It provides overall illumination of the entire
cockpit area. The circuit is protected by a 5-ampere
circuit breaker mounted beneath the battery and
connected to the emergency battery bus. Control is
provided by a rheostat switch placarded OVERHEAD
FLOODLIGHT-OFF-BRT on the overhead control panel
(fig. 2-18). Turning the control clockwise from OFF
turns the light on and increases its brilliance.
(3) Dome light. A dome light is installed in
the baggage area, in the overhead. The circuit is
protected by a 5-ampere circuit breaker mounted
beneath the battery and connected to the emergency
battery bus. Control is provided by a switch mounted
adjacent to the light.
(4) Cabin utility light. There is a cabin utility
light adjacent to each cabin light. Each utility light is
individually controlled by a rheostat placarded OFF-ONBRT on the back of the light. There is a momentary ON
switch in the center of the rheostat. Each light is
capable of producing a red or white spotlight by turning
the selector on the front of the light. To remove the light
from the stationary position, loosen the retaining screw
directly below the light escutcheon and pull down on the
light. The light is connected to the light housing by an
11 inch coiled cord that extends to approximately 50
inches. The CABIN LIGHTS switch must be on for utility
lamps to operate.
(5) Overhead panel lights. Lamps on the
overhead circuit breaker panel, control panel, and fuel
management panel are protected by a 7 1/2ampere
circuit breaker placarded LIGHTS OVHD on the
overhead circuit breaker panel (fig. 2-26). Control is
provided by a rheostat switch placarded OVERHEAD
PANEL LIGHTS-OFF-BRT on the overhead control
panel (fig. 2-18). Turning the control clockwise from off
turns the lights on and increases their brilliance.
(6) Subpanel and console lights. Lights on
the pilot's and copilot's subpanels, console edge lit
panels and pedestal extension panels are protected by a
7 1/2-ampere circuit breaker placarded LIGHTS
SUBPANEL & CONSOLE on the overhead circuit
breaker panel (fig. 2-26). Control is provided by two
rheostat switches placarded SUBPANEL or CONSOLE
LIGHTS-OFF-BRT on the overhead control panel (fig.
2-18). Turning the control clockwise from OFF turns the
lights on and increases their brilliance.
2-76. EMERGENCY LIGHTING.
a. Description. An independent battery operated
lighting system is installed. The system is actuated
automatically by shock, such as a forced landing. It
provides adequate lighting inside and outside the
fuselage to permit the crew to read instruction placards
and locate exits. An inertia switch, when subjected to a
2 3 G shock, will illuminate interior lights in the cockpit,
forward and aft cabin areas, and exterior lights aft of the
emergency exit and aft of the cabin door. The battery
power source is automatically recharged by the aircraft
electrical system.
(7) Free air temperature light. Two post lights
are mounted adjacent to the free air temperature gage
on the left cockpit sidewall trim panel. The circuit is
protected by a 7 1/2-ampere circuit breaker placarded
FLT INST on the overhead circuit breaker panel (fig. 226). Control is provided by a push button switch
adjacent to the gage. No intensity control is provided.
b. Operation.
An emergency lights override
switch, located on the overhead control panel (fig. 218),
is provided to turn the system off if it is accidentally
actuated. The switch is placarded EMERG LIGHTS
OVRD OFF-RESET-AUTO-TEST. Should the system
accidentally actuate, placing the switch in the
momentary OFF RESET position will extinguish
b. Cabin Lighting.
(1) Interior lights. Three cabin lights are
installed in the overhead trim. The circuit is protected
2-69
TM 55-1510-219-10
the lights. To test the system, place the switch in the
momentary TEST position. The lights should illuminate.
Moving the switch to the OFF-RESET position will turn
the system off and reset it.
Section XI. FLIGHT INSTRUMENTS
2-77. TURN-AND-SLIP INDICATORS.
2-80. COPILOT'S ALTIMETER.
Turn-and-slip indicators are installed separately
on the pilot and copilot sides of the instrument panel
(fig. 2-28). The pilot's indicator provides yaw damping
information to the autopilot.
These indicators are
gyroscopically operated. They use DC power and are
protected by 5-ampere circuit breakers placarded TURN
& SLIP PILOT or COPILOT on the overhead circuit
breaker panel (fig. 2-26).
The copilot's altimeter is located on the upper
right side of the instrument panel (fig. 2-28). It displays
altitude by means of a 10,000 foot counter, a 1000 foot
counter, a 100 foot counter, and a single needle pointer
that indicates on a circular scale marked in 50 foot
intervals. Below an altitude of 10, 000 feet, a diagonally
striped symbol covers the 10, 000 foot indicator. A knob
is provided at the bottom right corner of the altimeter for
setting readings in the pressure windows.
2-78. AIRSPEED INDICATORS.
2-81. VERTICAL VELOCITY INDICATORS.
Airspeed indicators are installed separately on the
pilot and copilot sides of the instrument panel (fig. 228). These indicators require no electrical power for
operation. The indicator dials are calibrated in knots
from 40 to 300. A striped pointer automatically displays
the maximum allowable airspeed (245 KIAS, 0.47 mach)
at the aircraft’s' present altitude.
Vertical velocity indicators are installed separately
on the pilot and copilot sides of the instrument panel
(fig. 2-28). They indicate the speed at which the aircraft
ascends or descends based on changes in atmospheric
pressure. The indicator is a direct reading pressure
instrument requiring no electrical power for operation.
2-79. PILOT'S ENCODING ALTIMETER.
2-82. ACCELEROMETER.
The altimeter is located on the upper left side of
the instrument panel (fig. 2-28). The altimeter is a selfcontained unit which consists of a precision pressure
altimeter combined with an altitude encoder. The
display indicates and the encoder transmits,
simultaneously, pressure altitude information to the
transponder. Altitude is displayed on the altimeter by a
10,000 foot counter, a 1000 foot counter, and a single
needle pointer which indicates hundreds of feet on a
circular scale in 50 foot increments. Below an altitude
of 10,000 feet, a diagonal warning symbol will appear on
the 10,000 foot counter. A barometric pressure setting
knob is provided to insert the desired altimeter setting in
inches Hg or millibars. A DC powered vibrator operates
inside the altimeter whenever aircraft power is on. If DC
power to the altitude encoder is lost, a warning flag
placarded CODE OFF will appear in the upper center
portion of the instrument face, indicating that the altitude
encoder is inoperative and that the system is not
reporting altitude to ground stations.
Operating
instructions
are
contained
in
Chapter
3.
The accelerometer, located on the instrument
panel registers and records positive and negative G
loads imposed on the aircraft. One hand moves in the
direction of the G load being applied while the other two,
one for positive G loads and one for negative G loads,
follow the indicating pointer to its maximum travel. The
recording pointers remain at the respective maximum
travel positions of the G's being applied, providing a
record of maximum G loads encountered. Depressing
the push-to-reset knob at the lower left corner of the
instrument allows the recording pointers to return to the
normal position.
2-83. FREE AIR TEMPERATURE (FAT) GAGE.
The free air temperature gage, mounted outboard
of the pilot's seat indicates the free air temperature in
degrees Celsius.
2-70
TM 55-1510-219-10
2-84. STANDBY MAGNETIC COMPASS.
(2)
The INST INDIRECT LIGHTS switch is
ON.
WARNING
(3) The MASTER PANEL LIGHTS switch is
OFF.
Inaccurate indications on the
standby magnetic compass will
occur while windshield heat
and/or air conditioning is being
used.
(4) The MASTER PANEL LIGHTS switch is
ON and the PILOT INST LIGHTS switch if OFF.
b. Master Warning Light (Red).
MASTER
WARNING light is located on each side of the
glareshield (fig. 2-28). Any time a warning light
illuminates, the MASTER WARNING light will
illuminate, and will remain illuminated until pressed. If a
new fault condition occurs, the MASTER WARNING
light will be reactivated, and the applicable annunciator
panel warning light will illuminate.
The standby magnetic compass is located below
the overhead fuel management panel to the right of the
windshield divider. It may be used in the event of failure
of the compass system, or for instrument cross check.
Readings should be taken only during level flight since
errors may be introduced by turning or acceleration. A
compass correction chart indicating deviation is located
on the magnetic compass.
c. Master Caution Light (Yellow). A MASTER
CAUTION light is located on each side of the glareshield
(fig. 2-28). Any time a caution light illuminates, the
MASTER CAUTION will illuminate, and will remain
illuminated until pressed. If a new fault condition
occurs, the MASTER CAUTION light will be reactivated,
and the applicable annunciator panel caution light will
illuminate. Annunciator Panels
2-85. MISCELLANEOUS INSTRUMENTS.
a. Annunciator Panels. Three annunciator panels
are installed. One is a warning panel with red fault
identification lights, and the others are caution/ advisory
panels with yellow and green identification lights. the
warning panel is mounted near the center of the
instrument panel below the glareshield (fig. 228) and
one caution/advisory panel is located on the center
subpanel. The mission annunciator panel is located on
the copilot's sidewall. Some normal flight operations
involve indications from the mission control panel,
described in Chapter 4. Illumination of a red warning
light signifies the existence of a hazardous condition
requiring immediate corrective action. A yellow caution
light signifies a condition other than hazardous requiring
pilot attention. A green advisory light indicates a
functional situation. Table 2-6 provides a list of causes
for illumination of the individual annunciator lights. In
frontal view both panels present rows of small, opaque
rectangular indicator lights. Word printing on each
indicator identifies the monitored function, situation, or
fault condition, but cannot be read until the light is
illuminated. The bulbs of all annunciator panel lights
are tested by activating the ANNUNCIATOR TEST
switch, located on the right side of the caution/ advisory
panel. The system is protected by two 5ampere circuit
breakers placarded ANN PWR and ANN IND on the
overhead circuit breaker panel (fig.
2-26).
The
annunciator system lights are dimmed when the
MASTER PANEL LIGHTS switch is actuated and the
pilot's flight instrument lights are on. The lights are
automatically reset to maximum brightness if:
(1) The main
generators) are OFF.
aircraft
power
(both
d. Pilot's Clock. One manual wind 8-day clock is
mounted in the center of the pilot's control wheel (fig. 216).
e. Copilot's Clock. A digital quartz chronometer is
mounted in the center of the copilot's control wheel (fig.
2-16). The quartz chronometer is a five function
clock/timer that is controlled by three pushbutton
switches located directly below the six digit liquid crystal
display.
(1) Mode selection. The MODE button is
pressed to select the desired mode of operation. The
mode annunciator is displayed above the mode
identifiers, and advances to indicate each of the
following modes:
• LC Local Time
• ZU Zulu or Greenwich Mean Time
• FT Trip or Flight Timer
• ET Elapsed Time
• DC Down counter with Alarm
(2) Local time mode (LC). Press the MODE
button to advance the annunciator to LC. To set the
hour, press the RST button once, then press and hold
the ADV button until the correct hour is displayed. To
set minutes, press the RST button again, then press and
hold the ADV button
DC
2-71
TM 55-1510-219-10
Figure 2-28. Instrument Panel (Sheet 1 of 2)
2-72
TM 55-1510-219-10
Figure 2-28. Instrument Panel (Sheet 2 of 2)
2-73
TM 55-1510-219-10
Table 2-6. Annunicator Panels
2-74
TM 55-1510-219-10
Table 2-6. Annunicator Panels (Continued)
2-75
TM 55-1510-219-10
Table 2-6. Annunicator Panels (Continued)
Press the RST button to set the time display to zero.
Press the ST-SP button one time. To stop the counting,
press the ST-SP button a second time. Ending time will
be displayed until the RST button is pressed to clear the
display. The clock may be used in other modes and the
elapsed time display will remain until cleared by
pressing the RST button. If the timer is counting when
the RST button is pressed, the display will reset to zero
and the count will begin again from zero.
until the correct minute is displayed. Press the SET
button once to display and hold selected time. Press the
ST-SP button to resume clock operation and/or
synchronize the display with a selected time standard.
(3) Zulu or Greenwich Mean Time mode
(ZU).
Press the MODE button to advance the
annunciator to ZU and set time as for local time shown
above. Minutes and seconds do not need to be reset if
local time is correctly set. Press the RST button to
display minutes/seconds, then press again to activate
the complete display.
(6) Downcounter mode (DC).
Press the
MODE button to advance the annunciator to DC. Press
the SET button twice to reset the hour display to zero.
Press and hold the ADV button until the desired hour is
displayed. Press the SET button again, to reset the
minute display to zero. Press and hold the ADV button
until the desired minute is displayed. Press the SET
button again, to reset the seconds display to zero.
Press and hold the ADV button until the desired second
is displayed. Press the SET button again, to arm the
counter. Press the STSP button to begin countdown.
When the countdown reaches zero, the display will flash
for approximately one minute and then reset. The
countdown may also be reset at any time by pressing
the ST-SP button.
When changing time zones, the hour may be
changed as above. It is not necessary to change the
minutes/seconds. Press the RST button twice to return
to the full time display.
(4) Trip/Flight timer mode (FT). Press the
MODE button to advance the annunciator to FT. Press
and hold the ST-SP button and verify that the display
shows zero. The timer will activate at takeoff and stop
at touchdown. To prevent an accidental reset of flight
time, the clock cannot manually reset during flight.
(5) Elapsed time mode (ET).
Press the
MODE button to advance the annunciator to ET.
Section XII. SERVICING, PARKING, AND MOORING.
2-86. GENERAL
The following paragraphs include the procedures
necessary to service the aircraft except lubrication. The
lubrication requirements of the aircraft are covered in
the aircraft maintenance manual. Tables 2-7, 2-8, 2-9
and 2-10 are used for identification of fuel, oil, etc. used
to service the aircraft.
The servicing instructions
provide procedures and precautions necessary to
service the aircraft.
WARNING
During warm weather open fuel
caps slowly to prevent being
sprayed with fuel.
WARNING
When aviation gasoline is used in a
turbine engine, extreme caution should
be used when around the combustion
chamber and exhaust area to avoid cuts
or abrasions.
The exhaust deposits
contain lead oxide which will cause lead
poisoning.
2-87. FUEL HANDLING PRECAUTIONS.
Table 2-2, Fuel Quantity Data, lists the quantity
and capacity of fuel tanks in the aircraft. Service the
fuel tanks after each flight to keep moisture out of the
tanks and to keep the bladder type cells from drying out.
Observe the following precautions:
2-76
TM 55-1510-219-10
of energized ground radar equipment installations.
CAUTION
Proper procedures for handling JP-4 and
JP-5 fuel cannot be over stressed. Clean,
fresh fuel shall be used and the entrance
of water into the fuel storage or aircraft
fuel system must be kept to a minimum.
CAUTION
k. Wear only nonsparking shoes near aircraft or
fueling equipment, as shoes with nailed soles or metal
heel plates can be a source of sparks.
2-88. FILLING FUEL TANKS.
WARNING
When conditions permit, the aircraft shall
be positioned so that the wind will carry
the fuel vapors away from all possible
sources of ignition. The fuel vehicle shall
be positioned to maintain a minimum
distance of 10 feet from any part of the
aircraft, while maintaining a minimum
distance of 20 feet between the fueling
vehicle and the fuel filler point.
Prior to removing the fuel tank filler cap,
the hose nozzle static ground wire shall
be attached to the grounding lugs that
are located adjacent to the filler opening.
Fill tanks as follows:
a. Attach bonding cables to aircraft.
a. Shut off unnecessary electrical equipment on the
aircraft, including radar and radar equipment. The master
switch may be left on, to monitor fuel quantity gages, but
shall not be moved during the fueling operation. Do not
allow operation of any electrical tools, such as drills or
buffers, in or near the aircraft during fueling.
b. Attach bonding cable from hose nozzle to
ground socket adjacent to fuel tank being filled.
CAUTION
Do not insert fuel nozzle completely into
fuel cell due to possible damage to
bottom of fuel cell.
b. Keep fuel servicing nozzles free of snow, water,
and mud at all times.
c. Carefully remove snow, frost, water, and ice from
the aircraft fuel filler cap area before removing the fuel
filler cap (fig. 2-29). Remove only one aircraft tank filler
cap at any one time, and replace each one immediately
after the servicing operation is completed.
d. Drain water from fuel tanks, filter cases, and
pumps prior to first flight of the day. Preheat, when
required, to insure free fuel drainage.
e. Avoid dragging the fueling hose where it can
damage the soft, flexible surface of the deicer boots.
c. Fill main tank before filling respective auxiliary
tanks unless less than a full fuel load is desired.
d. Secure applicable fuel tank filler cap.
sure latch tab on cap is pointed aft.
Make
e. Disconnect bonding cables from aircraft.
2-89. DRAINING MOISTURE FROM FUEL SYSTEM.
To remove moisture and sediment from the fuel
system, 12 fuel drains are installed (plus one for the
ferry system, when installed).
f: Observe NO SMOKING precautions.
g. Prior to transferring the fuel, insure that the hose
is grounded to the aircraft.
2-90. FUEL TYPES.
Approved fuel types are as follows:
h. Wash off spilled fuel immediately.
i. Handle the fuel hose and nozzle cautiously to avoid
damaging the wing skin.
j. Do not conduct fueling operations within 100 feet
of energized airborne radar equipment or within 300 feet
2-77
a. Army Standard Fuels. Army standard fuel is JP4.
b. Alternate Fuels. Army alternate fuels are JP-5
and JP-8.
TM 55-1510-219-10
Figure 2-29. Servicing Locations
2-78
TM 55-1510-219-10
Table 2-7. Fuel, Lubricants, Specifications and Capacities
Table 2-8. Approved Fuels
2-79
TM 55-1510-219-10
Table 2-8. Approved Fuels (Continued)
Table 2-9. Standard, Alternate and Emergency.
conforming to ASTM D-1655 specification may be used
when MIL-T-5624 fuels are not available. This usually
occurs during cross country flights where aircraft using
NATO F-44 (JP-5) are refueling with NATO F-40 (JP-4)
or Commercial ASTM Type B fuels. Whenever this
condition occurs, the engine operating characteristics
may change in that lower operating temperature, slower
acceleration, lower engine speed, easier starting, and
shorter range may be experienced. The reverse is true
when changing from F-40 (JP-4) fuel to F-44 (JP-5) or
Commercial ASTM Type A-1 fuels. Most commercial
turbine engines will operate satisfactorily on either
kerosene or JP-4 type fuel. The difference in specific
gravity may possibly require fuel control adjustments; if
so, the recommendations of the manufacturers of the
engine and airframe are to be followed.
c. Emergency Fuel. Avgas is emergency fuel and
subject to 150 hour time limit.
2-91. USE OF FUELS.
Fuel is used as follows:
a. Fuel Mixture. Standard and alternate fuels may
be mixed in any ratio. Emergency fuels may be mixed
in any ratio with standard and alternate fuels, however,
use of the lowest octane rating available is suggested.
Use of emergency fuel is subject to a 150 hour time
limit.
b. Use of Kerosene Fuels. The use of kerosene
fuels (JP-5 type) in turbine engines dictates the need for
observance of special precautions. Both ground starts
and air restarts at low temperature may be more difficult
due to low vapor pressure. Kerosene fuels having a
freezing point of minus 40 degrees C (minus 40 degrees
F) limit the maximum altitude of a mission to 28,000
feet under standard day conditions.
2-92. SERVICING OIL SYSTEM.
An integral oil tank occupies the cavity formed
between the accessory gearbox housing and the
compressor inlet case on the engine. The tank has a
calibrated oil dipstick and an oil drain plug. Avoid
spilling oil.
Any oil spilled must be removed
immediately. Use a cloth moistened in solvent to
remove oil. Overfilling may cause a discharge of oil
through the accessory gearbox breather until a
satisfactory level is reached. Service oil system as
follows:
c. Mixing of Fuels in Aircraft Tanks.
When
changing from one type of authorized fuel to another, for
example JP-4 to JP-5, it is not necessary to drain the
aircraft fuel system before adding the new fuel.
d. Fuel Specifications. Fuels having the same
NATO code number are interchangeable. Jet fuels
2-80
TM 55-1510-219-10
1. Open the access door on the upper
cowling to gain access to the oil
filler cap and dipstick.
1. Inflate nose wheel tires to a pressure
between 55 and.60 PSI.
2. Inflate main wheel tires to a pressure
between 73 and 77 PSI.
CAUTION
A cold oil check is unreliable. If possible,
check oil within 10 minutes after engine
shutdown.
If over 10 minutes have
elapsed, motor the engine (starter only) for
15-20 seconds, then recheck. If over 10
hours have elapsed, start the engine and
run for 2 minutes, then recheck. Add oil
as required. Do not overfill.
2-95. SERVICING THE ELECTRIC TOILET.
The toilet should be serviced during routine ground
maintenance of the aircraft following any usage. It is
more efficient and convenient to remove, clean and
recharge the toilet tank on a regular basis than to wait until
the tank is filled to capacity. Instructions for servicing are
provided on a decal applied to the front side of the
removable tank. Instructions are as follows:
2. Remove oil filler cap.
a. Tank Removal.
3. Insert a clean funnel, with a screen
incorporated, into the filler neck.
1. Open front access to the toilet, as
applicable, to remove the toilet tank.
4. Replenish with oil to within 1 quart
below MAX mark or the MAX COLD
on dipstick (cold engine). Fill to
MAX or MAX HOT (hot engine).
2. Depress the lock ring of the flush hose
quick disconnect coupling located on
the right side at the front of the tank
top.
5. Check oil filler cap for damaged
preformed
packing,
general
condition and locking.
3. Drain any residue of flush fluid in the
hose by partially disengaging the plug
from the quick disconnect and
manipulating the hose to assist
drainage.
CAUTION
4. Remove the flush hose from the quick
disconnect and place hose in the
retaining clip located on the underside
of the toilet mounting plate.
Insure that oil filler cap is correctly
installed and securely locked to prevent
loss of oil and possible engine failure.
6. If oil level is over 2 quarts low,
motor or run engine as required,
and service as necessary.
5. Install the cap attached to the quick
disconnect to seal the coupling.
7. Install and secure oil filler cap.
6. Close the knife valve at the bottom of
the toilet bowl by pushing the actuator
handle until the valve is fully closed.
8. Check for any oil leaks.
2-93. SERVICING HYDRAULIC
RESERVOIR.
BRAKE
7. Press the two fasteners on each side
of the knife valve actuator to unlock
the tank.
SYSTEM
8. Remove the tank by pulling the
recessed carrying handle on the tank
top.
b. Tank Cleaning.
1. Gain access to brake hydraulic
system reservoir.
2. Remove brake reservoir cap and fill
reservoir to washer on dipstick with
hydraulic fluid.
1. Dispose of tank contents by holding
the tank upside-down over a sewer or
toilet and pull the knife valve actuator
handle, opening the valve and
allowing the tank to drain.
3. Install brake reservoir cap.
2-94. INFLATING TIRES.
Inflate tires as follows:
2. Rinse the tank by filling one-half full
with water. Close the knife valve and
shake vigorously. Drain tank as in
previous step.
2-81
TM 55-1510-219-10
NOTE
Commercial
detergents
and
disinfectants can be included in
the rinse water if desired.
However, do not include these
materials in the tank precharge.
3. Rinse and drain the tank several
times to insure that the tank is
thoroughly clean.
4. Wipe the exterior surfaces of the
tank using a cloth moistened with
clear water and disinfectant.
2-96.
ANTI-ICING,
PROTECTION.
Charge the tank with a mixture of water and
chemical according to chemical manufacturer's
specification.
AND
DEFROSTING
The aircraft is protected in subfreezing weather by
spraying the surfaces (to be covered with protective covers)
with defrosting fluid. Spraying defrosting fluid on aircraft
surfaces before installing protective covers will permit
protective covers to be removed with a minimum of sticking.
To prevent freezing rain and snow from blowing under
protective covers and diluting the fluid, insure that protective
covers are fitted tightly. As a deicing measure, keep
exposed aircraft surface wet with fluid for protection against
frost.
NOTE
Do not apply anti-icing, deicing and
defrosting fluid to exposed aircraft surfaces
if snow is expected. Melting snow will
dilute the defrosting fluid and form a slush
mixture which will freeze in place and
become difficult to remove.
c. Tank Precharge.
Caution
During freezing temperatures,
toilet shall be serviced with
antifreeze solution to prevent
damage.
DEICING
Use undiluted anti-icing, deicing, and defrosting fluid
(MIL-A-8243) to treat aircraft surfaces for protection against
freezing rain and frost. Spray aircraft surface sufficiently to
wet area, but without excessive drainage. A fine spray is
recommended to prevent waste. Use diluted, hot fluid to
remove ice accumulations.
d. Tank Installation.
1. Remove frost or ice accumulations from
aircraft surfaces by spraying with diluted
anti-icing, deicing, and defrosting fluid
mixed in accordance with Table 2-10.
1. Install the tank by inserting the
slides located on each side of the
knife valve into the slide plate
assembly on the bottom of the
toilet and slide tank into place.
2. Spray diluted, hot fluid in a solid stream
(not over 15 gallons per minute).
Thoroughly saturate aircraft surface and
remove loose ice. Keep a sufficient
quantity of diluted, hot fluid on aircraft
surface coated with ice to prevent liquid
layer from freezing. Diluted, hot fluid
should be sprayed at a high pressure,
but not exceeding 300 PSI.
2. Press the two fasteners to the first
detent to secure the tank.
3. Remove the cap in the flush hose
quick disconnect and connect the
hose coupling to the quick
disconnect. Lock the disconnect
lock ring.
3. When facilities for heating are not
available and it is deemed necessary to
remove ice accumulations from aircraft
surfaces, undiluted defrosting fluid may
be used. Spray undiluted defrosting
fluid at 15 minute intervals to assure
complete coverage. Removal of ice
accumulations
using
undiluted
defrosting fluid in expensive and slow.
4. Pull the knife valve actuator to fully
open the valve.
5. Lift the toilet seat and shroud
assembly from the top of the toilet
and wipe with cloth moistened with
clear water and disinfectant. Wipe
the bowl and surrounding area.
4. If tires are frozen to ground, use
undiluted defrosting fluid to melt ice
around tire. Move aircraft as soon as
tires are free. Recommended Fluid
Dilution Chart
6. Check flushing operation of the
toilet and check for leaks.
7. Close access to the toilet.
2-82
TM 55-1510-219-10
Table 2-10. Recommended Fluid Dilution Chart
The oxygen system furnishes breathing oxygen to
the pilot, copilot and first aid position. Figure 2-19
shows the location of oxygen cylinder.
2-97. APPLICATION OF EXTERNAL POWER.
CAUTION
Before connecting the power
cables from the external power
source to the aircraft, insure that
the GPU is not touching the
aircraft at any point. Due to the
voltage drop in the cables, the
two ground systems will be of
different potentials. Should they
come in contact while the GPU is
operating, arcing could occur.
Turn off all external power while
connecting the power cable to, or
removing it from the external
power supply receptacle.
Be
certain that the polarity of the
external power source is the same
as that of the aircraft before it is
connected.
For GPU power
requirements,
consult
the
Maintenance Manual.
a. Oxygen System Safety Precautions.
WARNING
Keep fire and heat away from
oxygen equipment. Do not smoke
while working with or near oxygen
equipment, and take care not to
generate sparks with carelessly
handled tools when working on
the oxygen system.
(1) Keep oxygen regulators, cylinders, gages,
valves, fittings, masks, and all other components of the
oxygen system free of oil, grease, gasoline, and all
other readily combustible substances. The utmost care
shall be exercised in servicing, handling, and inspecting
the oxygen system.
(2) Do not allow foreign matter to enter oxygen
lines.
(3) Never allow electrical equipment to come in
contact with the oxygen cylinder.
(4) Never use oxygen from a cylinder without
first reducing its pressure through a regulator.
b. Replenishing Oxygen System.
1. Remove oxygen access door on outside of
aircraft (fig. 2-29).
2. Remove protective cap on oxygen system
filler valve.
3. Attach oxygen hose from oxygen servicing
unit to filler valve.
An external power source is often needed to supply
the electric current required to properly ground service
the aircraft electrical equipment and to facilitate starting
the aircraft's engines. An external DC power receptacle
is installed on the outboard side of the right engine
nacelle. An external AC power receptacle is installed on
the outboard side of the left engine nacelle.
2-98. SERVICING OXYGEN SYSTEM.
2-83
TM 55-1510-219-10
handled can be fatal. The applicable technical manuals
and pertinent directives should be studied for
familiarization with the aircraft, its components, and the
ground handling procedures applicable to it, before
attempting to accomplish ground handling.
b. Ground Handling Safety Practice.
Aircraft
equipped with turboprop engines require additional
maintenance safety practices. The following list of
safety practices should be observed at all times to
prevent possible injury to personnel and/or damaged or
destroyed aircraft:
(1) Keep intake air ducts free of loose articles
such as rags, tools, etc.
(2) Stay clear of exhaust outlet areas.
(3) During ground runup, make sure the brakes
are firmly set.
(4) Keep area fore and aft of propellers clear of
maintenance equipment.
(5)
Do not operate engines with control
surfaces in the locked position.
(6) Do not attempt towing or taxiing of the
aircraft with control surfaces in the locked position.
(7) When high winds are present, do not unlock
the control surfaces until prepared to operate them.
(8)
Do not operate engines while towing
equipment is attached to the aircraft, or while the aircraft
is tied down.
(9) Check the nose wheel position. Unless it is
in the centered position, avoid operating the engines at
high power settings.
(10) Hold control surfaces in the neutral position
when the engines are being operated at high power
settings.
(11) When moving the aircraft, do not push on
propeller deicing boots.
Damage to the heating
elements may result.
c. Moving Aircraft on Ground. Aircraft on the
ground shall be moved in accordance with the following:
(1) Taxiing. Taxiing shall be in accordance with
Chapter 8.
WARNING
If the oxygen system pressure is
below 200 PSI, do not attempt to
service system. Make an entry on
DA Form 2408-13.
4. Insure that supply cylinder shutoff valves
on the aircraft are open.
5. Fill system slowly to prevent overheating by
adjusting recharging rate with pressure
regulating valve on oxygen servicing unit.
6. Close pressure regulating valve on oxygen
servicing unit when pressure gage on
oxygen system indicates the pressure
obtained using the Oxygen System
Servicing Pressure Chart (fig. 2-30).
NOTE
Fill oxygen system to 1800 PSI at
70°F. For every degree above 70°,
increase pressure 3.5 PSI to a
maximum of 2000 PSI; and for
every degree F below 70°,
decrease pressure 3.5 PSI.
7.
Disconnect oxygen hose from oxygen
servicing unit and filler valve.
8. Install protective cap on oxygen filler valve.
9. Install oxygen access door.
2-99. GROUND HANDLING.
Ground handling covers all the essential information
concerning movement and handling (fig. 2-31) of the
aircraft while on the ground. The following paragraphs
give, in detail, the instructions and precautions
necessary to accomplish ground handling functions.
a. General Ground Handling Procedure. Accidents
resulting in injury to personnel and damage to
equipment can be avoided or minimized by close
observance of existing safety standard and recognized
ground handling procedures.
Carelessness or
insufficient knowledge of the aircraft or equipment being
2-84
TM 55-1510-219-10
Figure 2-30. Oxygen System Servicing Pressure
2-85
TM 55-1510-219-10
Figure 2-31. Parking, Covers, Ground Handling, & Towing Equipment
2-86
TM 55-1510-219-10
aircraft are covered by the various phases of the ground
handling procedures included in this section of general
ground handling instructions. (Refer to TM 55-1500204-25/1.)
CAUTION
When the aircraft is being towed,
a qualified person must be in the
pilot's seat to maintain control by
use of the brakes. When towing,
do not exceed nose gear turn
limits. Avoid short radius turns,
and always keep the inside or
pivot wheel turning during the
operation. Do not tow aircraft
with rudder locks installed, as
severe damage to the nose
steering linkage can result. When
moving the aircraft backwards, do
not apply the brakes abruptly.
Tow the aircraft slowly, avoiding
sudden stops, especially over
snowy, icy rough, soggy, or
muddy terrain. In arctic climates,
the aircraft must be towed by the
main gears, as an immense
breakaway load, resulting from
ice, frozen tires, and stiffened
grease in the wheel bearings may
damage the nose gear.
2-100. INSTALLATION OF PROTECTIVE COVERS.
The crew will insure that the aircraft protective
covers are installed.
2-101. MOORING.
The aircraft is moored to insure its immovability,
protection, and security under various weather
conditions. The following paragraphs give, in detail, the
instructions for proper mooring of the aircraft.
a. Mooring Provisions. Mooring points (fig. 2-32)
are provided beneath the wings and tail. Additional
mooring cables may be attached to each landing gear.
General mooring equipment and procedures necessary
to moor the aircraft, in addition to the following, are
given in TM 55-1500-204-25/1.
(1) Use mooring cables of 1/4 inch diameter
aircraft cable and clamp (clip-wire rope), chain or rope
3/8 inch diameter or larger. Length of the cable or rope
will be dependent upon existing circumstances. Allow
sufficient slack in ropes, chains, or cable to compensate
for tightening action due to moisture absorption of rope
or thermal contraction of cable or chain. Do not use slip
knots. Use bowline knots to secure aircraft to mooring
stakes.
(2) Chock the wheels.
b. Mooring Procedures for High Winds. Structural
damage can occur from high velocity winds; therefore, if
at all possible, the aircraft should be moved to a safe
weather area when winds above 75 knots are expected.
If aircraft must be secured use the following steps:
1. After aircraft is properly located, place nose
wheel in centered position. Head aircraft
into the wind, or as nearly so as is possible
within limits determined by locations of fixed
mooring rings.
When necessary, a 45
degree variation of direction is considered to
be satisfactory. Locate each aircraft at
slightly more than wing span distance from
all other aircraft. Position nose mooring
point approximately 3 to 5 feet downwind
from ground mooring anchors.
2. Deflate nose wheel shock strut to within 3/4
inch of its fully deflated position.
3. Fill all fuel tanks to capacity, if time permits.
4. Place wheel chocks fore and aft of main gear
wheels and nose wheel. Tie each pair of
chocks together with rope
CAUTION
Do not tow or taxi aircraft with
deflated shock struts.
(2) Towing. Towing lugs are provided on the
upper torque knee fitting of the nose strut. When it is
necessary to tow the aircraft with a vehicle, use the
vehicle tow bar.
In the event towing lines are
necessary, use towing lugs on the main landing gear.
Use towing lines long enough to clear nose and/or tail by
at least 15 feet. This length is required to prevent the
aircraft from overrunning the towing vehicle or fouling
the nose gear.
d.
Ground Handling Under Extreme Weather
Conditions. Extreme weather conditions necessitate
particular care in ground handling of the aircraft. In hot,
dry, sandy, desert conditions, special attention must be
devoted to finding a firmly packed parking and towing
area. If such areas are not available, steel mats or an
equivalent solid base must be provided for these
purposes. In wet, swampy areas, care must be taken to
avoid bogging down the aircraft. Under cold, icy, arctic
conditions, additional mooring is required, and added
precautions must be taken to avoid skidding during
towing operations.
The particular problems to be
encountered under adverse weather conditions and the
special methods designed to avoid damage to the
2-87
TM 55-1510-219-10
Figure 2-32. Mooring the Aircraft
2-88
TM 55-1510-219-10
5.
6.
7.
8.
9.
10.
11.
of damage from flying objects. Service
nose shock strut and reconnect battery.
or join together with wooden cleats nailed to
chocks on either side of wheels. Tie ice grip
chocks together with rope. Use sandbags in
lieu of chocks when aircraft is moored on
steel mats. Set parking brake as applicable.
Accomplish aircraft tiedown by utilizing
mooring points shown in figure 2-32. Make
tiedown with 1/4 inch aircraft cable, using
two wire rope clips or bolts, and a chair
tested for a 3000 pound pull.
Attach
tiedowns so as to remove all slack. (Use a
3/4-inch or larger manila rope if cable or
chain tiedown is not available.) If rope is
used for tiedown, use anti-slip knots, such
as bowline knot, rather than slip knots. In
the event tiedown rings are not available on
hard surfaced areas, move aircraft to an
area where portable tiedowns can be used.
Locate anchor rods at point shown in figure
2-31. When anchor kits are not available,
use metal stakes or deadman type anchors,
providing they can successfully sustain a
minimum pull of 3000 pounds.
In event nose position tiedown is considered
to be of doubtful security due to existing soil
condition, drive additional anchor rods at
nose tiedown position. Place padded work
stand or other suitable support under the aft
fuselage tiedown position and secure.
Place control surfaces in neutral position.
Place wing flaps in up position.
The requirements for dust excluders,
protective covers, and taping of openings
will be left to the discretion of the
responsible maintenance officer or the pilot
of the transient aircraft (fig. 2-31).
Secure propellers to prevent windmilling.
Disconnect battery.
During typhoon or hurricane wind conditions,
mooring security can be further increased
by placing sandbags along the wings to
break up the aerodynamic flow of air over
the wing, thereby reducing the lift being
applied against the mooring by the wind.
The storm appears to pass two times, each
time with a different wind direction. This will
necessitate turning the aircraft after the first
passing.
2-102. PARKING.
Parking is defined as the normal condition under
which the aircraft will be secured while on the ground.
This condition may vary from the temporary expedient
of setting the parking brake and chocking the wheels to
the more elaborate mooring procedures described under
Mooring. The proper steps for securing the aircraft must
be based on the time the aircraft will be left unattended,
the aircraft weights, the expected wind direction and
velocity, and the anticipated availability of ground and
air crews for mooring and/or evacuation.
When
practical head the aircraft into the wind, especially if
strong winds are forecast or if it will be necessary to
leave the aircraft overnight. Set the parking brake and
chock the wheels securely. Following engine shutdown,
position and engage the control locks.
NOTE
Cowlings and loose equipment
will be suitably secured at all
times when left in an unattended
condition.
a.
The parking brake system for the aircraft
incorporates two lever-type valves, one for each wheel
brake. Both valves are closed simultaneously by pulling
out the parking brake handle. Operate the parking
brake as follows:
1. Depress both brakes.
2. Pull parking brake handle out. This will
cause the parking brake valves to lock the
hydraulic fluid under pressure in the parking
brake system, thereby retaining braking
action.
3. Release brake pedals.
CAUTION
Do not set parking brakes when
the brakes are hot during freezing
ambient temperatures.
Allow
brakes to cool before setting
parking brakes.
4. To release the parking brakes push in on the
parking brake handle.
12. After high winds, inspect aircraft for visible
signs of structural damage and for evidence
2-89
TM 55-1510-219-10
2. Install elevator and aileron lockpin vertically
through pilot's control column to lock control
wheel.
3. Install rudder lock pin through flapper door
forward to pilot's seat, making sure rudder is
in neutral position.
4. Reverse steps 1 through 3 above to remove
control lock. Store control lock.
b. The control lock (fig. 2-17) holds the engine and
propeller control levers in a secure position, and the
elevator, rudder, and aileron in neutral position. Install
the control locks as follows:
1. With engine and propeller control levers in
secure position, slide lock onto control
pedestal to prevent operation of levers.
2-90
TM 55-1510-219-10
CHAPTER 3
AVIONICS
Section I. GENERAL
control panel. Individual system circuit breakers and the
associated avionics busses are shown in fig. 2-22.
With the switch in the ON position, the avionics power
relay is de-energized and power is applied through both
the AVIONICS MASTER POWER #1 and #2 circuit
breakers to the individual avionics circuit breakers on
the overhead circuit breaker panel (fig. 2-26). In the off
(aft) position, the relay is energized and power is
removed from avionics equipment. When external
power is applied to the aircraft, the avionics power relay
is normally energized, removing power from the
avionics equipment. To apply external power to the
avionics equipment, move the AVIONICS MASTER
POWER switch to the EXT PWR position. This will deenergize the avionics power relay and allow power to be
applied to avionics equipment.
b. AC Power. AC power for the avionics equipment
is provided by two inverters. The inverters supply 115volt and 26-volt single-phase AC power when operated
by the INVERTER #1 or #2 switches (fig. 2-18). Either
inverter is capable of powering all avionics equipment
requiring AC power. AC power from the inverters is
routed through fuses in the nose avionics compartment.
3-1. INTRODUCTION.
Except for mission avionics, this chapter covers all
avionics equipment installed in the RC-12D aircraft. It
provides a brief description of equipment covered, the
technical characteristics and locations.
It covers
systems and controls and provides the proper
techniques and procedures to be employed when
operating the equipment. For more detailed operational
information consult the vendor manuals that accompany
the aircraft loose tools.
3-2. AVIONICS EQUIPMENT CONFIGURATION.
The aircraft avionics covered is comprised of three
groups of electronic equipment. The Communication
group consists of the Interphone, UHF command, VHF
command and HF command systems. The Navigation
group provides the pilot and copilot with the
instrumentation required to establish and maintain an
accurate flight course and position, and to make an
approach
on
instruments
under
Instrument
Meteorological Conditions (IMC). The Navigation group
includes equipment for determining altitude, attitude,
position, destination, range and bearing, heading
reference, groundspeed, and drift angle.
The
Transponder and Radar group includes an identification,
position and emergency tracking system, and a radar
system to locate potentially dangerous weather areas
and a radar system to differentiate between friendly and
unfriendly search radar.
3-4. MICROPHONES, SWITCHES AND JACKS.
Boom, and oxygen mask microphones can be utilized in
the aircraft.
a. Microphone Switches. The pilot and copilot are
provided with individual MIC control switches, placarded
MIC/INTPH/XMIT, attached to respective control
wheels. A foot-actuated mic switch is also positioned on
the floorboards forward of each pilot's seat.
b. Controls and Functions. Microphone switches
and jack functions are as follows:
(1) Control wheel switches (fig. 2-16).
NOTE
All avionics equipment require a
3-minute warmup period.
The
weather radar has an automatic
time delay of 3 to 4 minutes.
CONTROL
MIC/INTPH
XMIT switch
MIC
3-3. POWER SOURCE.
a. DC Power. DC power for the avionics equipment
is provided by four sources: the aircraft battery, left and
right generators, and external power. Power is routed
through a 50-ampere circuit breaker to the avionics
power relay which is controlled by the AVIONICS
MASTER POWER SWITCH (fig. 2-18) on the overhead
3-1
FUNCTION
Keys selected facility.
(Not depressed) Microphone is
disconnected.
TM 55-1510-219-10
Released
Disconnects selected mic from
audio system.
(3) Subpanel jack selector switches.
CONTROL
FUNCTION
MIC HEADSET/
Selects microphone to connect
OXYGEN
to audio system switch
MASK
MIC HEADSET
Connects headset microphone to
audio system.
OXYGEN
Connects microphone of oxygen
MASK
mask to audio system.
INTPH
(depressed to first detent) Keys
interphone facility, disregards
position transmitter selector
switch.
XMIT
(depressed full down) Keys facility selected by transmitter select
switch.
(2) Floorboard switches.
CONTROL
FUNCTION
MIC foot switch
Controls connection of selected
mic to audio system.
Held depressed
Connects selected mic to audio
system.
Section II. COMMUNICATIONS
1
ON connects user's headset to
audio from VHF-AM transceiver
No. 1.
2
ON connects user's headset to
audio from VHF/AM/FM transceiver.
3
ON connects user's headset to
audio from command UHF
transceiver in use.
4
ON connects user's headset to
audio from HF or VOW transceiver in use.
5
ON connects user's headset to
audio from backup VOW transceiver.
NAV radios
Combination volume control
audio monitor
and "ON-OFF" switches for
controls
NAV receivers.
NAV-A
ON connects user's headset to
audio from VOR-1, VOR-2 or
set in use.
NAV-B
ON connects user's headset to
audio from TACAN or ADF set
in use.
(2) Microphone switches.
CONTROL
FUNCTION
Mic impedance
Two-position thumb-actuated
select switch (5
switch. Enables selection of inOhm/150 Ohm)
terface circuit with best impedence match to microphone used.
3-5. DESCRIPTION.
The communications equipment group is comprised
of an interphone system connected to individual Audio
Control Panels for the pilot and copilot which interface
with VHF, UHF and HF communication units.
3-6. AUDIO CONTROL PANELS (FIG. 3-1).
a. Description. Separate but identical audio control
panels serve the pilot and copilot. The controls and
switches of each panel provide the user with a means to
select desired reception and transmission sources, and
also a means to control the volume of audio signals sent
and received for interphone, communication and
navigation systems. The user selects between the UHF,
VHF and HF transceivers. Audio control panels are
protected by respective 2-ampere AUDIO PILOT and
AUDIO COPILOT circuit breakers positioned on the
overhead circuit breaker panel (fig. 2-26).
b. Controls and Indicators.
(1) Audio controls - Comm and Nav Radios (fig.
3-1).
CONTROL/
FUNCTION
INDICATOR
Master VOL
Controls sidetone volume to
control
headset.
Also serves as final
volume adjustment for received
audio from any source before
admission to headset.
Comm radios
Each is combination rotary conaudio controls
trol and "ON-OFF" push-pull
switch, permitting both receiver
selection and volume adjustment.
3-2
TM 55-1510-219-10
and that normally used avionic
circuit
breakers
remain
depressed. The circuit breakers
of routinely used avionic systems
are normally left depressed.
Microphone
select switches
HOT MIC
Controls activation of microphones.
Admits speech to interphone
system without need to key selected microphone.
NORM
Blocks speech from interphone
system unless selected microphone is keyed.
ICS OFF
Deactivates microphones
(3) Transmitter select switch (fig. 3-1).
CONTROL
FUNCTION
Transmit
Connects microphone and headinterphone
set to selected radio transmitter
selector switch
or interphone line routing received audio to headset. Bypasses control of respective receiver
audio switch.
PVT
Position not used.
ICS
Activates pilot-to-copilot intercom.
1
Permits audio reception from
VHF-AM No. 1 transceiver.
Routes key and mic signals to
VHF-AM No. 1 transceiver.
2
Permits audio reception from
VHF/AM/FM transceiver.
Routes key and mic signals to
VHF/AM/FM transceiver.
3
Permits audio reception from
UHF transceiver. Routes key
and mic signals to UHF transceiver.
4
Permits audio reception from
HF or VOW transceiver. Routes
key or mic signals to transceiver.
5
Permits audio reception from
backup VOW. Routes key and
mic signals to transceiver.
(4) CIPHER light - not used.
c. Normal Operation.
(1) Turn-on. Both Audio Control Panels are
activated when electrical power is applied to aircraft.
(2) Receive.
1. Receiver audio switches (Audio panel) As required.
2. Master VOL control (Audio panel) - As
required, in combination with volume
control of system being utilized.
NOTE
Audio select switches and volume
controls are routinely left in
positions of normal use.
3.
Move each receiver audio switch ON
then OFF, separately, to verify audio
presence in headphones for each
system. Adjust volume.
4. Mic select switch - As desired.
(3) Transmit.
1. Transmitter-interphone selector (audio
panel) - Set for transceiver desired.
2. MIC HEADSET/OXYGEN MASK switch
(instrument panel) -As desired.
3. MIC/INTPH/XMIT switch (control wheel)
- XMIT.
4.
MIC
switch (headset/oxygen
mask/floorboards) -Depress to transmit.
(4) Intercommunication.
1. Transmitter-interphone selector (audio
panel) -ICS.
2. MIC HEADSET/OXYGEN MASK switch
(instrument panel) -As desired.
3. Mic select switch (audio panel - As
desired (NORM/HOT MIC).
a. If HOT MIC selected - Talk when
ready.
b. If NORM selected - Depress MIC
switch and transmit.
c.
Master VOL control (selected
transceiver) -Set for comfort.
NOTE
It is presumed the AVIONICS
MASTER POWER switch is on,
3-3
TM 55-1510-219-10
Figure 3-1. Audio Control Panel (C-499) (Typical pilot, copilot)
control
con radio signals received.
(2) Automatic Direction Finding (ADF).
FILTER V/OFF
Selects filter to block voice
switch
transmissions from ADF ground
station.
FILTER R/OFF
Selects filter to block range
switch
transmissions from ADF ground
station.
d. Emergency Operation. Not applicable.
e. Shutdown.
1.
AVIONICS MASTER POWER switch
(overhead panel) -OFF.
2. Leave avionic controls and circuit
breakers positioned for normal operation.
3. Aircraft DC power - OFF.
3-7. RADIO CONTROL PANEL (FIG. 3-2).
a. Description. The radio control panel, located on
the pedestal extension allows the pilot or copilot to
selectively monitor audio signals from the High
Frequency (HF), Marker Beacon (MKR BCN), or ADF
systems. It also has controls for the selection of ADF
voice or range filters. Controls and functions are as
follows:
b. Controls and Indicators.
CONTROL
FUNCTION
HF VOL control
Adjusts volume of highfrequency radio signals received.
MKR BCN HI/
LO switch
MKR BCN VOL
3-8. UHF COMMAND SET (AN/ARC-164).
a. Description. The UHF command set is a line-ofsight radio transceiver which provides transmission and
reception of amplitude modulated (AM) signals in the
ultra high frequency range of 225.000 to 399.975 MHz
for a distance range of approximately 50 miles.
Channel selection is spaced at 0.025 MHz. A separate
receiver is incorporated to provide monitoring capability
for the UHF guard frequency (243.0 MHz). UHF audio
output is applied to the audio panel where it is routed to
the headsets.
Selects sensitivity of marker beacon receiver.
Adjusts volume of marker bea-
3-4
TM 55-1510-219-10
Figure 3-2. Radio Control Panel
(hundreds)
Manual
frequency
selector (tens)
Manual
frequency
selector (units)
Manual
frequency
selector (tenths)
Manual
frequency
selector
(hundredths and
thousandths)
Preset channel
selector
Mode selector
NOTE
The PRESET channel selector and
manual frequency selectors are
inoperative when the mode
selector is set to GUARD position.
The receiver-transmitter will be
set to the emergency frequency
only.
The transmit and receive sections of the UHF unit
operate independently but share the same power supply
and frequency control circuits.
Complete provisions only are installed for voice
security device KY-28 or KY-58 to locate on the
pedestal extension near the radio set.
The UHF
command set is protected by the 7 1/2 ampere UHF
circuit breaker on the overhead circuit breaker panel
(fig. 2-26). Figure 3-3 illustrates the UHF command
set. The associated blade type antenna is shown in
figure 2-1.
b. Controls and Indicators.
(1) UHF control panel fig. 3-3.
CONTROL
FUNCTION
Manual
Selects hundreds digit of frefrequency
quency (either 2 or 3) in MHz.
selector
MANUAL
PRESET
GUARD
3-5
Selects tens digit of frequency (O
through 9) in MHz.
Selects units digit of frequency
(O through 9) in MHz.
Selects tenths digit of frequency
(O through 9) in MHz.
Selects hundredths and thousandths digits of frequency (00,
25, 50, or 75) in MHz.
Selects one of 20 preset channel
frequencies.
Selects operating mode and
method of frequency selection.
Enables the manual selection of
any one of 7,000 frequencies,
Enables selection of any one of
20 preset channels. preset channel selector switch.
Selection automatically tunes the
main receiver and transmitter to
TM 55-1510-219-10
Figure 3-3. UHF Control panel (AN/ARC- 164)
It is presumed the AVIONICS
MASTER POWER switch is on,
and that normally used avionic
circuit
breakers
remain
depressed.
the guard frequency and the
guard receiver is disabled.
SQUELCH
Turns main receiver squelch on
switch
or off.
VOL control
Adjusts volume.
TONE
When pressed, transmits a 1,020
pushbutton
Hz tone on the selected frequency.
Function selector Selects operating function.
OFF
Turns set off.
MAIN
Selects normal transmission with
reception on main receiver.
BOTH
Selects normal transmissions
with reception on both the main
receiver and the guard frequency
receiver.
ADF
Activates ADF or homing system (if installed) and main receiver.
c. Normal Operation.
(1) Turn on.
1. Insure aircraft power is on.
2.
AVIONICS MASTER POWER switch
(overhead panel) - ON.
3. Function switch (UHF panel) MAIN or
BOTH position, as required.
NOTE
If function selector is at MAIN
setting, only the normal UHF
communications will be received.
If selector is at BOTH position,
emergency communications on
the guard channel and normal
UHF communications will both be
received.
(2) Receive.
1. UHF audio monitor switch (#3, audio
panel) - ON or transmit/interphone
selector (audio panel) - 3 position.
NOTE
3-6
TM 55-1510-219-10
NOTE
Disregard operating procedures
involving the voice security
(CIPHONY) control-indicator, if
unit is not installed.
2. VOL control (UHF panel) - Mid position.
(3) To use preset frequency (UHF panel).
1. Mode selector - PRESET position.
2. PRESET channel selector - Rotate to
channel desired.
(4) To use non-preset frequency (UHF panel).
1. Mode selector - MANUAL position.
2. Manual frequency selectors (5) - Rotate
each knob to set desired frequency
digits.
(1) Turn-on.
1. POWER ON switch (Voice Security
panel, fig 3-6) -ON.
2. Function switch (UHF panel) - BOTH.
(2) Receive (UHF panel).
1. Mode selector - As required.
2. Transmitter-interphone selector (audio
panel, fig. 3-1) - #3 position. or Audio
monitor control #3 - ON.
3. Set required frequency using preset
channel control or manual frequency
selector.
4. Adjust volume.
NOTE
The PRESET channel selector and
manual frequency selectors are
inoperative when the mode
selector is set to GUARD position.
(5) Volume - Adjust.
NOTE
To adjust volume when radio is
not
being
received,
turn
SQUELCH switch OFF, adjust
volume for comfortable noise
level, then turn squelch disable
switch ON.
NOTE
To adjust volume when audio is
not being received, turn squelch
switch OFF, adjust volume for
comfortable noise level, then turn
squelch switch ON.
(6) Squelch - As desired.
(7) Transmit.
1. Transmitter-interphone selector (audio
panel) - 3 position.
2. UHF panel - Set required frequency
using either preset channel control or
manual frequency select controls.
3. MIC HEADSET/OXYGEN MASK switch
(instrument panel) -As desired.
4. MIC switch - Depress to transmit.
5. Squelch - As required.
(3) Transmit (PLAIN).
1.
Transmit-interphone selector (audio
panel) - No. 3 position.
2. PLAIN/CIPHER switch (Voice Security
panel) - PLAIN.
3. Microphone switch - Press.
(4) Transmit (CIPHER).
1.
Transmit-interphone selector (audio
panel) - No. 3 position.
2. PLAIN/CIPHER switch (Voice security
panel) -CIPHER. (CIPHER indicator will
be on as long as switch is in CIPHER
position.)
3. RE-X/REG switch (Voice security panel)
- REG.
(8) Shutdown.
1. Function selector (UHF panel) - OFF.
d. UHF Command and Voice Security Operation
(KY-28).
3-7
TM 55-1510-219-10
3-10. VHF-AM COMMUNICATIONS (VHF-20B).
a. Description. VHF-AM communications provide
transmission and reception of amplitude modulated
signals in the very high frequency range of 116.000 to
151.975 MHz for a range of approximately 50 miles,
varying with altitude. A dual head control panel (COMM
1, fig. 3-4) is mounted on the pedestal (fig. 2-10)
accessible to both the pilot and copilot. The panel
provides two sets of control indicators, frequency
indicators, frequency select knobs, a single volume
control, and a single selector switch for quick frequency
changing. Transmission audio is routed by pilot and
copilot #1 transmitter select switches. Received audio is
routed by pilot and copilot # receiver audio switches to
the respective headsets. The VHF radio is protected by
one 10-ampere VHF circuit breaker on the overhead
control circuit breaker panel (fig. 2-26). The associated
antenna is shown in figure 2-1.
b. Controls and Indicators.
(1) VHF-AM control panel (fig. 3-4).
CONTROL
FUNCTION
Left control
Indicates set operating frequency
Frequency
(control TRANS switch left posiindicator
tion.)
Frequency
Selects desired set operating freselectors
quency (control TRANS switch
left position.)
VOL-OFF
Adjusts volume of received aucontrol
dio, turns set ON or OFF.
CONTROL
Illuminates, if control TRANS
indicator
switch in left position.
TRANS switch
Selects right or left control head
to control operating frequency of
set.
Frequency
Indicates set operating frequency
indicator
(control TRANS switch right position.)
Frequency
Selects desired set operating freselectors
quency (control TRANS switch
right position.)
CONTROL
Illuminates, if control TRANS
indicator
switch at right position.
COMM TEST
Overrides automatic squelch cirswitch
cuit.
c. VHF-AM Set - Normal Operation.
(1) Turn-on procedure. VOL control - Turn
clockwise.
4. Microphone switch - Press momentarily
(interrupted tone from voice security
unit should no longer be heard).
NOTE
No traffic will be passed if the
interrupted tone is still heard after
pressing
and
releasing
the
microphone switch.
5. Microphone switch - Press (do not talk).
Wait until beep is heard then speak into
microphone.
(5) Shutdown.
1. Function switch (UHF panel)-OFF.
2. POWER ON switch (Voice security
panel) - OFF.
e. UHF Command Set - Emergency Operation.
NOTE
Transmission
on
emergency
frequencies (guard channels) is
restricted to emergencies only.
An emergency frequency of 121.
500 MHz is also available on the
VHF command radio set.
1.
Transmit/interphone selector (audio
panel) - No. 3 position.
2. Mode selector switch (UHF panel) GUARD.
3. Microphone switch - Press.
3-9. BACKUP VOW (AN/ARC-164).
A radio set identical in type and performance to the
UHF command set (fig. 3-3) is installed on the pedestal
extension (fig. 2-10) to serve as backup VOW (Voice
Order Wire). This set provides the pilot and copilot with
secure 2-way voice communications.
Complete
provisions only are provided for a KY-58 voice security
device. The backup VOW set is protected by a 7 1/2
ampere BU VOW circuit breaker on the overhead circuit
breaker panel (fig. 2-26). The backup VOW and
Transponder share an antenna which is mounted on the
aircraft belly (fig. 2-1).
3-8
TM 55-1510-219-10
Figure 3-4. VHF-AM Control Panel (VHF-20B)
1. Transmitter-interphone selector (audio panel, fig. 3(2) Receiver operating procedure.
1) - #1 position.
1. Transmitter-interphone selector (audio
2. Frequency selector (VHF panel) - 121.
panel, fig. 3-1) - #1 position. or Audio
500 MHz (emergency frequency).
monitor control #1 - ON.
3. Microphone switch - Press.
2.
Frequency selector - Set desired
frequency.
3-11. VHF AM/FM COMMAND SET (AN/ARC-186).
3. VOL control - As required.
(3) Transmitter operating procedure.
a. Description. The VHF AM/FM Command Set
1. Transmitter-interphone selector (audio
provides for normal and secure 2-way voice
panel, fig. 3-1) - #1 position.
communication: AM in the very high frequency range of
2. Microphone switch - Press.
116.000 to 151.975 MHz and FM in the 30. 000 to
(4)
Shutdown procedure.
