TM-55-1510-215-10
TM 55-1510-215-10
TECHNICAL MANUAL
OPERATOR'S MANUAL
FOR
ARMY U-21G
AIRCRAFT
WARNING DATA
TABLE OF CONTENTS
INTRODUCTION
DESCRIPTION AND
OPERATION
AVIONICS
MISSION EQUIPMENT
OPERATING LIMITS AND
RESTRICTIONS
WEIGHT/BALANCE AND
LOADING
PERFORMANCE DATA
NORMAL PROCEDURES
EMERGENCY PROCEDURES
HEADQUARTERS
DEPARTMENT OF THE ARMY
29 DECEMBER 1982
REFERENCES
ABBREVIATIONS AND TERMS
This copy is a reprint which includes current pages
from Changes 1 through 10.
ALPHABETICAL INDEX
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WASHINGTON, D.C., 29 April 1994
NO. 10
Operator’s Manual
For
ARMY U-21G AIRCRAFT
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.
TM 55-1510-215-10, 29 December 1982, is changed as follows:
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By Order of the Secretary of the Army:
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Chief of Staff
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Operator's Manual
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ARMY U-21G AIRCRAFT
TM 55-1510-215-10, 29 December 1982, is changed as follows:
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Chief of Staff
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ARMY U-21G AIRCRAFT
TM 55-1510-215-10, 29 December 1982, is changed as follows:
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ARMY U-21G AIRCRAFT
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The Adjutant General
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ARMY U-21G AIRCRAFT
TM 55-1510-215-10, 29 December 1982, is changed as follows:
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The Adjutant General
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General, United States Army
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ARMY U-21G AIRCRAFT
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Operator's Manual
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TM 55-1510-215-10, 29 December 1982, is changed as follows:
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Chapter 5
Chapter 7
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Official:
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General, United States Army
Chief of Staff
ROBERT M. JOYCE
Major General, United States Army
The Adjutant General
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TM 55-1510-215-10
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 injury, or death.
STARTING AIRCRAFT
Operating procedures or practices defined in this Technical Manual must be followed correct. Failure to do so may result
in personal injury or loss of life. Exposure to exhaust gasses shall be avoided since exhaust gasses are an irritant to
eyes, skin and respiratory system.
NOISE LEVELS
Sound pressure levels in this aircraft during some operating conditions exceed the Surgeon General's hearing
conservation criteria as defined in TB MED 501. Hearing protection devices, such as the aviator helmet or ear plugs are
required to be worn by all personnel in and around the aircraft during its operation.
OPERATION OF AIRCRAFT ON GROUND
At all times during a towing operation, be sure there is a person in the cockpit to operate the brakes. Engines will be
started and operated by authorized personnel (AR95-1).
USE OF FIRE EXTINGUISHERS IN CONFINED AREAS
Monobromotrifluoromethane (CF 3Br) is very volatile, but is not easily detected by its odor. Although nontoxic, 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 light should be turned off during flight through clouds to prevent sensations of vertigo as a result of
reflections of the light on the clouds. The external rear view mirror should also be retracted or adjusted so as to prevent
any reflected light from entering the cockpit.
CARBON MONOXIDE
When smoke, suspected carbon monoxide fumes, or symptoms of lack of oxygen (hypoxia) exist, all personnel shall
immediately don oxygen masks, (if available) 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 longer than necessary.
a
TM 55-1510-215-10
SERVICING AIRCRAFT
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 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.
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.
Wear rubber gloves, apron, and face shield when handling batteries. If 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. The battery shall be serviced in accordance with
TM 11-6140-203-14-2 by qualified personnel only.
JET BLAST
Occasionally, during staring, 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.
PROPELLER FAILURE
While operating with the PROP GOV IDLE STOP circuit breaker pulled, the secondary low pitch system is inoperative.
Should the primary hydraulic low pitch stop and the primary governor fail, the propeller may reverse.
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. If seal is suspected to have been broken, notify Radioactive
Protective Officer.