VOL control 87.975 MHz band. Twenty channels may be preset.
Audio signals are applied through pilot and copilot
Counter-clockwise (OFF).
transmitter select switches No. 2 position and through
d. VHF-AM Set - Emergency Operation.
the pilot and copilot No. 2 receiver audio switches to
their respective headsets. Complete provisions only are
NOTE
installed for voice security device KY-28 or KY-58.
Transmissions
on
frequency
Circuits are protected by a 10-ampere VHF AM/FM
(121.500 MHz) are restricted to
circuit breaker on the overhead circuit breaker panel
emergencies only.
Emergency
(fig. 2-26). Figure 3-5 illustrates the VHF AM/FM
frequency 243.000 MHz (guard
control panel. The associated antenna is shown in
channel) is also available on the
figure 2-1.
UHF command radio.
3-9
TM 55-1510-219-10
Figure 3-5. VHF AM/FM Control Panel (AN/ARC- 186)
0.025 MHz
indicator
Indicates manually selected receiver-transmitter frequency in
0.025 MHz increments.
0.025 MHz
Selects receiver-transmitter freselector
quency in 0.025 MHz increments. Clockwise rotation increases frequency.
CHAN indicator Selects preset channel from 1 to
20. Clockwise rotation increases
number selected.
LOCKOUT FM/
Screwdriver adjustable threeAM switch
position switch. Warning tone
announces lockout.
Center
Selects AM or FM band.
AM
Disables AM band.
FM
Disables FM band.
FM SQUELCH
Screwdriver adjustable potenticontrol
ometer. Squelch fully overdriven
at full counter-clockwise position. Clockwise rotation increases input signal required to
open squelch.
WB/NB MEM
Three-position switch.
LOAD switch
b. Controls and Indicators.
(1) VHF AM/FM control panel (fig. 3-5).
CONTROL
FUNCTION
10 MHz selector
Selects receiver-transmitter frequency in increments of 10 MHz
from 30 to 150 MHz. Clockwise
rotation increases frequency.
10 MHz
Indicates manually selected reindicator
ceiver-transmitter frequency in
10 MHz increments from 30 to
150 MHz.
1.0 MHz selector Selects receiver-transmitter frequency in 1.0 MHz increments.
Clockwise rotation increases frequency.
1.0 MHz
Indicates manually selected reindicator
ceiver-transmitter frequency in
1.0 MHz increments.
0.1 MHz
Indicates manually selected reindicator
ceiver-transmitter frequency in
0.1 MHz. increments.
0.1 MHz selector Selects receiver-transmitter frequency in 0.1 Mhz increments.
Clockwise rotation increases frequency.
3-10
TM 55-1510-219-10
(2) Receive.
1. Frequency control emergency select
switch (fig. 3-5) - MAN or PRE, as
desired.
2. Transmitter-interphone selector (audio
panel, fig. 3-1) - #2 position. or Audio
monitor control #2 - ON.
3.
Manual frequency/preset channel
selectors - Set desired frequency.
4. VOL control - As required.
(3) Transmit.
1. Transmit/interphone select switch (audio
panel) -Position 2.
2. Microphone switch - Press.
(4) Shutdown.
1. Mode select switch (fig. 3-5) - OFF.
d. VHF AM/FM - Emergency Operation.
(1) Emergency AM Mode.
1. Transmit/interphone select switch (audio
panel) -Position 2.
2. Mode select switch - TR.
3.
PRE-MAN-FM/AM
EMER switch EMER AM.
4.
Manual frequency/preset channel
selector - Set emergency frequency.
5. Microphone switch - Press.
NB
Limits selectivity to narrowband intermediate frequency.
WB
Limits selectivity to wide-band
intermediate frequency of FM
band.
MEM LOAD
Momentary switch. If pressed,
loads manually selected frequency in preset channel memory
AM SQUELCH
Screwdriver adjustable potenticontrol
ometer. Squelch fully overridden
at full counter-clockwise position. Clockwise rotation increases input signal required to
open squelch.
Mode select
Three-position rotary switch.
switch
OFF
Shuts off receiver-transmitter.
TR
Selects transmit/receive modes.
DF
Not used.
SQ/DIS/TONE
Three-position switch.
select switch
Center
Selects squelch function,
SQ/DIS
Shuts off squelch function.
TONE
Transmits tone of approx. 1000
Hz.
Frequency
Four-position switch.
control
emergency/select
switch
PRE
Enables preset channel selection.
MAN
Enables manual frequency selection.
EMER AM or
Selects a prestored guard chanFM
nel.
VOL control
Clockwise rotation increases
volume.
c. Normal Operation.
(1) Turn-on.
1. Mode select switch (VHF AM/FM panel) TR.
(2) Emergency FM Mode.
1.
Transmit-interphone selector (audio
panel, fig. 3-1) - #2 position.
2. Mode select switch - TR or DF, as
desired.
3.
PRE-MAN-FM/AM
EMER switch EMER FM.
4.
Manual frequency/preset channel
selector - Set emergency frequency.
5. Microphone switch - Press.
6. Shutdown Mode select switch - OFF.
3-11
TM 55-1510-219-10
3-12. VOICE SECURITY
(PROVISIONS ONLY).
SYSTEM
phered communications at a distant location.
TSEC/KY-28
REG
Enables normal cipher or plain
communications.
ZEROIZE switch
Normally OFF. Place in ON position during emergency situations to neutralize and make inoperative the associated cipher
equipment.
CIPHER
Illuminates when PLAIN/
indicator
CIPHER switch is in CIPHER
position.
c. VHF/AM/FM Set and Voice Security Operation.
(1) Turn-on.
1. POWER ON switch (fig. 3-6) - ON.
NOTE
Voice security system TSEC/KY58 may be installed in lieu of
voice security system TSEC/KY28.
Complete provisions are
provided to install either the KY28 or the KY-58 voice security
system on the pedestal extension
(fig. 2-10).
a. Description. The KY-28 voice security system
provides
secure
(ciphered)
two-way
voice
communications for the pilot and copilot in conjunction
with the UHF and VHF/AM/FM command sets, and the
backup VOW set. System circuits are protected by the
VHF, VHF/AM/FM, RADIO RELAY, and BU VOW circuit
breakers on the overhead circuit breaker panel (fig. 226). Figure 3-6 illustrates the KY-28 voice security
(CIPHONY) control indicator.
b. Controls and Indicators.
(1) Voice security control/indicator. (fig. 3-6).
CONTROL
FUNCTION
INDICATOR
POWER ON
Turns set on or off.
switch
NOTE
The POWER ON switch must be in
ON position, regardless of the
mode of the operation, whenever
the voice security (CIPHONY) KY28 is installed in the aircraft.
(2) Receive.
1. SQUELCH
control (VHF/AM/ FM
panel) - As required.
2. Transmitter-interphone selector (audio
panel, fig. 3-1) - #2 position. or Audio
monitor control #2 - ON.
3. Mode selector (VHF/AM/FM panel) TR.
4. Frequency selectors (VHF/AM/FM
panel) - As required. PLAIN/CIPHER
switch (voice security panel) - As
required.
(3) Transmit (PLAIN).
1.
Transmit/interphone selector (audio
panel) - No. 2 position.
2. PLAIN/CIPHER switch (Voice security
panel) - PLAIN.
3. Microphone switch - Press.
(4) Transmit (CIPHER).
1.
Transmit/interphone selector (audio
panel) - No. 2 position.
2. PLAIN/CIPHER switch (Voice security
panel) -CIPHER. Indicator will be on
while switch is in CIPHER position.)
NOTE
The POWER ON switch must be in
ON position for FM liaison or
secure mission operations in
either the plain or cipher mode.
POWER ON
indicator
PLAIN indicator
PLAIN/CIPHER
switch
PLAIN
CIPHER
RE-X/REG
switch
RE-X
Illuminates when POWER ON
switch is placed in ON (up) position.
Illuminates when PLAIN/
CIPHER switch is in PLAIN position.
Enables unciphered communications on FM liaison set.
Enables ciphered communications on FM liaison set.
Enables re-transmission of ci3-12
TM 55-1510-219-10
Figure 3-6. Voice Security Control Indicator (C-8157/ARC) (KY-28)
NOTE
Voice security system TSEC/KY58 may be installed in lieu of
voice
security
TSEC/KY-28.
Complete provisions are provided
to install either the KY-28 or the
KY-58 voice security system on
the pedestal extension (fig. 2-10).
3. RE-X/REG switch (Voice security panel)
- As required. (Set RE-X position only if
distant station is using re-transmitting
equipment.)
4. Microphone switch - Press momentarily
(interrupted tone from voice security
unit should no longer be heard).
NOTE
No traffic will be passed if the
interrupted tone is still heard after
pressing
and
releasing
the
microphone switch.
The KY-58 voice security system provides secure
(ciphered) two-way voice communications for the pilot
and copilot in conjunction with the VHF/AM/FM
command set, UHF command set and the backup VOW
set. Voice security circuits are protected by the VHF,
VHF/AM/FM, RADIO RELAY, and BU VOW circuit
breakers on the overhead circuit breaker panel fig. 226). For KY-58 operating instructions, refer to TM 115810-262-OP and TM 11-5810-262-20.
5. Microphone switch- Press (do not talk).
Wait until beep is heard, then speak into
microphone.
(5) Shutdown.
1. Mode selector (VHF/AM/FM panel) OFF.
2. POWER ON switch (Voice security
panel) - OFF.
3-13. VOICE SECURITY SYSTEM TSEC/KY-58
(PROVISIONS ONLY).
3-14. HF COMMAND SET (718 U-5).
a. Description. The HF command set provides
long-range voice communications within the frequency
range of 2.000 to 29.999 MHz and employs
3-13
TM 55-1510-219-10
either standard amplitude modulation (AM), lower
sideband (LSB), or upper sideband (USB) modulation.
The distance range of the set is approximately 2,500
miles and varies with atmospheric conditions. The unit
is protected by a 3-ampere HF RCVR and the 25ampere HF PWR circuit breaker on the overhead circuit
breaker panel (fig. 2-26). The control panel is located
on the pedestal extension (fig. 2-10).
b. Controls and Indicators.
(1) HF control panel fig. 3-7).
CONTROL/
INDICATOR
Frequency
selector
Frequency
Aircraft can be configured for
either HF or VOW on position 4 of
the audio control panels.
2.
AVIONICS MASTER POWER switch
(overhead panel) - ON.
(2) Receive.
1. HF audio switch (#4, audio panel) - ON.
2. Function selector (HF panel, pedestal
extension) -Set to USB, LSB, or AM.
3. Volume knob (audio switch #4, audio
panel) - Set to mid position.
4. VOL knob (audio panel) - Adjust.
5. SQL knob (HF panel) - Adjust to just
quiet noise when no signal is being
received.
6. Tuning knob (HF panel) - Set desired
frequency.
(3) Transmit.
1. Transmitter-interphone selector (audio
panel) -No. 4 position.
2. Function selector (HF panel) - Set to
USB, LSB, or AM.
3. MIC HEADSET/OXYGEN MASK switch
(instrument panel) -As desired.
4. SQL knob (HF panel) - Adjust to just
quiet noise when no signal is being
received.
5. MIC switch - Depress to transmit.
(4) Shutdown.
1. Mode selector knob (HF panel) - OFF.
d. RF Test.
(1) Receive check.
1. Set SQL control fully clockwise and tune
to WWV. Check receive operation.
FUNCTION
Selects the desired operating frequency.
Displays the selected frequency.
indicator
SQL control
Adjusts squelch level.
RF TEST
Indicates operational status of
indicator
set when RT TEST is selected
with the function selector.
Function selector Turns set off and determines operating mode.
OFF
Turns set off.
USB
Selects upper sideband modulation.
LSB
Selects lower sideband modulation.
AM
Selects amplitude modulation.
CW
Not used in this installation.
SVU
Not used in this installation.
SVL
Not used in this installation.
RF TEST
Check operational status of the
system.
c. Normal Operation.
(1) Turn-on.
1. Insure aircraft power is on.
NOTE
Radio station WWV operates
continuously ton 2.500 Mhz,
5.000Mhz, 10.000 Mhz, 15.000 Mhz,
and 20.000 Mhz.
Select the
frequency that provides the best
received signal strength.
NOTE
It is presumed the AVIONICS
MASTER POWER switch is on and
that normally used avionic circuit
breakers remain depressed.
NOTE
3-14
TM 55-1510-219-10
Figure 3-7.HF Control Panel (718 U-5)
lamp during the tune cycle. Normal
(2) RF TEST lamp check (system not keyed).
tune time is 3 to 8 seconds.
1. Set the function selector to RF TEST.
Do not key the system. Observe the RF
TEST lamp.
NOTE
a. If the RF TEST lamp does not light,
RF TEST lamp indications, during
the lamp is defective or the
tune, are only valid for 3 to 8
receiver-exciter is faulty.
seconds after initial key is
b. If the RF TEST lamp blinks for more
applied.
To repeat the tune
than 1 minute, the receiver- exciter
check, key the HF system with the
is faulty.
microphone switch.
c.
If the RF TEST lamp lights
immediately or after an initial period
a. If the RF TEST lamp stays on
of blinking, the RF TEST lamp and
continuously, the receiver-exciter is
fault circuits are operational.
faulty.
d. Repeat RF TEST lamp check, as
b. If the RF TEST lamp blinks on and
required, by selecting RF TEST with
off, the PA-coupler is faulty.
the function selector.
c. If the RF TEST lamp extinguishes
(3) Tune check (system keyed).
after a nominal tune time, and
1. Set the function selector to RF TEST.
sidetone is heard when audio input
Change frequency selection to any
is supplied through the microphone,
frequency and key the system
the system is tuned and operating
momentarily. Observe the RF TEST
correctly for the selected frequency.
3-15
TM 55-1510-219-10
OFF and ON) also on the transmitter case, may be
actuated by inserting one finger through a small, round,
spring-loaded door on the right side of the aft fuselage.
The transmitter unit is accessible through a service
panel located on the bottom of the aft fuselage.
b. Controls - ELT Transmitter Case.
CONTROL
FUNCTION
RESET switch
When pressed, resets transmitter
Function switch
Selects operating mode of set.
ARM
Arms set to be actuated by impact switch (normal mode).
OFF
Turns set off.
ON
Activates transmitter for test
purposes.
e. HF Command Set - Emergency Operation.
Not applicable.
3-15.
EMERGENCY LOCATOR TRANSMITTER
(ELT).
a. Description. An emergency locator transmitter
(fig.3-8) provided to assist in locating an aircraft and
crew in the event an emergency landing is necessitated.
The output frequency is 121.5 and 243 MHz
simultaneously. Range is approximately line of-sight.
The transmitter unit has separate function control
switches located on one end of the case. In the event
the impact switch has been inadvertently actuated, the
beacon can be reset by firmly pressing the pushbutton
RESET switch on the front of the case. The RESET
switch and a 3-positiof toggle switch (placarded ARM,
3-16
TM 55-1510-219-10
Figure 3-8. Emergency Locator Transmitter (ELT-10)
3-17
TM 55-1510-219-10
Section III. NAVIGATION
circuit breaker on the overhead circuit breaker panel and
the 1.5-ampere F9 fuse on the No. 1 junction box.
b. Controls and Indicators.
(1) Switches (fig. 2-28).
INDICATOR
FUNCTION
Pilot's
Selects desired source of magnetCOMPASS No.
ic heading information for dis1/No. 2 switch
play on pilot's HSI and copilot's
RMI.
No. 1
Selects compass system No. 1
for display control.
No. 2
Selects compass system No. 2
for display control.
Copilot's
Selects desired source of magnetCOMPASS No.
ic heading information for dis1/No. 2 switch
play on co-pilot's HSI and pilot's
RMI.
No. 1
Selects compass system No. 1
for display.
No. 2
Selects compass system No. 2
for display.
RMI select
Selects which of two signals will
switch
be displayed on respective RMI
single needle pointer, if single
3-16. DESCRIPTION.
The navigation equipment group provides the pilot
and copilot with instrumentation required to establish
and maintain an accurate flight course and position, and
to make an approach on instruments under Instrument
Meteorological Conditions (IMC).
The navigation
configuration includes equipment for determining
attitude, position, destination range and bearing,
heading reference and groundspeed.
3-17. RADIO MAGNETIC INDICATORS (RMI).
a. Description. Two identical Radio Magnetic
Indicators (RMI) are installed, one for the pilot and one
for the copilot (fig. 3-9). Each unit serves as a
navigational aid for the respective user and, by means
of individual source select switches will display aircraft
magnetic or directional gyro heading and VOR, TACAN,
INS or ADF bearing information. The pilot's RMI is
protected by the 1-ampere #1 RMI circuit breaker on the
overhead circuit breaker panel (fig. 2-26) and the 1.5ampere F13 fuse on the No. I junction box. The
copilot's RMI is protected by the 1-ampere #2 RMI
Figure 3-9. Radio Magnetic Indicator (RMI) (332C-10)
3-18
TM 55-1510-219-10
Single needle
switch
INS position
needle switch is in the VOR/
TACAN position.
VOR 1
Selects VOR 1 bearing signals
for display.
TACAN
Selects TACAN bearing signal
for display.
(2) RMI (fig. 2-28).
CONTROL/
FUNCTION
INDICATOR
Double needle
Indicates bearing selected by
pointer
double needle switch.
Compass card
Indicates aircraft heading at top
of dial.
Heading index
Reference point for aircraft
heading.
Warning flag
Indicates loss of compass signal.
Double needle
Selects desired signal to be disswitch
played by double needle pointer.
ADF position
Selects ADF bearing information.
VOR position
Selects VOR 2 bearing information.
Single needle
Indicates bearing selected by sinpointer
gle needle switch.
VOR/TACAN
position
Selects desired signal to be displayed on single needle pointer.
Selects the bearing to waypoint
position.
Selects VOR 1 or TACAN bearing information.
3-18. HORIZONTAL SITUATION INDICATORS.
a. Description. The pilot and copilot have separate
HSI instruments on respective instrument panel
sections. Each HSI indicator combines displays to
provide a map-like presentation of the aircraft position.
Each indicator displays aircraft heading, course
deviation, and glideslope data. The pilot's HSI allows
the desired course and autopilot input to be set
manually. Either HSI will display back localizer sensing,
when the front course is selected and back course is
flown. Course deviation is supplied to the HSI by the
VOR 1 or VOR 2 systems, the TACAN (Tactical Air
Navigation System), or the INS (Inertial Navigation
System). Glideslope data is supplied by the VOR 1 or
VOR 2 systems. The HSI displays warning flags when
the VOR/LOC, INS, TACAN, heading or glideslope
signals are lost or become unreliable.
Figure 3-10. Pilot's Horizontal Situation Indicator (331A-8G)
3-19
TM 55-1510-219-10
markings. The symbol shows aircraft position and heading with
respect to a radio course and the
rotating dial.
Course deviation This bar represents the centerbar
line of the selected VOR or localizer course. The aircraft symbol shows pictorially, actual
aircraft position in relation to
this selected course.
HDG control
The heading marker is positioned on the rotating dial by
the heading knob and displays
preselected compass heading.
The heading marker rotates with
the heading dial so the difference between the marker and the
fore lubber line index is the
amount of heading error computer. In the heading mode, the
flight director vertical command
bar will display the required
bank commands to bring the aircraft onto and maintain the selected heading.
Course deviation In VOR operation, each dot repdots
resents five degrees deviation
from centerline. In ILS operation, each dot represents one degree deviation from centerline.
In INS position, each dot represents 3.75 NM linear deviation.
GS flag
Indicates that the information
displayed by the glideslope
pointer is invalid and should not
be used.
Glideslope
Indicates deviation from correct
pointer
glideslope during ILS approach.
Azimuth marks
Azimuth marks are fixed at 45°
bearings throughout 360 degrees
of compass card for quick reference.
(2) Pilot's VOR Switches (fig. 2-28).
CONTROL
FUNCTION
Pilot's course
Selects desired source of data for
indicator selector display on pilot's HSI and input
switch
to autopilot flight computer.
VOR I
Selects data from VOR I system.
VOR 2
Selects data from VOR 2 system.
TACAN
Selects data from TACAN system.
INS
Selects data from INS.
b. Controls and Indicators - Pilot's HSI.
(1) Pilot's HSI (fig. 3-10).
CONTROL/
FUNCTION
INDICATOR
Rotating heading The rotating heading dial, which
dial
rotates with the aircraft throughout 360 degrees, displays gyro
stabilized magnetic compass information. The azimuth ring is
graduated in five-degree increments.
Heading marker
Indicates heading selected by
HDG knob.
Lubber line
This pointed fixed index mark
aligns with the aircraft symbol to
indicate aircraft magnetic heading on the compass card.
Course arrow
Indicates VOR or TACAN radial course selected by COURSE
knob.
HEADING flag
Indicates that the heading information displayed is invalid and
should not be used.
COURSE
Provides a digital readout of the
indicator
selected course.
TO-FROM
Two to-from pointers are situatpointers
ed 180 degrees apart. The one
which is visible points toward
the station along the selected
VOR/TACAN radial.
NAV flag
The NAV flag indicates that the
information derived from the selected navigational beacon is invalid and should not be used.
COURSE
The yellow course arrow is posicontrol
tioned on the heading dial by
the course knob to select a magnetic bearing that coincides with
the desired VOR/TACAN radial
or localizer course. Like the
heading marker, the course arrow rotates with the heading dial
to provide a continuous readout
of course error to the computer.
When one of the radio modes is
selected, the vertical command
bar in the flight director will display bank commands to intercept and maintain the selected
radio source.
Aircraft symbol
A fixed aircraft symbol corresponds to the longitudinal axis
of the aircraft and lubber line
3-20
TM 55-1510-219-10
Heading marker
Pilot's compass
No. 1/No. 2
switch
No. 1
Selects desired source of magnetic heading data for display on
HSI compass card.
Selects data from No. 1 compass.
No. 2
Selects data from No. 2 compass.
c. Controls and Indicators - Copilot's HSI.
COMPASS flag
Lubber line
Course pointer
NOTE
A PILOT SELECT annunciator,
located on the copilot side of the
instrument
panel,
illuminates
whenever the setting of the
copilot's COURSE INDICATOR
switch matches that of the pilot's
COURSE INDICATOR switch. In
this condition the copilot's HSI is
"slaved" to the pilot's HSI, and the
copilot has no control over
course select functions of the
selected NAV receiver. The PILOT
SELECT annunciator does not
illuminate when INS is selected.
Rotating heading
dial
COURSE
indicator
VOR LOC flag
COURSE
control
(1) Copilot's HSI (fig. 3-11).
CONTROL/
FUNCTION
INDICATOR
Indicates heading selected by
HDG knob.
Indicates that the heading information displayed is invalid and
should not be used.
This pointed fixed index mark
aligns with the aircraft symbol to
indicate aircraft magnetic heading on the compass card.
Indicates VOR/TACAN radial
course selected by COURSE
knob.
The rotating heading dial, which
rotates with the aircraft throughout 360 degrees, displays gyro
stabilized magnetic compass information. The azimuth ring is
graduated in five-degree increments.
Provides a digital readout of the
selected course.
Flag indicates that information
derived from the selected navigational beacon is invalid and
should not be used.
The yellow course arrow is positioned on the heading dial by
Figure 3-11. Copilot's Horizontal Situation Indicator (331A-6P)
3-21
TM 55-1510-219-10
with
Course deviation
bar
Aircraft symbol
TO-FROM
pointers
HDG control
Course deviation
dots.
3.75 NM deviation from centerline.
GS flag
Indicates that the information
displayed by the glideslope
pointer is invalid and should not
be used.
Glideslope
Indicates deviation from correct
pointer
glideslope during ILS approach.
Azimuth marks
Azimuth marks are fixed at 45°
bearings throughout 360 degrees
of compass card for quick reference.
(2) Copilot's VOR Switches (fig. 2-28).
Copilot's Course
Selects desired source of data for
Indicator
display on copilot's HSI and inSelector Switch
put to autopilot.
VOR 1
Selects data from VOR 1 system.
VOR 2
Selects data from VOR 2 system.
TACAN
Selects data from TACAN system.
INS
Selects data from INS.
Copilot's
Selects desired source of magnetcompass No. 1/
ic heading data for display on
No. 2 switch
HSI compass card.
No. 1
Selects data from No. 1 compass.
No. 2
Selects data from No. 2 compass.
the course knob to select a
magnetic bearing that coincides
the desired VOR radial or localizer
course. Like the heading marker,
the course arrow rotates with the
heading dial to provide a continuous
readout of course error to the
computer. When one of the NAV
modes is selected, the vertical
command bar in the flight director
will display bank commands to
intercept and maintain the selected
radio source.
This bar represents the centerline of the selected VOR or localizer course. The aircraft symbol shows pictorially, actual
aircraft position in relation to
this selected course.
A fixed aircraft symbol corresponds to the longitudinal axis
of the aircraft and lubber line
markings. The symbol shows aircraft position and heading with
respect to a radio course and the
rotating dial.
Two to-from pointers are situated 180 degrees apart. The one
which is visible points toward
the station along the selected
VOR radial.
The heading marker is positioned on the rotating dial by
the heading knob and displays
preselected compass heading.
The heading marker rotates with
the heading dial so the difference between the marker and the
fore lubber line index is the
amount of heading error applied
to the flight director computer.
In the heading mode, the flight
director vertical command bar
will display the required bank
commands to bring the aircraft
onto and maintain the selected
heading.
In VOR operation, each dot represents five degrees deviation
from centerline. In ILS operation, each dot represents one degree deviation centerline. In INS
operation, each dot represents
3-19. HORIZON REFERENCE INDICATOR.
a. Description. The horizon reference indicator
(HRI) (fig. 3-12) is the pilot's basic attitude horizon
indicator and the attitude direction instrument for the
automatic flight control system.
b. Controls and Indicators.
CONTROL
FUNCTION
Crossed needles
Display computed commands to
autopilot.
Lateral deviation
Displays localizer deviation inindicator
formation from VOR receiver.
Vertical
Displays glideslope deviation indeviation
formation from VOR receiver.
indicator
Bank angle
Indicates aircraft bank angle.
pointer
Bank angle index Reference indicating zero-degree
bank.
Bank angle scale Allows measurement of aircraft
bank angle from zero to 60 degrees.
3-22
TM 55-1510-219-10
Figure 3-12. Horizon Reference Indicator (329B-9A)
Horizon line
Miniature
Sphere
GYRO flag
CMPTR flag
LOC flag
Affixed to sphere, remains parallel to the earth's horizon at all
times.
Indicates attitude of aircraft
aircraft with respect to the earth's horizon.
Remains oriented with the
earth's axis at all times.
Presence indicates loss of power
to, or low rotational speed of,
vertical gyro.
Presence indicates a malfunction
within the autopilot computer.
Presence indicates that localizer
information is not available or
not reliable.
GS flag
TEST
Presence indicates glideslope information is not being presented
on indicator.
When pressed, display indicates
pushbutton an additional 10°nose up, 20°
right roll and the GYRO flag is
visible.
3-20. PILOT'S TURN AND SLIP INDICATOR.
a. Description. The pilot's turn and slip indicator
(fig. 3-13) is used to provide automatic yaw damping
information to the autopilot in addition to performing the
functions of a turn -and slip indicator. It is protected by
the 5-ampere PILOT TURN & SLIP circuit breaker
located on the overhead circuit breaker panel (fig 2-26).
NOTE
When flying coupled to the INS
system, the CMPTR flag will be in
view
anytime
the
steering
information is invalid or a
malfunction exists in the autopilot
computer.
b. Controls and Indicators.
INDICATOR
FUNCTION
Turn rate
Deflects to indicate rate of turn.
indicator
Two-minute turn
Fixed markers indicate twomarks
minute turn rate when covered
by turn rate indicator.
GYRO warning
Presence indicates loss of power
to instrument.
3-23
TM 55-1510-219-10
Figure 3-13. Pilot's Turn and Slip Indicator (329T- 1)
Inclinometer
(
Indicates lateral acceleration
side slip) of aircraft.
Miniature
aircraft
3-21. COPILOT'S GYRO HORIZON INDICATOR.
a. Description. The copilot's gyro horizon indicator
(fig. 3-14) is a flight aid which indicates the aircraft's
attitude. The indicator is designed to operate through all
attitudes. There are no front panel fuses or circuit
breakers provided for the copilot's gyro horizon
indicator.
G flag
Sphere
Inclinometer
b. Controls and Indicators.
INDICATOR
FUNCTION
Bank angle index Reference indicating zero-degree
bank.
Bank angle
Indicates aircraft bank angle.
pointer
Bank angle scale Indicates aircraft bank angle
from zero to 90 degrees with
marks at 10, 20, 30, 45, 60, and
90 degrees.
Horizon line
Affixed to sphere, remains parallel to the earth's horizon at all
times.
Indicates attitude of aircraft
with respect to the earth's horizon.
Presence announces loss of power.
Indicates orientation with
earth's axis at all times.
Assists the copilot in making coordinated turns.
3-22. GYRO MAGNETIC COMPASS SYSTEMS.
a. Description. Two identical compass systems
provide accurate directional information for the aircraft
at all latitudes of the earth. As a heading reference, two
modes of operation are used: directional gyro (FREE)
mode, or slaved (SLAVE) mode. in polar regions of the
earth where magnetic heading references are not
reliable, the system is operated in the FREE mode. In
this mode, the system furnishes an inertial heading
reference, with latitude corrections introduced manually.
In areas where magnetic heading references are
reliable, the system is operated in the SLAVE mode. In
this mode, the directional gyro is slaved to the magnetic
flux detector, which supplies long-term magnetic
reference to correct the apparent drift of the gyro.
Magnetic heading information from both systems is
3-24
TM 55-1510-219-10
Figure 3-14. Copilot's Gyro Horizon Indicator (GH-14)
applied to various aircraft systems through pilot and
copilot COMPASS No. 1/No. 2 switches (fig. 2-28).
There are no circuit breakers for the gyro magnetic
compass systems. The circuits are protected by the 2ampere F2 and F6 fuses on the No. 1 junction box.
b. Vertical Gyro. A vertical gyro provides line-of
sight stabilization to the weather radar and roll and pitch
information to the autopilot. No controls are required or
provided for operation of the vertical gyro system. The
circuit is protected by the 3ampere F22 fuse in the No.
1 junction box.
1/No. 2 switch
No. 1
No. 2
COMPASS
SLAVE
annunciator
GYRO SLAVE/
FREE switch
SLAVE
c. Controls and Indicators.
CONTROL/
FUNCTION
INDICATOR
Pilot's
Selects desired source for magCOMPASS No.
netic heading information to dis1/No. 2 switch
play on pilot's HSI and copilot's
RMI.
No. 1
Selects compass system No. 1
for display.
No. 2
Selects compass system No. 2
for display.
Copilot's
Selects desired source for magCOMPASS No.
netic heading information to dis-
FREE
INCREASE/
DECREASE
INCREASE
DECREASE
3-25
play on copilot's HSI and pilot's
RMI.
Selects compass system No. 1
for display.
Selects compass system No. 2
for display.
Provides visual indication of
system synchronization operation.
Selects system mode of operation.
Selects slaved mode of operation.
Selects free mode of operation.
Provides manual fast synchronization of the system.
switch
Causes gyro heading output to
decrease (move in counterclockwise direction).
Causes gyro heading output to
increase (move in clockwise direction).
TM 55-1510-219-10
Either VOR can direct input signals to the flight
director indicator. Controls are shown in figure 315.
d. Normal Operation.
(1) To align.
1. Compass GYRO
switch SLAVE.
Each VOR system includes independent receiver
units for VOR/LOC and glideslope (GS). Each VOR
receiver provides a VOR input to a respective RMI
indicator and VOR and localizer data to the flight
director. Each glideslope receiver sends GS flag and
pointer deviation information to the flight director.
SLAVE/FREE
2. Compass INCREASE/ DECREASE
switch Hold switch momentarily in
the direction desired, and then
release. This will place system in
fast erect mode. The gyro will then
erect at approximately 30 degrees
per minute. While in the fast erect
mode, the HEADING flag (HSI) will
be in view. When the HEADING
flag retracts from view, the heading
displayed will be the magnetic
heading.
VOR/LOC indicators may be used for navigation
during manual control of the aircraft, or the autopilot
may be coupled to the VOR system, accepting VOR
inputs to the autopilot computer.
The pilot's unit (VOR 1) is a navigation radio system
which receives and interprets VHF Omnidirectional
Radio Range (VOR) signals, localizer (LOC) signals,
glideslope signals, and marker beacon signals. The
system operates in a VOR/LOC frequency range of
108.00 to 117.95 Mhz. Glideslope frequencies (329.15
to 335.00 Mhz) are paired with LOC frequencies, and
are automatically selected when LOC frequencies are
selected.
LOC frequencies are those frequencies
between 108.00 and 112.00 Mhz that end in odd tenths
(108.1, 109.3, 109.5, etc.). The marker beacon receiver
operates at 75 Mhz and is not tuneable.
(2) To determine magnetic heading.
1. Compass GYRO/SLAVE
switch SLAVE.
FREE
2. RMI rotating heading dial (compass
card) Read heading.
VOR 2 is similar to VOR 1 except VOR 2 cannot
receive or interpret marker beacon signals.
(3) To determine directional gyro heading.
1.
Compass GYRO/SLAVE
switch-FREE.
Each VOR system provides course deviation and
glide path data, which can be switched either to the
copilot's HSI or to the autopilot flight computer and
pilot's HSI, or both. The audio outputs of VOR 1 and
VOR 2 systems are supplied to the NAV control, on the
audio control panels. VOR 1 bearing data is supplied to
the single needle pointer on both Radio Magnetic
Indicators. VOR 2 bearing data is supplied to the
double-needle pointer on both Radio Magnetic
Indicators.
FREE
2. Compass INCREASE/ DECREASE
switch Hold until the RMI compass
card aligns with the magnetic
heading, then release.
3. Read heading. The heading will
agree with the appropriate HSI.
e. Shutdown. Both compass systems are shut
down when the inverter switch is turned off. (If either
system is ON, both compass sets will be energized.)
Both systems use a single VOR/LOC antenna,
located on the vertical stabilizer, and a single glideslope
antenna, located in the radome. VOR 1 uses a marker
beacon antenna located on the underside of the forward
fuselage.
3-23. VOR/LOC NAVIGATION SYSTEM.
The VOR 1 system is protected by the 2-ampere
VOR #1 and 35-ampere AVIONICS MASTER PWR #1
circuit breaker on the overhead circuit breaker panel
(fig. 2-26). The VOR 2 system is protected by the 2ampere VOR #2 and 35-ampere AVIONICS MASTER
PWR #2 circuit breakers on the overhead circuit breaker
panel.
a. Description. The aircraft is equipped with two
VOR systems which are controlled by the NAV-I NAV-2
control panel located in the pedestal (fig. 2-11).
3-26
TM 55-1510-219-10
Figure 3-15. Dual NAV 1 - NAV 2 Control Panel (VIR-30AGM, VIR-30AG)
b. Controls and Indicators.
Frequency
indicator
NAV TEST
switch
(1) NA V 1 Panel Section (VOR 1).
CONTROL/
FUNCTION
INDICATOR
Frequency
Displays selected frequency of
indicator
VOR 1 receiver.
Frequency
Selects operating frequency of
control
VOR 1 receiver.
VOL/OFF
Activates VOR 1 receiver. Percontrol
mits monitoring VOR 1 audio
and adjust volume of signals received.
NAV TEST
Activates self test of VOR 1 navswitch
igation system.
Displays selected frequency of
VOR 2 receiver.
Activates self test of VOR 2 navigation systems.
(3) Marker Beacon Indicators. (fig. 2-28).
INDICATOR
FUNCTION
"A" indicators
Illuminate when aircraft passes
(white)
over an inner marker beacon.
"O" indicators
Illuminate when aircraft passes
(blue)
over an outer marker beacon.
"M" indicators
Illuminate when aircraft passes
(amber)
over a middle marker beacon.
(4) Switches (fig. 2-28).
CONTROL
FUNCTION
Pilot's COURSE Selects VOR receiver to control
INDICATOR
pilot's HSI.
switch
VOR 1
VOR 1 controls pilot's HSI.
VOR 2
VOR 2 controls pilot's HSI.
Copilot's
Selects VOR receiver to control
COURSE
copilot's HSI.
INDICATOR
switch
(2) NA V 2 Panel Section (VOR 2).
CONTROL/
FUNCTION
INDICATOR
VOL/OFF
Activates VOR 2 receiver. Percontrol
mits monitoring VOR 2 audio
and adjust volume of signals received.
Frequency
Selects operating frequency of
control
VOR 2 receiver.
3-27
TM 55-1510-219-10
VOR 1
VOR 2
NAV-A switch
MKR BCN HILO switch
VOR 1 controls copilot's HSI.
VOR 2 controls copilot's HSI.
Applies VOR audio to respective
headsets.
Controls sensitivity of marker
beacon receiver.
(4) Glideslope operation.
1. Frequency selectors (NAV panel)
Set desired localizer frequency.
2. VOR switch (instrument
Select VOR source.
panel)
3. Glideslope pointer (HSI) Fly aircraft
to center pointer.
c. Operation.
(1) Turn-on.
(5) Self-test.
1. Insure aircraft DC and AC power is
on.
(a) NA V 1.
2. AVIONICS
MASTER
POWER
switch (overhead panel) ON.
1. Frequency select knob (NAV panel)
Select a VOR frequency.
3. Frequency controls (NAV panel) Set
for both receivers.
2. NAV TEST switch (NAV panel)
Press and hold.
4. VOL knobs (NAV panel) Turn
clockwise to activate sets, adjust
volume.
3. RMI Observe that single needle
indicates approximately 005°.
4. VOR/LOC flag Check that flag is out
of view.
5. NAV A, audio switch ON. Confirm
proper signal, then OFF.
5. TO/FROM pointer
pointer indicates TO.
6. Confirm proper indications RMI and
HSI.
Check
that
6. HSI course deviation bar Check for
centered bar.
(2) Normal operation.
7. Marker beacon lights Check that
all three lamps are illuminated and
flickering at approximately a 30 Hz
rate.
(a) Pilot, copilot COURSE INDICATOR
switches (instrument panel): Select
VOR source.
(b) To determine course to station on
pilot's HSI: Rotate course knob until
course deviation pointer is centered
and TO-FROM flag reads TO.
8. VOR frequency knob (NAV panel)
Select a LOC frequency.
9. VOR/LOC flag Check that flag is out
of view.
(c) To determine bearing from station
on pilot's HSI.: Rotate course knob
until course deviation pointer is
centered and TO-FROM flag reads
FROM.
10. HSI course deviation bar Check that
bar indicates a deflection of
approximately one dot right of
center.
(d) To determine course to station on
RMI.: Select VOR, verify needle
points course to station.
11. HSI glideslope pointer Check that
pointer indicates a deflection of
approximately one dot below center.
12. Marker beacon lights Check that
all three lamps are illuminated and
flickering at approximately a 30 Hz
rate.
(3) Localizer (LOC) operation.
1. VOR frequency knob (NAV panel)
Select frequency.
2. Pilot's,
copilot's
COURSE
INDICATOR switches (instrument
panel) Select VOR source.
(b) NA V 2. All NAV 2 self-test procedures are the same
as those used for NAV 1, with the
exception of the marker beacon
test. There is no marker beacon
receiver in the NAV 2 system.
3. Course deviation bar (HSI) Steer
aircraft to center bar.
3-28
TM 55-1510-219-10
aircraft position, or monitor conventional low frequency
AM radio transmissions. The system is designed to
provide reliable reception of a 400-watt radio station at a
range of 65 nautical miles throughout a 360-degree turn
of the aircraft. It operates in a frequency range of 190
to 1750 kilohertz.. Bearing indications are displayed
visually on the RMI's and aural signals are applied to the
audio control panels. Refer to Audio Control Panels (fig.
3-1) for controls used to monitor ADF audio signals.
(6) Shutdown. VOL knobs (NAV panel) OFF.
3-24. MARKER BEACON.
a. Description. A marker beacon receiver module
located inside the No. 1 VOR receiver, detects a
marker beacon signal as the aircraft passes over a
marker beacon transmitting antenna. The detected
signal is selectively filtered to activate the appropriate
marker beacon lamp on the instrument panel. In
addition, the detected signal is coupled to the aircraft
audio system to annunciate marker beacon passage.
The marker beacon receiver operates on a fixed
frequency of 75 Mhz. Volume and sensitivity of the
system are adjusted on the radio control panel (fig. 3-2)
located on the pedestal extension (fig. 2-11).
The ADF consists of a receiver, located on the
forward side of the aft cabin bulkhead inside the
pressure vessel; a control unit, located on the pedestal
extension;, a non-directional sense antenna, installed in
the aircraft dorsal fin; a directional loop antenna, located
on the underside of the fuselage; and a quadrangle error
corrector, installed on the loop antenna (to compensate
for the deflection of arriving radio signals by the wings
and fuselage of the aircraft). The system is protected by
the 1-ampere ADF, the 5-ampere RADIO RELAY, and
the 35ampere AVIONICS BUS #2 circuit breakers on
the overhead circuit breaker panel (fig. 2-26).
NOTE
Keying the HF radio set while
operating the ADF set will cause a
momentarily
unreliable
ADF
signal.
b. Controls and Indicators.
b. Controls and Indicators.
CONTROL
MKR BCN HILO switch
MKR BCN VOL
control
FUNCTION
Controls sensitivity of marker
beacon receiver.
Adjusts volume of marker beacon audio.
(1) Marker beacon indicators (fig. 2-28).
INDICATOR
FUNCTION
"A" indicators
Illuminate when aircraft passes
(white)
over an inner marker beacon.
"O" indicators
Illuminate when aircraft passes
(blue)
over an outer marker beacon.
"M" indicators
Illuminate when aircraft passes
(amber)
over an middle marker beacon.
(1) ADF control panel (fig. 3-16).
CONTROL/
FUNCTION
INDICATOR
LOOP control
Operative only when the function switch is in the LOOP or
ADF position. Center position
removes rotation signals from
the loop antenna and the ADF
pointer on the RMI's. First position L (left) or R (right) of center
applies slow speed rotation signals to loop antenna and ADF
pointer on RMI's for 360-degree
rotation left or right. Second position L(left) or R (right) of center applies fast speed rotation
signals to loop antenna and ADF
pointer on RMI's for 360-degree
rotation left or right. Center position holds antenna position.
BFO/OFF
At BFO (on) setting, permits
fine tuning with Beat Frequency
Oscillator (BFO). Also provides
audio tone when receiving
modulated CW. OFF turns BFO
off.
c. Operation.
1. Marker beacon indicator lights
(instrument panel) -Confirm beacon
indication.
2. Marker beacon HI-LO switch (radio
control panel) -As required.
3. Select NAV A on audio control
panel to monitor marker beacon
audio.
3-25. AUTOMATIC DIRECTION FINDER (DF-203).
a. Description. The Automatic Direction Finder
(ADF) is a radio navigation system which provides a
visible and audible indication of aircraft bearing relative
to a selected ground radio station. It may also be used
to home on a selected station, find
3-29
un-
TM 55-1510-219-10
Figure 3-16. ADF Control Panel (DF-203)
Tuning meter
TUNE control
Range switch
FREQUENCY
indicator
Mode selector
OFF
ADF
Indicates relative strength of received signals.
Selects operating frequency.
Selects operating frequency
band.
Indicates selected frequency.
(3) Switches - radio panel (fig. 3-2).
CONTROL
FUNCTION
FILTER V-OFF Selects whether voice filter will
switch
be used with ADF audio.
FILTER R-OFF Selects whether range filter will
switch
be used with ADF audio.
c. Operation.
Selects operating mode.
Turns set off.
Permits automatic direction
finding or homing operation.
ANT
Permits reception using sense
antenna.
LOOP
Permits audio-null homing and
manual direction finding operations.
GAIN control
Adjusts volume of received signal.
(2) Switches - audio panel (fig. 3-1).
CONTROL
FUNCTION
NAV B volume
control
NAV B volume control applies
ADF audio to aircraft audio system.
(1) To operate set as automatic direction
finder.
1. Mode selector - ADF.
2. BFO-OFF switch - BFO.
3. Range switch - Select.
4. TUNE control - Rotate for maximum
reading on tuning meter and zero
BFO beat.
5. GAIN control - As required.
6. Double needle switches (RMI, fig. 39) - As required.
3-30
TM 55-1510-219-10
Their range, though limited to line-of-sight, is designed
to provide reliable reception of a TACAN ground station
at a distance of 170 nautical miles at an aircraft altitude
of 20,000 feet. The normal time required for the
systems to lock on to a selected ground station signal is
three seconds.
The avionics TACAN system is
protected by the 2ampere TACAN circuit breaker
located on the overhead circuit breaker panel (fig. 226).
7. Double needle on RMI Read course
to station.
(2) To operate set for sense antenna direction
finding: 1. Mode selector ANT.
2. Range switch Select.
3. TUNE control rotate for maximum
reading on tuning meter.
4. GAIN control As required.
b. Avionics TACAN System. The AN/ARN136(V)
avionics TACAN system consists of the RT1321
receiver-transmitter and the CP-1398 azimuth computer,
both of which are located in the right nose avionics
compartment; the ID-2218 range indicator, located on
the instrument panel, the C-11265 control panel, located
in the pedestal extension; and an antenna located on
the top surface of the aircraft fuselage. The avionics
TACAN system operates in conjunction with TACAN and
VORTAC ground stations to provide distance, ground
speed, time-to-station, and bearing-to-station data. It
operates in the L band frequency range on one of 252
pre-selected frequencies, 126 X mode and 126 Y mode
channels. Course deviation from TACAN stations is
displayed on the HSI. Distance, time-to-station, and
ground speed are displayed on the TACAN digital
display (fig. 3-17). The ground speed and time-tostation are accurate only if the aircraft is flying directly
toward the ground station at a sufficient distance that
the slant range and ground range are nearly equal.
(3) To operate set for audio-null direction
finding:
1. Mode selector ANT.
2.
BFO-OFF switch BFO.
3. Range switch select.
4. TUNE control Tune desired station.
5. GAIN control Adjust for minimum
audio output.
6. Double needle switches (RMI, fig. 39) As required.
7. BFO-OFF switch OFF.
8. Mode selector LOOP.
9. LOOP switch L or R. Turn left or
right until a null is reached
(minimum sound in headsets).
10. Double needle on RMI (fig.
Read course to station.
NOTE
The true null and direction to the
radio station may be indicated by
either end of the single needle.
This ambiguity must be solved to
determine proper direction to the
station.
(4) Shutdown. Mode selector OFF.
3-26. TACAN SYSTEMS.
3-9)
The avionics TACAN system may be operated by
the flight director system or connected to and used with
the autopilot system. When employed as the primary
means of navigation, aircraft flight may be controlled
manually or by the autopilot. Indications of aircraft
heading and bearing to ground stations are displayed on
the course deviation indicators (HSIs). Relative bearing
to a station is displayed by the RMI bearing pointer.
TACAN distance, ground speed, and time-to-station are
all displayed on the TACAN indicator located on the
copilot's instrument panel (fig. 2-28).
The TACAN control panel (fig. 3-17) enables
selection of the TACAN frequency (channel) to be used,
and provides for self-test of TACAN circuits. X or Y
channel is selected by the X/Y switch. Most TACAN and
VORTAC stations are operated on the X mode. When Y
mode stations are operational, air navigation charts will
designate the Y mode stations. A toggle switch provides
system power ON/OFF control.
Audio control is
provided by a rotary control placarded VOL.
c. INS TACAN System. The INS TACAN system
is of the same type as the avionics TACAN system.
This set, inaccessible to pilot control, is coupled directly
a. Description.
Two Tactical Air Navigation
(TACAN) systems are provided. One is dedicated to the
INS and is used only for position updating; the other is
used in conjunction with other avionics systems,
including the flight director system and the autopilot.
TACAN is a radio navigation system which provides
aircraft distance and bearing information relative to a
TACAN ground station. Both systems operate in the Lband frequency range of 962 to 1213 MHz.
3-31
TM 55-1510-219-10
to INS circuits. It is dedicated only to updating the INS,
is activated when the INS is operational, and is
controlled only by the INS. The INS TACAN consists of
a range unit and a distance indicator, both located on
the INS equipment rack and both identical to their
counterparts in the avionics TACAN, and an antenna
located on the underside of the fuselage. No controls
are required or provided for the INS TACAN system.
The system is protected by a circuit breaker on the INS
junction box.
d. Controls and Indicators.
CONTROL/
FUNCTION
INDICATOR
TEST switch
Activates system self-test.
Channel
Displays TACAN channel selectindicator
ed.
X-Y switch
Selects X or Y mode for TACAN channels.
VOL control
Adjusts TACAN volume.
Channel
Dual knob for manual selection
selectors
of operating channel.
Outer knob
Selects tens and hundreds part
of channel number.
Inner knob
Selects units part of channel
number.
ON-OFF switch
NM indicator
Activates or deactivates system.
Displays slant range distance in
nautical miles from aircraft to
selected TACAN ground station.
KT indicator
Displays ground speed in knots.
MIN indicator
Displays time to TACAN station
in minutes.
e. Avionics TACAN Operation.
(1) Turn-On.
1. ON-OFF switch (TACAN panel) ON.
2. VOL knob As required.
3. Course
indicator
switches
(instrument panel) -Select TACAN.
NOTE
The pilot and copilot should not
select TACAN simultaneously for
test.
(2) Normal operation.
1. Pilot's
single
VOR/TACAN.
Figure 3-17. TACAN Control Panel and Range Indicator (ID-2218)
3-32
needle
switch
TM 55-1510-219-10
2. RMI select switch (side panels)
TACAN.
3. Pilot's Course
TACAN.
indicator
(3) System test.
1. TEST pushbutton Press and hold.
switch
4. Mode switch (X/Y) As required.
2. Range indicator Check for an
indication of 0.0 ± 0.1 nautical
miles.
5. TACAN control panel Select desired
channel.
3. Pilot's COURSE selector switch
Select TACAN.
6. Wait
5
seconds
for
acquisition and lock-on.
signal
4. Pilot's RMI selector switch Select
TACAN.
7. Insure
that
audio
station
identification signal is correct for the
ground station selected.
5. RMI double needle Check for an
indication of 180°± 2°.
8. Bearing pointer on RMI
bearing heading to station.
6. HSI course selector Turn to 180°
and adjust slowly until the course
deviation bar is centered. The bar
should center between a selected
course of 178°to 182°.
Read
9. COURSE knob (pilot's HSI) Set
course desired.
7. HSI course selector Turn the
selector + 10° from the setting
achieved in step 6, and check that
course deviation bar is located over
the far left 10°dot.
10. Course
deviation
bar (pilot's
HSI)Read deviation from selected
course. Course arrow will show
wind correction angle when the
course deviation bar is centered and
the aircraft is tracking the selected
course.
8. HSI course selector Turn the
selector -10° from setting achieved
in step 6, and check that course
deviation bar is located over the far
right 10°dot.
11. TACAN indicator Read range (NM).
12. To
determine course TO
or
course FROM a TACAN station,
rotate course knob (pilot's HSI) until
course deviation bar is centered and
the TO/FROM flag reads TO or
FROM.
9. TO-FROM indicator Check that TO
is indicated.
10. TEST pushbutton Release.
13. To use TACAN during pilotcontrolled flight, control aircraft by
manual controls, responding to
information displayed on the flight
director, RMI, TACAN, and other
instruments.
(4) Shutdown. Turn ON/OFF switch OFF.
3-27. AUTOMATIC FLIGHT CONTROL SYSTEM.
a. Description.
The AP-106 is an integrated
autopilot/flight director system that provides the
following:
14. To use TACAN with the autopilot,
depress A/P ENGAGE and monitor
autopilot performance on flight
director,
RMI,
and
TACAN
indicators.
Verify adherence to
preset heading and course, and
confirm
the
execution
of
displayed steering commands.
NOTE
The TACAN ground speed reading
will be accurate only when the
aircraft is on a course directly to
or from the TACAN station. When
headed away from the station, the
TACAN indicator minutes reading
will be in error.
Heading mode
Navigation mode
Approach mode
capture and track
with
automatic
Altitude hold mode
Back course localizer mode
Go-around mode
3-33
glideslope
TM 55-1510-219-10
Selection of a mode causes the legend of that
pushbutton switch to illuminate. The self-test switch on
the lower right of the autopilot control panel acts as a
lamp test when depressed. For operation at night,
overall illumination of the autopilot mode selector and
switches is adjusted by the PILOT'S INSTRUMENT light
control.
Synchronized control wheel steering
Indicated airspeed hold mode
All-angle adaptive capture for VOR, LOC, and LOC B/C
Attitude display
Heading display
Mode selection indicators
CONTROL/
INDICATOR
HDG switch
Elevator trim indicator
System integrity warning flags
Automatic yaw damping
Turn and slip indicator
NAV switch
The flight director and autopilot have a common
computer system. When the autopilot is engaged, the
flight control system controls the aircraft and the pilot
monitors the flight path by observing the information
displayed on the pilot's horizon reference indicator (HRI)
and the pilot's horizontal situation indicator (HSI) (flight
director system indicators).
Autopilot/flight director commands are selected at
the autopilot mode selector panel (fig. 3-18) on the
pilot's side of the instrument panel. Manual roll rate and
pitch commands are inserted at the autopilot pitch-turn
panel (fig.
3-21).
Autopilot operational status is
indicated by the autopilot/flight director annunciator
positioned above the pilot's horizon reference indicator
on the instrument panel (fig. 2-28). Two autopilot
switches are also provided on each control wheel (fig.
2-16). One switch is placarded PITCH SYNC & CWS
(pitch synchronize and control wheel steering), and the
other is placarded DISC TRIM/AP YD (disconnect
trim/autopilot yaw damp).
FUNCTION
Engages heading mode. Commands aircraft to acquire the
heading indicated by heading
marker on pilot's HSI.
Engages navigation mode. VORI/VOR-2 or TACAN selected,
commands aircraft to intercept
and track VOR radial selected
by course knob on pilot's HSI.
INS selected, commands aircraft
to track steering signals from
INS system. Intercept of approximately 45°and tracking will be
computed by the INS system.
NOTE
APPR cannot be selected with INS
selected.
APPR switch
ALT switch
IAS switch
Power for the system is provided through a
10ampere AP PWR circuit breaker located on the
overhead circuit breaker panel (fig. 2-26).
b. Controls and Indicators. The following controls
and indicators are provided for operation of the system.
B/C switch
ENG-DIS switch
(1) Instrument Panel (Pilot's).
ENG
(a) Autopilot Mode select panel (fig. 318). The autopilot/flight director commands are selected
by the autopilot mode selector.
Selection is
accomplished by pressing the face of the appropriate
push-on/push-off switch. The lateral modes are HDG,
NAV, APPR and B/C. When not in a lateral mode, the
flight director command bars are biased out of view.
The vertical modes are ALT, IAS, and pitch (all hold
modes). If a vertical mode is not selected, the pitch
hold mode is automatically operational.
DIS
TRIM UP
TRIM DN
indicator
3-34
Engages approach mode. Commands aircraft to intercept and
track ILS inbound course.
Engages altitude hold mode.
Commands aircraft to maintain
pressure altitude.
Engages indicated airspeed hold
mode. Commands aircraft to
maintain airspeed.
Engages backcourse mode. Commands aircraft to intercept back
course ILS.
Controls coupling of the automatic pilot.
Engages autopilot and illuminates engaged indicator.
Disengages autopilot and illuminates disengaged indicator.
Illuminates when autopilot is
indicator driving trim servo in up
direction.
Illuminates when autopilot is
driving trim servo in down direction.
TM 55-1510-219-10
Figure 3-18. Autopilot Mode Selector Panel (614E-42A)
Self-test switch
Tests display and selector indicator circuits when depressed.
(b) Autopilot trim test switch (fig 3- 18).
AUTOPILOT
TRIM TEST
switch
Simulates trim system malfunctions and illuminates AP TRIM
FAIL warning annunciator light.
(c) Aileron high torque mode test switch and annunciator
(fig. 3-19).
CONTROL/
INDICATOR
AIL HI
TORQUE
annunciator
TEST switch
FUNCTION
Illumination is automatic from
ground to 10,000 feet to show
aileron servo is set to operate in
high torque mode. Light extinguishes automatically above 10,
000 feet to indicate aileron servo
has terminated high torque
mode operation.
Normally off. Used only below
10,000 feet (TEST position) to
confirm operability of aileron
servo high torque mode, when
AIL HI TORQUE annunciator
light extinguishes.
Figure 3-19. Aileron High Torque Test Switch and
Annunciator
3-35
TM 55-1510-219-10
(d) Autopilot/flight director annunciator
panel (fig.
3-20).
The autopilot/flight director
incorporates its own annunciator panel located just
above the flight director display on the instrument panel.
The modes and indications given on the annunciator
panel are placarded on the face of the lenses and
illuminate when the respective conditions are indicated.
Dimming of the annunciator panel lights is provided by a
switch adjacent to the panel placarded DIM-BRT.
CONTROL/
INDICATOR
NAV ARM
indicator
NAV CAP
indicator
GS ARM
indicator
GS CAP
indicator
GA indicator
BACK LOC
indicator
ALT indicator
AP DISC
indicator
AP ENG
indicator
IAS indicator
FUNCTION
Illuminates when computer is
armed to accept navigation signals.
VOR-I/TACAN
illuminates
when selected radial is captured.
INS selected, illuminates when
INS is coupled to the flight director.
Illuminates when approach
mode is selected prior to glideslope capture. Extinguishes after
glideslope capture.
Illuminates when glideslope is
captured.
HDG indicator
LIN DEV
DIM BRT
control
Illuminates when go-around
mode is selected.
Illuminates when back course
mode is selected.
Illuminates when altitude hold
mode is selected.
Illuminates when autopilot is
disengaged.
Illuminates when autopilot is engaged.
Illuminates when airspeed hold
mode is selected.
Illuminates when heading mode
is selected.
Not Applicable
Adjusts intensity of illumination
of the flight director annunciator
(2) Pedestal - left power lever (fig. 2-11).
CONTROL/
FUNCTION
INDICATOR
GO AROUND
When pressed, autopilot disconswitch (outboard
nects, GA annunciator light (fig.
side left power
3-20) illuminates, and horizon
lever)
reference indicator commands
wings level, 7°nose up attitude.
Autopilot may be re-engaged to
follow the command.
Figure 3-20. Autopilot/Flight Director Annunciator Panel
3-36
TM 55-1510-219-10
(3) Pedestal extension.
(a)
21).
CONTROL
Turn control
knob
Pitch control
thumbwheel
Autopilot pitch-turn panel (fig.
faces, enabling the pilot to manually fly the aircraft to the desired attitude until button is
released.
c.
Turn-on. Power is applied to the system
anytime the aircraft avionics bus is energized.
d. Autopilot Modes of Operation.
3-
FUNCTION
Supplies roll rate commands to
autopilot. Spring loaded to center detent.
Supplies pitch rate commands to
autopilot. Spring loaded to center detent.
(1) Attitude. The autopilot is in the attitude
mode when the ENG-DIS switch (AP mode select panel)
is in the ENG position and no mode selector switches
(HDG, NAV, etc.) have been selected (fig 3-18). The
autopilot will fly the aircraft and accept pitch and roll rate
commands from the autopilot pitch-turn panel (fig. 321).
(4) Control wheel switch (fig. 2-16).
CONTROL/
FUNCTION
INDICATOR
DISC/TRIM/AP
When pressed to first detent, auYD pushbutton
topilot system and yaw damp
are disconnected. When pressed
to second detent, electric trim is
disconnected.
PITCH SYNC &
This button, on each control
CWS pushbutton wheel, may be used instead of
the pitch-turn control to establish the aircraft in a desired atti-
(a) Guidance mode.
When the
autopilot is in the attitude mode and a mode selector
switch (HDG, NAV, etc.) is pressed, the autopilot
accepts steering commands from the computer.
Depending on which selector switch (AP mode select
panel) is pressed, autopilot operation can be described
by the following subguidance modes.
(b) Heading mode. When HDG mode
is selected (AP mode select panel), with au topilot
engaged, the autopilot will fly the aircraft to, and then
maintain, the heading under the heading marker on the
pilot's HSI.
tude. Depressing the button
causes the autopilot servos to
disengage from the control sur-
Figure 3-21. Autopilot Pitch-Turn Control Panel
3-37
TM 55-1510-219-10
NOTE
The heading marker may be
preset to the go-around heading
after the localizer is captured.
After go-around airspeed and
power settings are established,
select the HDG mode to clear the
go-around mode. Pitch attitude
will remain at that used for goaround until changed with the
PITCH SYNC & CWS button or the
selection of a vertical mode.
(c) Navigation mode. When NAV mode
is selected (AP mode select panel), the system initially
switches to the NAV ARM heading hold submode, as
shown by illumination of NAV ARM and HDG indicators
(AP/flight director annunciator panel). The autopilot will
then command the aircraft to follow the heading under
the heading marker on the pilot's HSI (with the heading
marker set to produce the desired VOR or localizer
intercept angle). The flight computer will compute a
capture point based on deviation from the desired radio
beam, the rate at which the aircraft is approaching this
beam, and the course intercept angle. When beam
capture occurs, the HDG and NAV ARM indicator lamps
(AP/flight director annunciator panel) will extinguish and
the NAV CAP lamp will illuminate. The autopilot will
then track the selected radio course with automatic
crosswind correction.
(g) Pitch hold mode. The pitch hold
mode is selected by (1) selecting one of the vertical
mode selector switches, or (2) actuating the pitch
synchronize and control wheel steering switch (PITCH
SYNC & CWS), located on each control wheel.
(d) Back-course mode. When BACK
LOC mode is selected on the autopilot mode selector
panel, localizer capture is the same as in a frontcourse
approach in NAV or APPR mode.
Glideslope is
inhibited during a back-course approach. The HSI must
be set to the front-course heading so that lateral
deviation will be directional.
(h) Control wheel steering mode.
Pressing one of the PITCH SYNC & CWS switches
located on each control wheel disconnects the autopilot
servos from the control surfaces, allows the pilot to fly
the aircraft to a new pitch attitude, and synchronizes the
vertical command bar (pilot's horizon ref ind) to aircraft
attitude. The ALT or IAS mode will disengage (if
selected) when the PITCH SYNC & CWS button is
depressed. When the autopilot is coupled to the HDG,
NAV, APPR, or BACK LOC modes, releasing the PITCH
SYNC & CWS switch will cause the autopilot to couple
to the previously selected mode.
(e) Approach mode.
When APPR
mode is selected (AP mode select panel), localizer
capture is the same as in the NAV mode but glideslope
arm and capture functions are also provided. When the
APPR mode is selected the NAV ARM annunciator lamp
will illuminate, indicating that the system is armed for
localizer capture.
As the aircraft approaches the
localizer beam, the NAV CAP annunciator lamp will
illuminate. Once the localizer is being tracked, the GS
ARM annunciator lamp will illuminate.
Glideslope
capture is dependent on localizer capture and must
occur after localizer capture. The localizer is always
captured from a selected heading, but the glideslope
may be captured with the autopilot operating in any
vertical mode (pitch hold, altitude hold, or indicated
airspeed hold), and from above (not recommended) or
below the glideslope.
At the point of glideslope
intercept, the GS CAP annunciator lamp will illuminate
and all preselected vertical modes will be cleared.
(i) Altitude hold mode. Pressing the
ALT switch (AP mode select panel) when desired
altitude has been reached (with autopilot engaged) will
(1) cause the autopilot to fly the aircraft to maintain the
barometric altitude at which the aircraft was flying when
ALT switch was pressed, (2) illuminate the ALT
annunciator lamp (AP/flight director annunciator panel),
and (3) display the altitude hold commands on the
vertical command bar (pilot's horizon reference
indicator)
(j) Indicated airspeed hold mode.
Pressing the IAS switch (AP mode select panel) when
desired airspeed has been reached (with autopilot
engaged) will (1) cause the autopilot to fly the aircraft to
maintain the indicated airspeed at which the aircraft was
flying when IAS switch was pressed, (2) illuminate the
IAS annunciator lamp (AP/flight director annunciator
panel), and (3) display IAS hold commands on the
vertical command bar of the pilot's horizon reference
indicator.
(f) Go-around mode. Pressing the GO
AROUND button on the outboard side of the left power
lever selects the go-around mode. Go-around mode
may be selected from any lateral mode (HDG, NAV,
APPR, or BACK LOC). When go-around mode is
selected: (1)the autopilot is disengaged, (2) the GA
annunciator lamp will illuminate, and (3) a command
presentation for wings level and 3-38 7°nose up pitch
attitude will appear on the pilot's horizon reference
indicator.
3-38
TM 55-1510-219-10
2. In any function except "after
glideslope
capture",
use
the
autopilot pitch control for climbing
or descending. Movement of the
pitch control establishes a pitch rate
that is proportional to knob
displacement. If any vertical
mode button has been selected, it
will automatically release when the
pitch control knob is rotated.
e. Autopilot Operation.
(1) Engaging autopilot.
AP mode select
panel ENG/DIS switch to ENG position.
NOTE
When the autopilot is engaged,
the
yaw
damper
is
also
automatically
engaged
as
indicated by the lighted YAW
DAMP button (AP pitch/turn
control panel).
The autopilot and flight director are coupled when
both units are engaged. When coupled, the autopilot
accepts guidance commands from the flight director.
When the flight director is not engaged, the autopilot
accepts pitch and roll commands from the pitch/turn
control knobs as selected by the pilot.
The autopilot may be engaged in any reasonable
attitude and in either the coupled or uncoupled mode.
The autopilot will smoothly acquire the command
attitude. When uncoupled, the autopilot will maintain
the bank and pitch attitude at the time of engagement.
(2) Disengaging autopilot.
3. When HDG mode is selected, the
autopilot will command the aircraft
to execute a turn then maintain the
heading set by the heading marker.
4. Use the autopilot turn control to
command a roll rate when the
autopilot is engaged. At the time
control is returned to detent, the
autopilot maintains the bank angle
(up to approximately 30 degrees).
Rotating the turn control when the
autopilot is engaged and a lateral
mode is selected will cause the
selected lateral modes to release.
(a) The autopilot may be disengaged by
any of the following:
1. Actuating
compass
INCREASE/DECREASE switch.
(4) Control wheel synchronization.
The
PITCH SYNC & CWS button on the pilot's control wheel
can be used instead of the pitch/turn control to establish
the aircraft in a desired attitude. Depressing this button
causes autopilot elevator and aileron servos to
disengage from the control surfaces. The pilot then flies
the aircraft manually to a desired attitude, releases the
PITCH SYNC & CWS button to re-engage the servos,
and the autopilot holds the established attitude.
2. Pressing TEST button on AP mode
select panel.
3. Pressing GO
AROUND switch
on left power lever.
4. Pressing control wheel DISC switch
to AP YD position.
(b) The following functions will cause
the autopilot to automatically disengage:
The ALT and IAS mode will immediately disengage
(if selected) when the PITCH SYNC & CWS button is
pressed. Upon release of the PITCH SYNC & CWS
button, the autopilot will couple to the previously
selected lateral mode.
NOT
The APPR mode will not
disengage when the PITCH SYNC
& CWS button is depressed.
When the button is released, the
aircraft will return to the localizer
course and glideslope.
1. Vertical gyro failure.
2. Directional gyro failure.
3. Autopilot power or circuit failure.
4. Torque limiter failure.
(3) Maneuvering.
1. To change flight functions, press the
desired mode button on the AP
mode selector panel. The button
will illuminate along its edges and
autopilot annunciator lights on the
instrument panel will illuminate,
indicating the respective modes in
operation.
3-39
TM 55-1510-219-10
4. NAV switch (AP mode select panel)
Press.
(Observe illumination of
NAV ARM annunciator).
f. Takeoff and Climb Out.
(1) Before takeoff
1. Heading marker (pilot's HSI) Set to
runway heading.
5. NAV CAP annunciator (AP/flight
director annunciator panel) Monitor.
(At point of capture, NAV CAP
illuminates).
NOTE
Except as described below, do
not select a different VOR
frequency, TACAN channel, or
course once a course and
intercept have been programmed
or capture achieved. To select a
different course or VOR/TACAN
frequency, return to the HDG
mode, select the course or
frequency, return to the NAV
mode, then reset the desired
beam.
(2) To change course over a VOR station. To
change course over a VOR station while operating in
NAV mode, if course change is less than 30°: COURSE
knob (pilot's HSI) Set desired heading in COURSE
window.
2. HDG switch (AP mode select
panel) Press.
Do not engage
autopilot.
(2) Takeoff Pressing the PITCH SYNC &
CWS switch on control wheel will provide pitch sync,
and the cross-pointers on the pilot's horizon reference
indicator will command flight to the pitch attitude that
existed when the PITCH SYNC & CWS switch was
pressed.
(3) Climb out.
1. Establish climb profile.
2. ENG/DIS switch (AP mode select
panel) Set to ENG (above 200 feet
AGL).
3. IAS switch (AP mode select panel)
Press (if desired).
4. HDG knob (pilot's course indicator
selector) Move heading marker as
required for heading changes.
(3) To change course over a VOR station. To
change course over a VOR station while operating in
NAV mode, if course change is greater than 30°:
(4) Cruise altitude.
1. Vertical speed Reduce to approx.
500 feet per minute (just before
reaching cruise altitude).
1. HDG knob (AP mode select panel)
Set desired intercept heading under
heading marker.
2. ALT button (AP mode select panel)
Press (when reach cruise altitude).
2. HDG switch (AP mode select panel)
Press. (Observe HDG illuminates
(AP/flight
director
annunciator
panel).
g. VOR Operation.
(1) To establish aircraft on a desired VOR
radial, perform the following
1. VOR receiver
frequency.
Tune
3. COURSE knob (pilot's HSI) Set new
course in COURSE window.
appropriate
4. NAV switch (AP mode select panel)
Press
(Observe
NAV
ARM
illuminates
(AP/flight
director
annunciator panel).
2. COURSE knob (pilot's HSI) Set
desired course TO or FROM station
shown in COURSE window.
5. NAV
CAP
(AP/flight
director
annunciator
panel)
-Monitor.
(Illumination means capture of new
radial).
3. HDG knob (pilot's HSI) Set desired
beam intercept angle under heading
marker. (The intercept angle with
respect to the radio beam may be
any angle of 90°or less).
3-40
TM 55-1510-219-10
2. Disengages autopilot.
h. Automatic Approach Front Course.
NOTE
The localizer and glideslope are captured
automatically on the ILS front course
approach. The localizer must be captured
before glideslope capture can occur. The
localizer is always captured from a selected
heading, but the glideslope may be captured
from any of the vertical modes and from
above (not recommended) or below the
glideslope.
1. VOR receiver Tune appropriate
frequency.
3. Pilot's
HRI
shows
command
presentation for wings level and 7°
nose up climb attitude.
NOTE
The heading marker may be preset to goaround heading after the localizer is
captured.
After go-around airspeed and
power settings are established, selection of
the HDG mode will clear the go-around
mode. Pitch attitude will remain at that used
for go-around until changed with the PITCH
SYNC & CWS button or by selection of a
vertical mode.
j. Back Course Approach. As in front course
approach, the localizer is captured automatically. The
aircraft should be maneuvered into the approach area
by setting the heading marker and functioning in the
HDG mode.
2. COURSE knob (pilot's HSI) Set
inbound
runway
heading
in
COURSE window.
3. HDG knob (pilot's HSI) Set heading
marker to desired intercept angle.
4. HDG selector switch (AP mode
select panel) Press. (Observe HDG
illuminates on AP/flight director
annunciator panel).
1. VOR receiver (NAV panel) Tune
localizer frequency.
2. COURSE knob (pilot's HSI) Set
front course inbound runway
heading in COURSE window.
5. Vertical mode (AP mode select
panel) Select IAS or ALT.
3. HDG knob (pilot's HSI) Set heading
marker to desired intercept angle.
6. APPR switch (AP mode select
panel) Press. (Observe NAV ARM
illuminates on AP/flight director
annunciator panel).
4. HDG switch (AP mode select panel)
Press.
(Observe HDG lamp
illuminates on AP/flight director
annunciator panel).
7. NAV
CAP
(AP/flight
director
annunciator
panel)
-Observe
illumination (when capture of
localizer course is achieved).
5. B/C switch (AP mode select panel)
Press. (Observe BACK LOC and
NAV
ARM
lamps
illuminate
(AP/flight
director
annunciator
panel) indicating system is armed
for back localizer capture.
8. GS
ARM
(AP/flight
director
annunciator
panel)
-Observe
illumination (when autopilot is
armed for glideslope capture).
6. NAV CAP lamp will illuminate when
system has captured back localizer
course.
9. GS
CAP
(AP/flight
director
annunciator
panel)
-Observe
illumination (which confirms that all
vertical modes are cleared, and also
confirms that autopilot is tracking
glideslope).
7. PITCH control (AP/pitch-turn panel)
Use to establish and maintain
desired rate of descent.
i. Go-Around. If visual runway contact is not
made at decision height, go-around may be activated by
pressing the GA button on the left power lever, and may
be initiated from any lateral mode (HDG, NAV, APPR,
B/C) with the following results:
NOTE
The HDG mode should be used within one
mile of the runway due to the large radio
deviations encountered when flying over the
localizer transmitter.
1. Illuminates GA light on autopilot
annunciator panel.
3-41
TM 55-1510-219-10
1. Any interruption or failure of power.
k. Yaw Damper Operation.
2. Vertical gyro failure indication.
1. The rudder channel of the autopilot
may be selected separately for yaw
damping by depressing the YAW
DAMP switch on the pedestal. The
switch face will illuminate when the
yaw damper is engaged.
3. Flight control system power or
circuit failure.
4. Autopilot trim failure.
3-28. INERTIAL NAVIGATION SYSTEM.
2. To disengage the yaw damper,
press the disconnect button on the
pilot's or copilot's control wheel to
the first detent or press the YAW
DAMP switch on the pedestal.
a. Description. The Inertial Navigation System
(INS) is a self-contained navigation and attitude
reference system. It is aided by (but not dependent
upon) data obtained from the ground-based data link
system, its own TACAN system, the aircraft encoding
altimeter, the true airspeed computer, and the gyro
magnetic compass system. The position and attitude
information computed by the INS is supplied to the
automatic flight control system, weather radar system,
horizontal situation indicator, radio magnetic indicators,
and to mission equipment. The INS is independent of
aircraft maneuvers, weather conditions and terrain, and
in conjunction with other aircraft equipment permits
operations on instruments alone under Instrument
Meteorological Conditions (IMC). The INS provides a
visual display of present position data in Universal
Transverse Mercator (UTM) coordinates or conventional
geographic (latitude-longitude) coordinates during all
phases of flight. When approaching the point selected
for a leg switch, an ALERT lamp will light informing the
pilot of an eminent automatic leg switch or the need to
manually insert course change data. The INS may be
manually updated for precise aircraft present position
accuracy by flying over a reference point of known
coordinates. The INS may be updated automatically by
the TACAN system or the Data Link system. Altitude
information is automatically inserted into the INS
computer by an encoding altimeter whenever the INS is
operational.
3. Refer to Emergency Procedures for
other means of disconnecting the
yaw damper.
I. Disconnecting Autopilot. The autopilot may be
disconnected by any of the following actions:
1. Pressing the DISC TRIM/AP YD
switch to first detent. (Location:
outboard horn either control wheel.)
2. Placing the ENG-DIS switch to DIS
position. (Location: AP mode select
panel.)
NOTE After assuming
manual control, fly the aircraft using
the same heading, course, and
attitude displays to monitor autopilot
operation prior to assuming manual
control.
m. Emergency Procedures. The autopilot can be
disengaged by any of the following methods:
1. Press the AP/YD disconnect switch
(either control wheel).
2. Move the engage lever to DIS
position (either control wheel).
3. Engage the go-around mode (yaw
damper will remain on).
The Mode Selector Unit (MSU) (fig. 3-22) controls
system activation and selects operating modes.
4. AP PWR and AFCS DIRECT circuit
breakers (overhead panel) Pull.
The Control Display Unit (CDU) (fig. 3-23) provides
controls and indicators for entering data into the INS and
displaying navigation and system status information.
5. AVIONICS
MASTER
POWER
switch (overhead panel) -OFF.
6. Aircraft MASTER switch (overhead
panel) OFF.
The INS system is protected by the 10-ampere
PRIME POWER and the 5-ampere HEATER POWER
circuit breakers on the mission AC/DC power cabinet, by
the 5-ampere INS CONTROL circuit breaker on the
overhead circuit breaker panel and by the 20-ampere
circuit breaker on the front of the battery unit.
n. Automatic Disengagement.
The following
conditions will cause the autopilot to disengage
automatically.
3-42
TM 55-1510-219-10
Figure 3-22. Mode Selector Unit (C-IV-E)
Leveling starts after fast-warmup heaters are off.
b. Controls and Indicators.
(1) Mode selector unit (fig. 3-22).
CONTROL
FUNCTIONS
Mode select
Controls INS activation and seknob
lects operating modes.
OFF
Deactivates INS.
STBY (ground
To STBY from OFF mode.
use only)
Starts fast warmup of system to
operating conditions; activates
computer so information may be
inserted; all INS-controlled
warning flags will indicate warning.
To ALIGN from NAV mode:
INS is not downmoded but will allow automatic
shutdown if overtemperature is detected.
NAV
To NAV from STBY mode:
Causes INS to automatically sequence through
STBY and ALIGN to NAV mode, if present position is
inserted and aircraft is parked.
To STBY from any other mode:
INS operates as if in attitude reference mode; all
INS warning flags and lamps, except ATTITUDE and
PLATFORM HEADING, will indicate warning.
ALIGN (ground
use only, parked)
Activates normal navigation
mode after automatic alignment
is completed; must be selected
before moving aircraft.
Used to shorten time in STBY and to bypass battery
test, if stored heading is valid.
ATT
To ALIGN from STBY mode.
Starts automatic course alignment mode (if fast-warmup heaters are off); fine alignment will
not start until present position is
inserted into CDU.
BAT lamp
To ALIGN from OFF mode:
READY NAV
lamp
3-43
Activates attitude reference
mode. Used to provide only INS
attitude signals. Shuts down
computer and CDU leaving only
BAT and WARN lamps operative. Once selected, INS alignment is lost.
Lights to indicate INS shutdown
due to low battery unit voltage.
Lights to indicate INS high accuracy alignment is attained. If attained during ALIGN mode,
TM 55-1510-219-10
lamp remains illuminated until
NAV mode is selected.
Lamp lights momentarily during
alignment, if alignment accomplished while in NAV mode.
(2) INS control display unit 6fig. 3-23).
CONTROL
FUNCTIONS
INDICATOR
HOLD key
Used with other CDU controls
to stop present position display
from changing, in order to update position and to display recorded malfunction codes.
Lights when pressed first time;
goes out when pressed second
time or when inserted data is accepted by computer.
When pressed second time, allows displays to resume showing
changing current present position.
ROLL LIM key Allows selection of Limited Roll
steering mode. Press to select
mode, key lights. Roll steering
output is limited to 10 degrees.
Data display, left
and right
INSERT/
ADVANCE/ HI
PREC key
ALERT lamp
WARN lamp
Figure 3-23. INS Control Display Unit (C-IV-E)
3-44
Press second time to exit mode,
key light extinguishes. Roll steering output returns to normal
limit of 25 degrees.
Composed of lamps which illuminate to display numbers, decimal points, degree symbols, left
and right directions, and latitude
or longitude directions.
Allows insertion of loaded data
into computer. Enters displayed
data into INS. When pressed before pressing any numerical key,
alternates display of normal and
high precision data.
Illuminates amber to alert pilot
1.3 minutes before impending
automatic course leg change. Extinguishes when switched to new
leg, if AUTO/MAN switch is set
to AUTO.
Flashes on and off when passing
waypoint, if AUTO/MAN switch
is set to MAN. Light will extinguish if AUTO is selected or if a
course change is inserted.
Lights red to alert pilot INS selftest circuits have detected a system fault. Illumination may be
TM 55-1510-219-10
Keyboard
caused by continuous or intermittent condition. Intermittent
conditions light lamp until reset
by TEST switch. If continuous
condition does not degrade attitude operation, lamp goes out
when mode selector is set to
ATT.
Consists of 10 keys for entering
load data into data and FROMTO displays.
(TEST)
mode. Pilot must make
waypoint changes manually.
When pressed, performs test of
INS lamps and displays, remote
lamps and indicators controlled
by INS, and computer input/
output operations.
Used with other controls to activate display of
numerical codes denoting specific malfunctions and
resets malfunction warning circuits.
During alignment, activates the HSI test. Continued
pressing of switch provides constant INS outputs to
drive cockpit displays in a predetermined fashion.
NOTE
The INS can provide test signals to the
Horizontal Situation Indicator (HSI)and
connected displays. Pressing TEST switch
during Standby, Align, or NAV modes will
cause all digits on connected digital
displays to indicate "8's" and illuminates the
HSI and ALERT lamps. Additional HSI test
signals are provided when INS is in Align
and the data selector is at any position other
than DSRTK/STS. Under those conditions,
pressing TEST switch causes the HSI to
indicate heading, drift angle, and track angle
error all at "0°" or "30°". At the same time,
cross track deviation is indicated at "3.75"
nautical miles (one dot) right or left and INScontrolled HSI flags are retracted from view.
"N", "S", "E", and "W" (on keys 2,8,6 and 4) indicate
direction of latitude and longitude. TAC and DISP (on
keys "7" and "9") enable loading and display of TACAN
station data. MV and DISP (on keys "3" and "9") are
associated with loading and display of Magnetic
Variation and Magnetic Heading. GRID and DISP (on
keys "5" and "9") are associated with loading and display
of UTM coefficients.
CLEAR key
When pressed, illuminates and
erases data loaded into data displays or FROM-TO display.
Used to cancel erroneous data.
After clearing, data loading can
be resumed.
WYPT CHG key
When pressed, enables numbers
in FROM-TO display to be
changed. If INSERT/ADVANCE
key is pressed, computer will use
navigation leg defined by new
number in all navigation computations. If INSERT/ADVANCE
key is not pressed, computer will
continue using original numbers
in all navigation computations;
but distance/time information,
based on new leg, may be called
up and read in data displays (in
case of waypoints). When not in
TACAN mix mode, TACAN station number is inserted to display DIS/TIME information.
AUTO-MAN
This is a dual purpose control.
TEST switch
When the knob is pressed inward, the TEST switch function
is engaged. When the knob is rotated to either the AUTO or
as a selector between those
modes.
(AUTO)
Selects automatic leg switching
mode. Computer switches from
one leg to the next whenever
waypoint in "TO" side of the
FROM-TO display is reached.
(MAN)
Selects manual leg switching
Output test signal are also supplied to the autopilot
when INS steering is selected. Rotating AUTO/MAN
switch to AUTO and pressing TEST during Align
furnishes a 15°left bank steering command. A 15°right
bank steering command is furnished when the
AUTO/MAN switch is set to MAN.
FROM-TO
display
Display numbers defining
waypoints of navigation leg being flown or in the case of a
flashing display, displays TACAN station being used.
Waypoint numbers automatically change each
time a waypoint is reached. Unless flight plan
changes during flight, the automatic leg switching sequence will always be 1 2, 2 3, 3 4....8 9, 9 1, 1 2,
etc.
Data selector Selects data to be displayed in
data displays or entered into
INS. The rotary selector has 10
positions. Five positions (L/L
3-45
TM 55-1510-219-10
TK/GS
HDG/DA
XTK/TK
L/L POS
POS, L/L WYPT, UTM POS,
UTM WYPT and DSRTK/STS)
also allow data to be loaded into
data display then inserted into
computer memory.
Displays aircraft track angle in
left display and ground speed in
right display.
Displays aircraft true heading in
left display and drift angle in
right display.
Displays cross track distance in
left display and track angle error
in right display.
Displays or enters present aircraft position latitude in left
data display and longitude in
right data display. Both displays
indicate degrees and minutes to
nearest tenth of a minute.
The extra precision display
shows meters.
UTM WYPT
Displays or enters waypoint and
TACAN station data in UTM
coordinates. Also enables loading and display of spheroid coefficients if GRID and DISP keys
are pressed simultaneously.
Dim knob
Controls intensity of CDU key
lamps and displays.
Waypoint/DME Thumbwheel switch, used to seselector
lect waypoints for which data is
to be inserted or displayed.
Waypoint station "O" is for display only and cannot be loaded
with usable data.
c. INS - Normal Operating Procedures.
NOTE
The following data will be required prior to
operating the INS: Local magnetic variation,
latitude
and
longitude
or
Universal
Transverse Mercator (UTM) coordinates of
aircraft during INS alignment.
This
information is necessary to program the INS
computer during alignment procedure.
This position also enables the insertion of present
position coordinates during alignment and present
position updates.
L/L WY PT
Displays or enters waypoint and
TACAN station data, if used in
conjunction with the waypoint/
TACAN selector.
NOTE
When inserting data into INS computer,
always start at left and work to right. The
first digit inserted will appear in right
position of applicable display. It will step to
left as each subsequent digit is entered. The
degree sign, decimal point, and colon (if
applicable) will appear automatically.
(1) Preflight procedures.
CAUTION
Insure that cooling air is available to
navigation unit before turning the INS on.
This position will also cause display of inertial
present position data when the HOLD key is illuminated.
DIS/TIME
WIND
DSRTK/STS
UTM POS
Displays distance from aircraft
to TACAN station or any
waypoint, or between any two
waypoints in left display. Displays time to TACAN station or
any waypoint, or between any
two waypoints, in right display.
Displays wind direction in left
display and wind speed in right
display, when true airspeed is
greater than the Air Data System
lower limit (115-400 KIAS).
Displays desired track angle to
nearest degree in the left data
display, and INS system status
in right data display.
Displays or enters aircraft position in Universal Transverse
Mercator (UTM) coordinates,
with northing data in kilometers
in left display and easting data
in kilometers in right display.
NOTE
The INS requires an aircraft compass system
to provide accurate navigational data and
mission data. The compass system selected
with the copilot's compass select switch will
provide information to the INS platform.
3-46
TM 55-1510-219-10
NOTE
Aircraft must be connected to a
ground power unit if INS
alignment is performed prior to
engine starting. In this event, the
engines must not be started until
after the INS is placed in the NAV
mode.
1. Applicable circuit breakers Check
depressed.
b. (MSU) Mode selector NAV.
(Observe (CDU) FROM-TO display indicates "1 2",
and INSERT/ADVANCE pushbutton illuminates.)
NOTE
Aircraft must not be towed or
taxied during INS alignment.
Movement of this type during
alignment causes large navigation
errors. If aircraft is moved during
alignment, restart alignment by
setting mode selector to STBY,
then
back
to
ALIGN
and
reinserting present position.
2. INS Mode selector (MSU) ALIGN.
Confirm following:
a. (CDU) FROM-TO display indicates
"1 2".
b. (CDU)
INSERT/ADVANCE
pushbutton light illuminates.
Passenger or cargo loading in the aircraft could
cause the type of motion which affects the accuracy of
alignment. Any activity which causes the aircraft to
change attitude shall be avoided during the alignment
period.
c. (CDU) DIM knob Adjust for
maximum brightness of CDU
displays.
c. (CDU) BAT lamp illuminates for
approx.
12 seconds, then
extinguishes when in alignment
state 8.
NOTE
Avoid passenger or cargo loading
or any activity which may cause
aircraft to change position or
attitude during alignment.
If
aircraft
is
moved
during
alignment, it will be necessary to
restart alignment by setting the
mode selector to STBY, then back
to ALIGN and reinserting present
position data.
d. (CDU) AUTO/MAN/TEST
AUTO.
e. (CDU) Data selector UL POS or
UTM POS.
(Observe coordinates of last present position prior to
INS turn-off appear in data displays.)
f. (CDU) AUTO/MAN/TEST switch
Press and hold for test. Confirm
following on CDU:
(Left and right data displays indicate "88°88.8 N/S"
and "88°88.8 E/W" respectively.)
3. (CDU) DIM knob Adjust for optimum
brightness of CDU displays.
4. (CDU)
AUTO.
switch
AUTO/MAN/TEST switch
(FROM-TO display indicates "8.8".)
5. (CDU) Data selector L/L POS or
UTM POS, as desired.
(Observe coordinates of last present position prior to
INS shutdown appear in data displays.) NOTE The
stored heading procedure should have been performed
within the last week.
6. Perform recommended stored
heading
alignment
preflight
procedure:
(The following pushbuttons and lamps illuminate:
ROLL LIM, HOLD, INSERT/ADVANCE, WYPT CHG,
ALERT, BAT (on CDU and MSU), WARN, and READY
NAV.)
g. (CDU) AUTO/MAN/TEST switch
Release. Confirm following:
(Data displays indicate coordinates in computer
memory; all lamps and pushbuttons illuminated in Step
F except INSERT/ADVANCE pushbutton light go out.)
a. Appropriate
circuit
breakers
Depress
(including
those
for
magnetic compass system.)
h. If UTM coordinates are to be used
Verify
that
appropriate
Grid
coefficients have been loaded.
3-47
TM 55-1510-219-10
NOTE
To achieve best accuracy, engine
start and heavy loading activity
should be delayed until entry into
NAV mode.
i. Present position data Insert.
NOTE
After present position 'has been
inserted
and
computer
has
advanced to state "7", present
position cannot be reinserted
without downmoding to STBY
and restarting alignment.
j. Data selector DSRTK/STS
NOTE
With the data selector in either the
L/L WY PT or UTM WY PT
position, waypoint and TACAN
station data may be loaded at any
time after turn-on. The operator
should occasionally switch the
data selector to DSRTK/STS to
monitor system status indicators
in the right data display.
(Observe that the left data display indicates the
desired track stored in computer memory and that the
right data display indicates status "-194.")
NOTE
If the fourth digit from the right, in
the right data display is blank,
then a valid heading has not been
stored.
Proceed with normal
preflight procedure.
k. Monitor the right data display for a
change from alignment state "9" to
alignment state "8". This will be
indicated by a reading of "--184".
(2) (CDU) AUTO/MAN/TEST switch Press
and hold. Confirm:
1. Left and right data displays indicate
"88°88.8 N/S" and "88°88.8 E/W"
respectively.
2. FROM-TO display is "8.8".
3. Following
pushbuttons
and
lamps illuminate: ROLL LIM, HOLD,
INSERT/ADVANCE, WYPT CHG,
ALERT, BAT (on CDU and MSU),
WARN and READY NAV.
l. Monitor the right data display for
malfunction codes (table 3-5).
(Loss of 26 VAC is indicated by an illuminated
WARN lamp and a reading of ".03184" in the right-hand
display. An inoperative magnetic compass system is
indicated by an extinguished WARN lamp and a reading
of ".03184" in the right-hand display.)
(3) TEST switch Release. (Observe data
displays indicate coordinates in computer memory; also
all lamps and pushbuttons lighted in step 7 except
INSERT/ADVANCE pushbutton extinguish.)
(4) Appropriate grid coefficients Verify loaded
(if UTM coordinates are to be used.)
NOTE
If above displays appear, a stored
heading alignment may not be
possible.
(5) Verify UTM grid coefficients.
1. Data selector UTM WYPT.
m. If there are malfunction codes,
proceed
to
ABNORMAL
PROCEDURES in this chapter.
2. Keys "5" and "9" Press simultaneously.
Observe FROM-TO display is blank. Earth flatness
coefficient appears in left display. The relative earth
radius, in meters, appears in right display.)
NOTE
These values are retained from
turn-on to turn-on unless changed
by operator.
3. Verify that values correspond to
those required for spheroid being
used.
n. When alignment state "7" is
reached, the INS will advance to the
NAV mode.
(Observe that "---05" will appear in the STS display.)
3-48
TM 55-1510-219-10
Table 3-1. Various Values for UTM Grid Coefficients
NOTE
Values for various spheroids are
listed in Table 3-1.
4. If values are correct, return CDU to
normal
display
mode
by
momentarily setting data selector to
any position except UTM WYPT. If
values are to be changed, continue
with following steps:
NOTE
The INS geographic position, as
read in L/L displays, will not be
affected by any changes in these
coefficients.
5. Keys "2"or "8" Press to indicate
following is the flatness coefficient.
(Observe INSERT/ADVANCE pushbutton light
illuminates.)
6. Load earth flatness coefficient by
pressing
keyboard
keys
in
sequence.
(Example: 2 9 4 9 8 = 29498. Observe number
appears in left display as keys are pressed.)
7. INSERT/ADVANCE
pushbutton
Press.
(Observe INSERT/ADVANCE pushbutton light goes
out.)
8. Keys "4" or "6" Press, to indicate
following load is relative earth
radius.
(Observe INSERT/ADVANCE pushbutton light
illuminates.)
10. INSERT/ADVANCE
pushbutton
Press.
(Observe INSERT/ADVANCE pushbutton light
extinguishes.)
11. Data selector UTM POS.
(Observe coordinate will reflect values of new
spheroid.)
(6) Insert present position.
9. Load relative earth radius by
pressing
keyboard
keys
in
sequence.
(Example: 8 2 0 6 m = 8206. Observe numbers
appear in the right display as keys are pressed.)
(7) Perform abbreviated INS interface test As
required:
NOTE
Assuming a level aircraft, attitude
indicators will become level
during alignment State "8" and
remain level in all modes until INS
is shut down. Warning indicators
for INS attitude signals from INS
are valid while attitude sphere
display is level.
NOTE
Zone symbol is to be ignored.
3-49
TM 55-1510-219-10
NOTE
The INS can provide test signals to the
Horizontal Situation Indicator (HSI) and
connected displays. Pressing TEST switch
during Standby, Align, or NAV mode causes
all digits on connected digital displays to
indicate "8's," and lights the HSI and the
ALERT lamp. Additional HSI test signals are
provided when INS is in Align and data
selector is at any position other than
DSRTK/STS.
Under those conditions,
pressing TEST switch causes HSI to indicate
heading, drift angle, and track angle error all
at "0°" or "30°." At the same time, cross
track deviation is indicated at "3.75" nautical
miles (one dot) right or left and INScontrolled HSI flags are retracted from view.
MSU
CDU
HSI
Flight Director/
Autopilot
Mission Control
Panel
All lamps illuminated.
All lamps illuminated. All "8's"
displayed.
All angles 30°. Cross-track deviation bar is one dot right. All INS
flags are retracted.
A 15°steering command is issued.
LINK UPDATE annunciated.
TACAN UPDATE annunciated.
7. CDU TEST switch Hold depressed,
and rotate AUTO/MAN switch to
AUTO.
(Observe all indications are as in step (6) except a
15' left steering command is issued. On HSI, All angles
are "0°" and cross-track deviation bar is one dot left.)
NOTE
Output test signals are supplied to the
autopilot when INS steering is selected.
Rotating AUTO/MAN switch to AUTO and
pressing TEST during Align furnishes a 15°
left bank steering command. A 15° right
bank steering command is furnished when
AUTO/MAN switch is set to MAN.
8. CDU TEST switch Release.
desired, decouple INS.
(Observe operation returns to normal.)
If
NOTE
Prior
to
pressing
INSERT/ADVANCE
pushbutton, any incorrectly loaded data can
be corrected by pressing CLEAR pushbutton
and reloading correct data.
NOTE
The quick test procedure may be performed
any time after Alignment State "8" is reached
and prior to entry into NAV.
NOTE
While parked aircraft is undergoing
alignment, encoding altimeter will supply
the field elevation (aircraft altitude) into INS.
1. Mode selector ALIGN.
(Observe CDU displays are illuminated.)
2. Data
selector
DSRTK/STS
(monitor right data display until
State "8" (or lower) is reached.)
(Observe right data display is ---N4,
where "N" is not "9".)
NOTE
Once present position has been inserted
and computer has advanced to alignment
state "7", present position cannot be
reinserted without downmoding to STBY
and restarting alignment.
3. AUTO/MAN switch MAN.
NOTE
If longitude and latitude coordinates are
being used, skip step 8 (a) and proceed with
step 8 (b).
(a) Insert UTM coordinates of aircraft
present position:
4. INS Couple to flight director and
autopilot, as applicable.
5. Data selector Set to any position
except DSRTK/STS.
6. TEST switch Press and hold.
1. Data selector UTM POS.
(Observe that prior to initial load,
ADVANCE pushbutton light illuminates.)
NOTE
After performing the preceding
step, observe:
3-50
INSERT/
TM 55-1510-219-10
(Observe latitude and longitude data is displayed in
UTM and INSERT/ ADVANCE pushbutton light
extinguishes.)
NOTE
The
computer
will
convert
coordinates in the overlap area;
however display values will
reference appropriate zone.
2. To load zone and easting values
Press keys in sequence, starting
with"E".
(Example: Zone 16, 425 km East = E16 425.
Observe that zone and easting in kilometers appear in
right data display as keys are pressed.)
3. INSERT/ADVANCE
pushbutton
Press.
(Observe pushbutton light remains illuminated.)
4. To load northing data Press keys
in sequence, starting with "N" or "S"
to
indicate
north
or
south
hemisphere.
(Example: 4749 km North = N 4749) Observe
northing kilometers appear in left data displays as keys
are pressed.)
5. INSERT/ADVANCE
pushbutton
Press.
(Observe pushbutton light remains illuminated.)
6. INSERT/ADVANCE
pushbutton
Press.
(Observe extra precision display for present position
northing and easting, to the nearest meter, appears in
left and right data displays, respectively.)
7. To load extra precision easting data
Press keys in sequence, starting
with "E".
(Example: 297 m East = E 297) Observe that
easting meters appear in right data display as keys are
pressed.)
8. INSERT/ADVANCE
pushbutton
Press.
(Observe pushbutton light remains illuminated.)
9. To load extra precision northing
data Press keys in sequence,
starting with "N" or "S".
(Example: 901 m North = N 901) Observe that
northing meters appear in left data displays as keys are
pressed.)
NOTE
The "W" key may be used to
initiate easting entries; however
computer will always interpret
such entries as an "E" input. "E"
will be displayed in normal UTM
display.
NOTE
Extra precision values are always
added to normal values. As an
example, South 4, 476,995 m will
display "4476S" in normal display
and "995" in extra precision
display. There is no rounding
between the two displays.
(b) To insert geographic coordinates of
aircraft present position:
NOTE
Prior
to
pressing
INSERT/ADVANCE
pushbutton,
any incorrectly loaded data can be
corrected by pressing CLEAR
pushbutton and loading correct
data.
1. Data selector L/L POS.
(Observe that, prior to initial load, the INSERT/
ADVANCE pushbutton light is illuminated.)
2. To load latitude data Press keys in
sequence, starting with "N" or "S" to
indicate north or south.
(Example: 42°54.0' North = N 4 2 5 4 0. Observe
that latitude appears in left data display as keys are
pressed.)
3. INSERT/ADVANCE
pushbutton
Press.
(Observe pushbutton light remains illuminated.)
4. To load longitude data Press keys in
sequence, starting with "W" or "E"
to indicate west or east.
NOTE
The value is always added to
normal value regardless of which
key (N/S) is pressed to initiate the
entry.
The
normal
entry
establishes the hemisphere.
10. INSERT/ADVANCE
pushbutton
Press.
3-51
TM 55-1510-219-10
(Example: 87°54.9' West = W 8 7 5 4 9. Observe
that longitude appears in right data display as keys are
pressed.)
5. INSERT/ADVANCE
pushbuttonPress.
(Observe pushbutton light extinguishes.)
(8) To insert magnetic variation. (Only if
compass is not operational)
(9) Data selector DSRTK/STS. Confirm.
1. Left data display indicates desired
track angle in computer memory.
2. Right data display indicates ---84, --74, ---64, or ---54, depending on
which alignment state the computer
has reached.
(10) To
coordinates.
1. Data selector L/L WYPT.
2. Keys
"3"
and
"9"
Press
simultaneously.
(Observe that magnetic variation, to a tenth of an
arc-minute, appears in the right data display.)
program
destinations
or
TACAN
NOTE
If
latitude
and
longitude
coordinates are being used, skip
step I (a) and proceed with step 1
(b). Enter all of the data for a
given destination or TACAN
before starting to enter data for
another.
1. To insert waypoint coordinates:
NOTE
Magnetic variation is retained
from the last NAV mode to be
used for reasonableness tests
during stored heading alignment.
(a) Insertion
of
UTM
waypoint
coordinates:
NOTE
During NAV mode, magnetic
variation is computed when
aircraft roll is less than 9° and
magnetic heading input is valid
(absence of Error Code 17).
1. Data selector UTM WYPT.
Data displays will indicate last coordinates inserted
into related waypoint.
2.
Thumbwheel
Set
to waypoint
number to be loaded.
3. If magnetic variation
is to be
loaded, proceed with next two steps.
NOTE
UTM data may be loaded in any
order and, until final entry, a value
may be reloaded.
3. To load zone and easting Press
keys in sequence, starting with "E".
(Example: Zone 16, 425 km East = E16 425.
Observe that zone and easting in kilometers appear in
the right data display as keys are pressed.)
4. Magnetic
variation
Load
by
pressing
keyboard
keys
in
sequence, starting with "E" or "W"
to indicate east or west.
(Example: 14'53.6' East = E 1 4 5 3 6) Observe that
INSERT/ADVANCE pushbutton light illuminates when
first key is pressed, and magnetic variation appears in
right data display).
5. INSERT/ADVANCE
pushbutton
Press.
(Observe pushbutton light extinguishes.)
4. INSERT/ADVANCE
pushbutton
Press.
(Observe pushbutton light is illuminated.)
5. To load northing Press keys in
sequence, starting with "N" or "S"
to
indicate
north
or
south
hemisphere.
(Example: 4749 km North = N 4749. Observe that
northing kilometers appear in the left data display as
keys are pressed.)
NOTE
If in NAV mode and magnetic
heading input is valid, computer
will reject loaded value.
6. To return INS to normal display
mode,
momentarily
set
data
selector to any position except L/L
WYPT.
3-52
TM 55-1510-219-10
6.
INSERT/ADVANCE
button - Press.
push-
example,
South 4,476,995 m will
display "4476 S" in the normal display
and "995" in extra precision display. In
other words, there is no rounding
between the two displays.
(Observe pushbutton light remains illuminated.)
7.
INSERT/ADVANCE
button - Press.
push-
12.
(Observe that an extra precision display related to
resident value of northing and easting, to the nearest
meter, appears in left and right data displays,
respectively.)
8.
NOTE
A load cycle may be terminated prior
to insertion of all four values by
moving data selector or thumbwheel.
To load extra precision
easting value - Press keys in
sequence starting with "E".
(b)
coordinates:
(Example: 297 m East = E 297. Observe that
easting meters appear in the right data display as keys
are pressed.)
9.
INSERT/ADVANCE
button - Press.
1.
push-
To load extra precision
northing value - Press keys in
sequence, starting with "N" or
"S".
(Example: 901 m North = N 901. Observe that
northing meters appear in left data display as keys are
pressed. The value is always added to the normal value
regardless of which key (N/S) is pressed to initiate the
entry. The normal entry establishes the hemisphere.)
11.
INSERT/ADVANCE
button - Press.
Insertion
of
geographic
waypoint
Data selector - L/L WYPT.
Data displays indicate last coordinates inserted into
the selected waypoint.
(Observe pushbutton light remains illuminated.)
10.
Repeat steps 2 through 11 for
each waypoint to be loaded.
2.
Thumbwheel
Set
waypoint number to
loaded.
to
be
3.
To load latitude - Press keys
in sequence, starting with "N"
or "S" to indicate north or
south.
(Example: 42°54.0' North = N 4 2 5 4 0. Observe
that INSERT/ADVANCE pushbutton light illuminates
when first key is pressed, and latitude appears in left
data display as keys are pressed.)
push-
4.
(Within 3 seconds computer converts input into
latitude and longitude for storage in memory. The
stored value is again converted to UTM for display. The
INSERT/ADVANCE pushbutton light extinguishes.
Conversion routines may cause displays to change by
up to 10 m.)
INSERT/ADVANCE
button - Press.
push-
(Observe pushbutton light extinguishes.)
5.
NOTE
To load longitude - Press
keyboard keys in sequence,
starting with "W" or "E"
indicating west or east.
(Example: 87°54.9' West = W 8 7 5 4 9. Observe
that INSERT/ADVANCE pushbutton light illuminates
when first key is pressed, and longitude appears in
display as keys are pressed.)
The
computer
will
convert
coordinates in overlap area;
however, data display values will
reference appropriate zone.
6.
NOTE
INSERT/ADVANCE
button - Press.
push-
(Observe pushbutton light extinguishes.)
The "W" key may be used to
initiate easting entries; however,
the computer will always interpret
such entries as an "E" input. "E"
will be displayed in normal UTM
data display.
7.
If desire to insert extra
precision coordinate data Press INSERT/ ADVANCE
pushbutton.
(Observe that arc-seconds for loaded latitude and
longitude, to nearest tenth of a second, appear in left
and right data displays, respectively.)
NOTE
The extra precision values are always
added to normal values.
As an
3-53
TM 55-1510-219-10
8.
To load related arc-second
values for latitude -Press keys
in sequence, starting with "N".
(a) Insertion of UTM
NOTE
(Example: 35.8" North = N 358.)
9.
INSERT/ADVANCE
button- Press.
Prior
to
pressing
the
INSERT/ADVANCE pushbutton, any
incorrectly loaded data can be
corrected
by
pressing
CLEAR
pushbutton and loading correct data.
push-
(Observe pushbutton light extinguishes.)
10.
To load related arc-second
values for longitude -Press
keys in sequence, starting
with "E".
(Example: 20.1" East = E 201.)
11.
INSERT/ADVANCE
button - Press.
push-
1.
Data
selector
WYPT.
-
UTM
2.
Keys "7" and "9" - Press
simultaneously.
(Observe that number of TACAN station being used
for navigation flashes on and off in "TO" display and
data displays indicate coordinates of station selected by
thumbwheel.)
(Observe pushbutton light extinguishes.)
12.
TACAN station
data:
Repeat steps 2 through 11 for
each waypoint to be loaded.
3.
Thumbwheel - Set to number
of station to be loaded.
Confirm:
NOTE
In above example, if INSERT/
ADVANCE pushbutton was pressed,
the following normal display would
appear: "42°54.5 (N)" and 87°54.3(W).
The extra precision values are added
to normal values and normal data
displays are not rounded off.
a.
Thumbwheel is in detent.
b.
Station "0"
loaded.
cannot
be
(Observe that if number of station to be loaded is
same as number of the TACAN station currently being
used, number in "TO" display will be set to "0" when
TACAN data is loaded.)
NOTE
The normal geographic coordinates
must always be loaded prior to extra
precision values.
c.
NOTE
To load zone and easting
- Press keys in sequence,
starting with "E".
(Example: Zone 16, 425 km East = E16 425.
Observe that zone and easting in kilometers appear in
the right display as keys are pressed.)
The directions "N" or "S" and "E" or
"W" are established during normal
coordinate entry. Either key may be
used to initiate entry during extra
precision loads and values will be
added to the extra precision value
without affecting direction.
d.
INSERT/ADVANCE
pushbutton - Press.
(Observe that pushbutton light is illuminated.)
To load northing - Press keys in sequence, starting
with "N" or "S" to indicate north or south hemisphere.
NOTE
(Example: 4749 km North = N 4749. Observe that
Northing kilometers appear in left data display as keys
are pressed.)
It is characteristic of the computer
display routine to add "0.2" arcseconds to any display of "59.9" arcseconds. The value in computer is as
loaded by operator.
e.
INSERT/ADVANCE
pushbutton - Press.
(Observe pushbutton light remains illuminated.)
(11) To insert TACAN coordinates:
3-54
TM 55-1510-219-10
f.
INSERT/ADVANCE
pushbutton - Press.
ways interpret such entries as an E
input. "E" will be displayed in normal
UTM data display.
(Observe that extra precision display related to the
resident value of northing and easting, to nearest meter,
appears in left and right data displays, respectively.)
NOTE
The extra precision values are always
added to normal values.
As an
example, South 4,476.995 m will
display "4476 S" in normal display and
"995" in extra precision display. In
other words, there is no rounding
between the two displays.
NOTE
UTM data may be loaded in any order.
Until final fourth entry, actuation of
INSERT/ADVANCE pushbutton without
a prior data entry will cause normal
and extra precision UTM data to be
alternately displayed.
g.
k.
To load extra precision
easting value - Press
keys in sequence, starting
with "E".
(Observe right data display indicates last previously
inserted altitude, and left data display is blank.)
l.
(Example: 297 m East = E 297. Observe that
easting meters appear in right data display as keys are
pressed.)
h.
(Observe
illuminates.)
INSERT/ADVANCE
pushbutton - Press.
m.
To load extra precision
northing value - Press
keys in sequence, starting
with "N" or "S".
pushbutton
light
To load altitude in feet Press keys in sequence.
(Example: 1230 ft = 1230. Observe that numbers
appear in right data display as keys are pressed.)
NOTE
Altitude inputs are limited to 15,000 feet.
(Example: 901 m North = N 901. Observe that
northing meters appear in left data display as keys are
pressed. The value is always added to normal value
regardless of which key (N/S) is pressed to initiate entry.
The normal entry establishes hemisphere.)
j.
To indicate the following
load is altitude - Press
keys "4" or "6".
INSERT/ADVANCE
(Observe pushbutton remains illuminated.)
i.
INSERT/ADVANCE
pushbutton - Press.
n.
INSERT/ADVANCE
pushbutton - Press.
(Observe pushbutton light extinguishes.)
o.
INSERT/ADVANCE
pushbutton - Press.
INSERT/ADVANCE
pushbutton - Press.
(Observe that left data display indicates last
previously inserted channel number, and right display is
blank.)
(Observe that during the next 1 to 3 seconds, the
computer converts input into latitude and longitude for
storage in memory. The stored value is again converted
back to UTM for display to operator.
The
INSERT/ADVANCE pushbutton light extinguishes. The
conversion routines may cause data displays to change
by up to 10 m.)
p.
(Observe
illuminates.)
NOTE
INSERT/ADVANCE
q.
The computer will convert coordinates
in overlap area; however, data display
values will reference appropriate zone.
To indicate following load
is channel number -Press
keys "2" or "8"
pushbutton
light
To load channel number Press keys in sequence.
(Example: 109 = 109. Observe number appears in
left
data
display
as
keys
are
pressed.)
NOTE
The "W" key may be used to initiate
easting
entries;
however,
the
computer will al-
3-55
TM 55-1510-219-10
r.
displays indicate coordinates of station selected via
thumbwheel.)
INSERT/ADVANCE
pushbutton - press.
3.
(Observe pushbutton light extinguishes.)
NOTE
Any number will be accepted by INS;
however, only stations with a channel
number within range of "1" through
"126" will be used for TACAN mixing.
NOTE
Station "0" cannot be loaded.
4.
NOTE
Channel number has an implied "X" suffix.
NOTE
(Observe
reappear.)
station
INSERT/ADVANCE
pushbutton - Press.
northing,
zone,
and
Repeat steps 3 through
20 for each TACAN
station.
u.
To return INS to normal
mode, momentarily set
data selector to UTM
POS.
(b) Insertion
of
geographic
5.
easting
t.
6.
push-
To load longitude - Press
keys in sequence, starting
with "W" or "E" indicating
west or east.
(Example: 87°54.9' West = W 8 7 5 4 9. Observe
that INSERT/ADVANCE pushbutton light illuminates
when first key is pressed, and longitude appears in data
display as keys are pressed.)
TACAN
7.
NOTE
INSERT/ADVANCE
button - Press.
push-
(Observe pushbutton light extinguishes.).
Prior to pressing INSERT/ADVANCE
pushbutton, any incorrectly loaded
data can be corrected by pressing
CLEAR pushbutton and loading
correct data.
8.
INSERT/ADVANCE
button - Press.
push-
(Observe that the arc-seconds related to loaded
latitude and longitude, to nearest tenth of a second,
appear in left and right data display, respectively.)
Data selector - L/L WYPT.
9.
NOTE
If number of station to be loaded is
same as number of TACAN station
currently being used, number in "TO"
display will be set to "O" when TACAN
data is loaded.
2.
INSERT/ADVANCE
button - Press
(Observe pushbutton light extinguishes.)
station data:
1.
To load latitude - Press keys
in sequence, starting with "N"
or "S" to indicate north or
south.
(Example: 42°54.0' North = N 4 2 5 4 0. Observe
that INSERT/ADVANCE pushbutton light illuminates
when first key is pressed.)
Degree symbol (°) should be disregarded when
reading altitude and data display.
s.
Thumbwheel - Set to number
of station being loaded.
(Insure thumbwheel is in
detent.)
If extra precision coordinate
data is to be inserted Press
keys in sequence, starting
with "N", to load related arcsecond values for latitude.
(Example: 35.8" North = N 358.)
10.
Keys "7" and "9" - Press
simultaneously.
INSERT/ADVANCE
button - Press.
push-
(Observe pushbutton light extinguished.)
(Observe that number of TACAN station being used
for navigation flashes on and off in "TO" display. Data
11.
3-56
To load related arc-second
values for longitude -Press
TM 55-1510-219-10
keys in sequence, starting with "E".
NOTE
(Example: 20.1" East = E 201.)
Altitude inputs are limited to 15,000 feet.
12. INSERT/ADVANCE push-button Press.
16. INSER/ADVANCE
Press.
(Observe pushbutton light extinguishes.)
17. INSERT/ADVANCE push-button Press.
In
above
example,
if
INSERT/ADVANCE
pushbutton
were pressed, the following
normal display would appear:
"42°54.5 N" and "87°54.3 W". The
extra precision values are added
to normal values and normal
displays are not rounded off.
(Observe that left data display indicates last
previously inserted channel number, and right data
display is blank.)
18. To indicate the following load is
channel number -Press key "2" or
"8".
(Observe
illuminates)
NOTE
The
normal
geographic
coordinates must always be
loaded prior to extra precision
values.
pushbutton
light
(Example: 109 = 109. Numbers appear in left data
display as keys are pressed.)
The directions "N" or "S" and "E"
or "W" are established during
normal coordinate entry. Either
key may be used to initiate entry
during extra precision loads and
the values will be added to extra
precision value without affecting
direction.
20. INSERT/ADVANCE push-button Press.
(Observe pushbutton light extinguishes.)
NOTE
Any number will be accepted by
the INS; however only stations
with a channel number within
range of 1 through 126 will be
used for TACAN mixing.
NOTE
It is characteristic of the computer
display routine to add 0.2 arcseconds to any display of 59.9
arc-seconds.
The value in
computer is as loaded by
operator.
NOTE
The channel number has an implied
"X" suffix.
NOTE
13. INSERT/ADVANCE push-button Press.
Decimal
points
and
degree
symbols should be disregarded
when
reading
altitude
and
channel number displays.
(Observe that right data display indicates last
previously inserted altitude, and left data display is
blank.)
21. INSERT/ADVANCE push-button Press.
14. To indicate the following load is
altitude - Press key "4" or "6".
pushbutton
INSERT/ADVANCE
19. To load channel number - Press
keys in sequence.
NOTE
INSERT/ADVANCE
-
(Observe pushbutton light extinguishes.
NOTE
(Observe
illuminates.)
push-button
(Observe station latitude and longitude reappear.)
light
22. Repeat steps 3 through 19 for each
TACAN station.
15. To load altitude - Press keys in
sequence.
23. To return INS to normal display
modes, momentarily set data
selector to L/L POS.
(Example: 1230 ft = 1230. Numbers appear in right
data display as keys are pressed.
3-57
TM 55-1510-219-10
must be between 5 and 150 nmi
and channel between 1 and 126.
(12) Designating fly-to-destinations or TACAN.
1.
Data selector L/L WYPT or
UTM WYPT, as required.
2.
Waypoint
thumbwheel
Select destination number.
(Observe number of destination waypoint appears in
"TO" part of FROM-TO display.)
3.
(Observe present aircraft heading appears, to
nearest tenth of degree, in left data display; also drift
angle, to nearest degree, appears in right data display.)
8.
3.
Horizontal Situation Indicators
(pilot's and/or copilot's HSI)
Steer toward indicators.
4.
CDU ALERT lamp Monitor.
(14) Aided TACAN operation.
Mode selector NAV.
2.
Data selector DSRTK/STS.
3.
Key "4" Press.
UPDATE"
9.
To return INS display to
normal. Set data selector to
any position except WYPT or
DIS/TIME.
10.
To monitor progress of update
Set data selector to DSRTK/
STS.
(Observe Accuracy
Index (AI) will decrement to
"0".)
NOTE
To insure favorable geometry
during the update process, the
following TACAN station criteria
should be observed:
(Observe right data display is "000004" and
INSERT/ADVANCE pushbutton light is illuminated.)
4.
Monitor "TACAN
annunciator.
Mixing will not be annunciated if:
(a)
TACAN
control
is
inappropriately set; (b) TACAN
station data loaded in error; (c)
aircraft look-down angle is greater
than 30°; (d) horizontal ground
distance is less than two times the
altitude.
(Observe illumination approx 1.3 minutes before
reaching point for automatic leg switch.
Indicator
flashes on and off after passing a waypoint, if
AUTO/MAN switch is in MAN.)
1.
To monitor station selection
Observe
FROM-TO
data
display.