RF BURNS
Do not stand near the antennas when they are transmitting.
b
TM 55-1510-215-10
TECHNICAL MANUAL
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D. C., 29 December 1982
Operator's Manual
ARMY U-21G AIRCRAFT
REPORTING ERRORS AND RECOMMENDING IMPROVEMENTS
You can help improve this manual. If you find any mistakes or if you know of a way to improve these procedures,
please let us know. Mail your letter or 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, US Army Aviation and
Troop Command, ATTN: AMSAT-I-MP, 4300 Goodfellow Blvd., St. Louis, MO 63120-1798. A reply will be
furnished directly to you.
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited
CHAPTER 1.
CHAPTER 2.
Section
CHAPTER 3.
Section
I.
II.
III.
IV.
V.
VI.
VII.
VIII.
IX.
X.
XI.
XII.
I.
II.
III.
IV.
CHAPTER 4.
CHAPTER 5.
Section
CHAPTER 6.
Section
I.
II.
III.
IV.
V.
VI.
VII.
VIII.
I.
II.
III.
IV.
V.
PAGE
INTRODUCTION
1-1
AIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATION
Aircraft
Emergency equipment
Engines and related system
Fuel system
Fight controls
Propellers
Utility systems
Heating, ventilation, cooling, and environmental control unit
Electrical power supply and distribution system
Lighting
Flight instruments
Servicing, parking, and mooring
2-1
2-1
2-14
2-14
2-21
2-33
2-35
2-37
2-54A
2-57
2-64
2-67
2-71
AVIONICS
General
Communications
Navigation
Transponder and radar
3-1
3-1
3-1
3-12
3-31
MISSION EQUIPMENT
4-1
OPERATING LIMITS AND RESTRICTIONS
General
System limits
Power limits
Loading limits
Airspeed limits
Maneuvering limits
Environmental restrictions
Other limitations
5-1
5-1
5-1
5-5
5-8
5-8
5-9
5-11
5-11
WEIGHT/BALANCE AND LOADING
General
Weight and balance
Fuel/oil
Personnel
Cargo loading
6-1
6-1
6-1
6-7
6-7
6-18
Change 8 i
TM 55-1510-215-10
CHAPTER 7.
Section
CHAPTER 8.
Section
CHAPTER 9.
Section
APPENDIX A.
APPENDIX B.
INDEX
ii Change 7
I.
II.
III.
IV.
V.
VI.
VII.
VIII.
IX.
X.
XI.
XII.
XIII.
XIV.
XV.
XVI.
XVII.
I.
II.
III.
IV.
V.
VI.
I.
PERFORMANCE DATA
Introduction
Performance planning
Crosswind-takeoff and landing
Idle fuel flow
Torque available for takeoff
Normal takeoff
Normal rotation/takeoff airspeed
Acceleration check distance
Accelerate-stop distance
Minimum single-engine control airspeed
Single-engine climb
Operation envelope
Cruise climb
Cruise
Climb/descent
Approach Speed
Landing
7-1
7-1
7-5
7-16
7-18
7-20
7-22
7-24
7-26
7-28
7-30
7-32
7-34
7-36
7-38
7-116
7-118
7-120
NORMAL PROCEDURES
Mission planning
Operating procedures and maneuvers
Instrument flight
Flight characteristics
Adverse environmental conditions
Crew duties
8-1
8-1
8-2
8-21
8-22
8-25
8-29
EMERGENCY PROCEDURES
Aircraft systems
9-1
9-1
REFERENCES
ABBREVIATIONS AND TERMS
A-1
B-1
INDEX-1
TM 55-1510-215-10
CHAPTER 1
INTRODUCTION
1-1. General.
1-4. Appendix A, References.
These instructions are for use by the operator. They
apply to the U-21G aircraft.
Appendix A is a listing of official publications cited
within the manual applicable to and available for flight
crews.
1-2. Warnings, Cautions, and Notes.
1-5. Appendix B, Abbreviations and Terms.
Warnings, cautions, and notes are used to
emphasize important and critical instructions and are
used for the following conditions:
Appendix B is a listing of abbreviations and terms
used throughout the manual.
1-6. Index.
WARNING
An operating procedure, practice,
etc., which if not correctly followed,
could result in personal injury or loss
of life.
CAUTION
An operating procedure, practice,
etc., which, if not strictly observed,
could result in damage to or
destruction of equipment.
NOTE
An operating procedure, condition,
etc., which is essential to highlight.
The index lists, in alphabetical order, every rifled
paragraph, figure, and table contained in this manual.
Chapter 7, Performance Data, has an additional index
within the chapter.
1-7. Army Aviation Safety Program.
Reports necessary to comply with the safety
program are prescribed in AR 385-40.
1-8. Destruction of Army Materiel to Prevent Enemy
Use.
For information concerning destruction of Army
materiel to prevent enemy use, refer to TM 750-24-1-5.
1-9. Phased Maintenance Checklist.
Phased maintenance checklist for the U-21G are
prescribed in TM 55-1510-200-PM.
1-3. Description.
1-10. Forms and Records.
This manual contains the best operating instructions
and procedures for the U-21G 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. The pilot's
flying experience is recognized, and therefore, basic
flight principles are not included.
THIS MANUAL
SHALL BE CARRIED IN THE AIRCRAFT AT ALL
TIMES.
Army aviators flight record and aircraft maintenance
records which are to be used by crew members are
prescribed in TM 38-750 and TM 55-405-9.
1-11. Explanation of Change Symbols.
Changes to the text and tables, including new
maternal on added pages shall be indicated by a vertical
bar in the outer margin extending close to the entire
1-1
TM 55-1510-215-10
area of the material affected. Pages with emergency
markings, which consist of black diagonal lines around
three edges, shall have the vertical bar or change
symbol placed along the outer margins between the text
and the diagonal lines. Change 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 printing
hands, is utilized when there have been extensive
changes made to an illustration. Change symbols are
not used to indicate changes in the following:
a. Introductory material.
b. Indexes and tabular data where the change
cannot be identified.
c. Blank space resulting from the deletion of text,
an illustration, or a table.
d. Correction of minor inaccuracies, such as
spelling, punctuation, relocation of material, etc., unless
such correction changes the meaning of instructive
information and procedures.
1-2
1-12. Aircraft Designation System.
The designation system prescribed by AR 70-50 is
used in aircraft designations as follows:
EXAMPLE U-21G
U
-
Basic mission and type symbol (utility)
21 -
Design number
G
Series symbol
-
1-13. Use of Words Shall, Will, Should and May.
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.
TM 55-1510-215-10
CHAPTER 2
AIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATION
Section I. AIRCRAFT
2-1. Introduction.
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 and 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.
2-3. Dimensions.
Overall aircraft dimensions (fig. 2-3) are as follows:
Wing Span
45 ft 10.5 in
Length
35 ft 6 in
Height (at rest)
14 ft 2.56 in
Tread (between center
lines of main wheels)
12 ft 9 in
2-4. Ground Turning Radius.
Minimum ground turning radius of the aircraft is 29 ft
8.75 in (fig. 2-4).
2-2. General.
2-5. Maximum Weights.
The U-21G is an unpressurized, low wing, all metal
aircraft, powered by two T74-CP-700 turbo-prop engines
and has all-weather capability (figs. 2-1, 2-2).
Distinguishable features of the aircraft are the slender,
streamlined engine nacelles, square-tipped wing and tail
surfaces, a swept-back vertical stabilizer and a ventral
fin below the empennage. The basic mission of the
aircraft is to provide a utility service in the combat zone
supporting field commanders and their staff in the
conduct of command and control functions, troop
transport, aero-medical evacuation, administration,
liaison, and inspection. Cabin accommodations include:
six passenger-controlled reading lights mounted in the
cabin cold air outlet panels along the ceiling; one relief
tube aft of the cabin entrance door. A relief tube is
installed under the pilot's seat. Special equipment in the
passenger-cargo (cabin) compartment is removable.
Cabin entrance is made through a stair-type door on the
left side of the fuselage. The pilot and copilot seats are
side-by-side and separated from the cabin by a
removable curtain. A minimum crew of one pilot is
required for normal aircraft operation.
Maximum takeoff gross weight is 9,650 pounds.
Maximum landing weight is 9,168 pounds. Maximum
ramp weight is 9,705 pounds.
2-6. 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. Spring-loaded locks secure the main gear in the
down position, while the jackscrew in the actuator
secures the nose gear in the down position. 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 50-ampere push-to-reset
type circuit breaker, placarded LANDING GEAR
POWER, located on the copilot's circuit breaker panel
(figs. 2-5, 2-18), protects against electrical overload.