NOTE
(13) To fly selected INS course.
Pilot's RMI select switch INS.
Keys "7" and "9" - Press
simultaneously.
(Observe only the number of stations eligible for
mixing will be displayed. A "0" indicates that none of the
9 stations are eligible for selection.)
Navigation information is now
available from INS for display on
pilot's RMI and on pilot's and
copilot's HSI's as determined by
COURSE INDICATOR switches.
2.
6.
7.
NOTE
Pilot's COURSE INDICATOR
switch INS.
Data selector L/L WYPT or
UTM WYPT.
(Observe channel number of the TACAN station being
used for navigation flashes on and off. Data displays
indicate coordinates of station selected via thumbwheel.)
Data selector HDG DA.
1.
5.
INSERT/ADVANCE
pushbutton Press.
* One station must be at least 15 nmi off course.
(Observe right data display is "1-XX4" and
INSERT/ADVANCE pushbutton light is extinguished.)
* For optimum single TACAN station updating,
update should continue until aircraft has passed the
station.
NOTE
* For optimum dual TACAN updating, use one "offtrack" TACAN station and one "on-track" station.
Every 15 seconds, the INS will
select next eligible TACAN station
in sequence for 3-58 updating. To
be eligible, TACAN station range
* For optimum multi-TACAN station updating, the
stations should be evenly distributed in azimuth around
the aircraft.
3-58
TM 55-1510-219-10
11.
Waypoint thumbwheel Set to
number of first TACAN station
to be used.
(15) Switching from aided to unaided inertial.
operation.
1.
Data selector DSRTK/STS.
2.
Key "5" Press.
(Observe
INSERT/ADVANCE
pushbutton
illuminates; 000005 appears in right data display.)
3
(f)
Ground track angle: Data selector
TK/GS.
(Observe ground track angle appears in left data
display to nearest tenth of a degree.)
(g) Drift angle: Data selector HDG DA.
(Observe drift angle appears in right data display to
nearest degree.)
light
(h) Wind speed and direction: Data
selector WIND.
INSERT/ADVANCE pushbutton
Press.
(Observe
INSERT/ADVANCE
pushbutton
light
extinguishes. Data display returns to normal with "5"
appearing in first digit of right display.)
NOTE
(Wind direction appears in left data display to nearest
degree and wind speed appears in right display to
nearest knot.)
(i) Desired track angle: Data selector
DSRTK/STS.
(Observe desired track angle in left data display to
nearest degree.)
Benefits of previous aiding are
maintained but no additional
automatic updates will be made.
(j) Track angle error: Data selector
XTK/TKE.
(16) To obtain readouts from INS.
(Observe track angle error appears in right data
display to nearest degree.)
NOTE
The computer is assumed to be in
the NAV mode for all data displays.
(k) Cross track distance: Data selector
XTK/TKE.
selector
DSRTK/STS.
(Observe cross track distance appears in left data
display to nearest nautical mile.)
(Observe numbers indicating system status appear in
right data display.)
(I) Distance and time to next waypoint:
Data selector DIS/TIME.
(b) Geographic present position: Data
selector L/L POS.
(Observe distance to next waypoint, shown in "TO"
side of FROM-TO display, appears in left data display to
nearest nautical mile.)
(a) System
status:
Data
(Observe latitude and longitude of present position
appear in left and right data displays, respectively. Both
displays are to tenth of a minute.)
(Observe time to reach next waypoint at present
ground speed appears in right data display to nearest
tenth of a minute.)
(c) UTM position: Data selector UTM
POS.
(Observe northing and zone with easting of present
position appear in left and right displays, respectively.
Both displays are in kilometers.)
(d) True heading: Data selector HDG DA.
(Observe aircraft heading appears in left data display
to nearest tenth of a degree.)
(e) Ground
speed:
Data
selector
TK/GS.
(Observe ground speed appears in right data display to
nearest knot.)
3-59
(m) Extra precision geographic present
position display:
1.
Data selector L/L POS.
(Latitude and longitude of present position, to nearest
tenth of a minute, appears in left and right data displays,
respectively.)
2.
INSERT/ADVANCE
pushbutton Press.
(Observe arc-seconds related to present position
latitude and longitude, to nearest tenth of a second,
appear in left and right data displays, respectively.)
TM 55-1510-219-10
(Observe
coordinates.)
(n) Geographic present inertial position
display.
1.
Data selector L/L WYPT.
2.
HOLD pushbutton Press.
coordinates
5.
HOLD pushbutton Press.
(p) Distance and time to waypoint other
than next waypoint:
1.
NOTE
WYPT
Press.
CHG
pushbutton
(Observe WYPT CHG and INSERT/ ADVANCE
pushbutton lights illuminate.)
While HOLD pushbutton light is
illuminated, TACAN and data link
updates are inhibited.
2.
INSERT/ADVANCE
pushbutton Press.
Key "0" Press.
(Observe FROM side of FROM-TO data display
changes to "0".)
(Observe arc-second related to present inertial position
latitude and longitude, to nearest tenth of a second,
appears in left and right data displays, respectively.)
4.
latitude-longitude
(Observe INS returns to normal operation and HOLD
pushbutton light extinguishes.)
(Observe HOLD pushbutton light illuminates, latitude
and longitude or present inertial position to a tenth of
degree appear in left and right data displays,
respectively.)
3.
in
3.
Key corresponding to desired
waypoint Press.
(Observe "TO" side of FROM-TO data display
changes to desired waypoint number.)
Data selector UTM WYPT.
(Observe coordinates in UTM grid.)
NOTE
5.
Do not press INSERT/ADVANCE
pushbutton. This would cause an
immediate flight plan change.
HOLD pushbutton Press.
(Observe INS returns to normal operation and HOLD
pushbutton light extinguishes.)
(o) UTM
present
4.
inertial position
display:
1.
Data selector UTM WYPT.
2.
HOLD pushbutton Press.
(Observe distance to desired waypoint appears in left
data display to nearest nautical mile. Time to reach
desired waypoint at present ground speed appears in
right data display to nearest tenth of a minute.)
(Observe HOLD pushbutton light illuminates.
Northing and zone with easting of the present inertial
position in kilometers appear in left and right data
displays, respectively.)
5.
(Observe INSERT/ADVANCE and WYPT CHG
pushbutton lights extinguish.
Waypoints defining
current navigation leg appear in FROM-TO display.)
While HOLD pushbutton light is
illuminated, TACAN and data link
updates are inhibited.
(q) Distance and time between any two
waypoints:
1.
INSERT/ADVANCE
pushbutton Press.
WYPT
Press.
CHG
pushbutton
(Observe WYPT CHG and INSERT/ADVANCE
pushbutton lights illuminate.)
(Observe extra precision values related to present
inertial position northing and easting, to nearest meter,
appear in left and right data displays, respectively.)
4.
CLEAR pushbutton Press.
(Observe INS returns to normal operation.)
NOTE
3.
Data selector DIS/TIME.
2.
Data selector L/L WYPT.
Keys corresponding to desired
waypoints Press in sequence.
(Observe desired waypoint numbers appear in
FROM-TO data display as keys are pressed.)
3-60
TM 55-1510-219-10
NOTE
NOTE
Do not press INSERT/ADVANCE
pushbutton. This would cause an
immediate flight plan change.
If wrong key is pressed, press
CLEAR; displays will revert to that
indicated in step (2).
3.
7.
Data selector DIS/TIME.
(Observe distance between desired waypoints
appears in left data display to nearest nautical mile.
Time to travel between desired waypoints at present
ground speed appears in right data display to nearest
tenth of a minute.)
4.
(Observe INSERT/ADVANCE and WYPT CHG
pushbutton lights extinguish. The loaded digit will
appear in right position of FROM-TO display and will be
flashing on and off. Distance to that station to nearest
nautical mile appears in left data display. The right
display continues to display time to next waypoint.)
CLEAR pushbutton Press.
(Observe INS returns to normal operation.)
8.
(Observe WYPT CHG and INSERT/ ADVANCE
pushbutton light extinguishes.
Waypoints defining
current navigation leg appear in FROMTO data display.)
If in aided TACAN operation and if the
desired station is not being selected,
exit aided operation per procedure:
"Switching From Aided to Unaided
Inertial Operation", perform steps (1)
(8), and then return to aided operation
per
procedure:
"Aided
TACAN
Operation".
Data selector DIS/TIME.
Keys "7" and
simultaneously.
"9"
Press
(s) Coordinates of any waypoint:
(Observe number of TACAN station being used for
navigation flashes on and off in FROM-TO display.
Distance to TACAN station to nearest nautical mile is in
left data display. Time to next waypoint is in right data
display.)
3.
5.
Data selector L/L WYPT or
UTM WYPT.
2.
Waypoint thumbwheel Set
desired waypoint. Observe
following:
(L/L WYPT: Latitude and longitude of desired
waypoint, to a tenth of a minute, appear in left and right
data displays, respectively.)
If not in aided TACAN
operation Perform steps (5)
through (7).
WYPT
Press.
CHG
(UTM WYPT: Northing and zone with easting of
desired waypoint, to a kilometer, appear in left and right
data displays, respectively.)
pushbutton
3.
(Observe INSERT/ADVANCE and WYPT CHG
pushbutton light illuminates. Station number flashing
discontinues.)
6.
1.
If in aided TACAN operation
Monitor display.
(Observe station number is selected every 15 seconds.)
4.
WIND,
NOTE
(Observe distance to next waypoint to nearest
nautical mile is in left data display. Time to next
waypoint to nearest tenth of a minute is in right data
display.)
2.
Data
selector
momentarily.
(Returns INS to normal display mode.)
(r) Distance to any TACAN station:
1.
INSERT/ADVANCE
pushbutton Press.
INSERT/ADVANCE
pushbutton Press. Observe
following:
(L/L WYPT: The arc-seconds related to desired
waypoint, latitude, and longitude, to a tenth of an arcsecond, appear in left and right data displays,
respectively.)
Key
indicating
desired
TACAN station number Press.
(UTM WYPT: The extra precision display related to
desired waypoint northing and easting, in meters,
appear in left and right data displays, respectively.)
(Observe number will appear in left digit location of
FROM-TO data display.)
3-61
TM 55-1510-219-10
UTM WYPT: A coordinate is the addition of values
for kilometers and meters.
NOTE
L/L WYPT: A coordinate is the
addition of values for degrees,
whole minutes and seconds.
EXAMPLE: S 2,474,706 m will be displayed as "2474
S" and "706".
5.
EXAMPLE: W87°54' 58.6" = 870 54.9W and 58.6.
UTM WYPT: A coordinate is the addition of the
values for kilometers and meters.
(Observe TACAN station altitude, in feet, will appear
in right data display; degree symbol and decimal points
should be disregarded. Left data display is blank.)
EXAMPLE: S 2,474,706m = 2474 S and 706.
(t) TACAN station data:
1.
6.
Data selector L/L WYPT or
UTM WYPT.
2.
Keys "7" and
simultaneously.
"9"
3.
Waypoint thumbwheel
desired TACAN station.
INSERT/ADVANCE
pushbutton Press.
(Observe TACAN station channel number, in whole
numbers, will appear in left data display; degree symbol
and decimal point should be disregarded. Right data
display is blank.)
Press
Set
NOTE
If INSERT/ADVANCE pushbutton
is
pressed,
the
normal
coordinates indicated in step (3)
will be displayed.
(Observe number of TACAN station being used for
navigation flashes on and off.)
(L/L WYPT: Latitude and longitude of desired
TACAN station, to tenth of minute, appears in left and
right data displays, respectively.)
NOTE
Waypoint thumbwheel may be
moved at any time and normal
coordinates for new TACAN
station will be displayed.
(UTM WYPT: Northing and zone with easting of
desired TACAN station, to a kilometer, appear in left
and right data displays, respectively.)
4.
INSERT/ADVANCE
pushbutton Press.
INSERT/ADVANCE
pushbutton Press. Observe
following:
7.
(L/L WYPT: The arc-seconds related to desired
TACAN station, to tenth of arc-second, appear in left
and right data displays, respectively.)
Data selector Momentarily to
any position other than L/L
WYPT, UTM WYPT or
DIS/TIME. (Returns INS to
normal operation.)
(u) Magnetic heading.
(UTM WYPT: The extra precision display related to
desired TACAN station northing and easting, in meters,
appear in left and right data displays, respectively.)
1.
Data selector HDG/DA.
(Observe true heading to nearest tenth of a degree
appears in left data display. Drift angle to nearest
degree appears in right data display.)
NOTE
Direction is indicated in normal
data displays.
2.
NOTE
Keys "3" and "9" Press
simultaneously and hold.
(Observe magnetic heading to nearest tenth of a
degree appears in left data display.
Drift angle
continues to be displayed in right data display.)
L/L WYPT: A coordinate is the
addition of values for degrees,
whole minutes, and seconds.
3.
EXAMPLE: W 87° 54' 58.6" will be displayed as
"87°54.9 W" and "58.6".
Keys "3" and "9" Release.
(Observe left data display reverts to true heading.)
3-62
TM 55-1510-219-10
4.
(17) INS updating.
(Observe INSERT/ADVANCE pushbutton light
remains illuminated, and previous value of latitude
reappears.)
(a) Data link updating.
1.
Mode selector NAV.
2.
Data selector DSRTK/STS.
3.
Key "3" Press.
5.
(Observe right data display is "000003" and
INSERT/ADVANCE pushbutton lamp is illuminated.)
4.
INSERT/ADVANCE
pushbutton Press.
(Observe INSERT/ADVANCE and HOLD pushbutton
lights remain illuminated. North position error and east
position error, in tenth of a nautical mile, will appear in
left and right data displays, respectively.)
Monitor "DAT LINK UPDATE"
annunciator
on
mission
control panel.
NOTE
If
WARN
lamp
illuminates,
proceed to step 8; otherwise
proceed to step 9.
(Observe illumination will be within 30 seconds, if
correction of less than 3 + AI nmi is being transmitted to
INS.)
6.
The Accuracy Index (AI) will
decrement to "O" as updating
progresses.
7.
Data selector LIL POS.
(Observe latitude and longitude of present position
appear in left and right data displays, respectively.)
2.
HOLD pushbutton Press until
illuminated.
8.
If displayed values are within
tolerance,
press
HOLD
pushbutton to return INS to
normal operation. If one or
both values are out of
tolerance, proceed the step
(10).
9.
Key "2" Press.
(Observe latitude and longitude in data displays
freeze at values present when HOLD pushbutton is
pressed.)
NOTE
While HOLD pushbutton light is
illuminated, TACAN and data link
updates are inhibited.
3.
Data selector DSRTK/STS.
(Observe action code "02" and malfunction code
"49". This indicates that the radial error between the
loaded position and the INS position exceeds 38
nautical miles. Operator must evaluate possibility that
either INS is in error or reference point position is
incorrect. It is possible to force INS to accept updated
position by setting data selector to L/L POS and
proceeding to step (h).
(b) Normal geographic present position
check and update:
1.
Keys Press in sequence to
load longitude of position
reference, starting with "W" or
"E" to indicate west or east.
(Example: 87°54.9' west = W 8 7 5 4 9. Observe
longitude appears in right data display as keys are
pressed.)
(Observe right data display is "1-XX3" and
INSERT/ADVANCE pushbutton light is extinguished.)
5.
INSERT/ADVANCE pushbutton Press.
(Observe left data display is "00000 N";
INSERT/ADVANCE and HOLD pushbutton lights are
illuminated.)
Keys Press in sequence to
load latitude of position
reference, starting with "N" or
"S" to indicate north or south.
10.
INSERT/ADVANCE
pushbutton Press
(Observe INSERT/ADVANCE and HOLD pushbutton
lights extinguish. Present position appears in data
displays.
Present position check and update is
complete.)
(Example: 42°54.0' north = N 4 2 5 4 0.)
(Observe INSERT/ADVANCE pushbutton light
illuminates when first key is pressed, and latitude
appears in left data display as keys are pressed.)
3-63
TM 55-1510-219-10
(Example: 87°54.9' West = W 8 7 5 4 9.)
NOTE
9.
Within 30 seconds, computer will
process correction and revised
present position will appear in
data display.
If AI prior to
position update is 1 or greater,
computer will accept over 95
percent of correction shown in
difference display. If AI is "0",
amount of correction accepted
will be less and is a function of
time in NAV mode and number of
updates which have been made.
10.
11.
Data selector DSRTK/STS.
2.
Key "2" Press.
(Observe INSERT/ADVANCE pushbutton light
illuminates, "000002" appears in right data display.)
12.
13.
INSERT/ADVANCE
pushbutton Press.
NOTE
Normal
latitude-longitude
coordinates must always be loaded
prior to extra precision values.
NOTE
Directions "N" or "S" and "E" or "W"
are established during normal
coordinate entry. Either key may be
used to initiate entry during extra
precision loads and values will be
added to extra precision values
without affecting direction.
Load latitude by pressing keys
in sequence, starting with "N"
or "S" to indicate north or
south.
(Example: 42°54.0' North = N 4 2 5 4 0. Observe
latitude appears in left data display as keys are
pressed.)
8.
Load related arc-second values
for longitude in sequence,
starting with "E".
Extra precision values are added to
normal values and normal displays
are not rounded off.
HOLD
pushbutton
Press
(when aircraft passes over
known position reference.)
NOTE
INSERT/ADVANCE
pushbutton Press
and
pushbutton
NOTE
Data selector L/L POS.
(Observe
INSERT/ADVANCE
pushbuttons remain illuminated.)
INSERT/ADVANCE
Press.
(Example: 20.1°East = E 201.)
(Observe HOLD pushbutton light illuminates.
Latitude and longitude in data displays freeze at values
present when HOLD pushbutton was pressed.)
7.
Load related arc-second values
for latitude in sequence, starting
with "N".
(Observe INSERT/ADVANCE and HOLD pushbuttons
remain illuminated.)
(Observe latitude and longitude of present position
appear in left and right data displays, respectively.)
6.
pushbutton
(Example: 35.8°North = N 358.)
(Observe
right
data
display
is
"1...XX2",
INSERT/ADVANCE pushbutton light is extinguished,
and any TACAN or data link updating is discontinued.)
5.
INSERT/ADVANCE
Press.
(Observe arc-seconds related to present position
latitude and longitude, to nearest tenth of a second, appear
in left and right data displays, respectively.)
1.
4.
pushbutton
(Observe INSERT/ADVANCE and HOLD pushbutton
lights remain illuminated.)
(c) Extra precision geographic present
position check and update:
3.
INSERT/ADVANCE
Press.
It is characteristic of the computer
display routine to add 0.2 arcseconds to any display of 59.9 arcseconds.
Value in computer is
loaded by operator.
HOLD
Load longitude by pressing
keys in sequence, starting
with "W" or "E" indicating
west or east.
14.
3-64
Proceed to step (6) in procedure:
"Normal Geographic
TM 55-1510-219-10
(Observe extra precision display related to present
position northing and easting, to nearest meter, appears
in left and right data displays, respectively.)
Present Position Check and
Update."
(d) UTM present position check and
8.
update:
NOTE
UTM data may be loaded in any
order and, until final entry, a value
may be reloaded.
1.
(Example: 297 m East = E 297. Observe easting
meters appear in right data display as keys are
pressed.)
Data selector UTM POS.
9.
(Observe UTM coordinates of present position
appear in data displays.)
2.
HOLD
pushbutton
Press
(when aircraft passes over
known position reference.)
10.
While HOLD pushbutton light is
illuminated, TACAN and data link
updates are inhibited.
5.
NOTE
The extra precision values are
always added to normal values.
INSERT/ADVANCE
pushbutton Press.
NOTE
light
Any data inserted when HOLD
pushbutton
lamp
is
not
illuminated will be rejected by
computer.
Load northing by pressing
keys in sequence, starting
with "N" or "S" to indicate
north or south hemisphere.
11.
(Example: North 4749 km = N 4749. Observe
northing kilometers appear in left data display as keys
are pressed.)
6.
7.
pushbutton
INSERT/ADVANCE
pushbutton Press.
(Observe INSERT/ADVANCE and HOLD pushbutton
lights remain illuminated. North position error and east
position error in kilometers will appear in left and right
data displays, respectively.)
INSERT/ADVANCE
pushbutton Press.
(Observe INSERT/ADVANCE
remains illuminated.)
Load extra precision northing
value by pressing keys in
sequence, starting with "N" or
"S".
The "W" key may be used to
initiate easting entries; however,
the computer will always interpret
such entries as an "E" input.
(Example: Zone 16, 425 km East = E16 425.
Observe zone and easting in kilometers appear in right
data display as keys are pressed.)
pushbutton
light
NOTE
Load zone and easting by
pressing keys in sequence,
starting with "E".
(Observe INSERT/ADVANCE
remains illuminated.)
pushbutton
push-
(Example: 901 m North = N 901. Observe
Northing meters appear in left data display as keys are
pressed. The value is always added to normal value
regardless of which key (N/S) is pressed to initiate entry.
Normal entry establishes the hemisphere.)
NOTE
4.
INSERT/ADVANCE
button Press.
(Observe INSERT/ADVANCE
remains illuminated.)
(Observe HOLD pushbutton light illuminates.
Coordinates in data display freeze at values present
when HOLD pushbutton was pressed.)
3.
Load extra precision easting
value by pressing keys in
sequence, starting with "E".
light
12.
INSERT/ADVANCE
pushbutton Press.
3-65
If WARN lamp illuminates,
proceed
to
step
(13);
otherwise proceed to step (9)
in procedure: "Extra Precision
Geographic Present Position
Check
and
Update."
TM 55-1510-219-10
13.
Data selector DSRTK/STS.
(Observe action code "02" and malfunction code
"49". This indicates radial error between loaded position
and INS position exceeds 62 kilometers. Operator must
evaluate possibility that INS is in error or reference point
position is incorrect. It is possible to force INS to accept
updated position by setting data selector to UTM POS
and proceeding to step (8) of procedure: "Extra
Precision Geographic Present Position Check and
Update.")
14.
NOTE
Selecting zero as FROM waypoint
will cause desired track to be
defined by computed present
position (inertial present position
plus fixes) and TO waypoint.
3.
If updating is to be rejected
Press HOLD pushbutton.
(Observe WYPT CHG and INSERT/ ADVANCE
pushbuttons extinguish.)
(Observe HOLD and INSERT/ADVANCE pushbutton
lights extinguish. INS returns to normal operation.) (e)
Position update eradication:
NOTE
Waypoint zero always contains
ramp coordinates if no manual
flight plan changes are made.
When a manual flight plan change
is made, present position at
instant of insertion is stored in
waypoint zero.
NOTE
This procedure is not considered
common. Its use is limited to
those times where an operational
error has resulted in an erroneous
position fix.
(19) After landing procedures.
CAUTION
1.
Data selector DSRTK/STS.
2.
Key "1" Press.
(Observe INSERT/ADVANCE pushbutton
illuminates, 000001 appears in right data display.)
3.
If INS will be unattended for an
extended period, it should be shut
down.
light
CAUTION
INSERT/ADVANCE
pushbutton Press.
Do not leave INS operating unless
aircraft or ground power and
cooling air are available to
system.
(Observe INSERT/ADVANCE pushbutton light
extinguishes. Within 30 seconds, data displays return to
normal with "0" (normal inertial mode) in last digit of
right display. AI will be set to approximately three times
the number of hours in NAV.)
NOTE
The INS may be shut down,
downmoded to STBY or Align
mode,
or
operated
in
the
navigation mode after landing.
The
determining
factor
in
choosing course of action is
expected length of time before the
next takeoff.
(18) Flight course changes.
(a) Manual flight plan change insertion:
1.
WYPT
Press.
CHG
pushbutton
(Observe WYPT CHG and INSERT/ ADVANCE
pushbuttons illuminate.)
2.
INSERT/ADVANCE
pushbutton Press.
NOTE
Select new FROM and TO
waypoints
by
pressing
corresponding
keys
in
sequence.
Do not tow or taxi aircraft during
INS alignment. Movement during
alignment
requires
restarting
alignment
(Observe new waypoint numbers appear in FROMTO data displays as keys are pressed.)
.
3-66
TM 55-1510-219-10
Perform position update using best estimate of
parking coordinates.
(20) Transient stops.
NOTE
(d) Downmoding to standby:
Action to be taken during a
transient stop depends upon time
available and on availability of
accurate
parking
coordinates
(latitude and longitude.)
NOTE
INS can be downmoded to
standby operation which will
maintain
navigation
unit
at
operating temperature with gyro
wheels
running.
INS
is
downmoded
to
standby
as
follows:
(a) Realignment
INS
operating
(recommended if sufficient time and accurate parking
coordinates are available.)
NOTE
1.
INS can be downmoded to perform a
realignment
and
azimuth
gyro
calibration. Alignment to produce an
alignment state number of "5" can be
accomplished in approximately 17
minutes. During the 17 minute period,
an
automatic
azimuth
gyro
recalibration is determined on basis of
difference between inertial present
position before downmoding and
inserted present position. To obtain
further refinement of azimuth gyro drift
rate, calculated on basis of newly
computed error data, INS can be left in
alignment mode for a longer period,
allowing the alignment state number
to attain some value lower than "5".
1.
Data selector STBY, then to
ALIGN.
2.
Present position coordinates
Insert, according to procedure:
"Geographic Present Position
Insertion"
or "UTM Present
Position Insertion."
Mode selector STBY.
Caution
Do not leave INS operating unless
aircraft or ground power and
cooling air are available to
system.
(e) Shutdown:
1.
Mode selector OFF.
NOTE
INS will retain inertial present
position data computed at time INS
is downmoded.
This value is
compared with present position
inserted for next alignment and
difference is used to determine
azimuth gyro drift rate.
d. Abnormal Procedures.
(1) General.
INS contains self-testing
features which provide one or more warning indications
when a failure occurs. The WARN lamp on the CDU
provides a master warning for most malfunctions
occurring in the navigation unit. Malfunctions in the
MSU or CDU will normally be obvious because of
abnormal indications of displays and lamps. A battery
unit malfunction will shut down INS when battery power
is used.
(b) Realignment INS shutdown. Perform
complete alignment procedures.
(c) Position update (recommended if
time is not available for realignment.)
(2) Automatic INS shutdown.
NOTE
(a) Overtemperature. An
overtemperature in navigation unit will cause INS to shut down
(indicated by blank CDU displays) when mode selector
is at STBY or ALIGN during ground operation. The
WARN lamp on CDU will illuminate and will not
extinguish until mode selector is rotated to OFF. The
cooling system should be checked and corrected if
faulty. If cooling system is satisfactory, navigation unit
should be replaced.
Perform position update using parking
coordinates
in
accordance
with
procedure: "Insertion of Geographic
Waypoint Coordinates." If parking
coordinates are not available, proceed
as follows:
Continue operation in NAV, if INS accuracy appears
acceptable.
3-67
TM 55-1510-219-10
Table 3-2. Malfunction Code Check
Table 3-3. Malfunction Indications and Procedures
3-68
TM 55-1510-219-10
Table 3-3. Malfunction Indications and Procedures (Continued)
(b) Low battery charge. A low battery
unit charge will cause INS to shut down when INS is
operating on battery unit power. Both WARN lamp on
CDU and BAT lamp on MSU will illuminate and not
extinguish until the mode selector is set to OFF. The
battery unit should be replaced when this failure occurs.
NOTE
It is not possible to load displays from
the keyboard. A temporary failure of a
numerical key may prevent data
loading. If a number cannot be loaded
into latitude or longitude displays, after
pressing/ wiggling the key several
times, the cause may be the momentary
hang-up of another key. To identify the
faulty key, rotate the data selector to
DSRTK/STS. The right digit on right
display will indicate suspect key. Press
and release suspect key several times.
To test whether the keyboard problem
is corrected, try pressing any other
numerical key. Its number should now
appear as the right digit. If this test is
successful, press the CLEAR key and
return data selector to original data
loading position. Otherwise, a CDU
failure is indicated.
(c) Interpretation of failure indications.
It is important to be able to correctly interpret failure
indications in order to take effective action. Failure
indications are listed below under two main categories:
WARN light illuminated, and WARN light extinguished.
Under each of these categories are listed other
indications which will give the operator sufficient
information to take action.
1
WARN light on/off
DISPLAY
MALFUNCTION/ RECOMMENDED
ACTION
Action codes 01,
02, 03, 04, 05
See Table 3-4.
No action or
malfunction
codes displayed
Indicates computer failure.
Improper
displays
Indicates NU computer failure.
(d) CDU BAT indicator is illuminated.
2
CAUTION
Operation on battery is an indication
that there may be no aircraft power to
blower motor with resultant loss of
cooling. The INS can operate only a
limited time (normally 15 minutes) on
battery power before a low voltage
shutdown will occur. Then, immediate
corrective action must be taken.
WARN light off
DISPLAY
MALFUNCTION/ RECOMMENDED
ACTION
CDU displays
blank, incorrect
or frozen
CDU failure indicated.
3-69
TM 55-1510-219-10
NOTE
Table 3-4. Action Codes and Recommended Action
CDU BAT indicator will illuminate
for 12 seconds in alignment State
"8" (about 5 minutes after turnon). This is normal and indicates
a battery test is in progress. No
corrective action is required.
NOTE
Ground operation on battery
power should not exceed 5
minutes.
(e) To determine corrective action:
(Monitor CDU displays while rotating the CDU selector
switch.)
1.
If displays are frozen (do not
change while data selector is
being rotated) problem is
normally in the navigation
unit.
2.
NOTE
CDU
BAT
indicator
should
extinguish after above corrective
action. If it remains illuminated,
INS will eventually shut down
when battery voltage drops below
approximately 19VDC. Flight crew
should prepare for shutdown.
If displays respond normally
to the data selector, the
problem is normally absence
of 115V AC power to INS.
(f) For corrective action: Check to
assure proper settings of following switches and circuit
breakers essential to INS operation:
1 Overhead circuit breaker panel
(fig. 2-26): Circuit breakers In:
1.
*
AVIONICS MASTER
CONTR
2.
* INS CONTROL
3.
*
AVIONICS MASTER
PWR #1
4.
*
AVIONICS MASTER
PWR #2
(3) Malfunction indications and procedures.
Table 3-2 details the procedure for a Malfunction Code
Check.
Table 3-3 lists a number of malfunction
indications which occur under given modes of operation.
Follow procedure given. Table 3-4 details action codes
and recommended action. Table 3-5 lists failed test
symptoms by malfunction codes and lists codes for
recommended actions.
(4) Read computer memory through CDU
(look routine).
1.
WYPT CHG key Press.
2
Overhead control panel (fig.
2-18): INVERTER No. 1 or INVERTER No.2 switch On
(either).
3
Mission control panel (fig. 41):
1.
3-phase AC BUS switch
On.
2.
INS POWER & AC CONT
circuit breakers - In.
4 Mission AC/DC power cabinet:
INS PWR circuit breaker In.
3-70
2.
Keys Enter "99". Note FROM
TO displays "99".
3.
INSERT/ADVANCE
pushbutton Press.
4.
Data selector LIL WYPT.
5.
CDU Enter "N" followed by
octal address of desired (even
numbered) memory location.
TM 55-1510-219-10
Table 3-5. Malfunction Codes
3-71
TM 55-1510-219-10
NOTE
approximately 120° around the nose of the aircraft,
extending to a distance of 240 nautical miles. The
presentation on the screen shows the location of
potentially dangerous areas, in terms of distance and
azimuth with respect to the aircraft. The radar is also
capable of ground mapping operations. The weather
radar set is protected by a 5ampere RADAR circuit
breaker located on the overhead circuit breaker panel
(fig. 2-26).
Program prevents entry of an
address higher than 13777.
6.
Insure waypoint wheel is at
"0".
7.
INSERT/ADVANCE
pushbutton Press. (Observe
address will appear in both
data displays.)
8.
9.
b. Controls and Indicators.
(1) Weather radar control-indicator switch.
Data selector DST/TIME.
(Observe most significant half
of desired data appears in left
display and least significant
half appears in right display.)
To obtain data at next higher
memory locations -Advance
waypoint thumbwheel. (For
example: If address 400 was
entered with thumbwheel at
"0", address 402 will be
available when thumbwheel is
set to "1", "404" when
thumbwheel is set to "2", etc.)
3-29. WEATHER RADAR SET (AN/APN-215).
CONTROL
FUNCTION
GAIN control
Used to adjust radar receiver gain in
the MAP mode only.
STAB OFF
Push type on/off switch.
switch
control
stabilization signals.
Range switches
Momentary action type switches.
When pressed, clears the screen
and increases or decreases the
range depending on switch pressed.
TILT control
Varies the elevation angle of radar
antenna a maximum of 15
a. Description. The weather radar set (fig. 3-24)
provides a visual presentation of the general sky area of
Figure 3-24. Weather Radar Control-Indicator (AN/APN-215)
3-72
Used to
antenna
TM 55-1510-219-10
crease or decrease depending on
switch pressed. When either top or
bottom mode is reached, the
opposite switch must be pressed to
further change the mode.
degrees up or down from horizontal
attitude of aircraft.
60°switch
Push type on/off switch. When
activated. reduces antenna scan
from 120°to 60°.
TRACK switches
Momentary action type switches.
When activated, a yellow track line
extending from the apex of the
display through top range mark
appears and moves either left or
right, depending on the switch
pressed. The track line position will
be displayed in degrees in the upper
left corner of the screen. The line
will disappear approximately 15
seconds after the switch is released.
It will then automatically return to
"0" degrees.
HOLD switch
Controls operation of the radar set.
OFF
Turns set off.
STBY
Places set in standby mode. This
position also initiates a 90 second
warmup delay when first turned on.
TEST
Displays test pattern to check for
proper operation of the set. The
transmitter is disabled during this
mode.
ON
Places set in normal operation.
MODE switches
Momentary action type switches.
Pressing and holding either switch
will display an information list of
operational data on the screen. The
data heading will be in blue, all data
except present data will be in
yellow, and present selected data
will show in blue.
The three
weather levels will be displayed in
red, yellow, and green. If WXA
mode has been selected, the red
bar will flash on and off. If the
switch is released and immediately
pressed again, the mode will in-
Interfaces radar indicator with INS
navigational information to show
INS
waypoint
positions,
superimposed on the weather
display, relative to aircraft position.
BRT control
Used to adjust screen brightness.
c. Normal Operation.
WARNING
Do not operate the weather radar
set
while
personnel
or
combustible materials are within
18 feet of the antenna reflector
When the weather radar set is
operating,
high-power
radio
frequency energy is emitted from
the antenna reflector, which can
have harmful effects on the
human body and can ignite
combustible materials.
Push type on/off switch. When
activated, the last image presented
before pressing the switch is
displayed and held.
The word
HOLD will flash on and off in the
upper left corner of the screen.
Pressing the switch again will
update the display and resume
normal scan operation.
Function switch
NAV switch
CAUTION
Do not operate the weather radar
set in a confined space where the
nearest metal wall is 50 feet or
less from the antenna reflector.
Scanning such surfaces may
damage receiver crystals.
(1) Turn-on procedure.
Function switch
TEST or ON, as required. (Information will appear after
time delay period has elapsed.)
(2) Initial adjustments operating procedure.
3-73
1.
BRT control As required.
2.
MODE switches Press and
release as required.
3.
RANGE switches Press and
release as required.
4.
TILT control Move up or down
to observe targets above or
below aircraft level. The echo
display will change in shape
and location only.
TM 55-1510-219-10
Refer to TM 1-230 for weather
observation, interpretation and
application.
(3) Test procedure.
(a)
procedure:
1.
Function switch TEST.
2.
RANGE switches Press and
release as required to obtain
80 mile display.
3.
BRT control As required.
4.
Screen Verify proper display.
(The test display consists of
two green bands, two yellow
bands, and a red band on a
120-degree scan. The word
TEST will be displayed in the
upper right corner.
The
operating mode selected by
the MODE switches, either
MAP, WX, or WXA, will be
displayed in the lower left
corner. If WXA has been
selected, the red band in the
test pattern will flash on and
off.
The range will be
displayed in the upper right
corner beneath the word
TEST and appropriate range
mark distances will appear
along the right edge of the
screen.)
Weather
observation
Function switch ON.
2.
MODE switches Press and
release as required to select
WX.
3.
BRT control As required.
4.
TILT control Adjust until
weather pattern is displayed.
Include the areas above and
below the rainfall areas to
obtain a complete display.
6.
operating
2.
MODE switches Press and
release as required to select
MAP.
3.
BRT control As required.
4.
GAIN control As required to
present usable display.
(4) Standby procedure.
Function switch
STBY.
d. Shutdown procedure: Function switch OFF.
e. Emergency operation. Not applicable.
3-30. TRANSPONDER SET (APX-100).
a. Description. The transponder system receives,
decodes, and responds to interrogations from Air Traffic
Control (ATC) radar to allow aircraft identification,
altitude reporting, position tracking, and emergency
tracking. The system receives a radar frequency of
1030 MHz and transmits preset coded reply pulses on a
radar frequency of 1090 MHz at a minimum peak power
of 200 watts. The range of the system is limited to lineof-sight.
operating
1.
5.
(b)
Ground
mapping
procedure: 1. Function switch ON.
The transponder system consists of a combined
receiver/transmitter/ control panel (fig. 3-25) located on
the pedestal extension; a pair of remote switches, one
on each control wheel; and two antennas, located on the
underside and top of the fuselage. The system is
protected by the 3-ampere TRANSPONDER and the 35ampere AVIONICS MASTER PWR #1 circuit breakers
on the overhead circuit breaker panel (fig. 2-26).
b. Transponder Control Panel fig. 3-25).
MODE switches Press and
release to select WXA. Areas
of intense rainfall will appear
as flashing red. These areas
must be avoided.
TRACK switches Press to
move track line through area
of least weather intensity.
Read relative position in
degrees in upper left corner of
screen.
NOTE
3-74
CONTROL
FUNCTION
TEST/GO
indicator
Illuminates to indicate successful
completion of built-in-test (BIT).
TEST/MON
Illuminates to indicate system indicator
malfunction or interrogation by a ground
station.
ANT switch
Selects desired antenna for signal input.
TOP
Selects upper antenna.
DIV
Selects diverse (both) antennas.
BOT
Selects lower antenna.
RAD TEST
Enables reply to TEST mode in-
OUT switch
terrogations from test set.
TM 55-1510-219-10
Figure 3-25. Transponder Control Panel (AN/APX - 100)
MASTER
CONTROL
OFF
STBY
NORM
EMER
STATUS ANT
indicator
STATUS KIT
indicator
STATUS ALT
indicator
IDENT MIC
switch
IDENT
MIC
MODE 4 reply
indicator
Selects system operating mode.
Deactivates system.
Activates system warm-up
(standby) mode.
Activates normal operating
mode.
Transmits emergency reply code.
Illuminates to indicate the BIT
or MON fault is caused by high
VSWR in antenna.
Illuminates to indicate the BIT
or MON fault is caused by external computer.
Illuminates to indicate BIT or
MON fault is caused by to Altitude Digitizer.
Selects source of aircraft identification signal.
Activates transmission of identification (IP) pulse.
Enables either control wheel
POS IDENT switch to activate
transmission of ident signal
from transponder.
MODE 4
AUDIO OUT
switch
AUDIO
Illuminates to indicate a reply
has been made to a valid Mode
4 interrogation.
Selects monitor mode for Mode
4 operation.
Enables sound and sight monitoring of Mode 4 operation.
LIGHT
Enables monitoring REPLY indicator for mode 4 operation.
OUT
Deactivates monitor mode.
MODE 3/A code Select desired reply codes for
selectors
Mode 3/A operation.
MODE 1 code
Select desired reply codes for
selectors
Mode 1 operation.
MODE
4 TEST Selects test mode of Mode 4
OUT switch
operation.
TEST
Activates built-in-test of Mode 4
operation.
ON
Activates mode 4 operation.
OUT
Disables Mode 4 operation.
MODE 4 code
Selects preset Mode 4 code.
control
3-75
TM 55-1510-219-10
M-C, M-3A,
M-2, and M-1
switches
TEST
ON
OUT
MODE 2 code
selectors
POS IDENT
pushbutton
(control wheels,
fig. 2-16)
8.
MODE 4 code control - A.
Set a code in the external
computer.
9.
MODE 4 AUDIO OUT switch OUT.
(3) Modes 1, 2, 3/A, and/or 4 operating
procedure.
NOTE
If the external security computer
is not installed, a NO GO light will
illuminate any time the Mode 4
switch is moved out of the OFF
position.
1.
MASTER control - NORM.
2.
M-1, M-2, M-3/A, and/or
MODE 4 ON-OUT switches ON.
Actuate only those
switches corresponding to the
required
codes.
The
remaining switches should be
left in the OUT position.
3.
MODE 1 code selectors - Set
(if applicable).
4.
MODE 3/A code selectors Set (if applicable).
5.
MODE 4 code control - Set (if
required).
6.
MODE 4 REPLY indicator
- Monitor to determine when
transponder set is replying to
a SIF interrogation.
7.
MODE 4 AUDIO OUT switch Set (as required to monitor
Mode 4 interrogations and
replies).
8.
MODE
4
audio
and/or
indicator - Listen and/or
observe
(for
Mode
4
interrogations and replies).
9.
IDENT-MIC switch - Press to
IDENT momentarily.
10.
MODE 4 TEST/OUT switch TEST.
11.
Observe that the TEST GO
indicator lamp illuminates.
12.
MODE 4 TEST/OUT switch ON.
13.
ANT switch - BOT.
14.
Repeat steps 4, 5, and 6.
Observe that the TEST GO
indicator illuminates.
Select test or reply mode of respective codes.
Activates self-test of selected
code. Transponder can also reply to ground interrogations in
the selected mode during test.
Activates normal operation.
Deactivates operation of selected
code.
Select desired reply codes for
Mode 2 operation. The cover
over mode select switches must
be slid forward to display the selected Mode 2 code.
When pressed, activates transponder identification reply.
c. Transponder - Normal Operation.
(1) Turn-on procedure. MASTER switch STBY. Depending on the type of receiver installed, the
NO GO indicator may illuminate. Disregard this signal.
(2) Test procedure.
NOTE
Make no checks with the master
switch in EMER, or with M-3/A codes
7600 or 7700 without first obtaining
authorization from the interrogating
station(s).
1.
Allow set two minutes to
warm up.
2.
Select codes assigned for use
in modes 1 and 3/A by
depressing and releasing the
pushbutton for each switch
until the desired number
appears in the proper window.
3.
Lamp indicators - Operate
press-to-test feature.
4.
M-1 switch - Hold in TEST.
Observe that no indicator
lamps illuminate.
5.
M-1 switch - Return to ON.
6.
Repeat steps 4 and 5 for the
M-2, M-3/A and M-C mode
switches.
7.
MASTER control - NORM.
3-76
TM 55-1510-219-10
15.
16.
17.
18.
19.
ANT switch - TOP.
Repeat step 14.
ANT switch - DIV.
Repeat step 14.
When possible, obtain the
cooperation
of
an
interrogating
station
to
exercise the TEST mode.
Execute the following steps:
a.
RAD TEST OUT switch RAD TEST.
b.
Obtain verification from
interrogating station that a
TEST MODE reply was
received.
c.
RAD TEST OUT switch OUT.
ton must be depressed to transmit
identification pulses.
(5) Shutdown procedure.
(a) To retain Mode 4 code in external
computer during a temporary shutdown:
1.
MODE 4 CODE switch Rotate to HOLD.
2.
Wait 15 seconds.
3.
MASTER control - OFF.
4.
To zeroize the Mode 4 code
in the external computer turn
MODE 4 CODE switch to
ZERO.
5.
MASTER control - OFF.
This will automatically zeroize
the external computer unless
codes have been retained
(step 1. above).
(4) Transponder set identification-position
operating procedure: The transponder set can make
identification-position replies while operating in code
Modes 1, 2, and/or 3/A, in response to ground station
interrogations. This type of operation is initiated by the
operator as follows:
1.
Modes 1, 2, and/or 3/A - On,
as required.
2.
IDENT-MIC switch - Press
momentarily to IDENT, when
directed.
3-31. PILOT'S ENCODING ALTIMETER.
a. Description. The encoding altimeter (fig. 326),
provides the pilot with an indication of present aircraft
altitude above sea level. It also supplies information to
the transponder for Mode C (altitude reporting) operation
and to the aided inertial navigation system. The circuit
is protected by the 5-ampere PILOT'S ALT ENCD circuit
breaker on the overhead circuit breaker panel and the 1ampere F21 fuse in the No. 1 junction box.
b. Controls and Indicators.
CONTROL
FUNCTION
NOTE
Holding
circuits
within
the
transponder
receiver-transmitter
will transmit identification-position
signals for 15 to 30 seconds. This
is normally sufficient time for
ground control to identify the
aircraft's position. During the 15
to 30 second period, it is normal
procedure to acknowledge via the
aircraft communications set that
identification position signals are
being generated.
MILLIBARS
indicator
IN HG Indicator
Drum indicator
Needle indicator
NOTE
CODE OFF flag
(pilot only)
ALT indicator
Test button
Set any of the M1, M2, M3/A, M-C,
or MODE 4 switches to OUT to
inhibit transmission of replies in
undesired modes.
NOTE
With the IDENT/OUT/MIC switch set
to the MIC position, the POS IDENT
but3-77
Indicates local barometric pressure in millibars. Adjusted by
use of set knob.
Indicates local barometric metric pressure in inches of mercury. Adjusted by use of set knob.
Indicates aircraft altitude in tenthousands, thousands, and hundreds of feet above sea level.
Indicates aircraft altitude in
hundreds of feet with subdivisions at twenty-foot intervals.
Presence indicates loss of power
to instrument.
Not used.
When pressed, reading should
decrease by 500 feet.
TM 55-1510-219-10
Figure 3-26. Pilot's Encoding Altimeter
NOTE
c. Normal Operation.
(1) Turn-on procedure. Encoding altimeter
will operate when transponder is operating with M-C
switch set to center position.
If the altimeter does not read
within 70 feet of field elevation,
when the correct local barometric
setting is used, the altimeter
needs calibration or internal
failure has occurred. An error of
greater than 70 feet also nullifies
use of the altimeter for IFR flight.
(2) Operating procedure.
1.
Barometric set knob - Set desired
altimeter setting in IN.
HG.
window.
2.
CODE OFF flag - Check not visible.
3.
Needle indicator - Check operation.
d. Emergency Operation. Altimeter circuit breaker
- Pull (if encoder fault occurs).
3-78
TM 55-1510-219-10
CHAPTER 4
MISSION EQUIPMENT
Section I. MISSION AVIONICS
4-1. MISSION
TIONS.
AVIONICS
OPERATING
INSTRUC-
a. The top section contains the mission
caution/advisory annunciator panel. Refer to Table 4-1.
Operating instructions for mission avionics equipment
are published in Chapter 3.
b. The center section of the mission control panel
contains the instruments used to monitor the mission
equipment.
These instruments are the DC
volt/ammeter, AC loadmeter, the antenna azimuth
indicator and controller.
4-2. MISSION CONTROL PANEL.
The mission control panel (fig. 4-1), mounted on the
copilot's sidewall, contains three sections.
c. The bottom section of the mission control panel
contains the mission equipment control switches and the
mission equipment circuit breakers. Refer to Table 4-2.
Table 4-1. Mission Annunciators
MSN OVERTEMP
CRYPTO ALERT
PWR SPLY FAL/LT
CALL
3 O AC OFF
BATT FEED OFF
MISSION POWER
LINK MODE
RADOME HOT
LINK SYNC
SPCL EQPT OVRD
DIPLEXER PRESS
TWTA STANDBY
ANT MALF
ANT STOWED
ANT OPERATE
RADOME HEAT
MISSION AC ON
DAT LINK UPDATE
TACAN UPDATE
MISSION DC ON
WAVE GUIDE
EXT AC PWR ON
EXT DC PWR ON
BT00996
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Green
Green
Green
Green
Green
Green
Green
Green
Green
Green
Mission equipment is over heating.
Coded messages being received.
Mission power out of tolerance.
Receiving transmission on VOW.
Three phase AC power fault.
Ground fault detected in battery.
Mission power is off.
WBDL fault in link or contact.
Radome heat is too high.
WBDL has synchronization fault.
Mission power switch is in override.
Diplexer has lost pressurization.
WBDL is in standby mode.
Boom antenna is out of position.
Boom antenna is in horizontal position.
Boom antenna is in vertical position.
Radome heat is on.
Mission AC is on.
INS is updating with information from Data Link.
INS is updating with information from TACAN.
Mission DC is on.
Wave guide is pressurized.
External AC power is on.
External DC power is on.
4-1
TM 55-1510-219-10
Figure 4-1. Mission Control Panel
4-2
TM 55-1510-219-10
Table 4-2. Mission Control Switches
3 ∅ AC BUS
RESET
ON
OFF
Resets and Energizes 3 0 AC BUS.
MSN CONT
OVERRIDE
AUTO
OFF
Overrides or turns off the mission DC POWER.
1 MSN INV
OFF
2 MSN INV
Selects 1 or 2 mission inverter or turns
inverters off.
DATA LINE
HV ON
STBY
OFF
Turns data link high voltage on, off or standby.
ANT SELECT
NOSE
AUTO
TAIL
Selects nose or tail data link antennas. In
auto mode, direction and signal
strength of received data link
determines selection of nose or tail data link antenna.
ANT OVRD
OPR POS
AUTO
ROTATE
Selects automatic rotating boom operation or
manual selection of antenna
position.
AC EXT
POWER
OFF
Turns external AC power on or off.
MISSION AC VOLT
FREQ & LOAD
BT00997
Allows monitoring of mission AC circuit.
Section II. AIRCRAFT SURVIVABILITY EQUIPMENT
WARNING
Right engine nacelle dispenser is
for chaff only.
(1) Dispenser
assemblies.
Two
interchangeable dispenser assemblies are mounted on
the aircraft. One is located in the aft portion of the right
nacelle and the other is mounted on the right side of the
fuselage. On this aircraft the dispenser in the nacelle
will be used for chaff only while the dispenser mounted
on the fuselage can be used for either flares or chaff.
The selector switch (placarded C-F) on the dispenser
can be set for either chaff or flares. The unit also
contains the sensor for the flare detector. The dispenser
assembly breech plate has the electrical contact pins
which fire the impulse cartridges. The unit also contains
the sequencing mechanism.
4-3.
M-130 FLARE AND CHAFF DISPENSING
SYSTEM.
a. Description.
The M-130 flare and chaff
dispensing
system
provides
effective
survival
countermeasures against radar guided weapons
systems and infrared seeking missile threats. The
system consists of two dispenser assemblies with
payload module assemblies, a dispenser control panel,
a flare dispense switch, two control wheel mounted chaff
dispensing switches, an electronic module assembly,
and associated wiring. The flare and chaff dispensing
system is protected by a 5-ampere circuit breaker,
placarded M130 POWER, located on the mission
control panel (fig. 4-1).
4-3
(2) Payload
module
assemblies.
A
removable payload module assembly is provided for
each dispenser assembly. Each payload module has 30
chambers which will accept either flares or chaff. Flares
or chaff are loaded into the rear-end (studded end) of
the payload module, and secured in place by a retaining
plate.
(3) Electronic module assembly (EM). The
electronic module assembly contains the programmer,
the flare detector and a safety switch. The unit is
located behind the pilot's seat.
(a) Flare detector. The flare detector is
provided to insure that a flare is burning when it is
ejected from the dispenser payload module. If the initial
flare fails to ignite, the detector automatically fires
another flare within 75 milliseconds. If the second flare
fails to ignite, the detector will fire a third flare. If the
third flare ignition is not detected, the detector will not
fire another flare until the system is activated again by
pressing the FLARE DISPENSE switch.
(b) Programmer. The programmer is
used for the chaff mode only. It has four switches for
setting count and interval of salvo and burst.
(c) Safety switch. The safety switch
(with safety pin and yellow flag) prevents firing of chaff
TM 55-1510-219-10
or flares when the safety pin is inserted. The safety pin
shall be removed only while the aircraft is in flight or
during test of the system.
(4) Flare dispenser switch.
A single
pushbutton switch placarded FLARE DISPENSE,
located on the control pedestal, will fire a flare from the
dispenser payload module each time it is pressed. If the
FLARE DISPENSE switch is held down, it will dispense
a flare every 2.3 seconds.
(5) Control wheel mounted chaff dispense
switches. Two pushbutton switches placarded CHAFF
DISP, one located on top left portion of the pilot's control
wheel and the other located on the top right portion of
the copilot's control wheel, activates the chaff
dispensing system when pressed.
(6) Wing mounted safety switch. A wing
mounted safety switch (with safety pin and red flag)
located on top of the right wing, just aft of the nacelle,
prevents the firing of chaff or flares when the pin is
inserted. This safety pin shall be inserted while the
aircraft is on the ground and removed prior to flight or
during system test.
(7) Dispenser control panel (DCP).
The
dispenser control panel (fig. 4-2) is mounted in the
control pedestal. Control functions are as follows:
Figure 4-2. Chaff/Flare Dispenser Control Panel
4-4
CONTROL
ARM-SAFE switch
ARM light
RIPPLE FIRE switch
FLARE counter
TM 55-1510-219-10
0.97 inches square. The base of the chaff case is
flanged to provide one-way assembly into the dispenser
payload module. The chaff consists of aluminum
coated fiberglass strands.
FUNCTION
When in the SAFE position,
power is removed from the
M130 system. When in the
ARM position, power is applied
to the M-130 system.
(b) Countermeasure flare M206. These
units consist of an aluminum case 8 inches in length and
0.97 inches square. The base of the flare is flanged to
provide one-way assembly into the payload module.
The flare material consists of a magnesium and teflon
composition. A preformed packing is required in the
base of the flare unit prior to inserting the impulse
cartridge.
An amber press to test indicator
light placarded ARM illuminates
when the ARM-SAFE switch is
in the ARM position, when the
safety pins are removed from
the electronic module and the
wing safety switch. Clockwise
rotation will dim the indicator
light.
(c) Impulse cartridge M796.
This
cartridge fits into the base of either the flare or chaff and
is electrically initiated to eject flares or chaff from the
dispenser payload module.
A guarded switch placarded
RIPPLE
FIRE
fires
all
remaining flares when moved to
the up position. It is used in the
event of an inflight emergency
to dispense all flares from the
dispenser payload module.
b. Normal Operation.
NOTE
If aircraft is to be flown with flare
dispenser assembly removed,
fairing should be removed from
fuselage.
Indicates the number of flares
remaining in the dispenser
payload module.
FLARE counter
setting knob
Facilitates setting FLARE
counter to the number of flares
in the payload module before
flight.
CHAFF counter
Indicates the number of chaffs
remaining in the payload
module.
CHAFF counter
setting knob
Facilitates setting CHAFF
counter to the number of chaffs
in the payload module before
flight. SELECTOR SWITCH
MAN
Bypasses the programmer and
fires one chaff each time one of
the chaff dispense switches is
pressed.
(1) General. At the present time surface toair intermediate range guided missiles that are launched
against the aircraft must be visually detected by the
aircraft crew.
Crew members must insure visual
coverage over the ground area where a missile attack is
possible. The aircraft radar warning system will only
alert the pilot and copilot when the aircraft is being
tracked by radar-guided anti-aircraft weapons systems.
It will not indicate the firing of weapons against the
aircraft.
(2) Crew responsibilities.
The pilot or
designated crew member is responsible for removing
the safety pin from the right wing before flight, and for
replacing it immediately after flight. After the aircraft is
airborne the pilot is responsible for removing the safety
pin from the electronic module and moving the ARMSAFE switch on the dispenser control panel to ARM.
Before landing, he is responsible for re-inserting the
safety pin in the electronic module and moving the
ARM-SAFE switch to SAFE. While airborne the pilot
and copilot are responsible for scanning the terrain for
missile threats. When either pilot recognizes a missile
launch he will press the FLARE DISPENSE button to
eject flares.
PRGM
Chaff is fired in accordance with
the preset chaff program as set
into the electronic module
(count and interval of bursts and
salvo).
(8) Ammunition for dispenser. Ammunition
for the system consists of countermeasure chaff Ml,
countermeasure flares M206, and impulse cartridges
M796.
(3) Conditions for firing.
The dispenser
system should not be fired unless a missile launch is
observed or radar guided weapons systems is detected
and locked on. If a system malfunction is suspected,
aircraft commander may authorize attempts to dispense
flares or chaff as a test in a non-hostile area.
(a) Countermeasure chaff M1. These
units consist of a plastic case 8 inches in length and
4-5
TM 55-1510-219-10
hexagonal wrench provided in test set carrying
case.
4. Obtain test set power cable from the M-91 test
set carrying case and connect cable between
exterior connection J1 (28V DC) on aircraft and
aircraft power + 28V DC (J1) of test set.
5. Remove safety pin from EM and in the top skin
of the right wing.
WARNING
Aircraft must be in flight to
dispense flares.
(a) Firing procedure.
1. Flares. Upon observing a missile
launch the designated crewmember will fire a flare. If
more than one missile launch is observed, the firing
sequence should be continued until the aircraft has
cleared the threat area.
2. Chaff. Upon receiving an alert
from the aircraft radar warning system, the designated
crewmember will fire the chaff and initiate an evasive
maneuver. The number of burst/salvo and number of
salvo/program and their intervals as established by
training doctrine will be set into the programmer prior to
take-off (refer to TM 9-1095206-13 for information on
setting programmer). If desired, the operator may
override the programmed operational mode and fire
chaff countermeasures manually by moving the
dispenser function selector switch to MANUAL and
pressing a dispenser switch.
4-4.
CAUTION
On DCP, assure that RIPPLE FIRE
switch guard is in down position.
6. Provide aircraft power to DCP by setting M130
POWER circuit breaker to ON position.
7. On DCP, press ARM lamp.
Lamp will
illuminate. Release ARM lamp. Lamp will
extinguish.
8. On DCP, set FLARE counter to 30 CHAFF
COUNTER to 30 and MAN-PRGM switch to
MAN position.
9. On DCP, set ARM-SAFE switch to ARM. ARM
lamp will illuminate.
NOTE
When the test set is installed on
the dispenser assembly and 28
volts DC aircraft power has been
applied, the sequencer switch
inside of dispenser assembly
resets, making an audible sound
as it rotates. There will be no such
sound if the sequencer switch has
been previously reset or if switch
is in position 12 or 24.
SYSTEM DAILY PREFLIGHT/RE-ARM TEST.
The following test procedures shall be conducted prior to
the first flight of each day and prior to each re-arming of
the dispensers. The first dispenser tested shall be the
one used to dispense flares and the second one shall be
the one used to dispense chaff. Notify AVUM if any
improper indications occur during the tests.
WARNING
Assure payload module is not
connected to dispenser assembly
at any time during the following
test procedure.