Gear doors are opened and closed through a
mechanical linkage connected to the landing gear. The
nose wheel steering mechanism is automatically
Change 7 2-1
TM 55-1510-215-10
NOTE
The navigation and IFF antennas are
not typical as shown.
These
antennas are located in either of the
positions shown but not both.
1.
2.
3.
4.
5.
* 6.
7.
* 8.
9.
*10.
11.
*12.
13.
14.
Radar antenna
Exhaust stubs
Accessory section exhaust vent
Free air temperature sensor
Stall warning vane
Navigation antenna
UHF/VHF (COMM 2) antenna
Navigation antenna
Left static air source
IFF antenna
Radar altitude antenna
IFF antenna
DME antenna
VHF (COMM1) antenna
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
ADF antenna
Marker beacon antenna
XPDR antenna
Cockpit emergency entrance/exit hatch
Propeller
Glideslope antenna
Engine air intake
Ice light
Landing light
Cabin air vent outlet
Right static air source
* Location on some aircraft
Figure 2-1. Typical General Exterior Arrangement
2-2 Change 5
AP011650
TM 55-1510-215-10
LIGHT CARGO VERSION
CONVERSION
TO
LIGHT
CARGO
VERSION
CONSISTS
OF
REMOVING
THE
EXISTING
AMBULANCE OR STAFF SEATING INSTALLATIONS
NO CARGO LOADING OR UNLOADING EQUIPMENT
IS PROVIDED. REFER TO CHAPTER 6, AIRCRAFT
LOADING FOR CARGO HANDLING INFORMATION
AND INSTRUCTIONS.
1.
2.
3.
4.
5.
6.
7.
8.
Cockpit
Pilots and copilots sears
Cabin
Cabin emergency exit hatch
Passenger seats & stretchers
Cargo door
Tools & accessories
Cabin entrance doors
Figure 2-2. General Interior Arrangement
2-3
TM 55-1510-215-10
Figure 2-3. Principal Dimensions
2-4
TM 55-1510-215-10
GROUND CLEARANCES
VERTICAL STABILIZER
14 FT 2.56 IN
WING TIPS
6 FT 8 IN
PROPELLER
UPPER BLADE
LOWER BLADE
8 FT 9 IN
12 IN
WINGSPAN
45 FT 10.5 IN
LENGTH
35 FT 6 IN
Figure 2-4. Turning Radius and Ground Clearance
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
extension or retraction time should take no longer than 8
seconds.
a. Landing Gear Control Switch. Landing gear
system operation is controlled by a manually actuated
wheel-shaped switch placarded LDG GEAR CONT, UP
and DN, on the right subpanel (fig. 2-19). The control
switch and associated relay circuits are protected by a 5ampere circuit breaker, placarded LDG GR CONTROL,
on the right subpanel (fig. 2-19).
b. Landing Gear Down Position-Indicator Lights.
Landing gear down position is indicated by three green
lights on the control pedestal, placarded GEAR DOWN
(fig. 2-6). These lights have a press-to-test feature. The
circuit is protected by a 5-ampere circuit breaker,
placarded LG IND, on the right subpanel.
c. Landing Gear Position Warning Lights. Two red
bulbs, wired in parallel, are positioned inside the clear
plastic grip on the landing gear control handle (fig. 2-7).
These lights illuminate whenever the landing gear
handle 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 50%
of lever travel when the landing gear is not down and
locked. To turn the handle lights OFF, during singleengine 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 burns out. The circuit is protected by a 5ampere circuit breaker, placarded LDG GR CONTROL,
on the right subpanel (fig. 2-7).