NOTE
On test set, TS PWR ON lamp
(clear) illuminates and remains
illuminated throughout the test
sequence until aircraft power to
test set (via test set power cable)
is disconnected or shut off.
1. On flare dispenser assembly, assure the C-F
selector switch is in F (flare) position.
2. Obtain M-91 test set and ensure that TEST
SEQUENCE switch is in START/HOME
position.
3. Connect base plate of test set to breech of
dispenser assembly. Secure both mounting
studs uniformly hand tight, using 5/32 inch
10. Set mission chaff program on EM.
11. Perform the following operations on the M-91
test set: a. Press to test the remaining three
lamps on test set. Each lamp will illuminate.
4-6
TM 55-1510-219-10
14. Perform the following operations on the M-91
test set:
NOTE
Replace any lamp that does not
illuminate when pressed. If none
of the indicating lamps illuminate,
return test set to AVUM.
a. Rotate
TEST
SEQUENCE
switch counter-clockwise to SYS
NOT RESET position. SYS NOT
RESET
lamp
(amber)
will
illuminate.
DISPENSER
COMPLETE lamp (green) will
remain illuminated.
b. Press
and
release
MANUAL
SYSTEM RESET switch. SYS NOT
RESET
lamp
(amber)
will
extinguish.
b. Rotate
TEST
SEQUENCE
switch clockwise to the next
position, TS RESET. No visual
indication will occur.
c. Rotate
TEST
SEQUENCE
switch clockwise to SV SELF TEST
position. STRAY VOLTAGE lamp
(red) will illuminate.
d. Rotate
TEST
SEQUENCE
switch clockwise to next position,
TS RESET. STRAY VOLTAGE
lamp (red) will extinguish.
e. Rotate
TEST
SEQUENCE
switch clockwise to next position,
STRAY VOLT. STRAY VOLTAGE
lamp (red) should not illuminate.
f. Rotate
TEST
SEQUENCE
switch clockwise to next position,
SYS NOT RESET.
SYS NOT
RESET lamp (amber) should not
illuminate.
If lamp illuminates,
press
and
release
MANUAL
SYSTEM RESET switch and SYS
NOT RESET lamp should then
extinguish.
NOTE
When the MANUAL SYSTEM
RESET switch is pressed and
released, and 28 volts DC power
has been applied, the sequencer
switch inside the dispenser
assembly resets, making an
audible sound as it rotates. If the
sequencer switch has been
previously reset or if the switch is
in position 12 or 24, there will be
no such sound.
NOTE
When the MANUAL SYSTEM
RESET switch is pressed and
released, and 28 volts DC power
has been applied, the sequencer
switch inside the dispenser
assembly resets, making an
audible sound as it rotates. If the
sequencer switch has been
previously reset or if the switch is
in position 12 or 24, there will be
no such sound.
c. Rotate
TEST
SEQUENCE
switch counterclockwise to STRAY
VOLT position. STRAY VOLTAGE
lamp (red) should not illuminate.
d. Rotate
TEST
SEQUENCE
switch
counterclockwise
to
START/HOME position.
NOTE
When the TEST SEQUENCE
switch
is
turned
to
the
START/HOME
position,
the
DISPENSER COMPLETE lamp will
extinguish, the STRAY VOLTAGE
lamp will illuminate and then will
extinguish when passing through
the TS RESET position.
g. Rotate TEST SEQUENCE switch
clockwise to next position, DISP
COMP.
15. On CHAFF dispenser assembly, assure that C-F
selector switch is in C (chaff) position.
16. Remove M-91 test set from first dispenser
assembly.
17. Connect M-91 test set to breech assembly of
second dispenser assembly.
Secure both
mounting studs uniformly hand tight using ball
hexagonal key screwdriver provided in test set
carrying case.
12. Press FLARE DISP switch once. For each
depressing, the FLARE counter on DCP should
count down in groups of three.
13. On DCP, raise RIPPLE FIRE switch guard and
set toggle switch to up position until FLARE
counter counts down to 00. Return switch guard
to down position.
On DCP, reset FLARE
counter back to 30. DISPENSER COMPLETE
lamp (green) on test set will illuminate.
4-7
TM 55-1510-219-10
released, and 28 volts DC power
has been applied, the sequencer
switch inside the dispenser
assembly resets, making an
audible sound as it rotates. IF the
sequencer switch has been
previously reset or if the switch is
in position 12 or 24, there will be
no such sound.
NOTE
When the test set is installed on
the dispenser assembly and 28
volts DC aircraft power has been
applied, the sequence switch
inside the dispenser assembly
resets, making an audible sound
as it rotates. There will be no
such sound if the sequencer
switch has been previously reset
or if switch is in position 12 or 24.
g. Rotate TEST SEQUENCE switch
clockwise to next position, DISP
COMPL.
NOTE
On test set, TS PWR ON lamp
(clear) illuminates and remains
illuminated through the test
sequence until aircraft power to
test set (via test set power cable)
is disconnected or shut off.
19. Press pilot CHAFF DISP switch once. Press
copilot CHAFF DISP switch once. On DCP, for
each depressing, the CHAFF counter should
count down by an increment of one.
20. On DCP, set MAN-PRGM switch to PRGM
position.
21. Press any one of CHAFF DISP switches in
aircraft. In DCP, the number shown on CHAFF
counter should decrease in accordance with the
program set on the EM.
22. Repeatedly press other CHAFF DISPENSE
switch until CHAFF counter on DCP reads 00.
23. On test set, observe DISPENSE COMPLETE
lamp (green) is illuminated and then perform the
following operations:
18. Perform the following operations on the M-91
test set: a. Press to test all four lamps on test
set. Each lamp will illuminate.
NOTE
Replace any lamp that does not
illuminate when pressed. If none
of the indicating lamps illuminate,
return test set to AVUM.
a. Rotate TEST SEQUENCE switch
counter-clockwise to SYS NOT
RESET position. SYS NOT RESET
lamp (amber) will illuminate.
b. Press
and
release
MANUAL
SYSTEM RESET switch. SYS NOT
RESET
lamp
(amber)
will
extinguish.
b. Rotate TEST SEQUENCE switch
clockwise to TS RESET position.
No visual indication will occur.
c. Rotate TEST SEQUENCE switch
clockwise to SV SELF TEST
position. STRAY VOLTAGE lamp
(red) will illuminate.
d. Rotate TEST SEQUENCE switch
clockwise to next position, TS
RESET. STRAY VOLTAGE lamp
(red) will extinguish.
e. Rotate TEST SEQUENCE switch
clockwise to next position, STRAY
VOLT. STRAY VOLTAGE lamp
(red) should not illuminate.
f. Rotate TEST SEQUENCE switch
clockwise to next position, SYS
NOT RESET. SYS NOT RESET
lamp (amber) should not illuminate.
If lamp illuminates, press and
release MANUAL SYSTEM RESET
switch and SYS NOT RESET lamp
should then extinguish.
NOTE
When the MANUAL SYSTEM
RESET switch is pressed and
released, and 28 volts DC power
has been applied, the sequencer
switch inside the dispenser
assembly resets, making an
audible sound as it rotates. If the
sequencer switch has been
previously reset or if the switch is
in position 12 or 24, there will be
no such sound.
c. Rotate TEST SEQUENCE switch
counter-clockwise to STRAY VOLT
position. STRAY VOLTAGE lamp
(red) should not illuminate.
NOTE
When the MANUAL SYSTEM
RESET switch is pressed and
4-8
TM 55-1510-219-10
d. Rotate TEST SEQUENCE switch
counter-clockwise to START/HOME
position.
CAUTION
Prior to insertion of an impulse
cartridge, be sure there is
preformed packing in the flare
cartridge.
(There will be no
preformed packing in chaff
cartridges.)
Reinstall
any
preformed
packing
that
is
inadvertently removed with dust
cap.
The loading of impulse
cartridges into a flare or chaff
shall be accomplished one at a
time.
NOTE
When
the TEST
SEQUENCE
switch is turned to the OFF
position,
the
DISPENSER
COMPLETE lamp will extinguish,
the STRAY VOLTAGE lamp will
illuminate and then will extinguish
when the OFF position is reached.
24. Install safety pins.
25. Disconnect test set power cable.
26. Remove M-91 test set from dispenser assembly
and restore in carrying case along with the
power cable and hexagonal wrench.
27. On DCP, set ARM-SAFE switch to SAFE
position.
28. On DCP, reset CHAFF counter to 30.
29. Disconnect aircraft power by pulling the M130
POWER circuit breaker located on the mission
control panel (fig. 4-1).
30. Proceed immediately to ammunition loading
procedures.
4-5.
5. Insert one impulse cartridge into each flare (or
chaff).
6. Install retainer plate assembly by screwing to
two retainer bolts into payload module.
WARNING
The system must have been
tested to assure that there is no
stray voltage and all aircraft
power must be removed from the
system prior to unloading the
payload module.
AMMUNITION.
7. On the dispenser control panel, assure ARMSAFE switch is in SAFE position.
8. On the electronic module and right wing assure
safety pins and flag assemblies are installed.
9. Slide payload module assembly into dispenser
assembly and secure two stud bolts, hand tight,
using 5/32 inch hexagonal wrench.
a. Ammunition Loading Procedure.
WARNING
Only one shipping container is to
be opened at a time. If a shipping
container has been opened and
only
partially
emptied,
the
remaining
contents
will
be
secured in the container with an
appropriate type of packaging
material or filler to adequately
prevent jostling. All munitions in
storage must be in their original
shipping containers.
b. Ammunition Unloading Procedure.
WARNING
All aircraft power to the dispenser
system must be turned off prior to
removal of payload module from
dispenser assembly. Safety pin
flag shall be installed in the
electronic module prior to landing
and the safety pin flag shall be
installed in the wing-mounted
safety switch immediately after
landing.
1. Place payload module assembly on work
bench in approved safe area so that the
retaining plate is facing up.
2. Remove retaining plate by unscrewing two
retaining bolts.
3. Insert one flare (or chaff) at a time into each
chamber of payload module.
4. Remove plastic dust cap from each chaff or
flare.
1. On dispenser control panel, assure ARMSAFE switch is in SAFE position.
4-9
TM 55-1510-219-10
2. Assure safety pin and flag are inserted into
electronic module and in the wing mounted
safety switch.
4-6. RADAR SIGNAL DETECTING SETS AN/ APR39(V)1 OR AN/APR-39(V)2.
WARNING
If there is an indication that a
misfire occurred, notify EOD
personnel for disposition and
disposal.
The radar signal detecting system indicates the
relative position of search radar stations. Audio warning
signals are applied to the pilot's and copilot's headsets.
The radar signal detecting set is protected by the 7.5ampere circuit breaker placarded APR-39, located on
the mission control panel (fig. 4-1). The associated
antennas are shown in figure 2-1. For AN/APR-39(V)l
operating instructions, refer to TM 11-5841-283-20.
Refer to TM 11-584120-2 for the AN/APR-39 (V)2
operating instructions. Pattern #1, self test, shall be as
shown in figure 4-5.
3. Remove module from dispenser assembly by
unscrewing two stud bolts with a 5/32 inch
hexagonal wrench and sliding out of dispenser
assembly.
4. Remove retaining plate from payload module by
unscrewing two retaining bolts.
5. Remove expended and unexpended impulse
cartridges and flares (or chaff) from payload
module.
6. Repack unexpended items in original containers
and return to stores.
a. Radar Signal Detecting Set Control Panel
Functions (AN/APR-39(V)1) (fig. 4-3).
NOTE
It is not unusual for the case of a
chaff cartridge to crack when
fired.
It does not effect
performance of the item and
should not be reported as a
malfunction.
CONTROL
FUNCTION
PWR switch
Turns set On or Off.
SELF TEST
switch
Initiates self test.
DSCRM switch
Turns discriminate function On
or Off.
Figure 4-3. Radar Signal Detecting Set Control Panel (AN/APR-39(V)1
4-10
TM 55-1510-219-10
AUDIO control
DAY-NIGHT
control
Adjusts audio level.
b. Radar Signal Detecting Set Control Panel
Functions (AN/APR-39(V)2) (fig. 4-4).
CONTROL
FUNCTION
PWR switch
Turns set On or Off
ALT/HI Low
Selects mode of operation
switch
HI
Selects high altitude threat
mode
LOW
Selects low altitude threat mode
TEST switch
Initiates self test
AUDIO control
Adjust audio level
4-7. RADAR WARNING RECEIVER AN/APR-44()
(V)3.
The radar warning receiver (fig. 4-6) indicates the
presence of certain types of search radar signals.
The radar warning receiver is protected by the 5ampere
circuit breaker placarded APR-44, located on the
mission control panel (fig.
4-1).
For operating
instructions, refer to TM 11-5841-291-12.
c. Radar Signal Detecting Set Indicator Functions
(fig. 4-5).
CONTROL
MA indicator
BRIL control
Rotation adjusts intensity of
display.
FUNCTION
Illuminates to indicate the
presence of an MA threat.
CONTROL
Radar warning
indicator
FUNCTION
Illuminates to indicate
presence of an AI or SAM
threat.
VOLUME
control
POWER switch
Adjusts volume.
Turns set On or Off.
Adjusts brilliance.
Figure 4-4. Radar Signal Detecting Set Control Panel (AN/APR-39(V)2
4-11
the
TM 55-1510-219-10
Figure 4-5. Radar Signal Detecting Set Indicator
AP005715
Figure 4-6. Radar Warning Receiver Control Panel AN/APR-44( )(V)3
4-12
TM 55-1510-219-10
CHAPTER 5
OPERATING LIMITS AND RESTRICTIONS
5-1.
PURPOSE.
Section I. GENERAL
5-3. EXCEEDING OPERATIONAL LIMITS.
Anytime an operational limit is exceeded an
appropriate entry shall be made on DA Form 2408-13.
Entry shall state what limit or limits were exceeded,
range, time beyond limits, and any additional data that
would aid maintenance personnel in the maintenance
action that may be required.
This chapter identifies or refers to all operating
limits and restrictions that shall be observed during
ground and flight operations.
5-2.
GENERAL.
The operating limitations set forth in this chapter
are the direct result of design analysis, tests, and
operating experiences. Compliance with these limits will
allow the pilot to safely perform the assigned missions
and to derive maximum utility from the aircraft. Limits
concerning maneuvers, weight, and center of gravity are
also covered in this chapter.
5-4.
MINIMUM CREW REQUIREMENTS.
The minimum crew required for flight is two pilots.
Additional crewmembers will be added as required, at
the discretion of the commander, in accordance with
pertinent Department of the Army regulations.
MACOMS may authorize maintenance test flights to be
conducted with one qualified pilot at the pilot's station,
and a trained technical observer at the copilot's station,
in day/visual meteorological conditions only.
Section II. SYSTEM LIMITS
5-5.
The blue marking on the airspeed indicator denotes best
rate of climb with one engine inoperative, at maximum
gross weight, maximum forward c.g., sea level standard
day conditions.
INSTRUMENT MARKINGS.
Instruments which display operating limitations
are illustrated in figure 5-1. The operating limitations
are color coded on the instrument faces. Color coding
of each instrument is explained in the illustration.
5-6.
5-7.
PROPELLER LIMITATIONS.
The maximum propeller overspeed limit is 2200
RPM. Propeller speeds above 2000 RPM indicate
failure of the primary governor. Propeller speeds above
2080 RPM indicate failure of both primary and
secondary governors. Torque is limited to 81% for
sustained operation above 2000 RPM.
INSTRUMENT MARKING COLOR CODES.
Operating limitations and ranges are illustrated by
the colored markings which appear on the dial faces of
engine, flight, and utility system instruments. Red
markings indicate the limit above or below which
continued operation is likely to cause damage or shorten
life. The green markings indicate the safe or normal
range of operation. The yellow markings indicate the
range when special attention should be given to the
operation covered by the instrument. Operation is
permissible in the yellow range, but should be avoided.
White markings on the airspeed indicator denotes flap
operating range.
5-8.
STARTER LIMITATIONS.
The starters in this aircraft are limited to an
operating period of 30 seconds ON, then 5 minutes
OFF, for two starter operations. After two starter
operations the starter shall be operated for 30 seconds
ON, then 30 minutes OFF.
5-1
TM 55-1510-219-10
Figure 5-1. Instrument Markings (Sheet 1 of 3)
5-2
TM 55-1510-219-10
Figure 5-1. Instrument Markings (Sheet 2 of 3)
5-3
TM 55-1510-219-10
Figure 5-1. Instrument Markings (Sheet 3 of 3)
5-4
TM 55-1510-219-10
5-9.
AUTOPILOT LIMITATIONS.
a. An autopilot preflight check must be conducted
and found satisfactory prior to each flight on which the
autopilot is to be used.
c. Fuel System Anti-Icing.
Icing inhibitor
conforming to MIL-I-27686 PRIST will be added to
commercial fuel, not containing an icing inhibitor, during
fueling operations, regardless of ambient temperatures.
The additive provides anti-icing protection and also
functions as a biocide to kill microbial growth in aircraft
fuel systems.
b. A pilot must be seated at the controls with the
seat belt fastened when the autopilot is in operation.
c. Operation of the autopilot and yaw damper is
prohibited during takeoff and landing, and below 200
feet above terrain.
Maximum speed for autopilot
operation is 247 knots/0.47 Mach.
5-11. BRAKE DEICE LIMITATIONS.
The following limitations apply to the brake deice
system:
d. During a coupled ILS approach do not operate
the propellers in the 1750 to 1850 RPM range.
a. The brake deice system shall not be operated at
ambient temperatures above + 15°C.
5-10. FUEL SYSTEM LIMITS.
b. The brake deice system shall not be operated
longer than 10 minutes (one timer cycle) with the
landing gear retracted.
If operation does not
automatically terminate approximately 10 minutes after
gear retraction, turn the brake deice switch OFF.
NOTE
Aviation
gasoline
(AVGAS)
contains a form of lead which has
an accumulative adverse effect on
gas turbine engines. The lowest
octane AVGAS available (less lead
content) should be used. If any
AVGAS is used the total operating
time must be entered on DA Form
2408-13.
c. Maintain 85% NI or higher during simultaneous
operation of the brake deice and surface deice systems.
If adequate pneumatic pressure cannot be provided for
simultaneous operation of the brake deice and surface
deice systems, turn OFF the brake deice system.
d. In order to maintain an adequate supply of
systems pneumatic bleed air, the brake deice system
shall be turned OFF during single engine operation.
a. Operating Limits. Operation with FUEL PRESS
light on is limited to 10 hours. Log FUEL PRESS light
on time on DA Form 2408-13. One standby boost pump
may be inoperative for takeoff. (Crossfeed fuel will not
be available from the side with the inoperative standby
boost pump.) Operation on aviation gasoline is time
limited to 150 hours between engine overhaul and
altitude limited to 20,000 feet with one standby boost
pump inoperative. Crossfeed capability is required for
climb, when using aviation gasoline above 20,000 feet.
5-12. PITOT HEAT LIMITATIONS.
Pitot heat should not be used for more than 15
minutes while the aircraft is on the ground.
5-13. PNEUMATIC
LIMITATIONS.
b. Fuel Management. Auxiliary tanks will not be
filled for flight unless the main tanks are full. Maximum
allowable fuel imbalance is 1000 lbs. Do not take off if
fuel quantity gages indicate in yellow arc (less than 265
lbs. of fuel in each main tank). Crossfeed only during
single engine operation.
SURFACE
DEICE
SYSTEM
The pneumatic surface deice system shall not be
operated at ambient temperatures below -40'C.
5-5
TM 55-1510-219-10
Section III. POWER LIMITS
used in any quantity with primary
5-14. ENGINE LIMITATIONS.
Operation of the engines is monitored by
or alternate fuel.
instruments, with the operating limits marked on the
face of each instrument.
5-15. OVERTEMPERATURE
AND
OVERSPEED
LIMITATIONS.
CAUTION
Engine operation using only the
a. Whenever the limiting temperatures are
engine driven fuel pump without
exceeded and cannot be controlled by retarding the
boost pump fuel pressure is
power levers, the engine will be shut down and a landing
limited to 10 cumulative hours.
made as soon as possible.
All time in this category shall be
b. During engine operation the temperatures,
entered on DA Form 2408-13 for
speeds
and time limits listed in the Engine Operating
the attention of maintenance
Limitations
chart (table 5-1) must be observed. When
personnel.
these limits are exceeded, the incident will be entered
as an engine discrepancy in the appropriate
CAUTION
maintenance forms. It is particularly important to record
Use of aviation gasoline is timethe amount and duration of over-temperature and/or
limited to 150 hours of operation
overspeed.
during
any
Time
Betweenc. Continuous engine operation above 725 °C will
Overhaul (TBO) period. It may be
reduce engine life.
Table 5-1. Engine Operating Limitations
5-6
TM 55-1510-219-10
5-16. POWER DEFINITIONS FOR ENGINE OPERATIONS.
The following definitions describe the engine
power ratings.
Table 5-2. Generator Limits
a. Takeoff Power. The maximum power available,
limited to periods of five minutes duration.
b. Maximum Continuous Power.
The highest
power rating not limited by time. Use of this rating is
intended for emergency situations at the discretion of
the pilot.
5-17. GENERATOR LIMITS.
Maximum generator load is limited to 100% for
flight and variable during ground operations. Observe
the limits shown in Table 5-2 during ground operation.
Section IV. LOADING LIMITS
5-18. CENTER OF GRAVITY LIMITATIONS.
Center of gravity limits and instructions for
computation of the center of gravity are contained in
Chapter 6. The center of gravity range will remain
within limits, providing the aircraft loading is
accomplished according to instructions in Chapter 6.
planning must be accomplished,
prior to takeoff, by analysis of
maximum
takeoff
weight,
accelerate-stop distance, positive
engine-out
climb
at
liftoff,
accelerate-go distance, takeoff
climb
gradient,
and
climb
performance.
5-19. WEIGHT LIMITATIONS.
WARNING
The ability to sustain loss of
engine power and successfully
stop, continue the takeoff, or
climb before or after gear
retraction is not assured for all
conditions. Thorough mission
The maximum designed gross weight is 14,200
pounds for takeoff. Maximum landing weight is 13, 500
pounds. Maximum ramp weight is 14,290 pounds.
Maximum zero fuel weight is 11,500 pounds.
Section V. AIRSPEED LIMITS MAXIMUM AND MINIMUM
5-20. AIRSPEED LIMITATIONS.
Airspeed indicator readings contained in
procedures, text, and illustrations throughout this
Operator's Manual are given as indicated airspeed
(IAS). Airspeed indicator markings (fig.
5-1) and
placarded airspeeds, located on the cockpit overhead
control panel (fig. 2-18), are calibrated airspeed (CAS).
Airspeed Calibration Charts are provided in Chapter 7.
5-22. LANDING GEAR EXTENSION SPEED.
5-21. MAXIMUM ALLOWABLE AIRSPEED.
The airspeed limit for retracting the landing gear is
166 KIAS.
The airspeed limit for extending the landing gear
and for flight with the landing gear extended is 184
KIAS.
5-23. LANDING GEAR RETRACTION SPEED.
Refer to figure 5-2 to determine limiting airspeeds
at maximum gross weight under various conditions.
The maximum allowable airspeed is 247 KIAS/0.47
Mach.
5-24. WING FLAP EXTENSION SPEEDS.
The airspeed limit for APPROACH extension
(40%) of the wing flaps is 202 KIAS. The airspeed
5-7
TM 55-1510-219-10
Figure 5-2. Flight Envelope
5-8
TM 55-1510-219-10
limit for full DOWN extension (100%) of the wing flaps is
157 KIAS. If wing flaps are extended above these speeds,
the flaps or their operating mechanisms may be damaged.
control airspeed (Vmc) at sea level standard conditions is 89
KIAS.
5-26.
5-25. MINIMUM SINGLE-ENGINE CONTROL
AIRSPEED (VMC).
MAXIMUM DESIGN MANEUVERING SPEED.
The maximum design maneuvering speed is 173 KIAS.
Chapter 7, Section X describes minimum singleengine control airspeed. The minimum single-engine
Section VI. MANEUVERING LIMITS
5-27. MANEUVERS.
a. The following maneuvers are prohibited.
G's with wing flaps up or a positive load factor of
2.0 G’s, or a negative 1.227 G’s with wing flaps
down.
1. Spins.
b. Recommended turbulent air penetration airspeed is
173 KIAS.
2. Aerobatics of any kind.
5-28. BANK AND PITCH LIMITS.
3. Abrupt maneuvers above 173 KIAS.
a. Bank limits are 60° left or right.
4. Any maneuver which results in a positive load factor of 3.10 G’s or a negative load factor of 1.227
b. Pitch limits are 30° above or below the horizon.
Change 3
5-8.1
TM 55-1510-219-10
SECTION VII. ENVIRONMENTAL RESTRICTIONS
5-29. ALTITUDE LIMITATIONS.
The maximum altitude that the aircraft may be operated
at is 31,000 feet. When operating with inoperative yaw
damp, the altitude limit is 17,000 feet.
5-30. TEMPERATURE LIMITS.
a The aircraft shall not be operated when the ambient
temperatures are warmer than ISA +37°C at SL to 25,000
feet, or ISA +31°C above 25,000 feet.
b. Engine ice vanes shall be retracted at +15°C and
above.
5-31. FLIGHT UNDER IMC (INSTRUMENT
METEOROLOGICAL CONDITIONS).
This aircraft is qualified for operational in instrument
meteorological conditions.
5-31A. ICING LIMITATIONS (TYPICAL).
WARNING
While in icing conditions, if there is an
unexpected 30% increase of torque needed
to maintain airspeed in level flight, a
cumulative total of two or more inches of ice
accumulation on the wing, an unexplained
decrease of 15 knots IAS, or an unexplained
deviation between pilot’s and copilot’s
airspeed indicators, the king environment
should be exited as soon as practicable. Ice
accumulation on the pitot tube assemblies
could cause a complete loss of airspeed
indication.
The following conditions indicate a possible
accumulation of ice on the pitot tube assemblies and
unprotected airplane surfaces. If any of these conditions
are observed, the icing environment should be exited as
soon as practicable.
1. Total ice accumulation of two inches or more on the
wing surfaces. Determination of ice thickness can be
accomplished by summing the estimated ice thickness on
the wing prior to each pneumatic boot deice cycle (e.g. four
cycles of minimum recommended ½-inch accumulation.
2. A 30 percent increase in torque per engine required
to minimum an desired airspeed in level flight (not to
exceed 85 percent torque) when operating at recommended
holding speed.
3. A decrease in indicated airspeed of 15 knots after
entering the icing condition (not slower than 1.4 power off
stall speed) if maintaining original power setting in level
flight. This can be determined by comparing pre-icing
condition entry speed to the indicated speed after a surface
and antenna deice cycle is completed.
4. Any variations from normal indicated airspeed
between the pilot’s and copilot’s airspeed indicators.
5-8.2
Change 3
TM 55-1510-219-10
5-31B. ICING LIMITATIONS (SEVERE).
WARNING
Severe icing may result from environmental
conditions outside of those for which the
airplane is certificated. Flight in freezing
rain, freezing drizzle, or mixed icing
conditions (supercooled liquid water and ice
crystals) may result is a build-up on
protective surfaces exceeding the capability of
the ice protection system, or may result in ice
forming aft of these protected surfaces. This
ice may not shed using ice protection systems,
and may seriously degrade the performance
and controllability of the airplane.
5-32. WIND LIMITATIONS.
The maximum demonstrated crosswind for landing is
25 KTS. Wind limitations are described in Chapter 7.
5-33. OXYGEN REQUIREMENTS.
For oxygen requirements, see AR 95-1.
5-34. CABIN PRESSURE LIMITS.
Maximum cabin differential pressure is 6.1 PSI. If a
crack in either the cabin window or windshield should
occur, refer to chapter 9.
a. During flight, severe icing conditions that exceed
those for which the airplane is certificated shall be
determined by the following visual cues. If one or more of
these visual cues exists, immediately request priority
handling from air traffic control to facilitate a route or an
altitude change to exit the icing conditions:
(1) Unusually extensive ice accreted on the
airframe in areas not normally observed to collect ice.
(2) Accumulation of ice on the upper (or lower, as
appropriate) surface of the wing aft of the protected area.
(3) Accumulation of ice on the propeller spinner
farther aft than normally observed.
b. Since the autopilot may mask tactile cues that
indicate adverse changes in handling characteristics, use of
the autopilot is prohibited when any of the visual cues
specified above exist, or when unusual lateral trim
requirements or autopilot trim warnings are encountered
while the airplane is in icing conditions.
NOTE
All icing detection lights must be operative prior
to flight into icing conditions at night. This
supersedes any relief provided by the master
minimum equipment list (MMEL) or equivalent.
Change 3
5-9
TM 55-1510-219-10
Section VIII. OTHER LIMITATIONS
5-35. INSTRUMENT LANDING SYSTEM LIMITS.
ing qualities of the aircraft, however ice
accumulation does affect the data link operation. Therefore radome anti-ice should
be used only in conjunction with the mission equipment.
During ILS approach do not operate the propellers in the 1750 to 1850 RPM range.
5-36. FERRY CHAIR.
A ferry chair may be installed in the cabin area
for use on ferry missions. The seat may be installed
in the forward or the aft facing positions. The side facing
lavatory is limited to 170 pounds.
5-37. INTENTIONAL ENGINE CUT SPEED.
1.
Radome anti-ice shall be off during takeoff,
landing, and any single engine operations.
2.
Maintain 85% N 1 or above during simultaneous operation of the radome anti-ice,
brake deice, and surface deice systems. If adequate pneumatic pressure cannot be provided for simultaneous operation of these
systems, turn off the radome anti-ice and
brake deice systems.
Inflight engine cuts below the safe one-engine inoperative speed (V sse - 104 KIAS) are prohibited.
5-39. CABIN DOOR.
5-38. RADOME ANTI-ICE OPERATION.
The cabin door is weight limited to 300 pounds
to prevent possible structural damage.
The following limitations apply to operation of
the radome anti-ice system:
5-40. MAXIMUM DESIGN SINK RATE.
NOTE
Ice accumulation on the forward data link
randome does not adversely affect the fly-
The maximum design sink rate below 13,500
pounds gross weight is 600 feet per minute. The
maximum design sink rate above 13,500 pounds
gross weight is 500 feet per minute.
Section IX. REQUIRED EQUIPMENT FOR VARIOUS CONDITIONS OF FLIGHT
5-41. REQUIRED EQUIPMENT LISTING
a. A Required Equipment for Various Conditions of Right (table 5-3) is provided to enable the
pilot to identify those systems/components required
for flight. For the sake of brevity, the listing does
not include obviously required items such as wings,
rudder, flaps, engines, landing gear, etc. The list also
does not include items which do not affect the airworthiness of the aircraft such as galley equipment,
entertainment systems, passenger convenience
items, etc. It is, however, important to note the ALL
ITEMS WHICH ARE RELATED TO THAT AIRWORTHINESS OF THE AIRCRAFT AND NOT
INCLUDED ON THE LIST ARE AUTOMATICALLY REQUIRED TO BE OPERATIVE
b. It is the final responsibility of the pilot to
determine whether the lack or inoperative status of
a piece of equipment on the aircraft will limit the
conditions under which the aircraft may be operated.
5-10
Change 2
(-)Indicates item may be inoperative for the
specified Right condition.
(*)Refers to remarks and/or exceptions column
for explicit information or reference.
Numbered items contained in ( ) are required
for flights by AR 95-1.
c. The pilot is responsible for exercising the
necessary operational control to assure that no aircraft is flown with multiple items inoperative, without first determining that any interface or interrelationship between inoperative systems or components
will not result in a degradation in the level of safety
and/or cause an undue increase in crew workload.
d. The exposure to additional failures during
continued operation with inoperative systems or
components must also be considered in determining
that an acceptable level of safety is being maintained. The REL may not deviate from requirements
of the operators manual limitations section emergency procedures or safety of flight messages.
TM 55-1510-219-10
Table 5-3. Required Equipment Listing
SYSTEM
and/or
COMPONENT
Number of items installed
VFR Day
VFR Night
IFR Day
IFR Night
Icing Conditions
REMARKS and/or Exceptions
AIR CONDITIONING
Bleed Air Fail Light
2
-
-
1
1
2
Pressurization Controller
Safety Valve
Outflow Valve
Altitude Warning
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
Cabin Rate of Climb
Differential Pressure/Cabin Altitude
Pressurization Air Source
Duct Overtemp Light
COMMUNICATIONS
Interphone System
VHF Communications Systems
Static Discharge Wicks
1
1
2
1
1
1
1
-
1
1
1
-
1
1
1
-
1
1
1
-
1
1
1
-
1
2
24
-
-
- - - - 24 24 24
ELECTRICAL POWER
Battery
Battery Charge Light
DC Generator
DC Loadmeter
1
1
2
2
1
1
1
2
1
1
1
2
1
1
2
2
1
1
2
2
1
1
2
2
DC Generator Caution Light
2
2
2
2
2
2
Inverter
Inverter Warning Light
2
1
1
-
1
-
2
1
2
1
2
1
AC Frequency/Voltmeter
EQUIPMENT FURNISHINGS
Seat Belts and Shoulder Harnesses;
Pilot and Co-Pilot
Emergency Locator transmitter
2
2
2
2
2
2
2
1
*
-
*
-
*
-
*
-
*
-
5-11
Provided bleed air is not used from side of failed
light.
May be inoperative provided airplane remains
unpressurized.
May be inoperative provided bleed air is not used.
Minimum required - one wick at the outboard end
of each control surface plus top of vertical
stabilizer
One may be inoperative provided corresponding
generator caution light Is monitored. One may be
inoperative provided corresponding loadmeter is
monitored.
May be inoperative provided both inverters
are operative.
* One per installed seat.
TM 55-1510-219-10
Table 5-3. Required Equipment Listing
SYSTEM
and/or
COMPONENT
FIRE PROTECTION
Fire Detector System
Engine Fire Extinguisher
Portable Fire Extinguisher
Number of items installed
VFR Day
VFR Night
IFR Day
IFR Night
Icing Conditions
REMARKS and/or Exceptions
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
FLIGHT CONTROLS
Trim Tab Indicators - Rudder, Aileron,
and Elevator
3
3
3
3
3
3
Flap Position Indicator
1
1
1
1
1
1
Flap System
Rudder Boost
Yaw Damp
1
1
1
1
1
1
1
1
Stall Warning
Autopilot
1
1
1
-
1
-
1
-
1
-
1
-
FUEL EQUIPMENT
Standby Fuel Boost Pump
2
1
1
1
1
1
Engine Driven Boost Pump
Firewall Shutoff Valve
Fuel Quantity Indicator
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
Crossfeed Valve
1
-
-
-
-
-
Crossfeed Light
1
1
1
1
1
1
5-12
May be inoperative provided that the
tabs are visually checked in the neutral
position prior to each takeoff and
checked for full range of operation.
May be inoperative provided that the
flap travel is visually inspected prior to
takeoff.
May be inoperative for flight at and be
low 17,000 feet.
Both required for operation on aviation
gasoline above 20,000 feet.
One may be inoperative provided other
side is operational and amount of fuel
on board can be established to be adequate
for intended flight. Fuel flow on
affected side must be operational and
monitored.
Required for (1) operation with aviation
gasoline above 20,000 feet; 12) when
operating with aviation kerosene when
one standby boost pump is Inoperative.
If takeoff with inoperative crossfeed is
planned, mission should be limited to
that range attainable with single engine
operation, one engine supplying fuel.
May be inoperative provided proper operation
of crossfeed system is checked
prior to takeoff. Both fuel pressure
lights must be operative.
TM 55-1510-219-10
Table 5-3. Required Equipment Listing
SYSTEM
and/or
COMPONENT
Number of items installed
VFR Day
VFR Night
IFR Day
IFR Night
Icing Conditions
REMARKS and/or Exceptions
Fuel Flow Indicator
2
2
2
2
2
2
Fuel Pressure Warning Light
2
2
2
2
2
2
Motive Flow Valve
Jet Transfer Pump
Fuel Quantity Gage Selector Switch
2
2
1
1
1
1
1
1
1
-
-
-
-
1
1
2
4
2
4
2
4
2
4
2
4
1
2
4
Windshield Heat, Left and Right
Windshield Wiper
Auto Ignition System and Lights
Pitot Heater
Alternate Static Air Source
Propeller Deice System (Auto)
Propeller Deice System (Manual)
Heated Fuel Vent
Stall Warning Heater
Brake Deicer System
2
2
2
2
1
1
1
2
1
1
2
1
-
2
1
-
2
1
1
-
2
1
1
-
1
2
1
1
1
1
2
1
-
LANDING GEAR
Landing Gear Motor
1
1
1
1
1
1
Landing Gear Position Indicator Lights
3
3
3
3
3
3
Gear Handle Lights
Landing Gear Aural Warning
2
1
2
1
2
1
2
1
2
1
2
1
ICE AND RAIN PROTECTION
Airfoil Deice System (Wing and Horizontal Stabilizer)
Antenna Deice System
Engine Inertial Ice Vanes
Ice Vane Lights
5-13
One may be inoperative provided fuel
quantity gages are operative.
One may be inoperative provided
standby boost pump operation is ascertained
using opposite light with cross
feed prior to engine start. Standby
boost pump on side of failed light must
be operated in flight to assure fuel
pressure, should the engine driven
boost pump fail.
Required if aux tanks contain fuel.
Required if aux tanks contains fuel.
May be inoperative provided MAIN
quantity indicators are operational.
May be inoperative provided manual ice
vane controls are operational and used.
Right side may be inoperative.
Right side may be inoperative.
May be inoperative provided operations
are continued only to a point where re
pairs can be accomplished.
One of three may be inoperative provided
gear handle light is monitored.
TM 55-1510-219-10
Table 5-3. Required Equipment Listing
SYSTEM
and/or
COMPONENT
Number of items installed
VFR Day
VFR Night
IFR Day
IFR Night
Icing Conditions
REMARKS and/or Exceptions
LIGHTS
Cockpit and Instrument Lights
*
-
*
-
*
-
Landing and Taxi Light
Strobe Beacon
Position Lights
Wing Ice Lights
Master Fault Warning Light
Master Fault Caution Lights
Cabin Door Caution Light
2
2
3
2
2
2
1
*
2
3
*
*
2
3
*
*
*
NAVIGATION INSTRUMENTS
Altimeter
Airspeed Indicator
Vertical Speed Indicator
Standby Magnetic Compass
Horizon Indicator
Outside Air Temperature
Turn and Bank Indicator
Directional Gyro
Clock
Transponder
Distance Measuring Equipment
Navigation Equipment
2
2
2
1
2
1
2
2
2
1
1
*
1
1
1
1
-
1
1
1
1
-
1
1
1
2
1
1
1
1
1
*
1
1
1
2
1
1
1
1
1
*
1
1
1
2
1
1
1
1
*
OXYGEN
Oxygen System
Oxygen Mask
1
*
1
-
1
-
1
-
1
-
1
-
PROPELLERS
Propeller Overspeed Governor
Propeller Governor Test Switch
Autofeathering System
Autofeathering Armed Light
Reverse Not Ready Light
Propeller Synchrophaser
2
2
1
2
1
1
2
2
1
-
2
2
1
-
2
2
1
-
2
2
1
-
2
2
1
-
5-14
*Lights must illuminate all instruments
and controls.
Per FAR 91.33
*One required for night Icing flight.
*May be Inoperative provided visual Indicators
are checked prior to each
takeoff.
Right side may be inoperative
Right side may be inoperative.
Right side may be inoperative.
Right side may be inoperative.
Right side may be inoperative.
*Per AR-95-1
*Refer to oxygen requirements in Section VI.
TM 55-1510-219-10
Table 5-3. Required Equipment Listing
SYSTEM
and/or
COMPONENT
Number of items installed
VFR Day
VFR Night
IFR Day
IFR Night
Icing Conditions
REMARKS and/or Exceptions
ENGINE INDICATING INSTRUMENTS
Propeller Tachometer indicator
Propeller Synchroscope
Gas Generator Tachometer Indicator
TOT Indicator
Torque Indicator
2
1
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
ENGINE OIL INDICATORS
Oil Pressure Indicator
Oil Temperature Indicator
Chip Detector Light
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
BT00226
5-15/(5-16 blank)
TM 55-1510-219-10
CHAPTER 6
WEIGHT / BALANCE AND LOADING
Section I. GENERAL
aircraft forms and records are contained in AR 95-3, TM
55-1500-342-23, and DA PAM 738-751.
Sufficient data has been provided so that, knowing the
6-3. AIRCRAFT COMPARTMENT AND STATIONS.
basic weight and moment of the aircraft, any
combination of weight and balance can be computed.
The aircraft is separated into two compartments
associated with loading. These compartments are the
6-2. CLASS.
cockpit and the cabin. Figure 6-1 shows the general
description of aircraft compartments.
Army RC-12D aircraft are in Class 1.
Additional
directives governing weight and balance of Class 1
6-1.
EXTENT OF COVERAGE.
6-1
TM 55-1510-219-10
Figure 6-1. Aircraft Compartments and Stations
6-2
TM 55-1510-219-10
Section II. WEIGHT AND BALANCE
and moment/100 entry is considered the current weight
PURPOSE.
and balance status of the basic aircraft.
The data to be inserted on weight and balance
charts and forms are applicable only to the individual
6-8. WEIGHT AND BALANCE CLEARANCE FORM
aircraft, the serial number of which appears on the title
page of the booklet entitled WEIGHT AND BALANCE
F, DD FORM 365-4 (TACTICAL).
DATA supplied by the aircraft manufacturer and on the
various forms and charts which remain with the aircraft.
Refer to TM 55-1500-342-23 for Form F
The charts and forms referred to in this chapter may
instructions. Refer to tables 6-1, 6-2, and 6-3 for
differ in nomenclature and arrangement from time to
moments.
time, but the principle on which they are based will not
change.
NOTE
The
maximum
baggage
6-5. CHARTS AND FORMS.
compartment weight is 410
pounds. Do not exceed the 100
The standard system of weight and balance
Lbs/Sq. Ft floor loading limit.
control requires the use of several different charts and
forms. Within this chapter, the following are used:
Table 6-1. Baggage Moment
6-4.
a. Chart C Basic Weight and Balance Record, DD
Form 365-3.
b. Form F Weight and Balance Clearance Form F,
DD Form 365-4 (Tactical).
6-6.
RESPONSIBILITY.
The aircraft manufacturer inserts all aircraft
identifying data on the title page of the booklet entitled
WEIGHT AND BALANCE DATA and on the various
charts and forms. All charts, including one sample
Weight and Balance Clearance Form F, if applicable,
are completed at time of delivery. This record is the
basic weight and balance data of the aircraft at delivery.
All subsequent changes in weight and balance are
compiled by the weight and balance technician.
6-7. CHART C BASIC WEIGHT AND BALANCE
RECORD, DD FORM 365-3.
Chart C is a continuous history of the basic weight
and moment resulting from structural and equipment
changes made in service. At all times, the last weight
6-3
TM 55-1510-219-10
Table 6-2. Fuel Moments
6-4
TM 55-1510-219-10
Table 6-3. Center of Gravity Moment Table - Moment / 100
6-5
TM 55-1510-219-10
Table 6-3. Center of Gravity Moment Table - Moment / 100 (Continued)
6-6
TM 55-1510-219-10
Table 6-3. Center of Gravity Moment Table - Moment/100 (continued)
6-7
TM 55-1510-219-10
Section III. FUEL/OIL
6-10. FUEL AND OIL DATA.
6-9. FUEL LOAD.
a. Fuel Moment Table. Table 6-2 shows fuel
moment/100 given US gallons or pounds for JP-4 and
JP-5.
Fuel loading imposes a restriction on the amount
of load which can be carried. The required fuel must
first be determined, then that weight subtracted from the
total weight of passengers, baggage and fuel. Weight
up to and including the remaining allowable capacity can
be subtracted directly from the weight of passengers,
baggage and fuel. As the fuel load is increased, the
loading capacity is reduced.
b. Oil Data. Total oil weight is 62 pounds and is
included in the basic weight of the aircraft. Servicing
information
is
provided
in
Chapter
2.
Section IV. CENTER OF GRAVITY
6-11. CENTER OF GRAVITY LIMITATIONS.
Center of gravity limitations are
expressed in ARM inches which
refers to a positive measurement
from the aircraft's reference datum.
The forward CG limit at 11,279 Lbs.
or less is 181.0 ARM inches. At
14,200 Lbs. or less, the aft CG limit
is 196.4 ARM inches. Tables 6-4 and
6-5 provide forward and aft CG
limitations.
WARNING
The forward Center of Gravity limit
may be exceeded when the mission
gear is removed.
Table 6-4. Center of Gravity Limits (Landing Gear Down) Below 12,500 Lbs.
Table 6-5. Center of Gravity Limits (Landing Gear Down)
Section V. CARGO LOADING
b. Floorboard structural capacity shall be
considered in the loading of heavy or sharp-edged
containers and equipment. Shorings shall be used to
The basic factors to be considered in any loading
distribute highly condensed weights evenly over the
situation are as follows:
cargo areas.
a. Cargo shall be arranged to permit access to all
emergency equipment and exits during flight.
6-12. LOAD PLANNING.
6-8
TM 55-1510-219-10
c. All cargo shall be adequately secured to prevent
damage to the aircraft, other cargo, or the item itself.
6-13. LOADING PROCEDURE.
Loading of cargo is accomplished through the cabin
door (21.5 in. X 50.0 in.) or the cargo door (52.0 in. X
52.0 in).
6-14. SECURING LOADS.
All cargo shall be secured with restraints strong
enough to withstand the maximum force exerted in any
direction. The maximum force can be determined by
multiplying the weight of the cargo item by the
applicable load factor. These established load factors
(the ratio between the total force and the weight of the
cargo item) are 1.5 to the side and rear, 3.0 up, 6.6
down, and 9.0 forward.
6-9/(6-10 blank)
TM 55-1510-219-10
CHAPTER 7
PERFORMANCE
Section I. INTRODUCTION TO PERFORMANCE
7-1. INTRODUCTION
Cruise Altitude:
22,000 feet
The graphs in this section present performance
information for takeoff, climb, cruise, and landing at
various parameters of weight, altitude, and temperature.
1 Source: NOAA Standard Instrument Departures for
Western United States, 9 Jun 1983.
2 Source: NOAA Enroute High Altitude -U.S.
Chart H-2, 9 Jun 1983.
3 MEA on NOAA Enroute Low Altitude -U.S.
Chart L-8, 9 Jun 1983.
4 Includes distance between airport and VORTAC, per
Source in Footnote 1.
The following example presents calculations for a
proposed flight from Denver to Reno using the
conditions listed below:
7-2. CONDITIONS.
At Stapleton International (DEN):
Free Air Temperature .......................... 28°C (82°F)
Field Elevation ....................................... 15333 feet
Altimeter Setting ..................................30.25 in. Hg
Wind ............................................. 210°at 13 knots
Runway 35R Length ........................... 112,000 feet
At Cannon International (RNO): Free
Air Temperature .................................. 32°C (90°F)
Field Elevation ....................................... 14412 feet
Altimeter Setting .........................129.83 in. Hg feet
Wind ............................................ 260°at 10 knots
Runway 25 Length ................................ 16101 feet
Route segment of trip2:
DEN J116 EKR J173 SLC J154
BAM J32 RNO
7-3. PRESSURE ALTITUDE.
To determine the approximate pressure altitude at
origin and destination airports, add 1000 feet to field
elevation for each 1.00 in. Hg that the reported
altimeter setting value is below 29.92 in. Hg, and
subtract 1000 feet for each 1.00 in. Hg above 29.92 in.
Hg. Always subtract the reported altimeter setting from
29.92 in. Hg, then multiply the answer by 1000 to find
the difference in feet between field elevation and
pressure altitude.
Pressure Altitude at DEN:
29.92 in. Hg 30.25 in. Hg = 0.33
0.33 x 1000 feet = 330 feet
The pressure altitude at DEN is 330 feet below field
elevation.
Pressure altitude at DEN = 5333 -330 = 5003 feet
Route Segment Data 2
7-1
TM 55-1510-219-10
At Rotation ...........................................99 knots
At 50 feet ...........................................116 knots
Pressure altitude at RNO:
29.92 in.
Hg29.83 in. Hg = 0.09
0.09 x 1000 feet = +90 feet
The pressure altitude at RNO is 90 feet above field
elevation.
7-8.ACCELERATE-GO FLIGHT PATH EXAMPLE.
The following example assumes the aircraft is loaded
so that takeoff weight is 10,000 pounds.
Pressure altitude at RNO = 4412 +90 = 4502 feet.
7-4. TAKEOFF WEIGHT.
Maximum takeoff weight (from limitations section)
= 14,200 pounds.
a.
Accelerate-Go Distance Over 50-foot Obstacle
(Flaps 0%o).
Enter the graph at 28°C, 5003 feet
pressure altitude, 10,000 pounds, and 10 knots head wind
component:
7-5. TAKEOFF WEIGHT TO ACHIEVE POSITIVE
ONE-ENGINE-INOPERATIVE CLIMB AT LIFTOFF
(FLAPS 0%).
Total Distance Over 50-foot Obstacle.......... 5710 feet
Speed at Rotation (VR) ..........................94 knots
Speed at 35 Feet Above Runway (Climb
Speed) ...............................................106 knots
b.
Takeoff Climb Gradient One Engine Inoperative
(Flaps 0%o).
Enter the graph at 28°C, 5003 feet
pressure altitude, and 10,000 pounds:
Enter the graph at 5003 feet to 28°C, to determine
the maximum weight at which the accelerating
procedure should be attempted.
Maximum Accelerate-Go Weight
7-6.
12,990 pounds
Climb Gradient ..........................................5.5%
Climb Speed ......................................106 knots
ACCELERATE-STOP (FLAPS 0%).
A 5.5% climb gradient is 55 feet of vertical height per
1000 feet of horizontal distance.
Enter the accelerate-stop graph at 28°C, 5003 feet
pressure altitude, 12,000 pounds, and 10 knots head
wind component:
NOTE
The graphs for take-off climb gradient
assume a zero-wind condition. Climbing
into a head wind will result in higher
angles of climb, and hence better
obstacle clearance capabilities.
Accelerate-Stop Distance ................... 4520 feet
Takeoff Decision Speed .......................99 knots
7-7
TAKE-OFF DISTANCE (FLAPS 0%).
Enter the graph at 28°C, 5003 feet pressure altitude,
12,000 pounds, and 10 knots head wind component:
Ground Roll ........................................ 2860 feet
Total Distance over 50-foot
obstacle ............................................. 4400 feet
Take-off Speed:
Calculations of the horizontal distance to clear an
obstacle 100 feet above the runway surface: (fig.7-1)
Distance from 50 feet to 100 feet = 50 feet
(100 50) (1000 + 55) = 910 feet
Total Distance = 5710 +910 = 6620 feet
7-2
TM 55-1510-219-10
Figure 7-1. Accelerate-Go Flight Path
Enter the TIME, FUEL, AND DISTANCE TO
CLIMB Graph at 28°C to 5003 feet, and to 12,000
pounds, and enter at -2°C to 22,000 feet, and to 12, 000
pounds, and read:
7-9. FLIGHT PLANNING.
Calculations for flight time, block speed, and fuel
requirements for a proposed flight, are detailed below
using the same conditions presented in paragraph 7-2,
and a takeoff weight of 12,000 pounds:
Time to Climb . .......... 18.4 -2.2 = 16.2 = 16 minutes
Fuel Used to Climb ................ 248 -45 = 203 pounds
Distance Traveled ..............50 -7 = 43 nautical miles
Enter the TIME, FUEL, AND DISTANCE TO
DESCEND Graph at 22,000 feet, and enter again at
4502 feet, and read:
Time to Descend ........ 14.7 -3.1 =11.6 = 12 minutes
Fuel Used to Descend ........... 162 -39 = 123 pounds
Distance Traveled . .......... 67 -12 = 55 nautical miles
DEN
Pressure Altitude ...................................... 5003 feet
FAT ................................................................. 28°C
ISA Condition ..........................................ISA +23°C
DEN-SLC
Pressure Altitude ....................................22,000 feet
FAT ................................................................ -20°C
ISA Condition ..........................................ISA +27°C
SLC-RNO
Pressure Altitude ....................................22,000 feet
FAT ................................................................ -12°C
ISA Condition ........................................ISA + 1 7°C
RNO
Pressure Altitude ...................................... 4502 feet
FAT ................................................................. 32°C
ISA Condition ..........................................ISA +27°C
An estimated average cruise weight of 11,200 pounds
was used for this example.
Enter the tables for MAXIMUM ENDURANCE
POWER @ 1700 RPM for ISA + 10°C, ISA +20°C, and
ISA + 30°C, and read the cruise speeds for 22, 000 feet
at 12,000 pounds and 11,000 pounds:
Interpolate between these speeds for ISA +27°C
and ISA + 17°C at 11,200 pounds:
7-3
TM 55-1510-219-10
Cruise True Airspeeds At FI 220
Enter the MAXIMUM ENDURANCE POWER @
1700 RPM Tables for ISA +10°C, ISA +20°C, and ISA
+30°C at 12,000 pounds and 11,000 pounds and
interpolate the recommended torque settings for ISA +
27°C and ISA +17°C.
Fuel Flow Per Engine ..............................232 Lbs/hr
Total Fuel Flow .......................................464 Lbs/hr
NOTE
For flight planning, enter these
charts at the forecasted ISA
condition; for enroute power settings
and fuel flows, enter at the actual
indicated FAT.
ISA +27°C ............................50% torque per engine
ISA + 17°C ...........................49% torque per engine
Time and fuel used were calculated at MAXIMUM
ENDURANCE POWER @ 1700 RPM as follows:
Cruise True Airspeed (ISA +27°C) ............ 200 knots
Cruise True Airspeed (ISA +17°C) ............ 197 knots
Enter the MAXIMUM ENDURANCE POWER @
1700 RPm, Tables for ISA + 10°C, ISA +20°C, and ISA
+30°C at 12,000 pounds and 11,000 pounds at 22,000
feet, and interpolate the fuel flows for ISA +27°C and
ISA + 17°C at 11,200 pounds.
ISA +27°C
Fuel Flow Per Engine ..............................237 Lbs/hr
Total Fuel Flow .......................................474 Lbs/hr
ISA +17°C
Results are as follows:
TIME - FUEL - DISTANCE
7-4
TM 55-1510-219-10
7-10. RESERVE FUEL.
Reserve Fuel is calculated for 45 minutes at Maximum
Range Power @ 1700 RPM. Use planned cruise
altitude (22,000 feet), forecasted ISA condition (ISA +
17°C), and estimated weight at end of planned trip
(10,091 pounds). (Since the lowest weight column in
the tables is 11,000 pounds, assume weight at the end
of the planned trip to be 11, 000 pounds, and use that
fuel flow value for this example.)
Enter the tables for MAXIMUM RANGE POWER
@ 1700 RPM for ISA + 10°C and ISA +20°C at 11,000
Lbs and 22,000 feet, and read the total fuel flows:
Zero Fuel Weight = (12,090) (2359) = 9731
pounds
Maximum zero fuel weight limitation (from
Chapter 5) = 11,500 pounds.
Maximum Zero Fuel Weight Limitation has not
been exceeded.
Anytime the Zero Fuel Weight exceeds the
Maximum Zero Fuel Weight Limit, the excess must be
off-loaded from PAYLOAD. If desired, additional FUEL
ONLY may then be added until the ramp weight equals
the Maximum Ramp Weight Limit of 14,290 Lbs.
ISA +10°C ...............................................468 Lbs/hr
ISA +20°C................................................484 Lbs/hr
Then interpolate for the fuel flow at ISA + 17°C as
follows: Change in Fuel Flow = 484 468 = 16
Lbs/hr.
Change in Temperature = (ISA +20°C) (ISA + 10
C) = 10°C.
Rate of Change in Fuel Flow = Change in Fuel
Flow + Change in Temperature.
Rate of Change in Fuel Flow = (16 Lbs/hr) +
(10°C).
Rate of Change in Fuel Flow = 1.6 Lbs/hr increase
per 1°C increase.
Temperature increase from ISA + 10°C to ISA
+17°C = 7°C.
Total Change in Fuel Flow = 7 X 1.6 Lbs/hr = 11.2
Lbs/hr.
Total Fuel Flow = (ISA +10°C Fuel Flow) + (Total
Change in Fuel Flow).
Total Fuel Flow = (468) + (11.2) = 479.2 Lbs/ hr.
Reserve Fuel = 45 minutes X Total Fuel Flow.
Reserve Fuel = (0.75) X (479.2 Lbs/hr) = 359.4 =
360 lbs/hr.
Total Fuel Requirement = 1999 +360 = 2359
pounds.
7-11. ZERO FUEL WEIGHT LIMITATION.
For this example, the following conditions were
assumed:
Ramp Weight ................................... 12,090 pounds
Weight of Usable Fuel Onboard ..........2359 Pounds
7-12. LANDING INFORMATION.
The estimated Landing Weight is determined by
subtracting the fuel required for the trip from the Ramp
Weight:
Ramp Weight.......................................... 12,090 Lbs
Fuel Required for Total Trip ................ 1999 pounds
Landing Weight (12,090 - 1999) . ...... 10,091 pounds
Enter the NORMAL LANDING DISTANCE
WITHOUT PROPELLER REVERSING FLAPS 100%
Graph at 32°C, 4502 feet, 10,091 pounds, and 10 knots
head wind component:
Ground Roll............................................... 1725 feet
Total Distance Over 50-foot Roll .............. 1725 feet
Total Distance Over 50-foot Obstacle ....... 3060 feet
Approach Speed ...................................... 100 knots
Enter the CLIMB - BALKED LANDING Graph at
32°C, 4502 feet, and 10,091 pounds:
Rate of Climb .........................................1550 ft/min
Climb Gradient .............................................. 12.5%
7-13. COMMENTS PERTINENT TO THE USE OF
PERFORMANCE GRAPHS.
a.
In addition to presenting the answer for a
particular set of conditions, the example on a graph also
presents the order in which the various scales on the
graph should be used. For instance, if the first item in
the example is FAT, then enter the graph at the existing
FAT.
b.
The reference lines indicate where to begin
following the guidelines. Always project to the reference
line first, then follow the guidelines to the next known
item by maintaining the same PROPORTIONAL
Zero Fuel Weight = Ramp Weight - Weight of
Usable Fuel Onboard
7-5
TM 55-1510-219-10
DISTANCE between the guide line above and the guide
line below the projected line. For instance, if the
projected line intersects the reference line in the ratio of
30% down/70% up between the guidelines, then
maintain this same 30%/70% relationship between the
guide lines and follow them to the next known item.
d.
The full amount of usable fuel is available
for all approved flight conditions.
e.
Notes have been provided on various
graphs and tables to approximate performance with ice
vanes extended. The effect will vary, depending upon
airspeed, temperature, altitude, and ambient conditions.
At lower altitudes, where operation on the torque limit is
possible, the effect of ice vane extension will be less,
depending upon how much power can be recovered
after the ice vanes have been extended.
c.
The associated conditions define the
specific conditions from which performance parameters
have been determined. They are not intended to be
used as instructions; however, performance values
determined from charts can only be achieved if the
specified conditions exist.
7-6
TM 55-1510-219-10
Index of Figures
Accelerate Go Flight Path ................................................................................................................................7-3
Airspeed Calibration - Normal System ............................................................................................................7-10
Airspeed Calibration - Alternate System .........................................................................................................7-11
Altimeter Correction - Normal System ............................................................................................................7-12
Altimeter Correction - Alternate System .........................................................................................................7-13
Indicated Outside Air Temperature Correction ................................................................................................7-14
ISA Conversion ..............................................................................................................................................7-15
Fahrenheit to Celsius Temperature Conversion ..............................................................................................7-16
Take-Off Weight - Flaps 0% - To Achieve Positive One-Engine Inoperative Climb at Lift-Off .........................7-17
Take-Off Weight - Flaps 40% - To Achieve Positive One-Engine Inoperative Climb at Lift-Off .......................7-18
Wind Components .........................................................................................................................................7-19
Take-Off Distance - Flaps 0% ........................................................................................................................7-20
Take-Off Distance - Flaps 40% ......................................................................................................................7-21
Minimum Take-Off Power at 2000 RPM .........................................................................................................7-22
Accelerate-Stop - Flaps 0%.............................................................................................................................7-23
Accelerate-Stop - Flaps 40% ..........................................................................................................................7-24
Accelerate-Go Distance Over 50-Ft Obstacle - Flaps 0% ...............................................................................7-25
Accelerate-Go Distance Over 50-Ft Obstacle - Flaps 40% .............................................................................7-26
Time, Fuel, and Distance to Cruise Climb ......................................................................................................7-27
Climb Two-Engines Flaps 0%..........................................................................................................................7-28
Climb Two-Engines Flaps 40%........................................................................................................................7-29
Take-Off Climb Gradient - One-Engine Inoperative - Flaps 0% ......................................................................7-30
Take-Off Climb Gradient - One-Engine Inoperative - Flaps 40% ....................................................................7-31
Climb One-Engine Inoperative .......................................................................................................................7-32
Climb - Balked Landing ...................................................................................................................................7-33
Service Ceiling One-Engine Inoperative..........................................................................................................7-34
Maximum Cruise Power @ 1900 RPM, ISA -30 .................................................................................... 7-35, 7-36
Maximum Cruise Power @ 1900 RPM, ISA -20 . ................................................................................... 7-37, 7-38
Maximum Cruise Power @ 1900 RPM, ISA -10° ................................................................................... 7-39, 7-40
Maximum Cruise Power @ 1900 RPM, ISA .......................................................................................... 7-41, 7-42
Maximum Cruise Power @ 1900 RPM, ISA +10° .................................................................................. 7-43, 7-44
Maximum Cruise Power @ 1900 RPM, ISA +20°. ................................................................................. 7-45, 7-46
Maximum Cruise Power @ 1900 RPM, ISA +30° .................................................................................. 7-47, 7-48
Maximum Cruise Power @ 1900 RPM, ISA +37° .................................................................................. 7-49, 7-50
Maximum Cruise Power @ 1900 RPM ...........................................................................................................7-51
Maximum Cruise Speeds @ 1900 RPM .........................................................................................................7-52
Fuel Flow At Maximum Cruise Power @ 1900 RPM, ......................................................................................7-53
7-7
TM 55-1510-219-10
Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA -30°C ............................................ 7-54,7-55
Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA -20°C ............................................ 7-56,7-57
Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA -10°C ............................................ 7-58,7-59
Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA ...................................................... 7-60,7-61
Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA +10°C ........................................... 7-62,7-63
Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA +20°C ........................................... 7-64,7-65
Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA +30°C ........................................... 7-66,7-67
Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA +37°C ........................................... 7-68,7-69
Maximum Cruise Power @ 1700 RPM, ISA -30°C .................................................................................. 7-70,7-71
Maximum Cruise Power @ 1700 RPM, ISA -20°C .................................................................................. 7-72,7-73
Maximum Cruise Power @ 1700 RPM, ISA -10°C .................................................................................. 7-74,7-75
Maximum Cruise Power @ 1700 RPM, ISA ........................................................................................... 7-76.7-77
Maximum Cruise Power @ 1700 RPM, ISA +10°C ................................................................................. 7-78,7-79
Maximum Cruise Power @ 1700 RPM, ISA +20°C ................................................................................. 7-80,7-81
Maximum Cruise Power @ 1700 RPM, ISA +30°C ................................................................................. 7-82,7-83
Maximum Cruise Power @ 1700 RPM, ISA +37°C ................................................................................. 7-84,7-85
Maximum Cruise Power @ 1700 RPM ...........................................................................................................7-86
Maximum Range Power @ 1700 RPM, ISA -30°C ................................................................................. 7-87,7-88
Maximum Range Power @ 1700 RPM, ISA -20°C ................................................................................. 7-89,7-90
Maximum Range Power @ 1700 RPM, ISA -10°C ................................................................................. 7-91,7-92
Maximum Range Power @ 1700 RPM, ISA ............................................................................................ 7-93,7-94
Maximum Range Power @ 1700 RPM, ISA +10°C ................................................................................ 7-95,7-96
Maximum Range Power @ 1700 RPM, ISA +20°C ................................................................................ 7-97,7-98
Maximum Range Power @ 1700 RPM, ISA +30°C .............................................................................. 7-99,7-100
Maximum Range Power @ 1700 RPM, ISA +37°C ............................................................................ 7-101,7-102
Maximum Endurance Power @ 1700 RPM, ISA -30°C ....................................................................... 7-103,7-104
Maximum Endurance Power @ 1700 RPM, ISA -20°C ....................................................................... 7-105,7-106
Maximum Endurance Power @ 1700 RPM, ISA -10°C ....................................................................... 7-107,7-108
Maximum Endurance Power @ 1700 RPM, ISA ................................................................................. 7-109,7-110
Maximum Endurance Power @ 1700 RPM, ISA + 10°C ..................................................................... 7-111,7-112
Maximum Endurance Power @ 1700 RPM, ISA + 20°C ..................................................................... 7-113,7-114
Maximum Endurance Power @ 1700 RPM, ISA +30°C ...................................................................... 7-115,7-116
Maximum Endurance Power @ 1700 RPM, ISA +37°C ...................................................................... 7-117,7-118
Range Profile - Maximum Cruise Power @ 1900 RPM ................................................................................. 7-119
Range Profile - Maximum Range Power @ 1700 RPM ................................................................................. 7-120
Range Profile - 542 Gallons Usable Fuel....................................................................................................... 7-121
Endurance Profile - 542 Gallons Usable Fuel ............................................................................................... 7-122
Time, Fuel, And Distance to Descend .......................................................................................................... 7-123
7-8
TM 55-1510-219-10
Landing Distance Without Propeller Reversing - Flaps 0% ........................................................................... 7-124
Landing Distance Without Propeller Reversing - Flaps 100% ...................................................................... 7 -125
Landing Distance With Propeller Reversing - Flaps 0% ................................................................................ 7-126
Landing Distance With Propeller Reversing - Flaps 100% ............................................................................ 7-127
Stopping Distance Factors ........................................................................................................................... 7-128
7-9
TM 55-1510-219-10
Figure 7-2. Airspeed Calibration - Normal System
7-10
TM 55-1510-219-10
Figure 7-3. Airspeed Calibration - Alternate System
7-11
TM 55-1510-219-10
Figure 7-4. Altimeter Correction - Normal System
7-12
TM 55-1510-219-10
Figure 7-5. Altimeter Correction - Alternate System
7-13
TM 55-1510-219-10
Figure 7-6. Indicated Outside Air temperature Correction
7-14
TM 55-1510-219-10
Figure 7-7. ISA Conversion
7-15
TM 55-1510-219-10
Figure 7-8. Fahrenheit to Celsius Temperature Conversion
7-16
TM 55-1510-219-10
Figure 7-9. Take-Off Weight - Flaps 0%, To Achieve Positive One-Engine Inoperative Climb at Lift-Off
7-17
TM 55-1510-219-10
Figure 7-10. Take-Off Weight - Flaps 40%, To Achieve Positive One-Engine Inoperative Climb at Lift-Off
7-18
TM 55-1510-219-10
Figure 7-11. Wind Components
7-19
TM 55-1510-219-10
Figure 7-12. Takeoff Distance - Flaps 0%
7-20
TM 55-1510-219-10
Figure 7-13. Take-Off Distance - Flaps 40%
7-21
TM 55-1510-219-10
MINIMUM TAKEOFF POWER AT 2000 RPM
(65 KNOTS)
NOTES:
1.
TORQUE INCREASES APPROXIMATELY 1% FROM 0 TO 65 KNOTS.
2.
THE PERCENT TORQUE INDICATED IN THIS FIGURE IS THE MINIMUM VALUE AT WHICH
TAKE-OFF PERFORMANCE PRESENTED IN THE SECTION CAN BE REALIZED. ANY EXCESS
POWER WHICH MAY BE DEVELOPED WITHOUT EXCEEDING ENGINE LIMITATIONS MAY BE
UTILIZED.
Figure 7-14. Minimum Take-Off Power at 2000, RPM
7-22
Figure 7-15. Accelerate-Stop - Flaps 0%
7-23
TM 55-1510-219-10
Figure 7-16. Accelerate-Stop - Flaps 40
7-24
TM 55-1510-219-10
Figure 7-17. Accelerate-Go Distance Over 50-Ft Obstacle - Flaps 0%
7-25
TM 55-1510-219-10
Figure 7-18. Accelerate-Go Distance Over 50-Ft Obstacle - Flaps 40%
7-26
TM 55-1510-219-10
Figure 7-19. Time, Fuel, and Distance to Cruise Climb
7-27
TM 55-1510-219-10
Figure 7-20. Climb - Two Engines - Flaps 0%
7-28
TM 55-1510-219-10
Figure 7-21. Climb - Two Engines - Flaps 40%
7-29
TM 55-1510-219-10
Figure 7-22. Take-Off Climb Gradient - One-Engine Inoperative - Flaps 0%
7-30
TM 55-1510-219-10
Figure 7-23. Take-Off Climb Gradient - One-Engine Inoperative - Flaps 40%
7-31
TM 55-1510-219-10
Figure 7-24. Climb - One Engine Inoperative
7-32
TM 55-1510-219-10
Figure 7-25. Climb - Balked Landing
7-33
TM 55-1510-219-10
Figure 7-26. Service Ceiling - One Engine Inoperative
TM 55-1510-219-10
Figure 7-27. Maximum Cruise Power 1900 RPM, ISA -30°C (Sheet 1 of 2)
7-35
TM 55-1510-219-10
Figure 7-27. Maximum Cruise Power 1900 RPM, ISA -30°C (Sheet 2 of 2)
7-36
TM 55-1510-219-10
Figure 7-28. Maximum Cruise Power 1900 RPM, ISA -20°C (Sheet 1 of 2)
7-37
TM 55-1510-219-10
Figure 7-28. Maximum Cruise Power 1900 RPM, ISA -20°C (Sheet 2 of 2)
7-38
TM 55-1510-219-10
Figure 7-29. Maximum Cruise Power 1900 RPM, ISA - 10°C (Sheet 1 of 2)
7-39
TM 55-1510-219-10
Figure 7-29. Maximum Cruise Power 1900 RPM, ISA -10°C (Sheet 2 of 2)
7-40
TM 55-1510-219-10
Figure 7-30. Maximum Cruise Power 1900 RPM, ISA (Sheet 1 of 2)
7-41
TM 55-1510-219-10
Figure 7-30. Maximum Cruise Power 1900 RPM, ISA (Sheet 2 of 2)
7-42
TM 55-1510-219-10
Figure 7-31. Maximum Cruise Power 1900 RPM, ISA + 10°C (Sheet 1 of 2)
7-43
TM 55-1510-219-10
Figure 7-31. Maximum Cruise Power 1900 RPM, ISA + 10°C (Sheet 2 of 2)
7-44
TM 55-1510-219-10
Figure 7-32. Maximum Cruise Power 1900 RPM, ISA +20°C (Sheet 1 of 2)
7-45
TM 55-1510-219-10
Figure 7-32. Maximum Cruise Power 1900 RPM, ISA +20°C (Sheet 2 of 2)
7-46
TM 55-1510-219-10
Figure 7-33. Maximum Cruise Power 1900 RPM, ISA +30°C (Sheet 1 of 2)
7-47
TM 55-1510-219-10
Figure 7-33. Maximum Cruise Power 1900 RPM, ISA +30°C (Sheet 2 of 2)
7-48
TM 55-1510-219-10
Figure 7-34. Maximum Cruise Power 1900 RPM, ISA +37°C (Sheet 1 of 2)
7-49
TM 55-1510-219-10
Figure 7-34. Maximum Cruise Power 1900 RPM, ISA +37°C (Sheet 2 of 2)
7-50
TM 55-1510-219-10
Figure 7-35. Maximum Cruise Speeds @ 1900 RPM
7-51
TM 55-1510-219-10
Figure 7-36. Maximum Cruise Power @ 1900 RPM
7-52
TM 55-1510-219-10
Figure 7-37. Fuel Flow at Maximum Cruise Power @ 1900 RPM
7-53
TM 55-1510-219-10
Figure 7-38. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA -30°C (Sheet 1 of 2)
7-54
TM 55-1510-219-10
Figure 7-38. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA -30°C (Sheet 2 of 2)
7-55
TM 55-1510-219-10
Figure 7-39. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA -20°C (Sheet 1 of 2)
7-56
TM 55-1510-219-10
Figure 7-39. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA -20°C (Sheet 2 of 2)
7-57
TM 55-1510-219-10
Figure 7-40. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA - -10°C (Sheet 1 of 2)
7-58
TM 55-1510-219-10
Figure 7-40. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA -10°C (Sheet 2 of 2)
7-59
TM 55-1510-219-10
Figure 7-41. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA (Sheet 1 of 2)
7-60
TM 55-1510-219-10
Figure 7-41. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA (Sheet 2 of 2)
7-61
TM 55-1510-219-10
Figure 7-42. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA + 10°C (Sheet 1 of 2)
7-62
TM 55-1510-219-10
Figure 7-42. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA + 10°C (Sheet 2 of 2)
7-63
TM 55-1510-219-10
Figure 7-43. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA +20°C (Sheet 1 of 2)
7-64
TM 55-1510-219-10
Figure 7-43. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA +20°C (Sheet 2 of 2)
7-65
TM 55-1510-219-10
Figure 7-44. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA +30°C (Sheet 1 of 2)
7-66
TM 55-1510-219-10
Figure 7-44. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA +30°C (Sheet 2 of 2)
7-67
TM 55-1510-219-10
Figure 7-45. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA +37°C (Sheet 1 of 2)
7-68
TM 55-1510-219-10
Figure 7-45. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA +37°C (Sheet 2 of 2)
7-69
TM 55-1510-219-10
Figure 7-46. Maximum Cruise Power 1700 RPM, ISA -30°C (Sheet 1 of 2)
7-70
TM 55-1510-219-10
Figure 7-46. Maximum Cruise Power 1700 RPM, ISA -30°C (Sheet 2 of 2)
7-71
TM 55-1510-219-10
Figure 7-47. Maximum Cruise Power 1700 RPM, ISA -20°C (Sheet 1 of 2)
7-72
TM 55-1510-219-10
Figure 7-47. Maximum Cruise Power 1700 RPM, ISA -20°C (Sheet 2 of 2)
7-73
TM 55-1510-219-10
Figure 7-48. Maximum Cruise Power 1700 RPM, ISA - 10°C (Sheet 1 of 2)
7-74
TM 55-1510-219-10
Figure 7-48. Maximum Cruise Power 1700 RPM, ISA - 10°C (Sheet 2 of 2)
7-75
TM 55-1510-219-10
Figure 7-49. Maximum Cruise Power 1700 RPM, ISA (Sheet 1 of 2)
7-76
TM 55-1510-219-10
Figure 7-49. Maximum Cruise Power 1700 RPM, ISA (Sheet 2 of 2)
7-77
TM 55-1510-219-10
Figure 7-50. Maximum Cruise Power 1700 RPM, ISA ± 10°C (Sheet 1 of 2)
7-78
TM 55-1510-219-10
Figure 7-50. Maximum Cruise Power 1700 RPM, ISA + 10°C (Sheet 2 of 2)
7-79
TM 55-1510-219-10
Figure 7-51. Maximum Cruise Power 1700 RPM, ISA +20°C (Sheet 1 of 2)
7-80
TM 55-1510-219-10
Figure 7-51. Maximum Cruise Power 1700 RPM, ISA +20°C (Sheet 2 of 2)
7-81
TM 55-1510-219-10
Figure 7-52. Maximum Cruise Power 1700 RPM, ISA +30°C (Sheet 1 of 2)
7-82
TM 55-1510-219-10
Figure 7-52. Maximum Cruise Power 1700 RPM, ISA +30°C (Sheet 2 of 2)
7-83
TM 55-1510-219-10
Figure 7-53. Maximum Cruise Power 1700 RPM, ISA +37°C (Sheet 1 of 2)
7-84
TM 55-1510-219-10
Figure 7-53. Maximum Cruise Power 1700 RPM, ISA +37°C (Sheet 2 of 2)
7-85
TM 55-1510-219-10
MAXIMUM CRUISE POWER
1700 RPM
WEIGHT: 12,000 LBS
NOTES:
1.
ISA DEVIATION LINES REFLECT ACTUAL TEMPERATURES FOR FLIGHT PLANNING. INDICATED
TEMPERATURES SHOULD BE USED FOR IN-FLIGHT CRUISE POWER SETTINGS.
2.
FOR OPERATION WITH ICE VANES EXTENDED, TORQUE WILL DECREASE APPROXIMATELY
20%. IF DESIRED, ORIGINAL POWER MAY BE RESET, PROVIDED ITT LIMIT IS NOT EXCEEDED.
Figure 7-54. Maximum Cruise Power @ 1700 RPM
7-86
TM 55-1510-219-10
Figure 7-55. Maximum Range Power @ 1700 RPM, ISA-30°C (Sheet 1 of 2)
7-87
TM 55-1510-219-10
Figure 7-55. Maximum Range Power @ 1700 RPM, ISA-30°C (Sheet 2 of 2)
7-88
TM 55-1510-219-10
Figure 7-56. Maximum Range Power @ 1700 RPM, ISA-200C (Sheet 1 of 2)
7-89
TM 55-1510-219-10
Figure 7-56. Maximum Range Power @ 1700 RPM, ISA-20°C (Sheet 2 of 2)
7-90
TM 55-1510-219-10
Figure 7-57. Maximum Range Power @ 1700 RPM, ISA-10°C (Sheet 1 of 2)
7-91
TM 55-1510-219-10
Figure 7-57. Maximum Range Power @ 1700 RPM, ISA- 10°C (Sheet 2 of 2)
7-92
TM 55-1510-219-10
Figure 7-58. Maximum Range Power @ 1700 RPM, ISA (Sheet 1 of 2)
7-93
TM 55-1510-219-10
Figure 7-58. Maximum Range Power @ 1700 RPM, ISA (Sheet 2 of 2)
7-94
TM 55-1510-219-10
Figure 7-59. Maximum Range Power @ 1700 RPM, ISA + 10°C (Sheet 1 of 2)
7-95
TM
55-1510-219-10
Figure 7-59. Maximum Range Power @ 1700 RPM, ISA + 10°C (Sheet 2 of 2)
7-96
TM 55-1510219-10
Figure 7-60. Maximum Range Power @ 1700 RPM, ISA +20°C (Sheet 1 of 2)
7-97
TM 55-1510-219-10
Figure 7-60. Maximum Range Power @ 1700 RPM, ISA +20°C (Sheet 2 of 2)
7-98
TM 55-1510-219-10
Figure 7-61. Maximum Range Power @ 1700 RPM, ISA+30°C (Sheet 1 of 2)
7-99
TM 55-1510-219-10
Figure 7-61. Maximum Range Power @ 1700 RPM, ISA +30°C (Sheet 2 of 2)
7-100
TM 55-1510-219-10
Figure 7-62. Maximum Range Power @ 1700 RPM, ISA +37°C (Sheet 1 of 2)
7-101
TM 55-1510-219-10
Figure 7-62. Maximum Range Power @ 1700 RPM, ISA +37°C (Sheet 2 of 2)
7-102
TM 55-1510-219-10
Figure 7-63. Maximum Endurance Power @ 1700 RPM, ISA-30°C (Sheet 1 of 2)
7-103
TM 55-1510-219-10
Figure 7-63. Maximum Endurance Power @ 1700 RPM, ISA-30°C (Sheet 2 of 2)
7-104
TM 55-1510-219-10
Figure 7-64. Maximum Endurance Power @ 1700 RPM, ISA-20°C (Sheet 1 of 2)
7-105
TM 55-1510-219-10
Figure 7-64. Maximum Endurance Power @ 1700 RPM, ISA-20°C (Sheet 2 of 2)
7-106
TM 55-1510-219-10
Figure 7-65. Maximum Endurance Power @ 1700 RPM, ISA-10°C (Sheet 1 of 2)
7-107
TM 55-1510-219-10
Figure 7-65. Maximum Endurance Power @ 1700 RPM, ISA- 10°C (Sheet 2 of 2)
7-108
TM 55-1510-219-10
Figure 7-66. Maximum Endurance Power @ 1700 RPM, ISA (Sheet 1 of 2)
7-109
TM 55-1510-219-10
Figure 7-66. Maximum Endurance Power @ 1700 RPM, ISA (Sheet 2 of 2)
7-110
TM 55-1510-219-10
Figure 7-67. Maximum Endurance Power @ 1700 RPM, ISA + 10°C (Sheet 1 of 2)
7-111
TM 55-1510-219-10
Figure 7-67. Maximum Endurance Power @ 1700 RPM, ISA + 10°C (Sheet 2 of 2)
7-112
TM 55-1510-219-10
Figure 7-68. Maximum Endurance Power @ 1700 RPM, ISA +20°C (Sheet 1 of 2)
7-113
TM 55-1510-219-10
Figure 7-68. Maximum Endurance Power @ 1700 RPM, ISA +20 °C (Sheet 2 of 2)
7-114
TM 55-1510-219-10
Figure 7-69. Maximum Endurance Power @ 1700 RPM, ISA +30°C (Sheet 1 of 2)
7-115
TM 55-1510-219-10
Figure 7-69. Maximum Endurance Power @ 1700 RPM, ISA +30°C (Sheet 2 of 2)
7-116
TM 55-1510-219-10
Figure 7-70. Maximum Endurance Power @ 1700 RPM, ISA +37°C (Sheet I of 2)
7-117
TM 55-1510-219-10
Figure 7-70. Maximum Endurance Power @ 1700 RPM, ISA +37°C (Sheet 2 of 2)
7-118
TM 55-1510-219-10
Figure 7-71. Range Profile-Maximum Cruise Power @ 1900 RPM
7-119
TM 55-1510-219-10
Figure 7-72. Range Profile-Maximum Range Power @ 1700 RPM
7-120
TM 55-1510-219-10
Figure 7-73. Range Profile-542 Gallons Usable Fuel
7-121
TM 55-1510-219-10
Figure 7-74. Endurance Proflie-542 Gallons Usable Fuel
7-122
TM 55-1510-219-10
Figure 7-75. Time, Fuel, and Distance to Descend
7-123
TM 55-1510-219-10
Figure 7-76. Landing Distance Without Propeller Reversing-Flaps 0%
7-124
TM 55-1510-219-10
Figure 7-77. Landing Distance Without Propeller Reversing-Flaps 100%
7-125
TM 55-1510-219-10
Figure 7-78. Landing Distance With Propeller Reversing-Flaps 0%
7-126
TM 55-1510-219-10
Figure 7-79. Landing Distance With Propeller Reversing-Flaps 100%
7-127
TM 55-1510-219-10
Figure 7-80. Stopping Distance Factors
7-128
TM 55-1510-219-10
CHAPTER 8
NORMAL PROCEDURES
Section I. MISSION PLANNING
8-2. CREW BRIEFINGS.
A crew briefing must be conducted for a thorough
understanding of individual and team responsibilities.
The briefing should include, but not be limited to,
copilot, crew chief, and ground crew responsibilities and
the coordination necessary to complete the mission
most efficiently. A review of visual signals is desirable
when ground guides do not have a direct voice
communications link with the crew. Refer to Section VI
for
crew
briefings.
8-1. MISSION PLANNING.
Mission planning begins when the mission is
assigned and extends to the preflight check of the
aircraft. It includes, but is not limited to, checks of
operating limits and restrictions; weight, balance, and
loading; performance; publications; flight plan; and crew
and passenger briefings. The pilot in command shall
insure compliance with the contents of this manual that
are applicable to the mission.
Section II. OPERATING PROCEDURES AND MANEUVERS
The symbol "O" shall be used to indicate "if installed."
The star symbol
indicates an operational check
contained in the performance section of the condensed
checklist.
The asterisk symbol "*" indicates that
performance of the step is mandatory for all thru-flights.
The asterisk applies only to checks performed prior to
takeoff. Placarded items appear in upper case.
8-3. OPERATING PROCEDURES AND MANEUVERS.
This section deals with normal procedures and
includes all steps necessary for safe and efficient
operation of the aircraft from the time, a preflight begins
until the flight is completed and the aircraft is parked
and secured. Unique feel, characteristics, and reaction
of the aircraft during various phases of operation and
the techniques and procedures used for taxiing, takeoff,
climb, etc., are described, including precautions to be
observed.
Only the duties of the minimum crew
necessary for the actual operation of the aircraft are
included.
8-6. BEFORE EXTERIOR CHECK.
1. Publications-Check DA Forms 2408-12, -13, 14, and -18, DD Form 365-4, locally required
forms and publications, and availability of
operator's manual (-10) and checklist (-CL).
8-4. CHECKLIST.
Normal procedures are given primarily in checklist
form and are amplified as necessary in accompanying
paragraph form when a detailed description of a
procedure or maneuver is required. A condensed
version of the amplified checklist, omitting all
explanatory text, is contained in the Operator's and
Crewmember's Checklist, TM 55-1510-219-CL.
To
provide for easier cross referencing, the procedural
steps are numbered to coincide with the corresponding
numbered steps in TM 55-1510-219-CL.
2. Oxygen system-Check as required.
a. Oxygen supply pressure gages-Check.
b. Supply control lever (green)-ON.
c. Diluter control lever-100% OXYGEN.
d. Emergency control lever (red)-Set to
TEST MASK position while holding mask
directly away from face, then return to
NORMAL.
e. Oxygen mask-Don and adjust.
8-5. SYMBOLS DEFINITION.
Items which apply only to night or only to
instrument flying shall have an "N" or "I", respectively,
immediately preceding the check to which it is pertinent.
f. Emergency control lever (red)-Set to
TEST MASK position and check mask for
leaks, then return lever to NORMAL.
8-1
TM 55-1510-219-10
Figure 8-1. Exterior Inspection
8-2
TM 55-1510-219-10
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
g. Flow indicator-Check, during inhalation
blinker appears, during exhalation blinker
disappears). Repeat a minimum of 3
times.
Flight controls-Unlock and check.
Parking brake-Set.
Manual trim-Zero.
Gear-Down.
Ice vanes-IN.
Overhead panels witches and circuit breakersSet as follows.
a. Circuit breakers-Check.
b. Light dimming controls-As required.
c. Cabin temperature mode-OFF.
d. Ice and rain switches-As required.
e. Exterior light switches-As required.
f. Master panel light switch-As required.
g. Inverter switches-OFF.
h. Avionics
master
power
switch-As
required.
i. Environmental switches-As required.
j. Autofeather switches-OFF.
k. #1 ignition and engine start switch-OFF.
l. Master switch-OFF.
m. #2 ignition and engine start switch-OFF.
n. Standby pump switches-OFF.
o. Auxiliary transfer switches-AUTO.
p. Crossfeed switch-OFF.
AC and DC GPU's-As required.
External power advisory lights-As required.
Keylock switch-ON.
Fuel pumps/crossfeed operation-Check as
follows:
a. Fire pull handles-Pull.
b. Standby pump switches-ON.
c. Battery switch-ON.
d. #1 and#2 fuel press warning lightsIlluminated.
e. Fire pull handles-IN.
13.
14.
15.
16.
8-3
f. #1 and#2 fuel pressure warning lightsExtinguished.
g. Standby
pump
switches-STANDBY
PUMP.
h. #1 and #2 fuel pressure warning lightsIlluminated.
i. Crossfeed-Check.
Check
system
operation
by
activating
switch
momentarily left then right, noting that
#1/#2 FUEL PRESS warning lights
extinguish
and
that
the
FUEL
CROSSFEED advisory light illuminates
as switch is energized.
DC power-Check (24 VDC minimum for battery,
28 maximum for GPU starts).
Lighting systems-Check.
Anti-ice systems-Check.
a. Stall warning heat switch-ON.
b. Pitot heat switches (2)-ON. Check cover
removed.
c. Fuel vent heat switches (2)-ON.
d. Left wing heated fuel vent-Check by feel
for heat and condition.
e. Stall warning vane-Check by feel for heat
and condition.
f. Left pitot tube-Check by feel for heat and
free of obstructions.
g. Right pitot tube-Check by feel for heat
and free of obstructions.
h. Right wing heated fuel vent-Check by feel
for heat and condition.
i. Stall warning heat switch-OFF.
j. Pitot heat switches (2)-OFF.
k. Heated fuel vent switches (2)-OFF.
Annunciator panels-Test as required.
a. MASTER
CAUTION,
MASTER
WARNING, #1 FUEL PRESSURE, #2
FUEL PRESSURE, GEAR DN, L BL AIR
FAIL, R BL AIR FAIL, ALT WARN, INST
AC, #1 DC GEN, #1 INVERTER, #1 NO
FUEL XFR, #2 NO FUEL XFR, #2
INVERTER, #2 DC GEN-Check on.
b. ANNUNCIATOR TEST switch-Press and
hold. Check that all lights in aircraft and
mission annunciator panels illuminate,
FIRE PULL handle lights, marker beacon
lights, MASTER CAUTION and MASTER
WARNING
TM 55-1510-219-10
m. Load present position.
(1) Select N or S degrees, minutes and
tenths
-INSERT/ADVANCE
PRESS.
(2) Select E or W degrees, minutes and
tenths -INSERT/ADVANCE-PRESS.
NOTE
Insure correct values for UTM grid
spheroid coefficients are loaded
using UTM coordinates.
WAYPOINT SELECTION
a. Data selector-L/L WY PT.
b. WYPT thumb wheel-DESIRED WY PT
(Do not use 0).
c. LAT/LONG waypoint (Deg., Min., Tenths)LOAD.
d. INSERT ADVANCE-PRESS AGAIN.
Latitude and longitude in arc-seconds relating to
tenths entered is shown.
e. ARC/SEC LAT & LONG (Sec.
and
Tenths)-LOAD.
Repeat a thru e for each WY PT.
f. Flight plan cross check-Data selector to
DIS/TIME (left display will indicate
distance between WY PTS TO-FROM).
Press WY PT change. Verify logical
distance between waypoints.
TACAN STATION SELECTION
a. Data selector-L/L WY PT.
b. Simultaneously press-Keys 7 and 9.
c. WY PT thumb wheel-DESIRED TACAN.
d. TACAN
station
position-LOAD
LAT/LONG.
e. INSERT/ADVANCE-PRESS.
Latitude and longitude in arc-seconds relating to tenths
entered is shown.
f. ARC/SEC and Tenths-LOAD.
g. INSERT/ADVANCE-PRESS.
h. TACAN station altitude-LOAD (select Key
4 or 6). Enter altitude.
i. INSERT/ADVANCE-PRESS.
lights are on. Release switch and check
that all lights except those in step (a.) are
extinguished.
c. MASTER CAUTION and MASTER
WARNING lights-Press.
Both lights
extinguish.
d. Stall and gear warning system-TEST.
Check that warning horn sounds and that
the LDG gear control handle lights (2)
illuminate.
17. Fire protection system-Check as follows:
a. Fire
detector
test
switch-Rotate
counterclockwise to check three DETR
positions. FIRE PULL handles should
illuminate in each position.
Reset
MASTER WARNING in each position.
b. Fire
detector
test
switch-Rotate
counterclockwise to check two EXTGH
positions. SQUIB OK light, associated
EXTGH DISCH caution light and
MASTER CAUTION LIGHT should
illuminate in each position.
18. INS alignment-As required.
a. Exterior power 28 VDC-Connected.
b. Key lock switch-ON.
c. Battery switch-ON.
d. Aircraft inverters-ON.
e. Mission inverters-ON.
f. Mission control switch-As required.
g. 3 phase A.C. bus-RESET.
Check inertial cooling for ON.
h. Aircraft
master
avionics
switchEXTERNAL POWER
i. Mode selector-ALIGN
j. Data selector-DSTRK/STS
Align condition is shown on 5 digit RH display.
Align will not progress beyond 8 until present position is
loaded.
k. Test button-PRESS AND HOLD.
All displays read 8, ROLL LIMIT, HOLD,
INSERT/ADVANCE, WY PT, CHG, ALERT, BAT,
WARN, and READY NAV LAMPS lit. When released,
all extinguish except INSERT/ADVANCE. Insure all
malfunction codes are cleared.
l. Data selector-L/L POS (UTM for grid
nav).
8-4
TM 55-1510-219-10
deflection
if
an
autopilot
command is allowed to remain
active for any appreciable length
of time. Move turn knob and pitch
thumbwheel only enough to
check operation, then return them
to the center position.
d. Select HDG mode-Check.
j. TACAN channel number-LOAD (select
Key 2 or 8). Enter channel.
Repeat steps a thru j for all TACAN stations.
HSI INTERFACE TEST
NOTE
Interface test must be performed
after alignment progresses from
state 8 but prior to switching to
NAV.
a. AUTO/MAN switch-MAN.
e. Horizontal situation indicator (HSI)
heading marker under lubber line-Set.
b. Couple INS to FD and engage AP.
f. Engage autopilot and check controls stiff,
and AIL HI TORQUE, HDG, and AP ENG
are illuminated-Check.
c. Data selector (any position but DSRTK).
d. CDU TEST-PRESS and HOLD.
MSU and CDU all lamps lit 8's. RMI all angles are
30 degrees. Cross track deviation bar is 1 dot right,
NAV flags are retracted, WX radar NAV display
indicates 1 dot right of course.
On HRI, a 15 degree RH steering command.
Aircraft panel LINK UPDATE/TACAN UPDATE
annunciated and INS light illuminated.
e. Continue holding TEST switch from MAN
to AUTO. RMI all angles are 0. Cross
track deviation is one dot left. A 15
degree left steering command is issued.
g. Move HSI heading marker 10° left and
right and verify that FD and control
wheels respond in the appropriate
direction-Check.
h. Press AP/YD disengage switch and verify
that autopilot disengages and that flight
controls are free -Check.
i. Engage autopilot-Check.
j. Command 5°trim UP with AP pitch wheel
and verify that manual trim wheel moves
nose UP and AP trim light indicates UP
trim-Check.
f. Release TEST switch-Operations return to
normal.
k. Press pitch trim switch nose down and
verify that autopilot disengages and
AUTOPILOT TRIM FAIL and MASTER
WARNING lights illuminate-Check.
NOTE
The AP TRIM FAIL annunciator
will extinguish by pressing the
AP/YD disconnect button on the
control wheel to the first detent.
l. Repeat steps i thru k above using
opposite commands.
19. Electric elevator trim and autopilot/flight director
operation-Check as follows:
a. Pilot and copilot PITCH TRIM switchesPress to NOSE UP and NOSE DN
positions, singularly and in pairs. Check
that trim wheel moves in proper direction
and operates only when trim switches are
pressed in pairs. Any deviation requires
that electric elevator trim be turned off
and flight conducted using manual trim.
b. TRIM DISC switch-Press and check that
electric trim disconnects and that ELEV
TRIM light extinguishes.
m. Engage autopilot-Check.
n. Move HSI heading marker to command a
bank on flight director-Check.
c. Flight director (FD) and radio magnetic
indicator (RMI) warning flags maskedCheck.
NOTE
Since the pressure of airflow that
normally opposes movement of
control surfaces is absent during
preflight check, it is possible to
get a hard over control surface
o. Press GO-AROUND switch and verify that
GA annunciator light illuminates, autopilot
disengages, and that flight director
commands a wings level, 7° nose-up
attitude-Check.
p. Press TEST switch (pilot's HRI) and verify
that attitude display indicates an
8-5
TM 55-1510-219-10
All NAV 2 self-test procedures are the same as
those used for NAV 1, with the exception of the marker
beacon test. There is no marker beacon receiver in the
NAV 2 system.
c. TACAN
(1) TEST pushbutton -Press and hold.
(2) Range indicator-Check for an
indication of 0.0 ± 0.1 nautical
miles.
(3) Pilot's
COURSE
SELECTOR
switch-Select TACAN.
(4) Pilot's RMI selector switch-Select
TACAN.
(5) RMI double needle-Check for an
indication of 180°± 2°.
(6) HSI course selector-Turn to 180°
and adjust slowly until the course
deviation bar is centered. The bar
should center between a selected
course of 178°to 182°.
(7) HSI course selector-Turn the
selector + 10° from the setting
achieved in step 6, and check that
course deviation bar is located over
the far left 10°dot.
(8) HSI course selector-Turn the
selector + 10°from the setting
(9) HSI course selector-Turn the
selector -10° from the setting
achieved in step 6, and check that
course deviation bar is located over
the far right 10°dot.
(10) TO-FROM indicator-Check that TO
is indicated.
(11) TEST pushbutton-Release.
21. Flaps-Check.
22. Battery switch-As required.
23. Toilet-Check.
24. Emergency equipment-Check.
25. Mission equipment and circuit breakers-Check
and set.
26. Parachutes-Check (as required).
additional 10°pitch up and 20°right bankCheck.
q. Engage autopilot command DN with AP
pitch wheel and engage and hold AUTO
PILOT TRIM TEST switch when elevator
trim wheel starts to rotate.
r. Verify that autopilot disengages and AP
TRIM FAIL and MASTER WARNING
lights illuminate within 10 seconds.
20. Avionics-Check as follows:
a. NAV 1
(1) Frequency select knob (NAV panel)Select a VOR frequency.
(2) NAV TEST switch (NAV PANEL)Press and hold.
(3) RMI-Observe that single needle
indicates approximately 005°.
(4) VOR/LOC flag -Check that flag is
out of view.
(5) TO/FROM
pointer-Check that
pointer indicates TO.
(6) HSI course deviation bar-Check for
centered bar.
(7) Marker beacon lights-Check that all
three lamps are illuminated and
flickering at approximately a 30 Hz
rate.
(8) VOR frequency knob (NAV panel)Select a LOC frequency.
(9) HSI course deviation bar-Check that
bar indicates a deflection of
approximately one dot right of
center.
(10) HSI glideslope pointer-Check that
pointer indicates a deflection of
approximately one dot below center.
(11) Marker beacon lights-Check that all
three lamps are illuminated and
flickering at approximately a 30 Hz
rate.
b. NAV 2
8-6
TM 55-1510-219-10
11. Outboard deice boot-Check for secure
bonding, cracks, loose patches, stall
strips, and general condition.
12. Stall warning vane-Check free.
13. Tiedown-Released.
14. Inboard dipole antenna set-Check for
security and cracks at mounting points.
Check bonding secure, boots free of cuts
and cracks.
15. Wing ice light-Check condition.
16. AC GPU access door-Secure.
17. Recessed and heated fuel vents-Check
free of obstruction.
18. Inverter inlet and exhaust louvers-Check
free of obstructions.
8-9. LEFT MAIN LANDING GEAR.
a. Left main landing gear-Check as follows:
1. Tires-Check for cuts, bruises, wear,
proper inflation and wheel condition.
2. Brake assembly-Check brake lines for
damage or signs of leakage, brake linings
for wear (0.25-inch), brake deice
assembly and bleed air hose for condition
and security.
3. Shock strut-Check for signs of leakage,
minimum strut extension, (5.50 inches)
and that left and right strut extension is
approximately equal.
4. Torque knee-Check condition.
5. Safety switch-Check condition, wire, and
security.
6. Fire
extinguisher
pressure-Check
pressure within limits.
7. Wheel well, doors, and linkage-Check for
signs of leaks, broken wires, security, and
general condition.
8. Fuel sump drains (forward)-Check for
leaks.
8-10. LEFT ENGINE AND PROPELLER.
a. Left engine-Check as follows:
8-7. FUEL SAMPLE.
NOTE
Fuel and oil quantity check may
be performed prior to BEFORE
EXTERIOR CHECK. During warm
weather open fuel cap slowly to
prevent being sprayed by fuel
under pressure due to thermal
expansion.
*1. Check collective fuel sample from all drains for
possible contamination. Thru-flight check is
only required if aircraft has been refueled.
8-8. LEFT WING, AREA 1.
a. Left wing area-Check as follows (fig. 8-1):
*1. General condition-Check for skin damage
such as buckling, splitting, distortion,
dents, or fuel leaks.
2. Flaps-Check
for
full
retraction
(approximately 0.25 inch play) and skin
damage such as buckling, splitting,
distortion, or dents.
3. Fuel sump drains (3)-Check for leaks.
4. Controls and trim tab-Check security and
trim tab rig.
NOTE
All static wicks (24) must be
installed for optimum radio
performance.
5. Static wicks-Check security and condition.
6. Wing pod, navigation lights and antennas
(2) -Check condition.
7. Recognition light-Check condition.
8. Outboard antenna set-Check condition.
*9. Main tank fuel and cap-Check fuel level
visually, condition of seal, and cap tight
and properly installed.
10. Outboard wing fuel vent-Check free of
obstructions.
8-7
TM 55-1510-219-10
3. Deice boot-Check for secure bonding,
cracks, loose patches, and general
condition.
*4. Auxiliary tank fuel gage and cap-Check
fuel level visually, condition of seal, and
cap tight and properly installed.
5. Monopole antenna-Check condition.
8-12. FUSELAGE UNDERSIDE.
a. Fuselage underside-Check as follows:
*1. General condition-Check for skin damage,
such as buckling, splitting, distortion,
dents, or fuel leaks.
2. Antennas-Check security and general
condition.
8-13. NOSE SECTION, AREA 2.
a. Nose section-Check as follows:
1. Free
air
temperature
probe-Check
condition.
2. Avionics door, left side-Check secure.
3. Air conditioner exhaust-Check free of
obstruction.
4. Wide band data link antenna pod-Check
for cracks and chips.
5. Wheel well-Check for signs of leaks,
broken wires and general condition.
6. Doors and linkage-Check condition,
security, and alignment.
7. Nose gear turning stop-Check condition.
*8. Tire-Check for cuts, bruises, wear,
appearance of proper inflation, and wheel
condition.
*9. Shock strut-Check for signs of leakage
and 3.0 inches minimum extension.
10. Torque knee-Check condition.
11. Shimmy damper and linkage-Check for
security and condition.
12. Landing and taxi lights-Check for security
and condition.
13. Pitot tubes-Check covers removed,
alignment,
security,
and
free
of
obstructions.
14. Radome-Check for cracks and chips.
CAUTION
A cold oil check is unreliable. Oil
should be checked within 10
minutes after stopping engine. If
more than 10 minutes have
elapsed, motor engine for 30
seconds, then recheck. If more
than 10 hours have elapsed, run
engine for 2 minutes, then
recheck. Add oil as required. Do
not overfill.
*1. Engine oil-Check oil level, oil cap secure,
locking tab aft, and access door locked.
NOTE
Secure front cowling latches first.
2. Engine compartment, left side-Check for
fuel and oil leaks, security of oil cap, door
locking pins, and general condition.
*3. Left cowl locks-Locked.
4. Left exhaust stack-Check for cracks,
security and free of obstructions.
*5. Propeller blades and spinner-Check blade
condition, boots, security of spinner and
free propeller rotation.
*6. Engine air inlets and ice vane-Check free
of obstruction and ice vane retracted.
7. Bypass door-Check condition.
*8. Right cowl locks-Locked.
9. Right exhaust stack-Check for cracks,
security and free of obstructions.
10. Engine compartment, right side-Check
for fuel and oil leaks, ice vane linkage,
door locking pins, and general condition.
Lock compartment access door.
8-11. CENTER SECTION, LEFT SIDE.
a. Center section-Check as follows:
1. Heat exchanger inlet and outlet-Check for
cracks and free of obstruction.
2. Auxiliary tank fuel sump drain-Check for
leaks.
8-8
TM 55-1510-219-10
seconds, then recheck. If more
than 10 hours have elapsed, run
engine for 2 minutes, then
recheck. Add oil as required. Do
not overfill.
*1. Engine oil-Check oil level, oil cap secure
(locking tab aft), and access door locked.
2. Engine compartment, left side-Check for
fuel and oil leaks, security of oil cap, door
locking pins, and general condition.
*3. Left cowl locks-Locked.
4. Left exhaust stack-Check for cracks,
security and free of obstructions.
*5. Propeller blades and spinner-Check blade
condition, boots, security of spinner, and
free propeller rotation.
*6. Engine air inlets and ice vane-Check free
of obstruction and ice vane retracted.
7. Bypass door-Check condition.
*8. Right cowl locks-Locked.
9. Right exhaust stack-Check for cracks,
security and free of obstructions.
10. Engine compartment, right side-Check
for fuel and oil leaks, ice vane linkage,
door locking pins, and general condition.
Lock compartment access door.
8-16. RIGHT MAIN LANDING GEAR.
a. Right main landing gear-Check as follows:
1. Fuel sump drains (forward)-Check for
leaks.
*2. Tires-Check for cuts, bruises, wear,
proper inflation and wheel condition.
3. Brake assembly-Check brake lines for
damage or signs of leakage, brake linings
for wear (0.25 inch maximum) and brake
deice assembly and bleed air hose
condition and security.
*4. Shock strut-Check for signs of leakage
and minimum strut extension (5.50
inches).
5. Torque knee-Check condition.
CAUTION
Do not move wipers on dry
windshield or clean windshield
with anything other than mild
soap and water.
15. Windshields and wipers-Check windshield
for cracks and cleanliness and wipers for
contact with glass surface.
16. Air conditioner inlet-Check free of
obstructions.
17. Avionics door, right side-Check secure.
8-14. RIGHT WING CENTER SECTION.
a. Right wing center section-Check as follows:
1. Deice boot-Check for secure bonding,
cracks, loose patches and general
condition.
2. Battery access panel-Secure.
3. Battery vents-Check free of obstruction.
*4. Auxiliary tank fuel and cap-Check fuel
level visually, condition of seal, and cap
tight and properly installed (locking tab
aft).
5. Battery compartment drain-Check free of
obstruction.
6. Battery ram air intake-Check free of
obstruction.
7. INS temperature probe-Check condition
and free of obstructions.
8. Auxiliary tank fuel sump drain-Check for
leaks.
9. Heat exchanger outlet and inlet-Check for
cracks and free of obstructions.
10. Monopole antenna-Check condition.
8-15. RIGHT ENGINE AND PROPELLER.
a. Right engine and propeller-Check as follows:
CAUTION
A cold oil check is unreliable. Oil
should be checked within 10
minutes after stopping engine. If
more than 10 minutes have
elapsed, motor engine for 30
8-9
TM 55-1510-219-10
6. Safety switch-Check condition, wire, and
security.
7. Fire
extinguisher
pressure-Check
pressure within limits.
8. Wheel well, doors, and linkage-Check for
signs of leaks, broken wires, security, and
general condition.
8-17. RIGHT WING, AREA 3.
a. Right wing-Check as follows:
1. Recessed and heated fuel vents-Check
free of obstructions.
2. Inverter inlet and exhaust louvers-Check
free of obstructions.
3. GPU access door-Secured.
4. Inboard dipole antenna set-Check for
security and cracks at mounting points,
bonding secure, free of cuts and cracks.
5. Wing ice light-Check condition.
6. Outboard deice boot-Check for secure
bonding, cracks, loose patches, stall
strips, and general condition.
*7. Tiedown-Released.
*8. Main tank fuel and cap-Check fuel level
visually, condition of seal, and cap tight
and properly installed.
9. Outboard wing fuel vent-Check free of
obstruction.
10. Outboard antenna set-Check for security
and cracks at mounting points, bonding
secure, free of cuts and cracks.
11. Recognition light-Check condition.
12. Wing pod, navigation lights and antennas
(2) -Check condition.
13. Static wicks-Check security and condition.
14. Controls-Check security and condition of
ground adjustable tab.
15. Fuel sump drains (3)-Check for leaks.
16. Flaps-Check
for
full
retraction
(approximately 0.25 inch play) and skin
damage, such as buckling, splitting,
distortion, or dents.
17. Chaff dispenser-Check number of chaffs
in payload module and for security.
*18. General condition-Check for skin damage,
such as buckling, splitting, distortion,
dents, or fuel leaks.
8-18. FUSELAGE RIGHT SIDE, AREA 4.
a. Fuselage right side-Check as follows:
* 1. General condition-Check for skin damage
such as buckling, splitting, distortion or
dents.
2. Flare/Chaff dispenser-Check number of
flares in payload module and for security.
3. Emergency light-Check condition.
4. Beacon-Check condition.
5. Aft access door-Check secure.
6. Oxygen filler door-Check secure.
7. Static ports-Check clear of obstructions.
8. ASE antennas (2)-Check.
9. Emergency locator transmitter ARMED.
10. Emergency locator transmitter antennaCheck condition.
8-19. EMPENNAGE, AREA 5.
a. Empennage-Check as follows:
1. Vertical stabilizer, rudder, and trim tabCheck for skin damage, such as buckling,
distortion, or dents, and trim tab rig.
2. Antennas-Check condition.
3. Deice boots-Check for secure bonding,
cracks, loose patches, and general
condition.
4. Horizontal stabilizer and elevator-Check
for skin damage, such as buckling,
distortion and dents.
8-10
TM 55-1510-219-10
3. Cargo door-Locked and checked. Insure
that the cargo door is closed and locked
as follows:
a. Upper handle position-Closed and
locked (the index marks on each of
the four rotary cam locks must align
within the sight indicators).
b. Lower pin latch handle positionClosed and latched (the indicator
must align with stripe on carrier
rod).
NOTE
The untapered shoulder of the
latching pins extend past each
attachment lug.
In addition, the following inspection and test shall
be performed prior to the first flight of the day:
c. Open cabin door-Check that the
"CABIN/CARGO
DOOR"
annunciator light is extinguished.
d. Latch cabin door but do not lockCheck that the "CABIN/CARGO
DOOR"
annunciator
light
illuminates.
e. Battery switch ON-Check that the
"CABIN/CARGO
DOOR"
annunciator light is still illuminated.
f. Close and lock the cabin doorCheck that the "CABIN/CARGO
DOOR"
annunciator
light
is
extinguished.
g. Battery switch-OFF.
4. Emergency exit-Check secure and key
removed.
5. Mission cooling ducts-Check open and
free of obstructions.
6. Flare/Chaff dispenser preflight testCompleted.
7. Crew briefing-As required.
8-22. BEFORE STARTING ENGINES.
* 1. Parking brake-Set.
2. Magnetic compass-Check for fluid,
heading and current deviation card.
*3. Pedestal
controls-Set
as
follows.
NOTE
Any difference between the
indicated position on the trim tab
position indicator and the actual
position of the elevator trim tab
signifies an unairworthy condition
and must be corrected prior to the
next flight of the aircraft.
5. Elevator trim tab-Verify "O" (neutral)
position.
WARNING
If
the
possibility
of
ice
accumulation on the horizontal
stabilizer or elevator exists,
takeoff will not be attempted.
6. Static wicks (16)-Check installed.
7. Position
and
beacon
lights-Check
condition.
8. Rotating boom dipole antenna-Check
condition and position.
9. Wide band data link antenna pod-Check
for cracks and chips.
8-20. FUSELAGE, LEFT SIDE, AREA 6.
a. Fuselage-Check as follows:
* 1. General condition-Check for skin damage,
such as buckling, distortion, or dents.
2.
Static ports-Check clear of obstructions.
3.
ASE antennas (2)-Check.
4.
Emergency light-Check condition.
5.
Cabin door-Check door seal and general
condition.
6.
Fuselage top side-Check general condition
and antennas.
*8-21. INTERIOR CHECK.
1. Cargo/loose equipment-Check secure.
2. Cabin door-Locked and checked. Insure
that the cabin door is closed and locked
as follows: Check position of safety arm
and diaphragm plunger (lift door step) and
that each of the six rotary cam locks align
within the sight indicators.
8-11
TM 55-1510-219-10
c.
d.
e.
f.
Gear-Recheck DN.
Cabin lights-As Required.
Pilot's static air source-NORMAL.
Pilot's and copilot's audio control panelsAs required.
NOTE
Do not use alternate static source
during takeoff and landing except
in an emergency.
Pilot's
instruments will show a variation
in airspeed and altitude.
11. AC and DC GPU's-As required.
12. External
power
advisory
lights-As
required.
*13. Battery-ON.
14. DC power-Check.
*8-23. FIRST ENGINE START (BATTERY START).
CAUTION
Mission control switch should be
OFF prior to start.
NOTE
The engines must not be started
until after the INS is placed into
the NAV mode or OFF as required.
Starting procedures are identical for both engines.
When making a battery start, the right engine should be
started first. When making a ground power unit (GPU)
start, the left engine should be started first due to the
GPU receptacle being located adjacent to the right
engine. A crew member should monitor the outside
observer throughout the engines
start.
1. Avionics master switch-OFF.
NOTE
Use the beacons NIGHT position
for
all
ground
operations,
changing to the DAY position only
when taking the active runway for
takeoff (when conditions permit).
2. Exterior light switches-As required.
CAUTION
Movement of power levers into
reverse range while engines are
shut down may result in bending
and damage to control linkages.
a. Power levers-IDLE.
b. Propeller levers-As required.
c. Condition levers-FUEL CUTOFF.
d. Flaps-UP.
4. Lower console switches-Set as follows.
a. Flare/Chaff dispenser control-SAFE.
b. Avionics-As required.
c. Rudder boost switch-ON.
5. Gear alternate engage and ratchet
handles-Stowed.
6. Free air temperature gage-Check, note
current reading.
7. Instrument panel-Check and set as
follows.
a. Pilot's and copilot's course indicator
switches -As required.
b. Pilot's and copilot's RMI switches-As
required.
c. Pilot's and copilot's MIC switch-As
required.
d. Pilot's and copilot's compass
switches-As required.
e. Gyro switches-SLAVE.
f Flight
instruments-Check
instruments for protective glass,
warning flags (10) pilot, 5 copilot),
static
readings,
and
heading
correction card.
g. Radar- OFF.
h. APR-39 and APR-44-OFF.
i. Engine
instruments-Check
for
protective glass and static readings.
8. Prop sync switch-OFF.
9. Mission panel switches and circuit
breakers-Set.
10. Subpanels-Check and set as follows.
a. Fire protection test switch-OFF.
b. Landing, taxi, and recognition lightsOFF.
8-12
TM 55-1510-219-10
4. TGT and N1-Monitor (TGT 1000IC
maximum, N1 52% minimum).
5. Oil pressure-Check (60 PSI minimum).
6. Ignition and engine start switch-OFF after
TGT stabilized.
7. Battery charge light-ON.
8. Second engine generator-RESET, then
ON.
9. Inverter switches-ON, check INVERTER
lights OFF.
10. Condition levers-As required.
8-25. ABORT START.
1. Condition lever-FUEL CUTOFF.
2. Ignition and engine start switch-STARTER
ONLY.
3. TGT-Monitor for drop in temperature.
4. Ignition and engine start switch-OFF.
8-26. ENGINE CLEARING.
1. Condition lever-FUEL CUTOFF.
2. Ignition and engine start switch-OFF (1
minute minimum).
CAUTION
Do not exceed starter limitation of
30 seconds ON and 5 minutes
OFF for two starting attempts and
engine clearing procedure. Allow
30 minutes off before additional
starter operation.
3. Ignition and engine start switch-STARTER
ONLY (15 seconds minimum, 30 seconds
maximum).
4. Ignition and engine start switch-OFF.
8-27. FIRST ENGINE START (GPU START).
1. INS-As required.
CAUTION
Mission control switch should be
OFF prior to start.
3. Ignition and engine start switch-ON.
Propeller should begin to rotate and
associated IGN ON light should
illuminate. Associated FUEL PRESS light
should extinguish.
4. Condition lever (after N1 RPM stabilizes,
12% minimum)-LOW IDLE.
CAUTION
Monitor TGT to avoid a hot start.
If there is a rapid rise in TGT, be
prepared to abort the start before
limits are exceeded.
During
starting, the maximum allowable
TGT is 1000°C for five seconds. If
this limit is exceeded, use ABORT
START
procedure
and
discontinue start. Enter the peak
temperature and duration on DA
Form 2408-13.
5. TGT and N1-Monitor (TGT 1000°C
maximum, N1 52% minimum).
6. Oil pressure-Check (60 PSI minimum).
7. Ignition and engine start switch-OFF after
TGT stabilized.
8. Condition lever-HIGH IDLE. Monitor
TGT as the condition lever is advanced.
9. Generator switch-RESET, then ON.
8-24. SECOND ENGINE START (BATTERY START).
1. First engine generator load-50% or less.
2. Ignition and engine start switch-ON.
Propeller should begin to rotate and
associated IGN on light should illuminate.
Associated FUEL PRESS light should
extinguish.
3. Condition lever (after N1 RPM passes
12% minimum)-LOW IDLE.
CAUTION
Monitor TGT to avoid a hot start.
If there is a rapid rise in TGT, be
prepared to abort the start before
limits are exceeded.
During
starting, the maximum allowable
TGT is 10000C for five seconds. If
this limit is exceeded, use ABORT
START
procedure
and
discontinue start. Enter the peak
temperature and duration on DA
Form 2408-13.
8-13
TM 55-1510-219-10
associated IGN ON light should
illuminate. Associated FUEL PRESS light
should extinguish.
2. Condition lever (after N1 RPM passes
12% minimum)-LOW IDLE.
CAUTION
Monitor TGT to avoid a hot start.
If there is a rapid rise in TGT, be
prepared to abort the start before
limits are exceeded.
During
engine
start,
the
maximum
allowable TGT is 1000°C for five
seconds. If this limit is exceeded,
use ABORT START procedure and
discontinue start. Enter the peak
temperature and duration on DA
Form 2408-13.
3. TGT and N1-Monitor (TGT 1000°C
maximum, N1 52% minimum).