d. Landing Gear Warning Light Test Button. A test
button, placarded HDL LT TEST, is located on the right
subpanel (fig. 2-7). Failure of the landing gear handle to
2-5
TM 55-1510-215-10
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
Shoulder harness lock lever
Shoulder harness inertia real
External rear view mirror
Storm window lock
Fuel management panel
Sun visor
Free air temperature gage
Magnetic compass
Overhead control panel
Windshield wipers
Control wheel
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
Copilot's circuit breaker panel
Rudder pedals
Oxygen regulator control panel
Clock
Seat belt
Utility pocket
Control pedestal
External mirror adjustment knob
Oxygen system gage
Oxygen system controls and regulator control panel
AP000513
Figure 2-5. Typical Cockpit
2-6 Change 5
TM 55-1510-215-10
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
Strobe beacon control switch
Strobe beacon select switch
Fire detection system test switch
Keylock switch
Wing flap position indicator
Landing gear down position indicator lights
Pneumatic pressure gage
Condition levers
Friction lock knobs
Wing flap switch
Rudder tab control and position indicator
Flight director mode controller
UHF control panel
Landing gear emergency clutch disengage lever
Landing gear emergency extension placard
Landing gear emergency extension handle
Aileron tab control and position indicator
Elevator tab control and position indicator
Propeller levers
Power levers
Go-around button
AP011648
Figure 2-6. Typical Control Pedestal and Landing Gear Emergency Extension Controls
Change 5 2-7
TM 55-1510-215-10
(Figure 2-7 sheet 1 of 2)
(Figure 2-7 sheet 2 of 2)
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
Parking brake handle
Ignition and engine start switches (2)
Engine autoignition arm lights (2)
Navigation lights switch
Inlet air separator switch
Propeller governor test switches (2)
Fuel vent heat switches (2)
Propeller deicer switch
Pitot heat switch
Landing lights switch
11.
12.
13.
14.
15.
16.
17.
18.
19.
Propeller deicer ammeter
Volt-loadmeters (2)
Left cockpit air control
Landing gear warning horn silence switch
Heater control switch
Cabin temperature control rheostat
Right cockpit air control
Cabin air control
Landing gear down lock release switch
Figure 2-7. Typical Subpanels (sheet 1 of 2)
2-8 Change 6
TM 55-1510-215-10
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
Master switch gang bar
Aircraft inverter switch
Autoignition switches (2)
Taxi lights switch
Deice cycle switch
Stall warning heat switch
Ice lights switch
Propeller autofeather switch
Windshield anti-ice switch
Engine lip boot heat switches (2)
30.
31.
32.
33.
34.
35.
36.
37.
38.
39.
Fuel control heat switches (2)
Engine ice vane control handles (2)
Generator switches (2)
Battery switch
Landing gear control handle
Vent blower switch
Windshield anti-ice circuit breakers (4)
Defrost air control
Landing gear handle light test switch
Landing gear handle lights (2) red (inside handle)
Figure 2-7. Typical Subpanels (sheet 2 of 2)
Change 6 2-9
TM 55-1510-215-10
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 LDG
GR CONTROL, on the right subpanel (fig. 2-7).
e. Landing Gear Warning Horn.
When either
power lever is retarded below approximately 80% N1
when the landing gear is not down and locked, a warning
horn behind the left side of the instrument panel will
sound intermittently.
The warning horn circuit is
protected by a 5-ampere circuit breaker, placarded LG
WARN HORN, on the right subpanel (fig. 2-7).
f. Landing Gear Warning Horn Silence Button.
The landing gear warning horn can be silenced during
flight (with power retarded, gear and flaps up), by
pressing the button, placarded WARN HORN SILENCE,
on the right subpanel (fig. 2-7). After silencing the
warning horn, it will remain silent until either the flaps are
extended or the power levers are advanced, then
retarded again. During single-engine operation, the
warning horn may be silenced by either pressing the
warning horn silence button (flaps up) or advancing the
power lever for the inoperative engine to a position
above the warning horn microswitch setting. If the
warning horn has been silenced by pressing the silence
button, the power lever for the inoperative engine must
again be advanced past the warning horn microswitch
setting to reset the switch. The horn will sound however,
if the power lever for the operative engine is retarded.
The circuit is protected by a 5-ampere circuit breaker,
placarded LG WARN HORN, on the right subpanel
(fig. 2-7).
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
deactivates the landing gear control circuits when the
strut is compressed. This switch also activates a downlock hook, preventing the landing gear handle 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 LCK REL located adjacent to the
landing gear handle (fig. 2-7). 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
2-10
will sound. The safety switch on the left main landing
gear strut activates the engine inlet air separator, the
nacelle lip ice boots and heater ram air intake boot when
the strut is extended, and the two ventilation air blowers
(for nose avionics and cockpit/cabin areas) when the
strut is compressed.