4. Oil pressure-Check (60 PSI minimum).
5. Ignition and engine start switch-OFF after
TGT stabilized.
6. Propeller levers-FEATHER.
7. GPU-Disconnect. (Check aircraft external
power and mission external power light
extinguished).
8. Propellers levers-HIGH RPM.
9. Generator switches-RESET, then ON.
10. Aircraft inverter switches-ON, check
INVERTER lights OFF.
11. Condition levers-As required.
8-29. BEFORE TAXIING.
*1. Brake deice-As required. To activate the
brake deice system proceed as follows:
a. Bleed air valves-As required.
b. Brake deice switch-DEICE. Check
BRAKE DEICE ON light illuminated.
c. Condition levers-HIGH IDLE.
CAUTION
Verify airflow is present from aft
cockpit eyeball outlets to insure
sufficient cooling for mission
equipment.
*2. Cabin temperature and mode-Set.
NOTE
The engines must not be started
until after the INS is placed into
the NAV mode or OFF as required.
2. Avionics master switch-As required.
NOTE
Use the beacons NIGHT position
for
all
ground
operations,
changing to the DAY position only
when taking the active runway for
takeoff (when conditions permit).
3. Exterior light switches-As required.
4. Ignition and engine start switch-ON.
Propeller should begin to rotate and
associated IGN ON light should
illuminate. Associated FUEL PRESS light
should extinguish.
5. Condition lever (after N1 RPM stabilizes,
12% minimum)-LOW IDLE.
CAUTION
Monitor TGT to avoid a hot start.
If there is a rapid rise in TGT, be
prepared to abort the start before
limits are exceeded.
During
engine
start,
the
maximum
allowable TGT is 1000°C for five
seconds. If this limit is exceeded,
use ABORT START procedure and
discontinue start. Enter the peak
temperature and duration on DA
Form 2408-13.
6. TGT and N1-Monitor (TGT 1000°C
maximum, NI 52% minimum).
7. Oil pressure-Check (60 PSI minimum).
8. Ignition and engine start switch-OFF after
TGT stabilized.
9. Condition lever-As required. Monitor TGT
as the condition lever is advanced.
10. DC GPU disconnect-As required.
11. Generator switch (GPU disconnected)RESET, then ON.
12. Condition lever-HIGH IDLE.
8-28. SECOND ENGINE START (GPU START).
1. Ignition and engine start switch-ON.
Propeller should start to rotate and
8-14
TM 55-1510-219-10
3. AC/DC power-Check.
WARNING
Do not operate radar in congested
areas.
Injury could result to
personnel in close proximity to
operating radar.
CAUTION
Do not operate the weather radar
or data link systems in an area
where
the
nearest
effective
surface is 50 feet or less from the
antenna reflector. Scanning such
surfaces within 50 feet of the
antenna reflector may damage
receiver crystals.
4. Avionics master power switch-ON as
required.
5. Mission panel-Set.
a. Mission control-AUTO.
b. Data link high voltage-Standby.
c. Antenna select-AUTO.
d. Antenna steering-AUTO (check
azimuth).
e. Antenna override-AUTO rotate.
f. Antenna circuit breaker-As required.
6. Electric elevator trim and autopilot/flight
director operation-Check as required.
7. Avionics-Check as required.
8. Flaps-Check as required.
9. Altimeters-Check and set.
*8-30. TAXIING.
1. Brakes-Check.
2. Flight instruments-Check for normal
operation.
3. Mission control panel-Set as required.
a. Data link high voltage-ON as
required.
8-31. ENGINE RUNUP.
1. Propeller manual feathering-Check by
pulling propeller levers aft through detent
to FEATHER. Check that propeller will
feather, then advance levers to the HIGH
RPM position.
2. Autofeather-Check as follows:
a. Condition levers-LOW IDLE.
b. Autofeather switch-Hold to TEST.
c. Power
levers-Advance
until
AUTOFEATHER
lights
are
illuminated (approximately 22%
torque).
d. #1 power lever-Retard.
(1) At approximately18% torque-#2
AUTOFEATHER light out.
(2) At approximately12% torque-Both
AUTOFEATHER
lights
out
(propeller starts to feather).
e. #1 power lever-Approximately 22%
torque.
f. Repeat steps b thru d for #2 engine.
3. Overspeed governors-Check as follows:
a. Power levers-Set approximately
1950 RPM (both engines).
b. #1 propeller governor test switchHold.
c. #1 propeller RPM 1830 to 1910Check.
d. Repeat steps b and c for #2 engine.
e. Power levers-Set 1800 RPM.
4. Primary governors-Check as follows:
a. Power 1800 RPM-Set/check.
b. Propeller levers aft to detent-Set.
c. Propeller RPM 1600-1640-Check.
d. Propeller levers to HIGH RPM-Set.
5. Ice vanes-Check as follows:
a. Ice vane switches to EXTEND.
Verify torque drop, TGT increase,
and illumination of ICE VANE EXT
light-Check.
b. Ice vane switches to RETRACT.
Verify return to original torque and
TGT, and ICE VANE light
extinguished-Check.
6. Anti-ice and deice systems-Check as
follows:
a. Left pitot switch ON-Check for
loadmeter rise, then OFF.
b. Right pitot switch ON-Check for
loadmeter rise, then OFF.
8-15
TM 55-1510-219-10
c. Stall warning switch ON-Check for
loadmeter rise, then OFF.
d. Fuel vent switch ON-Check for
loadmeter rise, then OFF.
e. Windshield
anti-ice
switches
NORMAL and HI-Check PILOT and
COPILOT
(individually)
for
loadmeter rise, then OFF.
f. Propeller- AUTO (check 14-18
amps).
g. Propeller
switches-INNER
and
OUTER (momentarily), check for
loadmeter rise.
h. Surface deice switch AUTO-Check
for a drop in pneumatic pressure
and wing deice boots inflation and
after 6 seconds for a second drop in
pneumatic pressure. Check manual
position for proper indications.
i. Antenna deice single cycle autoCheck for drop in pneumatic
pressure and boots inflated. Check
manual
position
for
proper
indications.
j. Radome anti-ice-ON, check for
proper indications.
k. Engine inlet lip heat switches-ON,
Check for proper indications.
l. Anti-ice
and
deice
systems
switches-As required.
7. Beacon-As required (DAY or NIGHT).
8. Pneumatic pressure-Check as follows:
a. Condition levers-HIGH IDLE.
b. Power levers-IDLE.
c. Left bleed air valve switch-PNEU &
ENVIRO OFF.
d. Pneumatic pressure-12 to 20 PSICheck.
e. Left
bleed
air
light-Check
illuminated.
f. Right bleed air valve switch-PNEU
& ENVIRO OFF.
g. Left and right bleed air off and left
and right bleed air fail lights-Check
illuminated.
h. Left bleed air valve switch-OPEN.
i. Left bleed air off, and left and right
bleed air fail lights off, and
pneumatic pressure-Check (12 to 20
PSI).
j. Right bleed air valve switch-OPEN.
k. Right
bleed
air
off
lightExtinguished.
9. Pressurization system-Check as follows:
a. Cabin
door
caution
light
extinguished-Check.
b. Storm windows closed-Check.
c. Bleed air valve switches OPENCheck.
d. Cabin altitude 500 feet lower than
field pressure altitude-Set.
e. Cabin pressurization switch-TEST
(hold).
f. Cabin climb gage descending
indication-Check,
then
release
TEST switch.
g. Aircraft altitude set to planned
cruise altitude plus 500 feet-Check
(if this setting does not result in a
CABIN ALT indication of at least
500 feet over takeoff field pressure
altitude, adjust as required).
h. Rate control set between 9 and 12
o'clock-Check.
10. Windshield anti-ice-As required.
NOTE
If windshield anti-ice is needed
prior to takeoff, use normal
setting for a minimum of 15
minutes prior to selecting high
temperature to provide adequate
preheating and minimize effects
of thermal shock.
8-32. BEFORE TAKEOFF.
1. Autofeather switch-ARM.
2. Bleed air valves-As required.
3. Ice & rain switches-As required.
4. Fuel panel-Check fuel quantity and switch
positions.
5. Flight and engine instruments-Check for
normal indications.
6. Cabin controller-Set.
7. Annunciator
panels-Check
(Note
indications).
8. Propeller levers-HIGH RPM.
9. Friction locks-Set.
10. Flaps-As required.
11. Trim-Set.
12. Avionics-Set.
13. Flights controls-Check.
14. Departure briefing-Complete.
8-33. LINE UP.
1. Transponder-As required.
2. Engine auto ignition switch- ARM.
8-16 Change 1
TM 55-1510-219-10
3. Power stabilized-Check approximately
25% minimum.
4. Condition levers-LOW IDLE.
5. Lights-As required (landing, taxi, beacon).
6. Mission control panel-Set.
8-34. TAKEOFF.
To aid in planning the takeoff and to obtain
maximum aircraft performance, make full use of the
information affecting takeoff shown in Chapter 7. The
data shown is achieved by setting brakes, setting
TAKEOFF POWER, and then releasing brakes. When
runway lengths permit, the normal takeoff may be
modified by starting the takeoff after power has been
stabilized at approximately 25% torque, then applying
power smoothly so as to attain full power. This will
result in a smoother takeoff but will significantly increase
takeoff distance.
8-35. AFTER TAKEOFF.
WARNING
Immediately after takeoff, the pilot
flying the aircraft should avoid
adjusting controls located on the
aft portion of the extended
pedestal to preclude inducing
spatial disorientation due to
coriolis illusion.
As cruise climb airspeed is attained, adjust power
to the climb power setting. The copilot then activates
the YAW DAMP and checks that the cabin is
pressurizing. Both pilots check the wings and nacelles
for fuel or oil leaks. The procedural steps after takeoff
are as follows:
1. Gear-UP.
2. Flaps-UP.
3. Landing lights-OFF.
(4.) Windshield anti-ice-As required.
NOTE
Turn windshield anti-ice on to
normal when passing 10,000 feet
AGL or prior to entering the
freezing level (whichever comes
first). Leave on until no longer
required during descent for
landing. High temperature may be
selected as required after a
minimum warm-up period of 15
minutes.
8-36. CUMB.
a. Cruise Climb. Cruise climb is performed at a
speed which is the best combination of climb, fuel bumoff, and distance covered. Set propellers at 1900 RPM
and torque as required.
Adhere to the following
airspeed schedule as closely as possible.
SL to 10,000 feet ...................... 140 KIAS
10,000 to 20,000....................... 130 KIAS
20,000 to 31,000 feet................ 120 KIAS
b. Climb-Maximum Rate. Maximum rate of climb
performance is obtained by setting propellers at 2,000
RPM, torque at 100% (or maximum climb TGT), and
maintaining best rate-of-climb airspeed.
Refer to
Chapter 7 for best rate-of-climb airspeed for specific
weights.
1. Climb power-Set.
2. Propeller sync-As required.
3. Autofeather-As required.
4. Yaw damp- As required.
5. Cabin pressurization-Check.
6. Wings and nacelles-Check.
7. ASE-As required.
a. Flare/chaff dispenser safety pin-As
required
b. Flare/chaff function selector switchAs required.
c. APR-39-As required.
d. APR-44-As required.
8-37. CRUISE.
Cruise power settings are entirely dependent upon
the prevailing circumstances and the type of mission
being flown. Refer to Chapter 7 for airspeed, power
settings, and fuel flow information.
The following
procedures are applicable to all cruise requirements.
1. Power-Set Refer to the cruise power
graphs contained in Chapter 7.
To
account for ram air temperature increase,
it is essential that temperature be
obtained at stabilized cruise airspeed.
2. Wings and nacelles-Check.
3. Ice & rain switches-As required. Insure
that anti-ice equipment is activated before
entering icing conditions.
NOTE
Ice vanes must be extended when
operating in visible moisture at
5°C or less. Visible moisture .is
moisture in any form, clouds, ice
crystals, snow, rain, sleet, hail, or
any combination of these.
4. Auxiliary fuel gages-Monitor. Insure that
fuel is being transferred from auxiliary
tanks.
(Chapter 2, Section IV.)
Change 1 8-17
TM 55-1510-219-10
5. Altimeters-Check. Verify that altimeter
setting complies with transition altitude
requirement.
6. Engine instrument indications-Noted.
Check all engine instruments for normal
indications and record on appropriate
forms for use in engine trend monitoring.
7. Recognition lights-As required.
8-38. DESCENT.
Descent from cruising altitude should normally be
made by letting down a cruise airspeed with reduced
power.
NOTE
CABIN pressure CONTROLLER
should be adjusted prior to
starting descent.
a. Descent-Max Rate (Clean).
To obtain the
maximum rate of descent in clean configuration,
perform the following:
1. Power levers-IDLE.
2. Propeller levers-HIGH RPM.
3. Flaps-UP.
4. Gear-UP.
5. Airspeed-Vmo.
6. Cabin pressurization-Set. Adjust CABIN
CONTROLLER dial as required and
adjust RATE control knob so that cabin
rate of descent equals one-third aircraft
rate of descent.
7. Ice & rain switches-As required.
8. Recognition lights-As required.
b. Descent-Max Rate (Landing Configuration). If
required to descend at a low airspeed (e.g., to conserve
airspace or in turbulence), approach flaps and landing
gear may be extended to increase the rate and angle of
descent while maintaining the slower airspeed. To
perform the maximum rate of descent in landing
configuration, perform the following:
1. Power levers-IDLE.
2. Propeller levers-HIGH RPM.
3. Flap switch-APPROACH.
4. Gear switch-DN.
5. Airspeed-184 KIAS maximum.
6. Cabin pressurization-Set. Adjust CABIN
CONTROLLER dial as required and
adjust RATE control knob so that cabin
descent rate equals one-third aircraft
descent rate.
7. Ice and rain switches-As required.
8. Recognition lights- As required
8-39. DESCENT-ARRIVAL
Perform the following checks prior to the final
descent for landing.
1. Cabin pressurization-Set. Adjust CABIN
CONTROLIER dial as required.
2. Ice & rain switches-As required.
(3.) Windshield anti-ice-As required.
NOTE
Set windshield anti-ice to normal
or high as required well before
descent into icing conditions or
into warm moist air to aid in
defogging. Turn off windshield
anti-ice
when
descent
is
completed to lower altitudes and
when heating is no longer
required.
This will preclude
possible wind screen distortions.
4. Lights-ON.
5. Altimeters-Set to current altimeter setting.
6. ASE-As required.
8-40. BEFORE LANDING.
1. Prop sync switch-OFF.
2. Autofeather switch-ARM.
3. Propeller levers-As required.
NOTE
During
approach,
propellers
should be set at 1900 RPM to
prevent glideslope interference
(ILS approach), provide better
power response during approach,
and minimize attitude change
when advancing propeller levers
for landing.
4. Flap
switch
(below
202
KIAS)APPROACH.
5. Gear switch (below 184 KIAS)-DN.
6. Rotating boom dipole antenna-Check
stowed.
7. Landing lights-As required.
8. Brake deice-As required.
8-41. LANDING.
Performance data charts for landing computations
assume that the runway is paved, level and dry.
Additional runway must be allowed when these
conditions are not met Refer to Chapter 7 for landing
data.
Do not consider headwind during landing
computations; however, if landing must be down
8-18 Change 1
TM 55-1510-219-10
wind, include the tailwind in landing distance
computations. Plan the final approach to arrive at 50
feet over the landing area at APPROACH SPEED (Vref)
plus 1/2 wind gust speed.
Perform the following
procedures as the aircraft nears the runway.
1. Autopilot and yaw damp-Disengaged.
2. Gear down lights-Check.
3. Propeller levers-HIGH RPM.
8-42. GO-AROUND.
When a go-around is commenced prior to the
LANDING check, use power as required to climb to, or
maintain, the desired altitude and airspeed. If the goaround is started after the LANDING check has been
performed, apply maximum allowable power and
simultaneously increase pitch attitude to stop the
descent. Retract the landing gear after insuring that the
aircraft will not touch the ground. Retract the flaps to
APPROACH, adjusting pitch attitude simultaneously to
avoid an altitude loss. Accelerate to best rate-of-climb
airspeed (Vy), retracting flaps fully after attaining the Vref
speed used for the approach. Perform the following
checks:
1. Power-Maximum allowable.
2. Gear- UP.
3. Flaps-UP.
4. Landing lights-OFF.
5. Climb power-Set.
6. Yaw damp-As required.
8-43. AFTER LANDING.
Complete the following procedures after the
aircraft has cleared the runway:
1. Condition levers-As required.
2. Engine auto ignition switch-OFF.
3. Ice and rain switches-OFF.
4. Flaps-UP.
5. Avionics-As required.
6. Lights-As required.
NOTE
Reverse should not be used as a
normal
practice
when
maneuvering or parking the
aircraft.
7. Mission control panel-Set (DL HV OFF).
8-44. ENGINE SHUTDOWN.
CAUTION
To prevent sustained loads on
rudder shock links, the aircraft
should be parked with the nose
gear centered.
1. Brake deice-OFF.
2. Parking brake-Set.
3. Landing/taxi lights-OFF.
4. Overhead floodlight-As required.
5. Cabin temperature mode switch-OFF.
6. Autofeather switch-OFF.
7. Vent and aft vent blower switches-AUTO.
8. INS-OFF.
9. Mission equipment-OFF, as required.
10. Inverter switches-OFF.
11. Battery condition-Check as required.
12. Avionics master switch-OFF.
13. TGT-Check. TGT must be 660°C or
below for one minute prior to shutdown.
14. Propeller levers-FEATHER.
CAUTION
Monitor TGT during shutdown. If
sustained
combustion
is
observed, proceed immediately to
ABORT START procedure.
15. Condition levers-FUEL CUTOFF.
WARNING
Do not turn exterior lights off until
propeller's rotation has stopped.
16. Exterior lights -OFF.
17. Master panel lights - OFF.
18. Master switch - OFF
19. Keylock switch - OFF.
20. Oxygen system - As required.
8-19
TM 55-1510-219-10
CAUTION
If strong winds are anticipated
while the aircraft is unattended,
the propellers shall be secured to
prevent their windmilling with
zero engine oil pressure.
9. Walk-around
inspection-Complete.
Conduct
a
thorough
walk-around
inspection, checking for damage, fluid
leaks, and levels. Check that covers,
tiedowns, restraints, safety pins and
chocks are installed as required.
10. Aircraft forms-Complete. In addition to
established requirements for reporting any
system defects, unusual and excessive
operation such as hard landings, etc., the
flight crew will also make entries on DA
Form 2408-13 to indicate when limits in
the Operator's Manual have been
exceeded.
11. Aircraft secured-Check; lock cabin door
as required.
NOTE
A cold oil check is unreliable. Oil
should be checked within 10
minutes after stopping engine.
8-45. BEFORE LEAVING AIRCRAFT.
1. Wheels-Chocked.
2. Parking brake-As required.
NOTE
Brakes should be released after
chocks are in place (ramp
conditions permitting).
3. Flight controls-Locked.
4. Fuel pumps-Set.
a. Standby pumps-OFF
b. Aux transfer-AUTO
c. Crossfeed-CLOSED
5. Emergency exit lock-As required.
6. Mode 4-As required.
7. Aft cabin light-OFF.
8. Door light-OFF.
Section III. INSTRUMENT FLIGHT
the Flight Director and airspeed indicator. The rate of
transition is directly proportional to the rate at which the
outside references deteriorate. Approaching rotation
speed (Vr), the cross-check should be totally committed
to the instruments so that erroneous sensory inputs can
be ignored. At rotation speed, establish takeoff attitude
on the Flight Director. Maintain this pitch attitude and
wings-level attitude until the aircraft becomes airborne.
When both the vertical velocity indicator and altimeter
show positive climb indications, retract the landing gear.
After the landing gear is retracted, adjust the pitch
attitude as required to attain best rate-of-climb airspeed
(Vy). Use PITCH-SYNC as required to reposition the
Flight Director pitch steering bar. Retract flaps after
obtaining best single-engine rate-of-climb speed (Vyse),
and readjust pitch as required.
Control the bank
attitudes to maintain the desired heading. Support
Flight Director indications throughout the maneuver by
cross-checking "raw data" information displayed on
supporting instruments.
8-46. GENERAL.
This aircraft is qualified for operation in instrument
meteorological conditions.
Flight handling, stability
characteristics and range are approximately the same
during instrument flight conditions as when under visual
flight conditions.
8-47. INSTRUMENT TAKEOFF.
Complete the BEFORE TAKEOFF check.
Engage the heading (HDG) mode on the autopilot
computer/control. (DO NOT ENGAGE AUTOPILOT.)
Set heading marker (HDG) to runway heading and
adjust pitch bar. Align the aircraft with the runway
centerline, insuring that nosewheel is straight before
stopping aircraft.
Hold brakes and complete the,
LINEUP check. Insure that the roll steering bar is
centered. Power application and copilot duties are
identical to those prescribed for a "visual" takeoff. After
the brakes are released, initial directional control should
be accomplished predominantly with the aid of outside
visual references.
As the takeoff progresses, the
crosscheck should transition from outside references to
8-20
TM 55-1510-219-10
NOTE
Due to possible precession error,
the pitch steering bar may lower
slightly
during
acceleration,
causing the pitch attitude to appear
higher than actual pitch attitude.
To avoid lowering the nose
prematurely,
crosscheck
the
vertical velocity and altimeter to
insure proper climb performance.
The
erection
system
will
automatically remove the error after
the acceleration ceases.
8-48. AUTOMATIC APPROACHES.
There are no special preparations required for
placing the aircraft under autopilot control. Refer to
Chapter 3 for procedures to be followed for automatic
approaches.
NOTE
The ILS localizer and glideslope
warning flags indicate insufficient
signal strength to the receiver.
Section IV. FLIGHT CHARACTERISTICS
immediately. With wing flaps down, there is little or no
roll tendency and stalling speed is much slower than with
wing flaps up. The Stall Speed Chart (Fig. 8-2) shows
the indicated power-off stall speeds with aircraft in various
configurations. Altitude loss during a full stall will be
approximately 800 feet.
8-49. STALLS.
A prestall warning in the form of very light
buffeting can be felt when a stall is approached. An aural
warning is provided by a warning horn. The warning horn
starts approximately five to ten knots above stall speed
with the aircraft in any configuration.
c. Accelerated Stalls. The aircraft gives noticeable
stall warning in the form of buffeting when the stall
occurs. The stall warning and buffet can be demonstrated
in turns by applying excessive back pressure on the
control wheel.
a. Power-On Stalls. The power-on stall attitude is
very steep and unless this high-pitch attitude is
maintained, the aircraft will generally "settle" or "mush"
instead of stall.
It is difficult to stall the aircraft
inadvertently in any normal maneuver. A light buffet
precedes most stalls, and the first indication of
approaching stall is generally a decrease in control
effectiveness, accompanied by a "chirping" tone from the
stall warning horn. The stall itself is characterized by a
rolling tendency if the aircraft is allowed to yaw. The
proper use of rudder will prevent the tendency to roll. A
slight pitching tendency will develop if the aircraft is held
in the stall, resulting in the nose dropping sharply, then
pitching up toward the horizon; this cycle is repeated until
recovery is made. Control is regained very quickly with
little altitude loss, providing the nose is not lowered
excessively. Begin recovery with forward movement of
the control wheel and a gradual return to level flight. The
roll tendency caused by yaw is more pronounced in
power-on stalls, as is the pitching tendency; however,
both are easily controlled after the initial entry. Power-on
stall characteristics are not greatly affected by wing flap
position, except that stalling speed is reduced in
proportion to the degree of wing flap extension.
8-50. SPINS.
Intentional spins are prohibited. If a spin is
inadvertently entered use the following recovery
procedure:
NOTE
Spin demonstrations have not been
conducted.
The
recovery
technique is based on the best
available information.
The first
three actions should be as nearly
simultaneously as possible.
1. Power levers IDLE.
2. Apply full rudder opposite the direction of spin
rotation.
3. Simultaneously with rudder application, push the
control wheel forward and neutralize ailerons.
b. Power-Off Stalls.
The roll tendency is
considerably less pronounced in power-off stalls (in any
configuration) and is more easily prevented or corrected
by adequate rudder and aileron control, respectively. The
nose will generally drop straight through with some
tendency to pitch up again if recovery is not made
4. When rotation stops, neutralize rudder.
8-21
TM 55-1510-219-10
Figure 8-2. Stall Speeds
8-22
TM 55-1510-219-10
conventional throughout a dive maneuver; however,
caution should be used if rough air is encountered
after maximum allowable dive speed has been
reached, since it is difficult to reduce speed in dive
configuration. Dive recovery should be very gentle
to avoid excessive aircraft stresses.
CAUTION
Do not pull out of the resulting
dive too abruptly as this could
cause excessive wing loads
and a possible secondary stall.
5. Pull out of dive by exerting a smooth, steady
back pressure on the control wheel, avoiding
an accelerated stall and excessive aircraft
stresses.
8-52. MANEUVERING FLIGHT.
The maximum speed (Va) at which abrupt
full control inputs can be applied without exceeding
the design load factor of the aircraft (Va 170 KIAS is
shown in Chapter 5). The data is based on 14,200
pounds. There are no unusual characteristics under
accelerated flight.
8-51. DIVING.
Maximum diving airspeed (red line) is 245
KIAS or 0. 47 Mach. Flight characteristics are
Section V. ADVERSE ENVIRONMENTAL CONDITIONS
advancing CONDITION levers to HIGH IDLE during
the starting procedure, place the . power lever in
BETA and the propeller lever in HIGH RPM before
advancing the condition lever to HIGH IDLE.
8-53. INTRODUCTION.
The purpose of this part is to inform the pilot
of the special precautions and procedures to be
followed during the various weather conditions that
may be encountered in flight. This part is primarily
narrative, only those checklists that cover specific
procedures characteristic of weather operations are
included. The checklist in Section II provides for
adverse environmental operations.
c. Warm-Up and Ground Test.
Warm-up
procedures and ground test are the same as those
outlined in Section II.
d. Taxiing. Whenever possible, taxiing in deep
snow, light weight dry show or slush should be
avoided, particularly in colder FAT conditions. If it is
necessary to taxi through snow or slush, do not set
the parking brake when stopped. If possible, do not
park the aircraft in snow or slush deep enough to
reach the brake assemblies. Chocks or sandbags
should be used to prevent the aircraft from rolling
while parked. Before attempting to taxi, activate the
brake deice system, insuring that the bleed air valves
are OPEN and that the condition levers are in HIGH
IDLE. An outside observer should visually check
wheel rotation to insure brake assemblies have been
deiced. The condition levers may be returned to
LOW IDLE as soon as the brakes are free of ice.
8-54. COLD WEATHER OPERATIONS.
Operational difficulties maybe encountered
during extremely cold weather, unless proper steps
are taken prior to or immediately after flight. All
personnel should understand and be fully aware of
the necessary procedures and precautions involved.
a. Preparation For Flight. Accumulations of
snow, ice, or frost on aircraft surfaces will adversely
affect takeoff distance, climb performance and stall
speeds to a dangerous degree. Such accumulations
must be removed before flight. In addition to the
normal exterior checks, following the removal of ice,
snow, or frost, inspect wing and empennage surfaces
to verify that these remain sufficiently cleared. Also,
move all control surfaces to confirm full freedom of
movement. Assure that tires are not frozen to wheel
chocks or to the ground. Use ground heaters, antiice
solution, or brake deice to free frozen tires. When
heat is applied to release tires, the temperature
should not exceed 71°C (160°F). Refer to Chapter 2
for anti-icing, deicing, and defrosting treatment.
e. Before Takeoff If icing conditions are
expected, activate all anti-ice systems before
takeoff, allowing sufficient time for the equipment to
become effective.
f. Takeoff.
Takeoff procedures for cold
weather operations are the same as for normal
takeoff. Taking off with temperatures at or below
freezing, with water, slush or snow on the runway,
can cause ice to accumulate on the landing gear and
can throw ice into the wheel well areas. Such
takeoffs
b. Engine Starting. When starting engines on
ramps covered with ice, propeller levers must be in
the FEATHER position to prevent the tires from
sliding. To prevent exceeding torque limits when
8-23
TM 55-1510-219-10
to the sensitive brakes. In order not to impair pilot
visibility, reverse thrust should be used with caution
when landing on a runway covered with snow or
standing water. Use procedures in Section II for
normal landing.
shall be made with brake deice on and with the ice
vanes extended to preclude the possibility of ice
build-up on engine air inlets.
Monitor oil
temperatures to insure operation within limits.
Before flight into icing conditions, the pilot and
copilot WSHLD ANTI-ICE switches should be set at
NORMAL position.
g. During Flight.
j. Engine Shutdown.
in Section II.
Use normal procedures
k. Before Leaving the Aircraft.
When the
aircraft is parked outside on ice or in a fluctuating
freeze-thaw temperature condition the following
procedures should be followed in addition to the
normal procedures in Section II. After wheel chocks
are in place, release the brakes to prevent freezing.
Fill fuel tanks to minimize condensation, remove any
accumulation of dirt and ice from the landing gear
shock struts, and install protective covers to guard
against possible collection of snow and ice.
(1) After takeoff from a runway covered
with snow or slush, it may be advisable to leave
brake deice ON to dislodge ice accumulated from the
spray of slush or water. Monitor BRAKE DEICE
annunciator for automatic termination of system
operation and then turn the switch OFF. During
flight, trim tabs and controls should also be exercised
periodically to prevent freezing. Insure that anti-icing
systems are activated before entering icing
conditions. Do not activate the surface deice system
until ice has accumulated one-half to one inch. The
propeller deice system operates effectively as an
anti-ice system and it may be operated continuously
in flight. If propeller imbalance due to ice does
occur, it may be relieved by increasing RPM briefly,
then returning to desired setting.
8-55. DESERT OPERATION AND HOT WEATHER
OPERATION.
Dust, sand, and high temperatures
encountered during desert operation can sharply
reduce the operational life of the aircraft and its
equipment. The abrasive qualities of dust and sand
upon turbine blades and moving parts of the aircraft
and the destructive effect of heat upon the aircraft
instruments will necessitate hours of maintenance if
basic preventive measures are not followed. In
flight, the hazards of dust and sand will be difficult to
escape, since dust clouds over a desert may be
found at altitudes up to 10,000 feet. During hot
weather operations, the principle difficulties
encountered are high turbine gas temperatures
(TGT) during engine starting, over-heating of brakes,
and longer takeoff and landing rolls due to the higher
density altitudes. In areas where high humidity is
encountered, electrical equipment (such as
communication equipment and instruments) will be
subject to malfunction by corrosion, fungi and
moisture absorption by nonmetallic materials.
Ice vanes must be extended when
operating in visible moisture or when freedom from
visible moisture cannot be assured, at 5°C FAT or
less. Ice vanes are designed as an anti-ice system,
not a deice system. After the engine air inlet screens
are blocked, lowering the ice vanes will not rectify
the condition. Ice vanes should be retracted at 15°C
FAT and above to assure adequate engine oil
cooling.
(2) Stalling airspeeds should be expected
to increase when ice has accumulated on the aircraft
causing distortion of the wing airfoil. For the same
reason, stall warning devices are not accurate and
should not be relied upon. Keep a comfortable
margin of airspeed above the normal stall airspeed.
Maintain a minimum of 140 knots during sustained
icing conditions to prevent ice accumulation on
unprotected surfaces of the wing. In the event of
windshield icing, reduce airspeed to 226 knots or
below.
a. Preparation For Flight. Check the position of
the aircraft in relation to other aircraft. Propeller
sand blast can damage closely parked aircraft.
Check that the landing gear shock struts are free of
dust and sand. Check instrument panel and general
interior for dust and sand accumulation. Open main
entrance door and cockpit vent storm windows to
ventilate the aircraft.
h. Descent. Use normal procedures in Section
II. Brake icing should be considered if moisture was
encountered during previous ground operations or
inflight in icing conditions with gear extended.
i. Landing. Landing on an icy runway should
be attempted only when absolutely necessary and
should not be attempted unless the wind is within 10
degrees of runway heading. Application of brakes
without skidding the tires on ice is very difficult, due
8-24
TM 55-1510-219-10
CAUTION
N1 speeds of 70% or higher may be required to
keep oil temperature within limits.
b. Engine Starting. Use normal procedures in Section
II. Engine starting under conditions of high ambient
temperatures may produce a higher than normal TGT
during the start. The TGT should be closely monitored
when the condition lever is moved to the LO IDLE position.
If overtemperature tendencies are encountered the condition lever should be moved to IDLE CUTOFF position
periodically during acceleration of gas generator RPM (N1).
Be prepared to abort the start before temperature limitations
are exceeded.
c. Warm-Up Ground Tests. Use normal procedures in
Section II. To minimize the possibility of damage to the
engines during dusty/sandy conditions, activate ICE
VANES if the temperature is below 15°C.
d. Taxiing. Use normal procedures in Section II.
When practical, avoid taxing over sandy terrain to minimize propeller damage and engine deterioration that results
front impingement of sand and gravel. During hot weather
operation, use minimum braking action to prevent overheating.
e. Takeoff. Use normal procedures in Section II.
Avoid taking off in the wake of another aircraft if the runway surface is sandy or dusty.
f. During Flight. Use normal procedures in Section II.
g. Descent. Use normal procedures in Section II.
h. Landing. Use normal procedures in Section II.
j. Before Leaving Aircraft. Use normal procedures in
Section II. Take extreme care to prevent sand or dust from
entering the fuel and oil system during servicing. During
hot weather, release the brakes immediately after installing
wheel chocks to prevent brake disc warpage.
8-56. TURBULENCE AND THUNDERSTORM
OPERATION.
CAUTION
Due to the comparatively light wing loading,
control in severe turbulence and thunderstorms
is critical. Since turbulence imposes heavy
loads on the aircraft structure, make all necessary changes in aircraft attitude with the least
amount of control pressures to avoid excessive
loads on the aircraft's structure.
Thunderstorms and areas of severe turbulence should be
avoided. If such areas are to be penetrated. it will be necessary to counter rapid changes in attitude and accept major
indicated altitude variations. Penetration should be of an
altitude which provides adequate maneuvering margins as a
loss or gain of several thousand feet of altitude may be
expected. The recommended penetration speed in severe
turbulence is 170 KIAS. Pitch attitude and constant power
settings arc vital to proper flight technique. Establish
recommended penetration speed and proper attitude prior to
entering turbulent air to minimize most difficulties. False
indications by the pressure instruments due to barometric
pressure variations within the storm make them unreliable.
Maintaining a preestablished attitude will result in a fairly
constant airspeed. Turn cockpit and cabin lights on to minimize the blinding effects of lighting. Do not use autopilot
altitude hold. Maintain constant power settings and pitch
attitude regardless of airspeed or altitude indications. Concentrate on maintaining a level attitude by reference to the
Flight Director/attitude indicator. Maintain original reading.
Make no turns unless absolutely necessary.
i. Engine Shutdown. Use normal procedures in
Section II.
CAUTION
During hot weather, if fuel tanks are completely
filled, fuel expression may cause overflow,
thereby creating a fire hazard.
Change 1
8-25
TM 55-1510-219-10
8-57. ICE AND RAIN (TYPICAL).
WARNING
While in icing conditions, if there is an
unexplained 30% increase of torque needed
to maintain airspeed in level flight, a
cumulative total of two or more inches of ice
accumulation on the wing, an unexplained
decrease of 15 knots IAS, or an unexplained
deviation between pilot’s and copilot’s
airspeed indicators, the icing environment
should be exited as soon as practicable. Ice
accumulation on the pitot tube assemblies
could cause a complete loss of airspeed
indication.
The following conditions indicate a possible
accumulation of ice on the pitot tube assemblies and
unprotected airplane surfaces. If any of these conditions
are observed, the icing environment should be exited as
soon as practicable.
(1) Total ice accumulation of two inches or more on
the wing surfaces. Determination of ice thickness can be
accomplished by summing the estimated ice thickness on
the wing prior to each pneumatic boot deice cycle (e.g. four
cycles of minimum recommendation ½-inch accumulation.
(2) A 30 percent increase in torque per engine
required to maintain a desired airspeed in level flight (not
to exceed 85 percent torque) when operating at
recommended holding/loiter speed.
(3) A decrease in indicated airspeed of 15 knots
after entering the icing condition (not slower than 1.4
power off stall speed) if maintaining original power setting
in level flight. This can be determined by comparing preicing condition entry speed to the indicated speed after a
surface and antenna deice cycle is completed.
(4) Any variations from normal indicated airspeed
between the pilot’s and copilot’s airspeed indicators.
a. Typical Ice. Icing occurs because of supercooled
water vapor such as fog, clouds or rain. The most severe
icing occurs on aircraft surfaces in visible moisture or
precipitation with a true outside air temperature between 5°C and +1°C; however, under some circumstances,
dangerous icing conditions may be encountered with
temperatures below -10°C. The surface of the aircraft
must be at a temperature of freezing or below before ice
will stick to the aircraft. If severe icing conditions are
encountered, ascend or descend to altitudes where these
conditions do not prevail. If flight into icing conditions is
8-26
Change 3
unavoidable, proper use of aircraft anti-icing and deicing
systems may minimize the problems encountered
Approximately 15 minutes prior to flight into temperature
conditions which could produce frost or icing conditions,
the pilot and co-pilot windshield anti-ice switches should
he set at normal or high temperature position (after
preheating) as necessary to eliminate windshield ice.
Stalling airspeeds should be expected to increase when ice
has accumulated on the aircraft causing distortion of the
wing airfoil. For the same reason, stall warning devices are
not accurate and should not be relied upon. Keep a
comfortable margin of airspeed above the normal stall
airspeed with ice on the aircraft. Maintain a minimum of
140 knots during sustained icing conditions to prevent ice
accumulation on unprotected surfaces of the wing. In the
event of windshield icing, reduce airspeed to 226 knots or
below.
b. Rain. Rain presents no particular problems other
than restricted visibility and occasional incorrect airspeed
indications.
c. Taxiing. Extreme are must be exercised when
taxiing on ice or slippery runways. Excessive use of either
brakes or power may result in an uncontrollable skid.
d. Takeoff. Extreme care must be exercised during
takeoff from ice or slippery runways. Excessive use of
either brakes or power may result in an uncontrollable skid.
e. Climb. Keep aircraft attitude as flat as possible and
climb with higher airspeed than usual, so that the lower
surfaces of the aircraft will not be iced by flight at a high
angle of attack.
f. Cruising Flight. Prevention of ice formation is far
more effective and satisfactory than attempts to dislodge
the ice after it has formed. If icing conditions are
inadvertently encountered, turn on the anti-icing systems
prior to the first sign of ice formation.
Do not operate deicer boots continuously. Allow at
least one-half inch of ice on the boots before activating the
deicer boots to remove the ice. Continued flight in severe
icing conditions should not be attempted. If ice forms on
the wing area aft of the deicer boots, climb or descend to an
altitude where conditions are less severe.
g. Landing. Extreme care must be exercised when
landing on ice or slippery runways. Excessive use of either
brakes or power may result in an uncontrollable skid. Ice
accumulation on the aircraft will result in higher stalling
airspeeds due to the change in aerodynamic characteristics
and increased weight of the aircraft due to ice buildup.
Approach and landing airspeeds must be increased
accordingly.
TM 55-1510-219-10
NOTE
When operating on wet or icy runways, refer to
stopping distance factors shown in Chapter 7.
8-57A. ICING (SEVERE).
a. The following weather conditions may be conducive
to severe in-flight icing:
(a) Immediately request priority handling from
air traffic control to facilitate a route or an altitude change
to exit the severe icing conditions in order to avoid
extended exposure to flight conditions more severe than
those for which the airplane has been certificated.
(b) Avoid abrupt and excessive maneuvering
that may exacerbate control difficulties.
(c) Do not engage the autopilot.
(1) Visible rain at temperatures below zero degrees
Celsius ambient air temperature.
(d) If the autopilot is engaged, hold the control
wheel firmly and disengage the autopilot.
(2) Droplets that splash or splatter on impact at
temperatures below zero degrees Celsius ambient air
temperature.
(e) If an unusual roll response or
uncommanded roll control movement is observed, reduce
the angle-of-attack.
b. The following procedures for exiting a severe icing
environment are applicable to all flight phases from takeoff
to landing.
(f) Do not extend flaps during extended
operation in icing conditions. Operations with flaps
extended can result in a reduced angle-of-attack, with the
possibility of ice forming on the upper surface further aft
on the wing than normal, possibly aft of the protected area.
(1) Monitor the ambient air temperature. While
severe icing may form at temperatures as cold as -18
degrees Celsius, increased vigilance is warranted at
temperatures around freezing with visible moisture present.
(2) Upon observing the visual cues specified in the
limitations section of the airplane flight manual (Military
Operations Manual) for the identification of severe icing
conditions (reference paragraph 5-31B.), accomplish the
following:
(g) If the flaps are extended, do not retract
them until the airframe is clear of ice.
(h) Report these weather conditions to air
traffic control.
Change 3
8-26.1
TM 55-1520-219-10
Section VI. CREW DUTIES
8-58. CREW/PASSENGER BRIEFING.
The following guide should be used in accomplishing
required passenger briefings. Items that do not pertain to a
specific mission may be omitted.
d. Normal Procedures.
1. Entry and exit of aircraft.
2. Seating and seat position.
a. Crew introduction.
3. Seat belts.
b. Equipment.
4. Movement in aircraft.
1. Personal, to include ID tags.
5. Internal communications.
2. Professional (medical equipment, etc.).
6. Security of equipment.
3. Survival.
7. Smoking.
c. Flight data.
8. Oxygen.
1. Route.
9. Refueling.
2. Altitude.
10. Weapons and prohibited items.
3. Time enroute.
11. Protective masks.
4. Weather.
12. Toilet.
e. Emergency Procedures.
1. Emergency Exits.
8-26.2
Change 3
TM 55-1510-219-10
2. Emergency equipment.
3. Vy (climb to 500' AGL).
3. Emergency landing/ditching
procedures.
4. Vyse
8-59. DEPARTURE BRIEFING.
8-60. ARRIVAL BRIEFING.
The following is a guide that should be used
as applicable in accomplishing the required crew
briefing prior to takeoff; however, if the crew has
operated together previously and the pilot is certain
that the copilot understands all items of the briefing,
he may omit the briefing by stating "standard
briefing," when the briefing is called for during the
BEFORE TAKEOFF CHECK.
The following is a guide that should be used
as applicable in accomplishing the required crew
briefing prior to landing; however, if the crew has
operated together previously and the pilot is certain
that the copilot understands all items of the briefing,
he may omit the briefing by stating "standard
briefing," when the briefing is called for during the
DESCENT-ARRIVAL CHECK.
a. ATC Clearance Review.
1. Routine.
a. Weather/Altimeter Setting.
b. Airfield/Facilities Review.
2. Initial altitude.
1. Field elevation.
2. Runway length.
b. Departure Procedure Review.
1. SID.
3. Runway condition.
2. Noise abatement procedure.
c. Approach Procedure-Review.
3. VFR departure route.
1. Approach plan/profile.
c. Copilot Duties Review.
1. Adjust takeoff power.
2. Altitude restrictions.
3. Missed approach.
2. Monitor engine instruments.
a. Point.
3. Power check at 65 knots.
b. Time.
4. Call out engine malfunctions.
c. Intentions.
5. Tune/ident all Nav/Com radios.
4. Decision height or MDA.
6. Make all radio calls.
7. Adjust transponder
required.
and
radar
5. Lost communications.
as
8. Complete flight log during flight (note
altitudes and headings).
d. Back Up Approach/Frequencies.
e. Copilot Duties Review.
1. Nav/Com set-up.
2. Monitor altitude and airspeeds.
9. Note departure time.
3. Monitor approach.
d. PPCReview.
1. Takeoff power.
2. Vr.
4. Callout visual/field in sight.
f. Landing Performance Data Review.
1. Approach speed.
2. Runway required.
8-27/8-28 blank
TM 55-1510-219-10
CHAPTER 9
EMERGENCY PROCEDURES
Section I. AIRCRAFT SYSTEMS
in the remarks section of DA Form 2408-13
describing the malfunction.
9-1. AIRCRAFT SYSTEMS.
This section describes the aircraft systems
emergencies that may reasonably be expected to
occur and presents the procedures to be followed.
Emergency procedures are given in checklist form
when applicable. A condensed version of these
procedures is in the Operator's and Crewmember's
Checklist, TM 55-1510-219-CL.
Emergency
operations of avionics equipment are covered when
appropriate in Chapter 3, Avionics, and are repeated
in this section only as safety of flight is affected.
9-5. EMERGENCY EXITS AND EQUIPMENT.
Emergency exits and equipment are shown
in figure 9-1.
9-6. EMERGENCY ENTRANCE.
Entry may be made through the cabin
emergency hatch. The hatch may be released by
pulling on its flush-mounted pull-out handle,
placarded EMERGENCY EXIT-PULL HANDLE TO
RELEASE. The hatch is of the nonhinged plug type
which removes completely from the frame when the
latches are released. After the latches are released,
the hatch may be pushed in.
9-2. IMMEDIATE ACTION EMERGENCY CHECKS.
Immediate action emergency items are
underlined for your reference and shall be committed
to memory. During an emergency, the checklist will
be called for to verify the memory steps performed
and to assist in completing any additional emergency
procedures.
9-7. ENGINE MALFUNCTION.
a. Flight Characteristics Under Partial Power
Conditions.
There are no unusual flight
characteristics during single-engine operation as long
as airspeed is maintained at or above minimum
control speed (Vmc) and power-off stall speeds. The
capability of the aircraft to climb or maintain level
flight depends on configuration, gross weight,
altitude, and free air temperature. Performance and
control will improve by feathering the propeller of the
inoperative engine, retracting the landing gear and
flaps, and establishing the appropriate single-engine
best rate-of-climb speed (Vyse). Minimum control
speed (Vmc) with flaps retracted is approximately 1
knot | higher than with flaps at takeoff (40%) position.
NOTE
The
urgency
of
certain
emergencies
requires
immediate action by the pilot.
The most important single
consideration
is
aircraft
control.
All procedures are
subordinate
to
this
requirement.
9-3. DEFINITION OF LANDING TERMS.
The term LANDING IMMEDIATELY is
defined as executing a landing without delay. (The
primary consideration is to assure the survival of
occupants.
) The term LAND AS SOON AS
POSSIBLE is defined as executing a landing at the
nearest suitable landing area without delay. The
term LAND AS SOON AS PRACTICABLE is defined
as executing a landing to the nearest suitable airfield.
b. Engine Malfunction During And After
Takeoff The action to be taken in the event of an
engine malfunction during takeoff depends on
whether or not liftoff speed (V1of) has been attained.
If an engine fails immediately after liftoff, many
variables such as airspeed, runway remaining,
aircraft weight, altitude at time of engine failure, and
single-engine performance must be considered in
deciding whether it is safer to land or continue flight.
9-4. AFTER EMERGENCY ACTION.
After
a
malfunction
has
occurred,
appropriate emergency actions have been taken, and
the aircraft is on the ground, an entry shall be made
c. Engine Malfunction Before Liftoff (Abort). If
an engine fails and the aircraft has not accelerated
9-1
TM 55-1510-219-10
Figure 9-1. Emergency Exits and Equipment
9-2
TM 55-1520-219-10
to recommend liftoff speed (V1of), retard power levers
immediately to IDLE and stop the aircraft with brakes and
reverse thrust. Perform the following:
1. Power - Maximum controllable.
2. Gear - UP.
1. Power levers - IDLE.
3. Flaps - UP.
2. Braking as required.
4. Landing lights - OFF.
NOTE
If insufficient runway remains for stopping,
perform steps 3. through 5.
5. Brake deice - OFF.
6. Engine cleanup - Perform.
7. Generator load - 100% Maximum.
3. Condition levers - FUEL CUT-OFF.
NOTE
4. Fire pull handles - Pull.
5. Master switch - OFF.
d. Engine Malfunction after Liftoff. If an engine fails
after becoming airborne, maintain single-engine best rateof-climb speed (Vyse) or, if airspeed is below Vyse maintain
whatever airspeed is attained between liftoff (V1of) and Vyse
until sufficient altitude is attained to trade altitude for
airspeed and accelerate to Vyse.
If the autofeather system fails to operate,
identify the affected engine, then feather the
propeller of the affected engine.
(3) Engine Malfunction after Liftoff (Flight
Continued without Autofeather). Perform the following:
1. Power - Maximum controllable.
NOTE
(1) Engine Malfunction after Liftoff (Abort).
Perform the following and land in a wings-level attitude:
1. Power levers - IDLE.
2. Gear - Down.
NOTE
If able to land on remaining runway, use brakes
and reverse thrust as required, then perform
steps 3. through 5.
3. Condition levers - FUEL CUT-OFF.
4. Fire pull handles - Pull.
5. Master switch - OFF.
(2) Engine Malfunction after Liftoff (Flight
Continued). Perform the following:
If airspeed is below Vyse, maintain whatever
airspeed has been attained (between V 1of and
Vyse) until sufficient altitude can be obtained to
trade off altitude for airspeed to assist in
acceleration to Vyse.
2. Dead engine - Identify.
3. POWER lever (dead engine) - IDLE.
4. PROP lever (dead engine) - FEATHER.
5. GEAR-UP.
NOTE
If takeoff was made with flaps extended, insure
that airspeed is above computed approach speed
(Vref) before retracting flaps.
6. FLAPS-UP.
NOTE
7. LANDING LIGHTS - OFF.
Do not retard the malfunctioning engine power
lever, or turn the autofeather system OFF until
propeller is completely feathered. To do so will
deactivate the autofeather circuit and prevent
automatic feathering.
8. BRAKE DEICE - OFF.
9. Engine cleanup - Perform.
10. Generator load - 100% Maximum.
(9-2.1 blank)9-2.2
Change 3
TM 55-1520-219-10
e. Engine Malfunction During Flight.
1.
Autopilot/yaw damper - Disengage.
2.
Power - As required.
3.
Dead Engine - IDENTIFY.
4.
Power lever (affected engine) - IDLE.
5.
Propeller lever (affected engine) - FEATHER.
g. Engine Malfunction (Second Engine). If the second
engine fails, do not feather the propeller if an engine restart
is to be attempted. Engine restart without starter assist can
not be accomplished with a feathered propeller, and the
propeller will not unfeather without the engine operating.
140 KIAS is recommended as the best all-around glide
speed (considering engine restart, distance covered,
transition to landing configuration, etc.), although it does
not necessarily result in the minimum rate of descent.
Perform the following procedure if the second engine fails
during cruise flight.
6 . Gear - As required.
1.
Airspeed - 140 KIAS.
7.
Flaps - As required.
2.
Power lever - IDLE.
8.
Power - Set for single engine cruise.
3.
Propeller lever - Do not FEATHER.
9.
Engine cleanup - Perform.
4.
Conduct engine restart procedure.
10.
Generator load - 100% Maximum.
f. Engine Malfunction During Final Approach. If an
engine malfunctions during a final approach (after
LANDING CHECK), the propeller should not be manually
feathered unless time and altitude permit or conditions
require it. Continue approach using the following
procedure:
9-8. ENGINE SHUTDOWN IN FLIGHT.
If it becomes necessary to shut an engine down during
flight, perform the following:
1. Power lever - IDLE.
1. Power - As required.
2. Gear - DN.
Change 3
9-3
TM 55-1510-219-10
2.
Propeller lever - FEATHER.
11.
Oil pressure - Check.
3.
Condition lever - FUEL CUTOFF.
12.
Ignition and engine start switch - OFF.
4.
Engine cleanup - Perform.
13.
Generator switch - RESET, then ON.
14.
Engine cleanup - Perform if engine restart
unsuccessful.
15.
Cabin temperature mode switch - As required.
9-9. ENGINE CLEANUP.
The cleanup procedure to be used after engine
malfunction, shutdown, or an unsuccessful restart is
as follows:
16.
Electrical equipment - As required.
1.
Condition lever - FUEL CUTOFF.
17.
Auto ignition switch - ARMED.
2.
Engine auto ignition switch - OFF.
18.
Propellers - Synchronized.
3.
Autofeather switch - OFF.
19.
Power - As required.
4.
Generator switch - OFF.
5.
Prop sync switch - OFF.
6.
Radome heat - OFF.
9-10. ENGINE RESTART DURING FLIGHT (usING STARTER).
Engine restarts may be attempted at all altitudes. If a restart is attempted, perform the following:
9-11. ENGINE RESTART DURING FLIGHT (NOT
USING STARTER).
A restart without starter assist may be accomplished provided airspeed is at or above 140 KIAS
altitude is below 20,000 feet, and the propeller is not
feathered. If altitude permits, diving the aircraft will
increase N1 and assist in restart. N1 required for airstart should be at or above 9%. If a start is attempted, perform following:
1.
Cabin temperature mode switch - OFF.
1.
Cabin temperature mode switch - OFF.
2.
Electrical load - Reduce to minimum.
2.
Electrical load - Reduce to minimum.
3.
Fire pull handle - In.
3.
Generator switch (affected engine) - OFF.
4.
Power lever - IDLE.
4.
Fire pull handle - Check in.
5.
Propeller lever - FEATHER.
5.
Power lever - IDLE.
6.
Condition lever - FUEL CUTOFF.
6.
Propeller lever - HIGH RPM.
7.
TGT (operative engine) - 700°C or less.
7.
Condition lever - FUEL CUTOFF.
8.
Ignition and engine start switch - ON.
8.
Airspeed 140 KIAS minimum - Check.
9.
Condition lever - LOW IDLE.
9.
Altitude below 20,000 feet - Check.
10.
Engine auto ignition switch - ARM.
11.
Condition lever - LOW IDLE.
12.
TGT - 1000°C 5 seconds maximum.
13.
Oil pressure - Check.
14.
Generator switch - RESET then ON.
15.
Engine cleanup - Perform if engine restart
unsuccessful.
16.
Cabin temperature mode switch - As required.
17.
Electrical equipment - As required.
18.
Auto ignition switch - ARMED.
19.
Propellers - Synchronized.
NOTE
If a rise in TGT does not occur within 10
seconds after moving the condition lever
to LOW IDLE, abort the start.
10.
TGT - 1000°C, 5 seconds maximum.
NOTE
If N1 is below 12%, starting temperatures
tend to be higher than normal. To preclude overtemperature (1000°C or above)
during engine acceleration to idle speed,
periodically move the condition lever into
FUEL CUTOFF position as necessary.
9-4
TM 55-1520-219-10
20. Power - As required.
9-12. MAXIMUM GLIDE.
In the event of failure of both engines, maximum
gliding distance can be obtained by feathering both
propellers to reduce propeller drag and by maintaining the
appropriate airspeed with the gear and flaps up. Figure 9-2
gives the approximated gliding distances in relation to
altitude.
9-13. LANDING WITH TWO ENGINES INOPERATIVE.
Maintain best glide speed (figure 9-2). If sufficient
altitude remains after reaching a suitable landing area, a
circular pattern will provide best observation of surface
conditions, wind velocity and direction. When the
condition of the terrain has been noted and the landing area
selected, set up a rectangular pattern. Extending
APPROACH flaps and landing gear early in the pattern
will give an indication of glide performance sooner, and
will allow more time to make adjustments for the added
drag. Fly the base leg as necessary to control Point of
touch-down. Plan to overshoot rather than undershoot,
then use flaps as necessary to arrive at the selected landing
point.
Keep in mind that, with both propellers feathered, the
normal tendency is to overshoot due to less drag. In the
event a positive gear-down indication cannot be
determined, prepare for a gear-up landing. Also, unless the
surface of the landing area is hard and smooth, the landing
should be made with the landing gear up. If landing on
rough terrain, land in a slightly tail-low attitude to keep
nacelles from possibly digging in. If possible, land with
flaps fully extended.
If the L CHIP DET or R CHIP DET warning
annunciator illuminates, and safe single-engine flight can
be maintained, perform the following:
1. Perform engine shutdown.
2. Land as soon as practicable.
9-16. DUCT OVERTEMP CAUTION LIGHT
ILLUMINATED.
Insure that the cabin floor outlets are open and
unobstructed, then perform the following steps in sequence
until the light is extinguished. After completion of steps 1.
through 4., if the light does not extinguish, allow
approximately 30 seconds after each adjustment for the
system temperature to stabilize. The overtemperature
condition is considered corrected at any point during the
procedure that the light goes out.
1. Cabin air control - In.
2. Cabin temperature mode switch - AUTO.
3. Cabin temperature rheostat - Full decrease.
4. Vent blower switch - HI.
5. Cabin temperature mode switch - MAN HEAT.
6. Manual temperature switch - DECREASE (hold).
7. Left bleed air valve switch - ENVIRO OFF.
8. Light still illuminated (30 seconds) - Left bleed air
valve switch - OPEN.
9. Right bleed air valve switch - ENVIRO OFF.
9-14. LOW OIL PRESSURE.
In the event of a low oil pressure indication, perform the
procedures below, as applicable:
10. Light still illuminated (30 seconds) - Right bleed air
valve switch - OPEN.
NOTE
1. Oil pressure below 105 PSI below 21,000 feet or 85
PSI at 21,000 feet and above, torque - 49%
maximum.
2. Oil pressure below 60 PSI - Perform engine
shutdown, or land as soon as practicable using
minimum power to ensure safe arrival.
9-15. CHIP DETECTOR WARNING
ANNUNCIATOR ILLUMINATED.
If the overtemperature light has not extinguished
after completing the above procedure, the
warning system has malfunctioned.
9-17. ICE VANE FAILURE.
Ice vane failure is indicated by VANE FAIL caution
light illumination. If an ice vane fails to operate
electrically, perform the following:
Change 3
9-5
TM 55-1510-219-10
Figure 9-2. Maximum Glide Distance
9-6
TM 55-1510-219-10
b. Excessive Differential Pressure. If cabin differential pressure exceeds 6.1 PSI, perform the following:
After the ice vanes have been manually
extended, they may be mechanically retracted. No electrical extension or retraction shall be attempted as damage to the
electric actuator may result. Linkage in
the nacelle area must be reset prior to operation of the electric system. Do not reset ice vane control circuit breaker.
1.
Airspeed - 160 KIAS or below.
2.
Ice vane control circuit breaker - Pull.
3.
Ice vane - Operate manually.
4.
Airspeed - Resume normal airspeed.
DO NOT RETRACT ice vanes electrically after manual extension.
9-18. ENGINE BLEED AIR SYSTEM MALFUNCTION.
a. Bleed Air Failure Light Illuminated. Steady
illumination of the warning light in flight indicates
a possible ruptured bleed air line aft of the engine
firewall. The light will remain illuminated for the remainder of flight. Perform the following:
NOTE
BLEED AIR FAIL lights may momentarily illuminate during simultaneous surface
deice and brake deice operation at low
N 1 . speeds.