CAUTION
Do not pump handle after GEAR
DOWN position indicator lights (3)
are illuminated. Further movement of
the handle could damage the drive
mechanism.
h. Landing Gear Emergency Clutch Disengage
Lever. Prior to emergency or manual landing gear
extension, the landing gear motor must be disengaged
from the landing gear drive mechanism.
This is
accomplished by a manually operated clutch disengage
lever (fig. 2-6) located adjacent to the landing gear
emergency extension handle. To disengage the clutch,
pull the clutch lever up and turn clockwise. To engage
the clutch, turn the clutch lever counter-clockwise and
release.
i. Landing Gear Emergency Extension Handle.
The landing gear emergency extension handle (fig. 2-6)
located on the cockpit floor to the right of the pilot's seat,
is used for manual extension of the landing gear. The
landing gear extension system is actuated by pumping
the handle up and down. This movement operates a
ratchet mechanism which drives the normal system to
extend the landing gear. The landing gear cannot be
retracted manually. Refer to chapter 9 for emergency
gear extension procedures.
2-7. Steerable Nose Wheel.
The aircraft can be maneuvered on the ground by
the steerable nose wheel system. Direct linkage from
the rudder pedals 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.
TM 55-1510-215-10
2-8. Wheel Brake System.
The main landing wheels are equipped with multipledisc hydraulic brakes actuated by master cylinders
attached to the rudder pedals at the pilot's and copilot's
position. A shuttle valve, adjacent to each set of pedals,
permits braking action changeover from one set of
pedals to the other. 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
corresponding wheels. No emergency brake system is
provided.
Repeated and excessive application of
brakes, without allowing sufficient time for cooling, will
cause loss of brake efficiency, possible failure of brake
or wheel structure, possible blowout of tires, and in
extreme cases may cause the wheel and brake
assembly to be destroyed by fire. Parking brakes shall
not be set during flight. The following precautions
should be observed insofar as is practical:
a. Extreme care should be used during any braking
application to prevent skidding the tires and causing flat
spots.
b. With the landing gear extended, approximately
10 minutes should be allowed to elapse between
landings where maximum braking has been applied.
c. With the landing gear retracted, approximately
30 minutes should be allowed to elapse between
maximum-brake application landings.
CAUTION
It is damaging to the braking system
if continuous braking drag is applied
over an appreciable distance while
taxiing at slow speeds.
d. For short landing rolls, reversing the propellers
and a single, smooth application of the brakes with
constantly increasing pedal pressure is most desirable.
2-9. Parking Brake Handle.
Dual parking brake valves are installed adjacent to
the rudder pedals between the master cylinders of the
pilot’s rudder pedals and the wheel brakes. Both valves
can be closed simultaneously by pulling out the handle
placarded PARKING BRAKE on the left subpanel
(fig. 2-7), after pressing the brake pedals on the pilot's
side to build up pressure. Parking brakes are released
when the brake handle is pushed in.
2-10. Entrance and Exit Provisions.
WARNING
Do not open door or attempt to lock
or secure during flight.
CAUTION
Structural damage may be caused if
more than one person is on the
entrance door at any time.
The main cabin entrance door is used for normal or
emergency exit (fig. 2-8). A removable window (cabin
emergency exit hatch) on the right side of the fuselage,
and the cockpit emergency entrance/exit hatch are used
for emergency exit only. The main cabin entrance door,
the cabin emergency exit hatch (removable window),
and the cockpit emergency entrance/exit hatch provide
emergency escape routes from the aircraft either on the
ground or when ditching.
Refer to chapter 9 for
emergency exit locations and bailout procedures.
a. Aircraft entrance door provisions consist of two
adjoining door assemblies on the left side of the fuselage
which open outward from the cabin compartment. The
forward, or cargo door is normally closed during egress
of the crew. The cargo door is secured in a closed
position by two hand-operated, locking slide-bolts,
located on the top and bottom aft corners of the door.
The door opens forward on two hinges and is held open
by a swivel rod, which inserts into a hold-fitting installed
in the wing fairing below the door opening. A cable
attaches to the door top and to inside fuselage structure
above the door opening, preventing damage due to door
overswing. The aft side of the cargo door is part of the
frame structure for the main entrance door, and has
striker plates for securing the door in the closed position.