1.
Brake deice switch - OFF.
2.
TGT and torque - Monitor (note readings).
Bleed air valve switch - PNEU & ENVIRO OFF.
3.
1.
Cabin controller - Select higher setting.
2.
If condition persists: Left bleed air
valve switch - ENVIRO OFF flight illuminated).
3.
If condition still persists: Right bleed
air valve switch - ENVIRO OFF (light
illuminated).
4.
If condition still persists - Descend immediately.
5.
If unable to descend:
masks - 100% and on.
6.
If unable to descend: CABIN PRESS
switch - DUMP.
7.
Bleed air valve switches - OPEN, if
cabin heating is required.
Crew
oxygen
9-19. LOSS OF PRESSURIZATION (ABOVE 10,
000 FEET).
If cabin pressurization is lost when operating
above 10,000 feet or the ALTITUDE warning light
illuminates, perform the following:
1. Crew oxygen masks - 100% and on.
9-20. CABIN DOOR CAUTION LIGHT ILLUMINATED.
Remain clear of cabin door and perform the following:
1.
Bleed air valve switches - ENVIRO OFF.
2.
Descend below 14,000 feet as soon as practicable.
3.
Oxygen - As required.
9-21. SINGLE-ENGINE DESCENT/ARRIVAL.
NOTE
NOTE
Approximately 85% N 1 is required to
maintain pressurization schedule.
Brake deice on the affected side, and rudder boost, will not be available with
BLEED AIR VALVE switch in PNEU &
ENVIRO OFF.
Perform the following procedure prior to the final descent for landing.
4.
Cabin pressurization - Check.
1. Cabin controller - Set.
2. Seat belts and harnesses - Secure.
9-7
TM 55-1510-219-10
3.
Ice and rain switches - As required.
4.
Altimeters - Set.
5.
Recognition lights - ON.
6.
Arrival briefing - Complete.
when close to grouud under conditions of
high gross weights and/or high density altitude.
1. Power - Maximum controllable.
2. Gear - UP.
9-22. SINGLE-ENGINE BEFORE LANDING.
1. Propeller lever - As required.
NOTE
During approach, propeller should be set
at 1900 RPM to prevent glideslope interference (ILS approach), provide better
power response during approach, and to
minimize attitude change when advancing
propeller levers for landing.
2.
Flaps - APPROACH.
3.
Gear - DN.
4.
Landing lights - As required.
5.
Yaw damp - OFF.
6.
Brake deice - OFF.
3 . Flaps - As required.
4. Landing lights - OFF.
5. Power - As required.
6. Yaw damp - As required.
9-25. PROPELLER FAILURE (OVER 2000 RPM).
If an overspeed condition occurs that cannot be
controlled with the propeller lever, or by reduciug
power, perform the following:
1. Power lever (affected engine) - IDLE.
2. Propeller lever - FEATHER.
3. Condition lever - As required.
4. Engine cleanup - As required,
9-23. SINGLE-ENGINE LANDING CHECK.
9-26. FIRE.
Perform the following procedure during final approach to runway.
The safety of aircraft occupants is the primary
consideration when a fire occurs; therefore, it is imperative that every effort be made by the flight crew
to put the fire out. On the ground it is essential that
the engines be shut down, crew evacuated, and fire
fighting begun immediately. If the aircraft is airborne when a fire occurs, the most important single
action that can be taken by the pilot is to land safely
as soon as possible.
1. Autopilot/yaw damp - Disengaged.
2. Gear lights - Check.
3. Propeller lever (operative engine) - HIGH
RPM.
NOTE
To insure constant reversing characteristics. the propeller control must be in the
HIGH RPM position.
9-24. SINGLE-ENGINE GO-AROUND.
a. Engine Fire. The following procedures shall
be performed in case of engine fire:
(1) Engine/nacelle fire during start or
ground operations. If engine/nacelle fire is identified
during start or ground operation, perform the following:
1. Propeller levers - FEATHER.
Once flaps are fully extended, a singleengine go-around may not be possible
9-8
Change 3
2 . Condition levers - FUEL CUTOFF.
3. Fire pull handle - Pull.
TM 55-1510-219-10
CAUTION
If fire extinguisher has been used to extinguish
an engine fire, do not attempt to restart until
maintenance personnel have inspected the
aircraft and released it for flight.
The extinguisher agent (Bromochlorodifluoromethane) in the fire extinguisher can produce
toxic effects if inhaled.
1. Fight the fire.
4. Push to extinguish switch - Push.
2. Land as soon as possible.
5. Master switch - OFF.
(2) Engine fire in flight (fire pull handle light
illuminated). If an engine fire is suspected in flight,
perform the following:
1. Power lever - IDLE.
c. Wing Fire. There is little that can be done to control
a wing fire except to shut off fuel and electrical systems
that may be contributing to the fire, or which could
aggravate it. Diving and slipping the aircraft away from
the burning wing may help. If a wing fire occurs, perform
the following:
2. Fire pull handle out - Advance power.
1. Perform engine shutdown on affected side.
3. Fire pull handle light still illuminated Perform engine fire in flight procedures
(identified).
2. Land as soon as possible.
NOTE
Flight into the sun at high aircraft pitch attitude
may actuate the fire warning system. Lowering
the nose and/or changing headings will confirm
a warning system failure caused by sun rays.
(3) Engine fire in flight (identified). If an engine
fire is confirmed in flight, perform the following:
d. Electrical Fire. Upon noticing the existence or
indications of an electrical fire, turn off all affected
electrical circuits, if known. If electrical fire source is
unknown, perform the following:
1. Crew oxygen - 100%.
2. Master switch - OFF (Visual Conditions Only).
3. All nonessential electrical equipment - OFF.
NOTE
Due to the possibilities of fire warning system
malfunctions, the fire should be visually
identified before the engine is secured and the
extinguisher actuated.
1. Power lever - IDLE.
2. Propeller lever - FEATHER.
3. Condition lever - FUEL CUT-OFF.
With loss of DC electrical power, the aircraft
will depressurize. All electrical instruments,
with the exception of the Prop RPM, N1 RPM,
and TGT gages will be inoperative.
4. Battery switch - ON.
5. Generator switches (individually) - RESET, then
ON.
6. Circuit breakers - Check for indication of
defective circuit.
4. Fire pull handle - Pull.
CAUTION
5. Fire extinguisher - Actuate as required.
6. Engine cleanup - Perform.
As each electrical switch is returned to ON, note
loadmeter reading and check for evidence of
fire.
b. Fuselage Fire. If a fuselage fire occurs, perform the
following:
Change 3
9-9
TM 55-1520-219-10
7. Essential electrical equipment - On (individually
until fire source is isolated).
8. Land as soon as practicable.
e. Smoke and Fume Elimination. To eliminate smoke
and fumes from the aircraft, perform the following:
c. Nacelle Fuel Leak. If nacelle fuel leaks are evident,
perform the following:
1. Perform engine shutdown.
2. Fire pull handle - Pull.
3. Land as soon as practicable.
1. Crew oxygen - 100% and on.
2. Bleed air valve switches - ENVIRO OFF.
3. Vent blower switch - AUTO.
4. Aft vent blower switch - OFF.
d. Fuel Crossfeed. Fuel crossfeed is normally used
only during single-engine operation. The fuel from the
dead engine side may be used to supply the live engine by
routing the fuel through the crossfeed system. During
extended flights, this method of fuel usage will provide a
more balanced lateral load condition in the aircraft. For
fuel crossfeed. use the following procedure:
5. Cabin temperature mode switch - OFF.
1. AUX TRANSFER switches - AUTO.
6. If smoke and fumes are not eliminated, CABIN
PRESS switch - DUMP.
NOTE
Opening storm window (after depressurizing)
will facilitate smoke and fume removal.
NOTE
With the FIRE PULL handle pulled, the fuel in
the auxiliary tank for that side will not be
available (usable) for crossfeed.
2. Standby pumps - OFF.
7. Engine oil pressure - Monitor.
3. Crossfeed switch - As required.
9-27. FUEL SYSTEM.
4. Fuel crossfeed light illuminated - Check.
a. Fuel Press Warning Light Illuminated. Illumination
of the #1 or #2 FUEL PRESS warning light usually
indicates failure of the respective engine-driven boost
pump. Perform the following:
1. Standby pump switch - ON.
With the FIRE PULL handle pulled, the FUEL
press light will remain illuminated on the side
supplying fuel.
2. Fuel pressure light out - Check.
5. Fuel pressure light extinguished - Check.
3. Fuel pressure light still on - Record unboosted
time.
6. Fuel quantity - Monitor.
b. No Fuel Transfer Caution Light Illuminated.
Illumination of the #1 or #2 NO FUEL XFR light with fuel
remaining in the respective auxiliary fuel tank indicates a
failure of that automatic fuel transfer system. Proceed as
follows:
1. AUX TRANSFER switch (affected side)
OVERRIDE.
2. Auxiliary fuel quantity - Monitor.
3. AUX TRANSFER switch (after respective
auxiliary fuel has completely transferred) AUTO.
9-10
NOTE
Change 3
TM 55-1510-219-10
e. NAC LOW Light Illuminated. Illumination of the #1
or #2 NAC LOW caution light indicates that the affected
tank has 20 minutes remaining at sea level, maximum
cruise power consumption rate. Proceed as follows:
WARNING
Failure of the fuel tank venting system will
prevent the fuel in the wing tanks from gravity
feeding into the nacelle tank. Fuel vent system
failure may be indicated by illumination of the
#1 or #2 NAC LOW caution light with greater
than 20 minutes of usable fuel indicated in the
main tank fuel system. The total usable fuel
remaining in the main fuel supply system with
the LOW FUEL caution light illuminated may
be as little as 114 pounds, regardless of the total
fuel quantity indicated. Continued flight may
result in engine flameout due to fuel starvation.
1. Twenty minutes fuel remaining - Confirm.
2. Land as soon as possible.
9-28. ELECTRICAL SYSTEM EMERGENCIES.
a. DC GEN Light Illuminated. Illumination of a #1 or
#2 DC GEN caution light indicates failure.
(Go to page 9-11)
Change 3
9-10.1/(9-10.2 blank)
TM 55-1510-219-10
2. Other engine instruments-Monitor.
of a generator or one of its associated circuits
(generator control unit). If one generator system
becomes inoperative, all nonessential electrical
equipment should be used judiciously to avoid
overloading the remaining generator. The use of
accessories which create a very high drain should be
avoided. If both generators are shut off due to either
generator system failure or engine failure, all
nonessential equipment should be turned off to
preserve battery power for extending the landing
gear and wing flaps.
When a DC GEN light
illuminates, perform the following:
1.
2.
3.
4.
f: Circuit Breaker Tripped. If the circuit breaker
is for a nonessential item, do not reset in flight. If the
circuit breaker is for an essential item, the circuit
breaker may be reset once. If a bus feeder circuit
breaker (on the overhead circuit breaker panel) trips,
a short is indicated. Do not reset in flight. If a circuit
breaker trips, perform as follows:
1. BUS FEEDER breaker tripped-Do not reset.
2. Nonessential circuit-Do not reset.
3. Essential circuit-Reset once.
Generator switch-OFF, RESET, then ON.
Generator switch (no reset)-OFF.
Mission control switch-OVERRIDE.
Operating loadmeter-100% maximum.
g. BATTERY CHARGE Light Illuminated.
1. Battery volt-amp meter Check.
b. Both DC GEN Lights Illuminated.
a. Amp reading above 7 amps and
decreasing Monitor.
1. All nonessential equipment-OFF.
2. Land as soon as practicable.
b. Amp reading above 7 amps and
increasing Battery switch OFF.
c. Excessive Loadmeter Indication (Over
100%). If either loadmeter indicates over 100%,
perform the following:
2. Battery switch ON (for landing prior to
gear/flap extension).
9-29. EMERGENCY DESCENT.
1. Battery switch-OFF (monitor loadmeter).
2. Loadmeter
over
100%-Nonessential
electrical equipment OFF.
3. Loadmeter under 100%-BATT switch ON.
Emergency descent is a maximum effort in
which damage to the aircraft must be considered
secondary to getting the aircraft down. The following
procedure assumes the structural integrity of the
aircraft and smooth flight conditions. If structural
integrity is in doubt, limit speed as much as possible,
reduce rate of descent if necessary, and avoid high
maneuvering loads.
For emergency descent,
perform the following:
d. Inverter Light Illuminated. Illumination of the
#1 or #2 AIRCRAFT INVERTER caution light
indicates failure of the affected inverter. When
either inverter fails, the total aircraft load is
automatically switched to the remaining inverter.
When a #1 or #2 AIRCRAFT INVERTER light
illuminates, perform the following:
NOTE
Windshield defogging may be
required.
1. Affected AIRCRAFT INVERTER switch-OFF.
e. INST AC Light Illuminated. Illumination of
the INST AC warning light indicates that 26 VAC
power is not available. All items connected to the 26
VAC bus will be inoperative. Under these conditions,
power must be controlled by indications of the N1
and TGT gages. Perform the following:
1.
2.
3.
4.
5.
1. N1 and TGT indications-Check.
9-11
Power lever-IDLE.
Propeller lever-HIGH RPM.
Flap lever-APPROACH.
Gear-DN.
Airspeed-184 KIAS maximum.
TM 55-1510-219-10
may be stowed and the landing
gear
retracted
electrically.
Rotate the alternate engage
handle counterclockwise and
push it down.
Stow the
handle, push in the LANDING
GEAR RELAY circuit breaker
on the overhead circuit breaker
panel and retract the gear in
the normal manner with the
landing gear switch.
9-30. LANDING EMERGENCIES.
WARNING
Structural damage may exist
after landing with brake, tire, or
landing gear malfunctions.
Under no circumstances shall
an attempt be made to inspect
the aircraft until jacks have
been installed.
a. Landing Gear Unsafe Indication. Should one
or more of the landing gear fail to indicate a safe
condition, the following steps should be taken before
proceeding manually to extend the gear.
c. Gear-up Landing (All Gear Up or Unlocked).
Due to decreased drag with the gear up, the
tendency will be to overshoot the approach. The
center-of-gravity with the gear retracted is aft of the
main wheels. This condition will allow the aircraft to
be landed with the gear retracted and should result in
a minimum amount of structural damage to the
aircraft, providing the wings are kept level. It is
recommended that the fuel load be reduced and the
landing made with flaps fully extended on a hard
surface runway. Landing on soft ground or dirt is not
recommended as sod has a tendency to roll up into
chunks, damaging the underside of the aircraft's
structure.
When fuel load has been reduced,
prepare for a gear-up landing as follows:
1. LANDING GEAR RELAY circuit
breaker Check in.
2. Gear lights Check.
3. Gear handle DN.
4. Manual gear extension As required.
NOTE
If gear continues to indicate
unsafe, attempt to verify
position visually by other
aircraft.
1.
2.
3.
4.
5.
b. Landing Gear Emergency Extension.
CAUTION
After an emergency landing
gear extension has been made,
do not stow the gear ratchet
handle or move any landing
gear controls or reset any
switches or circuit breakers
until
the
cause
of
the
malfunction
has
been
corrected.
6.
7.
8.
9.
10.
11.
1. Airspeed 130 KIAS.
2. LANDING GEAR RELAY circuit
breaker Out.
3. Gear handle DN.
4. Landing gear alternate engage handle
Lift and turn clockwise to the stop.
5. Alternate landing gear extension
handle Pump.
6. Gear lights illuminated Check.
12.
Crew emergency briefing Complete.
Loose equipment Stowed.
Bleed air valves ENVIRO OFF.
Cabin pressure switch DUMP.
Cabin emergency hatch Remove and
stow.
Seat belts and harnesses Secured.
Landing gear alternate engage handle
Disengaged.
Alternate landing gear extension
handle Stowed.
Gear relay circuit breaker In.
Gear handle UP.
Nonessential electrical equipment
OFF.
Flaps As required (DOWN for landing).
NOTE
Fly a normal approach to
touchdown.
After landing,
accomplish the following:
13. Condition levers FUEL CUTOFF.
14. Fire pull handles Pull.
15. Master switch OFF.
NOTE
After
a
practice
manual
extension, the alternate handle
9-12
TM 55-1510-219-10
d. Landing With Nose Gear Unsafe. If the LDG
GEAR CONTROL warning light is illuminated and
the nose GEAR DOWN LIGHT shows an unsafe
condition, the nose gear is probably not locked down,
and the gear position should be checked visually by
another aircraft, if possible. If all attempts to lock the
nose gear fail, a landing should be made with the
main gear down and locked. Hold the nose off the
runway as long as possible and do not use brakes.
Use the following procedures:
NOTE
Fly a normal approach to
touchdown.
After landing,
accomplish the following:
9. Condition levers FUEL CUTOFF.
10. Fire pull handle Pull.
11. Master switch OFF.
f Landing With Flat Tire(s). If aware that a
main gear tire(s) is flat, a landing close to the edge of
the runway opposite the flat tire will help avoid
veering off the runway. If the nose wheel tire is flat,
use minimum braking.
1.
2.
3.
4.
5.
Crew emergency briefing Complete.
Loose equipment Stowed.
Bleed air valves ENVIRO OFF.
Cabin pressure switch DUMP.
Cabin emergency hatch Remove and
stow.
6. Seat belts and harnesses Secured.
7. Nonessential electrical equipment
OFF.
9-31. CRACKED WINDSHIELD.
a. External Crack. If an external windshield
crack is noted, no action is required in flight.
NOTE
Heating elements may be
inoperative in areas of crack.
NOTE
Fly a normal approach to
touchdown.
After landing,
accomplish the following:
b. Internal Crack. If an internal crack occurs,
perform the following:
8. Condition levers FUEL CUTOFF.
9. Fire pull handle Pull.
10. Master switch OFF.
1. Descend Below 25,000 feet.
e. Landing With One Main Gear Unsafe. If one
main landing gear fails to extend, retract the other
gear and make a gear-up landing. If all efforts to
retract the extended gear fail, land the aircraft on a
hard runway surface, touching down on the same
edge of the runway as the extended gear. Roll on
the down and locked gear, holding the opposite wing
up and the nose gear straight as long as possible. If
the gear has extended, but is unsafe, apply brakes
lightly on the unsafe side to assist in locking the
gear. If the gear has not extended or does not lock,
allow the wing to lower slowly to the runway. Use the
following procedures:
2. Cabin Pressure Reset pressure
differential to 4 PSI or less within 10
minutes.
9-32. CRACKED CABIN WINDOW
If crack in a single ply of the external cabin
window occurs, unpressurized flight may be
continued. Proceed as follows:
1. Oxygen As required.
2. Cabin pressurization Depressurize.
3. Descend As required.
1.
2.
3.
4.
5.
Crew emergency briefing Complete.
Loose equipment Stowed.
Bleed air valves ENVIRO OFF.
Cabin pressure switch DUMP.
Cabin emergency hatch Remove and
stow.
6. Seat belts and harnesses Secured.
7. Nonessential
electrical equipment
OFF.
8. Touchdown On safe main gear first.
NOTE
If both plys of the external
cabin window have developed
cracks, the aircraft shall not be
flown once landed, without
proper
ferry
flight
authorization.
9-13
TM 55-1510-219-10
7.
8.
9.
10.
11.
9-33. DITCHING.
If a decision to ditch is made, immediately
alert all crewmembers to prepare for ditching. Plan
the approach into the wind if the wind is high and the
seas are heavy. If the swells are heavy but the wind
is light, land parallel to the swells. Set up a minimum
rate descent (power on or off, as the situation
dictates airspeed 110-120 KIAS). Do not try to flare
as in a normal landing, as it is very difficult to judge
altitude over water, particularly in a slick sea.
Leveling off too high may cause a nose low "drop in,"
while having the tail too low on impact may result in
the aircraft pitching forward and "digging in. " Expect
more than one impact shock and several skips
before the final hard shock. There may be nothing
but spray visible for several seconds while the
aircraft is decelerating. To prevent cartwheeling, it is
important that the wings be level when the aircraft
hits the water. After the aircraft is at rest, supervise
evacuation of passengers and exit the aircraft as
quickly as possible. In a planned ditching, the life
raft and first-aid kits should be secured close to the
cabin emergency hatch for easy access when
evacuating; however, do not remove the raft from its
carrying case inside the aircraft. After exiting the
aircraft, keep the raft away from any damaged
surfaces which might tear or puncture the fabric.
The length of time that the aircraft will float depends
on the fuel level and the extent of aircraft damage
caused by the ditching. Refer to Figure 9-3 for body
positions during ditching, and Table 9-1 for personnel
procedures.
Figure 9-4 shows wind swell
information. Perform the following procedures:
9-34. FLIGHT CONTROLS MALFUNCTION.
Use the following procedures, as applicable,
for flight control malfunctions.
a. Unscheduled Rudder Boost Activation.
Rudder boost operation without a large variation of
power between engines indicates a failure of the
system. Perform the following:
1. Rudder boost-OFF.
NOTE
The rudder boost system may
not operate when the brake
deice system is in use.
Availability of the rudder boost
system will be restored to
normal when the BRAKE
DEICE switch is turned OFF.
IF CONDITION PERSISTS:
2. Bleed air valve switches-PNEU &
ENVIRO-OFF.
3. Rudder trim-Adjust.
b. Unscheduled Electric Elevator Trim. In the
event of unscheduled electric elevator trim, perform
the following:
WARNING
Do not unstrap from the seat
until all motion stops. The
possibility
of
injury
and
disorientation requires that
evacuation not be attempted
until the aircraft comes to a
complete stop.
1.
2.
3.
4.
5.
6.
Gear-UP.
Flaps-DOWN.
Nonessential electrical equipment-OFF.
Approach-Normal, power on.
Emergency lights-As required. Ditching
1. ELEV TRIM switch OFF.
2. ELEC TRIM circuit breaker OUT.
9-35. BAILOUT.
When the decision has been made to
abandon the aircraft in flight. the pilot will give the
warning signal. Exit from the aircraft will be through
the main entrance door, and in the departure
sequence using the exit routes as indicated in Figure
9-1.
Proceed as follows if bailout becomes
necessary:
Radio calls/transponder-As required.
Crew emergency briefing-As required.
Bleed air valves-ENVIRO OFF.
Cabin pressurization switch-DUMP.
Cabin emergency hatch-Remove and stow.
Seat belts and harnesses-Secured.
1. Notify copilot to prepare to bail out.
2. Distress message Transmit.
3. Voice security ZEROIZE.
9-14
TM 55-1510-219-10
Table 9-1. Personnel Procedures During Ditching
4. Transponder-7700.
9. Cabin pressure switch-DUMP.
5. Flaps-DOWN.
10. Radio cord, oxygen hose, harnesses and
seat belt-Disconnect.
6. Airspeed-100 KIAS.
11. Parachute-Attach to harness.
7. Trim-As required.
12. Cabin door-Open.
8. Autopilot-Engage.
13. Abandon the aircraft.
9-15
TM 55-1510-219-10
BRACE POSITIONS
Figure 9-3. Emergency Body Positions
Figure 9-4. Wind Swell Ditch Heading Evaluation
9-16
TM 55-1510-219-10
APPENDIX A
REFERENCES
Reference information for the subject material contained in this manual can be found in the following
publications.
AR 70-50
AR 95-1
AR 95-3
AR 380-40
AR 385-40
AR 700-26
FAR Part 91
FM 1-230
TB 55-9150-200-24
TB AVN 23-13
TB MED 501
Designating and Naming Defense Equipment, Rockets, and Guided Missiles
Army Aviation - General Provisions and Flight Regulations
Weight and Balance - Army Aircraft
Safeguarding COMSEC Information
Accident Reporting and Records
Aircraft Designation System
General Operating and Flight Rules
Meteorology for Army Aviators
Engine and Transmission Oils, Fuels, and Additives for Army Aircraft
Anti-icing, Deicing and Defrosting Procedures for Parked Aircraft
Noise and Conservation of Hearing
TM 9-1095-206-13&P
Operator's Aviation Unit Maintenance and Aviation Intermediate
Maintenance Manual (Including Repair Parts and Special Tools List) to
Dispenser, General Purpose Aircraft: M-130
(C) TM 11-5825-252-15
Operator, Organizational, DS, GS, and Deport Maintenance Manual: RC12D Aircraft Mission Equipment, (V)
TM 11-5841-291-12
Operator and Organizational Maintenance Manual, Radar Warning System,
AN/APR-44(V) 1
TM 11-5841-283-20
Organizational Maintenance Manual for Detection Set, Radar Signal AN/
APR-39(V) 1.
TM 11-6140-203-14-2
Operator's Organizational, Direct Support, General Support and Depot
Maintenance Manual Including Repair Parts and Special Tools List: Aircraft
Nickel-Cadmium Batteries
TM 11-6940-214-12
Operator and Organizational Maintenance Manual, Simulator, Radar Signal,
SM-756/APR-44(V)
TM 55-410
Aircraft Maintenance, Servicing and Ground Handling Under Extreme
Environmental Conditions
TM 55-1500-314-25
Handling, Storage, and Disposal of Army Aircraft Components Containing
Radioactive Materials
TM 55-1500-204-25/1
General Aircraft Maintenance Manual
TM 750-244-1-5
Procedures for the Destruction of Aircraft and Associated Equipment to
Prevent Enemy Use
A-1/(A-2 blank)
TM 55-1510-219-10
APPENDIX B
ABBREVIATIONS AND TERMS
For the purpose of this manual, the following abbreviations and terms apply. See appropriate technical
manuals for additional terms and abbreviations.
AIRSPEED TERMINOLOGY.
CAS
Calibrated airspeed is indicated airspeed corrected for position and instrument error.
FT/MIN
Feet per minute.
GS
Ground speed, though not an airspeed, is directly calculable from true airspeed if
the true wind speed and direction are known.
IAS
Indicated airspeed is the speed as shown on the airspeed indicator and assumes no
error.
KT
Knots.
TAS
True airspeed is calibrated airspeed corrected for temperature, pressure, and compressibility effects.
Va
Maneuvering speed is the maximum speed at which application of full available
aerodynamic control will not overstress the aircraft.
Vf
Design flap speed is the highest speed permissible at which wing flaps may be actuated.
Vfe
Maximum flap extended speed is the highest speed permissible with wing flaps in
a prescribed extended position.
Vle
Maximum landing gear extended speed is the maximum speed at which an aircraft
can be safely flown with the landing gear extended.
Vlo
Maximum landing gear operating speed is the maximum speed at which the landing
gear can be safely extended or retracted.
Lift off speed (takeoff airspeed).
Vlof
Vmca
The minimum flight speed at which the aircraft is directionally controllable as determined in accordance with Federal Aviation Regulations. Aircraft Certification conditions include one engine becoming inoperative and windmilling; a 5°bank towards the operative engine; takeoff power on operative engine; landing gear up;
flaps up; and most rearward CG. This speed has been demonstrated to provide satisfactory control above power off stall speed (which varies with weight, configuration, and flight attitude).
Vmo
Maximum operating limit speed.
Vne
Never exceed speed.
Vr
Rotation speed.
Vs
Power off stalling speed or the minimum steady flight speed at which the aircraft
is controllable.
Vso
Stalling speed or the minimum steady flight speed in the landing configuration.
Vsse
The safe one-engine inoperative speed selected to provide a reasonable margin
against the occurrence of an unintentional stall when making intentional engine
cuts.
B-1
TM 55-1510-219-10
Vx
Vxse
Vy
Vyse
Best angle of climb speed.
Best single-engine angle of climb speed.
Best rate of climb speed.
The best single engine rate of climb speed.
METEOROLOGICAL TERMINOLOGY.
Altimeter Setting
°C
°F
FAT
Indicated Pressure
Altitude
Barometric pressure corrected to sea level.
Degrees Celsius.
Degrees Fahrenheit.
Free air temperature is the free air static temperature, obtained either from ground
meteorological sources or from inflight temperature indications adjusted for compressibility effects.
The number actually read from an altimeter when the barometric scale (Kollsman
window) has been set to 29.92 inches of mercury (1013 millibars).
ISA
International standard atmosphere in which:
a. The air is a dry perfect gas;
b. The temperature at sea level is 59 degrees Fahrenheit, 15 degrees Celsius;
c. The pressure at sea level is 29.92 inches Hg;
d. The temperature gradient from sea level to the altitude at which the temperature
is -69.7 degrees Fahrenheit is -0.003566 Fahrenheit per foot and zero above that altitude.
Pressure Altitude
(press alt)
Indicated pressure altitude corrected for altimeter error
SL
Sea level.
Wind
The wind velocities recorded as variables on the charts of this manual are to be understood as the headwind or tailwind components of the actual winds at 50 feet
above runway surface (tower winds).
Beta Range
The region of the power lever control which is aft of the idle stop and forward of
reversing range where blade pitch angle can be changed without a change of gas
generator RPM.
Cruise Climb
Is the maximum power approved for normal climb. This power is torque or temperature (ITT) limited.
High Idle
Obtained by placing the condition lever in the HIGH IDLE position.
HP
Horsepower.
Low Idle
Obtained by placing the condition lever in the LO IDLE position.
Maximum Cruise
Power
Is the highest power rating for cruise and is not time limited.
Maximum Power
The maximum power available from an engine for use during an emergency operation.
Normal Rated Climb
Power
The maximum power available from an engine for continuous normal climb operations.
Normal Rated Power
The maximum power available from an engine for continuous operation in cruise
(with lower ITT limit than normal rated climb power).
B-2
TM 55-1510-219-10
Reverse Thrust
Obtained by lifting the power levers and moving them aft of the beta range.
RPM
Revolutions per minute.
Takeoff Power
The maximum power available from an engine for takeoff, limited to periods of five
minutes duration.
CONTROL AND INSTRUMENT TERMINOLOGY.
Condition Lever
The fuel shut-off lever actuates a valve in the fuel control unit which controls the
(Fuel Shut-off Lever) flow of fuel at the fuel control outlet and regulates the idle range from LO to HIGH.
Interstage Turbine
Temperature (ITT)
Eight probes wired in parallel indicate the temperature between the compressor and
power turbines.
N1 Tachometer (Gas The tachometer registers the RPM of the gas generator with 100% representing a gas
Generator RPM)
generator speed of 37,500 RPM.
Power Lever (Gas
Generator N1 RPM)
Propeller Control
Lever (N2 RPM)
Propeller Governor
Torquemeter
This lever serves to modulate engine power from full reverse thrust to takeoff. The
position for idle represents the lowest recommended level of power for flight operation.
This lever requests the control to maintain RPM at a selected value and, in the maximum decrease RPM position, feathers the propeller.
This governor will maintain the selected propeller speed requested by the propeller
control lever.
The torquemeter system determines the shaft output torque. Torque values are obtained by tapping into two outlets on the reduction gear case and recording the differential pressure from the outlets.
GRAPH AND TABULAR TERMINOLOGY.
AGL
Above ground level.
Best Angle of Climb
The best angle-of-climb speed is the airspeed which delivers the greatest gain of altitude in the shortest possible horizontal distance with gear and flaps up.
Best Rate of Climb
The best rate-of-climb speed is the airspeed which delivers the greatest gain of altitude in the shortest possible time with gear and flaps up.
Clean Configuration
Gear and flaps up regardless of mission antenna installation.
Demonstrated
Crosswind
The maximum 90°crosswind component for which adequate control of the aircraft
during takeoff and landing was actually demonstrated during certification tests.
Gradient
The ratio of the change in height to the horizontal distance, usually expressed in
percent.
Landing Weight
The weight of the aircraft at landing touchdown.
Maximum Zero Fuel
Weight
Any weight above the value given must be loaded as fuel.
MEA
Minimum enroute altitude.
Obstacle Clearance
Climb Speed
Obstacle clearance climb speed is a speed near Vx and Vy, 1.1 times power off stall
speed, or 1.2 times minimum single-engine stall-speed, whichever is higher.
Ramp Weight
The gross weight of the aircraft before engine start. Included is the takeoff weight
plus a fuel allowance for start, taxi, run up and takeoff ground roll to liftoff.
B-3
TM 55-1510-219-10
Route Segment
A part of a route. Each end of that part is identified by:
a. A geographic location; or
b. A point at which a definite radio fix can be established.
Service Ceiling
The altitude at which the minimum rate of climb of 100 feet per minute can be attained for existing aircraft weight.
Takeoff Weight
The weight of the aircraft at liftoff from the runway.
WEIGHT AND BALANCE TERMINOLOGY.
Arm
The distance from the center of gravity of an object to a line about which moments
are to be computed.
Approved Loading
Envelope
Those combinations of aircraft weight and center of gravity which define the limits
beyond which loading is not approved.
Basic Empty Weight
The aircraft weight with unusable fuel, full oil, and full operating fluids.
Center-of-Gravity
A point at which the weight of an object may be considered concentrated for weight
and balance purposes.
CG Limits
CG limits are the extremes of movement which the CG can have without making
the aircraft unsafe to fly. The CG of the loaded aircraft must be within these limits
at takeoff, in the air, and on landing.
Datum
A vertical plane perpendicular to the aircraft longitudinal axis from which fore and
aft (usually aft) measurements are made for weight and balance purposes.
Engine Oil
That portion of the engine oil which can be drained from the engine.
Empty Weight
The aircraft weight with fixed ballast, unusable fuel, engine oil, engine coolant, hydraulic fluid, and in other respects as required by applicable regulatory standards.
Landing Weight
The weight of the aircraft at landing touchdown.
Maximum Weight
The largest weight allowed by design, structural, performance or other limitations.
Moment
A measure of the rotational tendency of a weight, about a specified line, mathematically equal to the product of the weight and the arm.
Standard
Weights corresponding to the aircraft as offered with seating and interior, avionics,
accessories, fixed ballast and other equipment specified by the manufacturer as
composing a standard aircraft.
Station
The longitudinal distance from some point to the zero datum or zero fuselage station.
Takeoff Weight
The weight of the aircraft at liftoff.
Unusable Fuel
The fuel remaining after consumption of usable fuel.
Usable Fuel
That portion of the total fuel which is available for consumption as determined in
accordance with applicable regulatory standards.
Useful Load
The difference between the aircraft ramp weight and basic empty weight.
B-4
TM 55-1510-219-10
MISCELLANEOUS ABBREVIATIONS
Deg
DN
FT
FT LB
GAL
HR
kH
RC
Degrees
Down
Foot or feet
Foot-pounds
Gallons
Hours
Kilohertz
Rate of climb
LB
MAX
MH
MIN
NAUT
NM
PSI
B-5/(B-6 blank)
Pounds
Maximum
Megahertz
Minimum
Nautical
Nautical miles
Pounds per square inch
TM 55-1510-219-10
INDEX
Subject
Paragraph, Figure,
Table Number
A
Accelerate-Go Flight Path Example ........... 7-8, F7-1
Accelerate-Stop (Flaps 0%) ...................... 7-6
Accelerometer .......................................... 2-82
AC Power Supply ...................................... 2-73
Action Codes and Recommended Action . . T3-4
ADF Control Panel (DF-203) ..................... F3-16
After Emergency Action ............................. 9-4
After Landing ............................................ 8-43
After Takeoff ............................................. 8-35
Aileron High Torque Test Switch
and Annunciator ....................................... F3-19
Air Conditioning System ........................... 2-68
Air Induction Systems - General ............... 2-18
Aircraft Compartments and Stations . ........ 6-3, F6-1
Aircraft Designation System ...................... 1-11
Aircraft Systems Emergencies .................. 9-1
Airspeed Indicators ................................... 2-78
Airspeed Limitations ................................. 5-20
Altitude Limitations ................................... 5-29
Ammunition ............................................... 4-5
Annunciator Panels.................................... T2-6
Antenna Deice System .............................. 2-53
Anti-Icing, Deicing and Defrosting ..............
Protection ................................................. 2-96
Appendix A, References............................. 1-4
Appendix B, Abbreviations and Terms........ 1-5
Application of External Power ................... 2-97
Approved Fuels ........................................ T2-8
Army Aviation Safety Program .................. 1-7
Arrival Briefing .......................................... 8-60
Audio Control Panels ................................ 3-6, F3-1
Autoignition System .................................. 2-31
Automatic Approaches .............................. 8-48
Automatic Direction Finder (DF-203) ......... 3-25
Automatic Flight Control System ............... 3-27
Autopilot/Flight Director Annunciator..........
Panel ........................................................ F3-20
Autopilot Limitations ................................. 5-9
Autopilot Mode Selector Panel ...................
(614E-42A) ............................................... F3-18
Autopilot Pitch-Turn Control Panel ............ F3-21
Auxiliary Fuel Tank Mechanical Gage ....... F2-13
Avionics Equipment Configuration ............ 3-2
Paragraph, Figure,
Table Number
Subject
B
Backup VOW AN/ARC-164 .................
Baggage Moments ..............................
Bailout ................................................
Bank and Pitch Limits .........................
Before Exterior Check .........................
Before Landing ....................................
Before Leaving Aircraft ........................
Before Starting Engines ......................
Before Takeoff ....................................
Before Taxiing .....................................
Brake Deice Limitations .......................
Brake Deice System ............................
3-8
T6-1
9-35
5-28
8-6
8-40
8-45
8-22
8-32
8-29
5-11
2-57
C
Cabin and Cargo Doors........................
Cabin Door Caution Light ....................
Cabin Door Caution Light Illuminated ..
Cabin Door Limitations ........................
Cabin Pressure Limits .........................
Center of Gravity Limitations ...............
Center of Gravity Limits (Landing .........
Gear Down) ........................................
Center of Gravity Moments .................
Chaff/Flare Dispenser Control Panel ...
Chart C - Basic Weight and Balance ....
Record DD Form 365-3 .......................
Charts and Forms ................................
Checklist .............................................
Chip Detector Warning Light Illuminated .
Cigarette Lighters and Ash Trays ........
Class ..................................................
Climb ..................................................
Cockpit ...............................................
Cold Weather Operations ....................
Comments Pertinent To the Use of.......
Performance Graphs ...........................
Communications - Description ............
Condition Levers .................................
Conditions - Performance ....................
Control Locks .....................................
Control Wheels ...................................
Copilot's Altimeter ...............................
Copilot's Gyro Horizon Indicator ..........
INDEX-1
F2-7
2-10
9-20
5-39
5-34
5-18,6-11
T6-4, T6-5
T6-3
F4-2
6-7
6-5
8-4
9-15
2-63
6-2
8-36
F2-9
8-54
7-13
3-5
2-23
7-2
2-17
2-317 F2-16
2-80
3-21, F3-14
TM 55-1510-219-10
Paragraph, Figure,
Table Number
Subject
C
Copilot's Horizontal Situation
Indicator (331A-6P) ..................................
Cracked Cabin Window ............................
Cracked Windshield .................................
Crew Briefings ..........................................
Crossfeed Fuel Flow ................................
Cruise .......................................................
Cylinder Capacity vs Pressure and
Temperature ............................................
F3-11
9-32
9-31
8-2, 8-58
F2-15
8-37
F2-20
D
DC Electrical System ...............................
DC Power Supply .....................................
Definition of Landing Terms .....................
Defrosting/Defogging System ...................
Departure Briefing ....................................
Descent ...................................................
Descent-Arrival ........................................
Desert Operation and Hot Weather
Operation .................................................
Destruction of Army Materiel To Prevent
Enemy Use ..............................................
Dimensions ..............................................
Ditching ...................................................
Diving ......................................................
Draining Moisture From Fuel System .......
Dual NAV I - NAV 2 Control Panel ............
Duct Overtemp Caution Light Illuminated .
F2-22
2-72
9-3
2-51
8-59
8-38
8-39
8-55
1-8
2-3
9-33, T9-1
8-51
2-89
F3-15
9-16
E
Electrical Power Supply and Distribution
System - Description ................................
Electrical System Emergencies ................
Electric Toilet ...........................................
Emergency Body Positions .......................
Emergency Descent .................................
Emergency Entrance ................................
Emergency Equipment - Description .........
Emergency Exits and Equipment...............
Emergency Lighting .................................
Emergency Locator Transmitter (ELT) ......
Engine Bleed Air System Malfunction .......
Engine Chip Detection System .................
Engine Cleanup ........................................
2-71
9-28
2-64
F9-3
9-29
9-6
2-13
9-5, F9-1
2-76
3-15, F3-8
9-18
2-29
9-9
Subject
Paragraph, Figure,
Table Number
E
Engine Clearing
Engine Compartment Cooling ..................
Engine Fire Detection System ..................
Engine Fire Extinguisher Gage Pressure ..
Engine Fire Extinguisher System ..............
Engine Fuel Control System .....................
Engine Ice Protection Systems .................
Engine Ignition System ............................
Engine Instruments ..................................
Engine Limitations ...................................
Engine Malfunction ...................................
Engine Operating Limitations ...................
Engine Restart During Flight ....................
Engine Runup ..........................................
Engines - Description ...............................
Engine Shutdown .....................................
Engine Shutdown In Flight .......................
Engine Starter-Generators .......................
Entrance and Exit Provisions ...................
Environmental Controls ...........................
Environmental System .............................
Exceeding Operational Limits ...................
Exhaust and Propeller Danger Areas ........
Exhaust Danger Area ...............................
Explanation of Change Symbols ..............
Extent of Coverage ..................................
Exterior Inspection ...................................
Exterior Inspection - Center Section,
Left Side ..................................................
Exterior Inspection - Empennage, Area 5 .
Exterior Inspection - Fuselage Left Side,
Area 6 .....................................................
Exterior Inspection - Fuselage Right Side,
Area 6 .....................................................
Exterior Inspection - Fuselage Underside .
Exterior Inspection - Left Engine and
Propeller ..................................................
Exterior Inspection - Left Main
Landing Gear ...........................................
Exterior Inspection - Left Wing, Area 1 ......
Exterior Inspection - Nose Section, Area 2
Exterior Inspection - Right Engine and
Propeller ..................................................
Exterior Inspection - Right Main
Landing Gear ...........................................
Exterior Inspection - Right Wing, Area 3 ...
Exterior Inspection - Right Wing,
Center Section ..........................................
INDEX-2
8-26
2-17
2-25
T2-1
2-26
2-21
2-20
2-30
2-33
5-14
9-7
T5-1
9-10, 9-11
8-31
2-16
8-44
9-8
2-32
2-9
2-70
F2-21
5-3
F2-5
2-6
1-10
6-1
F8-1
8-11
8-19
8-20
8-18
8-12
8-10
8-9
8-8
8-13
8-15
8-16
8-17
8-14
TM 55-1510-219-10
Paragraph, Figure,
Table Number
Subject
Paragraph, Figure,
Table Number
Subject
G
E
Exterior Lighting . . . . . . . . . . . . . . . . . . . . . . . . 2-74, F2-27
F
Feathering Provisions . . . . . . . . . . . . . . . . . . . . . . . . . 2-44
Ferry Chair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-36
Ferry Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-36
Filling Fuel Tanks . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-88
Fire . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9-26
First Aid Kits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-14
First Engine Start (Battery Start) . . . . . . . . . . . . . . . . . . 8-23
First Engine Start (GPU Start) ............................ 8-27
Flight Controls - Description . . . . . . . . . . . . . . . . . . . . . . . . 2-37
Flight Controls Lock . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-40
Flight Controls Malfunction ........................... 9-34
Flight Envelope . . . . . . . . . . . . . . . . . . . . . . . . . . F5-2
Flight Planning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-9
Flight Under IMC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-31
Foreign Object Damage Control . . . . . . . . . . . . . . . . 2-19
Forms and Records . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-9
Free Air Temperature (FAT) Gage . . . . . . . . . . . . . . 2-83
Friction Lock Knobs . . . . . . . . . . . . . . . . . . . . . . . . 2-24
Fuel and Oil Data . . . . . . . . . . . . . . . . . . . . . . . . . . 6-10
Fuel Moments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . T6-2
Fuel Handling Precautions . . . . . . . . . . . . . . . . . . . . . 2-87
Fuel Load . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-9
Fuel, Lubricants, Specifications
and Capacities . . . . . . . . . . . . . . . . . . . . . . . . . . . . . T2-7
Fuel Management Panel & Auxiliary
Tank Mechanical Gage . . . . . . . . . . . . . . . . . . . . . . F2-13
Fuel Quantity Data . . . . . . . . . . . . . . . . . . . . . . . . . . . T2-2
Fuel Sample . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-7
Fuel Sump Drain Locations . . . . . . . . . . . . . . . . . . . . . . . T2-3
Fuel Supply System. . . . . . . . . . . . . . . . . . . . . . . . . . . 2-34
Fuel System Anti-Icing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-58
Fuel System Emergencies . . . . . . . . . . . . . . . . . . . . . . . 9-27
Fuel System Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-10
Fuel System Management . . . . . . . . . . . . . . . . . . . . . . . 2-35
Fuel System Schematic ............................... F2-12
Fuel Types . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-90
G
General Exterior Arrangement . . . . . . . . . . . . . . . . . . F2-1
General Interior Arrangement . . . . . . . . . . . . . . . . . . . F2-2
Generator Limits . . . . . . . . . . . . . . . . . . . . . . . 5-17, T5-2
Go Around . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-42
Gravity Feed Fuel Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . F2-14
Ground Handling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-99
Ground Turning Radius . . . . . . . . . . . . . . . . . . . . . . . . . 2-4, F2-4
Gyromagnetic Compass Systems . . . . . . . . . . . . . . . . . 3-22
H
HF Command Set (718 U-5) . . . . . . . . . . . . . . 3-14, F3-7
Hand-Operated Fire Extinguisher . . . . . . . . . . . . . 2-15
Heating System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-67
Horizontal Reference Indicator . . . . . . . . . . . . . . . 3-19, F3-12
Horizontal Situation Indicators . . . . . . . 3-18, F3-10, F3-11
I
INS Control Display Unit (C-IV-E) . . . . . . . . . . . . . . . . . F3-23
INS Mode Selector Unit (C-IV-E) . . . . . . . . . . . . . . . . . F3-22
Ice and Rain (Typical) . . . . . . . . . . . . . . . . . . . . . . . . . 8-57
Icing Limitations (Typical) . . . . . . . . . . . . . . . . . . . . . . . . 5-31A
Icing Limitations (Severe) . . . . . . . . . . . . . . . . . . . . . 5-31B
Icing (Severe) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-57A
Ice Vane Failure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9-17
Immediate Action Emergency Checks . . . . . . . . . . . . . . 9-2
Index . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-6
Inertial Navigation System . . . . . . . . . . . . . . . . . . . . . 3-28
Inflating Tires . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-94
Installation of Protective Covers . . . . . . . . . . . . . . . . . . . 2-100
Instrument Landing System Limits . . . . . . . . . . . . . . . 5-35
Instrument Marking Color Codes . . . . . . . . . . . . . . . 5-6
Instrument Markings . . . . . . . . . . . . . . . . . . . . 5-5, F5-1
Instrument Panel. . . . . . . . . . . . . . . . . . . . . . . . . . . . . F2-28
Instrument Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-47
Intentional Engine Cut Speed . . . . . . . . . . . . . . . . . . 5-37
Interior Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-21
Interior Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-75
Introduction to Performance . . . . . . . . . . . . . . . . . . . . . 7-1
L
Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Landing Emergencies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Landing Gear Extension Speed .........................
Landing Gear Retraction Speed . . . . . . . . . . . . . . . . . . . . . .
Landing Gear System . . . . . . . . . . . . . . . . . . . . . . . . . .
Landing Information . . . . . . . . . . . . . . . . . . . . . . . . .
Landing With Two Engines Inoperative . . . . . . . . .
Line Up . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Change 3
8-41
9-30
5-22
5-23
2-7
7-12
9-13
8-33
INDEX-3
TM 55-1510-219-10
Paragraph, Figure
Table Number
Subject
L
Loading Procedure . . . . . . . . . . . . . . . . . . 6-13
Load Planning . . . . . . . . . . . . . . . . . . . . . 6-12
Loss of Pressurization (Above 10,000 Feet) . . . 9-19
Low Oil Pressure . . . . . . . . . . . . . . . . . . . .9-14
Paragraph, Figure,
Table Number
Subject
O
Oxygen
Oxygen
Oxygen
Oxygen
Oxygen
Duration In Minutes . . . . . . . . . . . T2-5
Flow Planning Rates vs Altitude . . . T2-4
Requirements . . . . . . . . . . . . . . . . 5-33
System . . . . . . . . . . . . . . . 2-61, F2-19
System Servicing . . . . . . . . . . . . . F2-30
M
P
M-130 Flare and Chaff Dispensing
System
. . . . . . . . . . . . . . . . . . . . . . . . 4-3
Malfunction Codes . . . . . . . . . . . . . . . . . . T3-5
Malfunction Codes Check . . . . . . . . . . . . . T3-2
Malfunction Indications and Procedures . . . T3-3
Maneuvering Flight . . . . . . . . . . . . . . . . . 8-52
Maneuvers . . . . . . . . . . . . . . . . . . . . . . . 5-27
Marker Beacon . . . . . . . . . . . . . . . . . . . . 3-24
Maximum Allowable Airspeed . . . . . . . . . . 5-21
Maximum Design Maneuvering Speed . . . . . 5-26
Maximum Design Sink Rate . . . . . . . . . . . 5-40
Maximum Glide . . . . . . . . . . . . . . . 9-12
Maximum Glide Distance . . . . . . . . . . . . . F9-2
Maximum Weights . . . . . . . . . . . . . . . . . . . 2-5
Microphones, Switches and Jacks . . . . . . . . . 3-4
Minimum Crew Requirements . . . . . . . . . . . 5-4
Minimum Single-Engine Control Airspeed. . . . 5-25
Miscellaneous Instruments . . . . . . . . . . . . . . . . 2-85
Mission Annunciators . . . . . . . . . . . . . . . . T4-1
Mission Avionics Operating Instructions . . . . 4-1
Mission Control Panel . . . . . . . . . . . . 4-2, F4-1
Mission Control Switches . . . . . . . . . . . . . T4-2
Mission Equipment Power System . . . . . . F2-25
Mission Planning . . . . . . . . . . . . . . . . . . . . 8-1
Mooring . . . . . . . . . . . . . . . . . . . . . . . . 2-101
Mooring the Aircraft . . . . . . . . . . . . . . . . F2-32
N
Navigation - Description . . . . . . . . . . . . . . . 3-16
O
Oil Supply System . . . . . . . . . . . . . . . . . . 2-27
Operating Procedures and Maneuvers . . . . . . 8-3
Overhead Circuit Breaker Panel . . . . . . . . F2-26
Overhead Control Panel . . . . . . . . . . . . . F 2 - 1 8
Overtemperature and Overspeed
Limitations . . . . . . . . . . . . . . . . . . . . . . 5-15
INDEX-4
Parking . . . . . . . . . . . . . . . . . . . . . . . . . 2-102
Parking Brake . . . . . . . . . . . . . . . . . . . . . . 2-8
Parking, Covers, Ground Handling,
and Towing Equipment . . . . . . . . . . . . F2-31
Pedestal . . . . . . . . . . . . . . . . . . . . . . . . F2-11
Personnel Procedures During Ditching . . . . . T9-1
Pilot and Copilot Seats . . . . . . . . . . . . . . . F2-8
Pilot’s Encoding Altimeter . . . . 2-79, 3-31, F3-26
Pilot’s Horizontal Situation
Indicator (331A-8G) . . . . . . . . . . . . . . F3-10
Pilot’s Turn and Slip Indicator . . . . . 3-20, F3-13
Pitot and Stall Warning Heat System . . . . . 2-55
Pitot and Static System . . . . . . . . . . . . . . . 2-56
Pitot Heat Limitations . . . . . . . . . . . . . . . 5-12
Pneumatic Surface Deice System Limits . . . 5-13
Placard Items . . . . . . . . . . . . . . . . . . . . . 1-13
Power Definitions for Engine Operations . . . 5-16
Power Levers . . . . . . . . . . . . . . . . . . . . . . 2-22
Pressure Altitude . . . . . . . . . . . . . . . . . . . . 7-3
Pressurization System . . . . . . . . . . . . . . . . 2-60
Principal Dimensions . . . . . . . . . . . . . . . . F2-3
Propeller Electrothermal Anti-Ice System . . . 2 - 5 4
Propeller Failure (Over 2080 RPM) . . . . . . . 9-25
Propeller Governors . . . . . . . . . . . . . . . . . . . . 2-45
Propeller Levers . . . . . . . . . . . . . . . . . . . . 2-48
Propeller Limitations . . . . . . . . . . . . . . . . . 5-7
Propeller Reversing . . . . . . . . . . . . . . . . . . 2-49
Propeller Synchrophaser . . . . . . . . . . . . . . 2-47
Propeller Tachometers . . . . . . . . . . . . . . . . 2 - 5 0
Propeller Test Switches . . . . . . . . . . . . . . . 2-46
Propellen - Description . . . . . . . . . . . . . . . 2-43
PT6A-41 Engine . . . . . . . . . . . . . . . . . . . F 2 - 1 0
R
Radar Signal Detecting Set Control Panel
APR-39(V)1 . . . . . . . . . . . . . . . . . . . . . F 4 - 3
Radar Signal Detecting Set Control
Panel APR-39(V)2) . . . . . . . . . . . . . . . . F4-4
TM 55-1510-219-10
Paragraph, Figure,
Table Number
Subject
T
R
Radar Signal Detecting Set Indicator ..................F4-5
Radar Signal Detecting Sets .....................4-6
Radar Warning Receiver Control
Panel (AN/APR-44( )(V3)) .........................4-7, F4-6
Radio Control Panel ..................................3-7, F3-2
Radio Magnetic Indicators (RMI) ...............3-17, F3-9
Radome Anti-ice Operation .......................5-38
Recommended Fluid Dilution Chart ...........T2-10
Relief Tube ...............................................2-66
Required Equipment Listing .......................5-41, T5-3
Reserve Fuel ............................................7-10
Responsibility - Weight and Balance .........6-6
Rudder System .........................................2-39
S
Seats ........................................................2-12
Second Engine Start (Battery Start) ..........8-24
Second Engine Start (GPU Start) ..............8-28
Securing Loads..........................................6-14
Servicing Hydraulic Brake System
Reservoir ..................................................2-93
Servicing Locations ...................................F2-29
Servicing Oil System ................................2-92
Servicing Oxygen System .........................2-98
Servicing the Electric Toilet .......................2-95
Single-Engine Before Landing ...................9-22
Single-Engine Descent/Arrival ...................9-21
Single Engine Go-Around .........................9-24
Single Engine Landing Check ...................9-23
Single Phase AC Electrical System ...........F2-23
Spins ........................................................8-50
Stalls ........................................................8-49
Stall Speed ...............................................F8-2
Stall Warning System ...............................2-56
Standard, Alternate and Emergency Fuel . ...........T2-9
Standby Magnetic Compass .....................2-84
Starter Limitations ....................................5-8
Subpanels ................................................F2-6
Sun Visors ................................................2-65
Surface Deicer System .............................2-52
Symbols Definition ....................................8-5
System Daily Preflight/Re-Arm Test ...................4-4
Paragraph, Figure,
Table Number
Subject
TACAN Control Panel and
Range Indicator ................................... F3-17
TACAN Control Panel and
Range Indicator ................................... F3-17
TACAN Systems .................................. 3-26
Takeoff ................................................ 8-34
Takeoff Distance (Flaps 0%) ................ 7-7
Takeoff Weight .................................... 7-4
Takeoff Weight To Achieve Positive
One-Engine-Inop Climb ....................... 7-5
Taxiing ................................................ 8-30
Temperature Limits ............................. 5-30
Three Phase AC Electrical System ......................F2-24
Torque Limiter ..................................... 2-28
Transponder Set (AN/APX-100) ........... 3-30, F3-25
Trim Tabs ............................................ 2-41
Turbulence and Thunderstorm
Operation ............................................ 8-56
Turn-And-Slip Indicator ........................ 2-77, F3-13
U
UHF Command Set (AN/ARC-164) ....................3-8, F3-3
Unpressurized Ventilation .................... 2-69
Use of fuels ......................................... 2-91
Use of Words Shall, Will, Should,
and May .............................................. 1-12
V
Various Values for UTM
Grid Coefficients .................................. T3-1
Vertical Velocity Indicators ................... 2-81
VHF AM Communications
(VHF-20B) ........................................... 3-10, F3-4
VHF AM-FM Command Set
(AN/ARC-186) ..................................... 3-10, F3-11
Voice Security System (TSEC/KY-28) .................3-12, F3-6
Voice Security System (TSEC/KY-58) .................3-13
VOR/LOC Navigation System .............. 3-23
W
Warning Annunciator Panel Legend.....................T2-6
Warnings, Cautions, and Notes ........... 1-2
Weather Radar Control-Indicator ......... F3-24
INDEX-5
TM 55-1510-219-10
Paragraph, Figure,
Table Number
Subject
W
W
Weather Radar Set (AN/APN-215) ............3-29
Weight and Balance Clearance Form F...............6-8
Weight Limitations ....................................5-19
Wind Limitations .......................................5-32
Windows ..................................................2-11
Windshield Electrothermal Anti-Ice System . .........2-59
Windshield Wipers ....................................2-62
Paragraph, Figure,
Table Number
Subject
Wind Swell Ditch Heading Evaluation ............ F9-4
Wing Flap Extension Speeds ............... 5-24
Wing Flaps .......................................... 2-42
Z
Zero Fuel Weight Limitation ................. 7-11
INDEX-6
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
PATRICIA P. HICKERSON
Colonel, United States Army
The Adjutant General
DISTRIBUTIUON:
To Be distributed in accordance with DA Form 12-31-E, block no.
requirements for TM 55-1510-219-10.
0121, -10 & CL maintenance
*U.S. GOVERNMENT PRING OFFICE: 1994-300-421/02023
ELECTRONIC DA FORM 2028 INSTRUCTIONS
The following format must be used if submitting an electronic 2028. The subject
line must be exactly the same and all fields must be included; however, only the
following fields are mandatory: 1, 3, 4, 5, 6, 7, 8, 9, 10, 13, 15, 16, 17, and 27.
From: ‘Whomever” <[email protected]>
Is-lp @ redstone.army.mil
To:
Subject: DA Form 2028
1. From: Joe Smith
2. Unit: home
3. Address: 4300 Park
4. City: Hometown
5. St: AL
6. Zip: 77777
7. Date Sent: 19-OCT-93
8. Pub No: 55-2840-229-23
9. Pub Title: TM
10. Publication Date: 04-JUL-85
11. Change Number: 7
12. Submitter Rank: MSG
13. Submitter Fname: Joe
14. Submitter Mname: T
15. Submitter Lname: Smith
16. Submitter Phone: 123-123-1234
17. Problem: 1
18. Page: 2
19. Paragraph: 3
20. Line: 4
21. NSN: 5
22. Reference: 6
23. Figure: 7
24. Table: 8
25. Item: 9
28. Total: 123
27. Text:
This is the text for the problem below line 27.
PIN: 057981-000
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