The main entrance door must be open before the cargo
door can be opened.
Change 7 2-11
TM 55-1510-215-10
Figure 2-8. Main Entrance and Cargo Doors
2-12
TM 55-1510-215-10
b. Normal aircraft entry and exit is accomplished
through the main entrance door. This door may be
opened, or closed from either the inside or outside of the
aircraft by clockwise/counter-clockwise rotation of the
door handles.
The stair-type steps are integrally
mounted on the inside of the door. A plastic encased
cable provides a stop to support the door in the open
position, serves as a hand assist during entry or exit,
and as a convenience for closing the door from the
inside.
NOTE
ONLY PILOT’S SEAT IS
SHOWN,
COPILOT’S
SEAT IS IDENTICAL
2-11. Cabin Door Warning Light.
WARNING
Do not open door or attempt to lock
or secure it during flight.
As a safety precaution, two flashing MASTER
CAUTION lights (fig. 2-22), on the instrument panel and
a steady illuminated DOOR OPEN yellow caution light
on the annunciator panel (fig. 2-22) indicate the main
entrance door is not closed and locked. This circuit is
protected by a 5-ampere circuit breaker, placarded ANN
PANEL, located on the right subpanel (fig. 2-7).
2-12. Windows.
Forward visibility from the cockpit is provided by a
two piece electro-thermal windshield made of laminated
plate glass. Side visibility for the pilot and copilot is
provided by a window made of acrylic on each side of
the cockpit. Immediately forward of each side window is
a triangular-shaped storm window which may be opened
during flight or ground operation. Windows are provided
on each side of the cabin.
2-13. Seats.
a. Pilot and Copilot Seats.
Aircraft seating
accommodations for the pilot and copilot are adjustable,
chair-type units (fig. 2-9). An entry aisle separates the
two seats. Both seats can be adjusted fore and aft,
vertically. 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
1.
2.
3.
4.
5.
6.
7.
Inertia reel
Shoulder harness
Inertia reel lock
Arm rest
Seat adjustment handle
Ash tray
Seat belt
Figure 2-9. Pilot’s and Copilot’s Seats
aft. The vertical adjustment handle is located beneath
the bottom front outboard corner of each seat. Pulling
up on the handle, allows the seat to move up or down.
b. Seat Belts and Shoulder Harnesses - Pilot and
Copilot Seats. Pilot and copilot seats are each equipped
with a seat belt and shoulder harness. Each belt
attaches to two floor disconnect fittings located aft of the
seat. Pilot and copilot shoulder harnesses incorporate
an inertia reel to provide crash restraint. The inertia reel
will lock automatically under a 2 "G" impact, or may be
manually locked if desired.
2-13
TM 55-1510-215-10
Section II. EMERGENCY EQUIPMENT
2-14. 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. For the operation of emergency exits and for
location of all emergency equipment, refer to chapter 9.
below the copilot's seat, and a second is located aft of
the main entrance door, (fig. 9-1). These 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.
2-15. Hand-Operated Fire Extinguisher.
NOTE
Engine fire extinguisher systems are
not installed.
WARNING
Exposure to high concentrations of
monobromotrifluoromethane (CF3 Br)
or decomposition products should be
avoided. The liquid should 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.
One hand-operated fire extinguisher is mounted
2-16. First Aid Kits.
Five first aid kits are provided in the aircraft (fig. 9-1).
One kit is installed on the backside of the pilot and
copilot seat respectively, and the remaining kits are
located in the cabin compartment.
2-17. Fire Axe.
The fire axe is located under the copilot's seat in the
cockpit.
Section III. ENGINES AND RELATED SYSTEMS
2-18. Description.
The aircraft is powered by two T74-CP-700
turboprop engines rated at 550 SHP (fig. 2-10). This
engine is a reverse flow, free turbine type, employing a
three-stage axial and a single-stage centrifugal
compressor assembled as an integral unit. A large area
circular steel screen around the air intake at the rear of
the gas generator case precludes foreign object
ingestion by the compressor. The combustion chamber
is of the reverse flow type and consists primarily of an
annular heat-resistant steel liner open at one end. The
single-stage compressor turbine and free power turbine
are in line with each other and are counter-rotating. The
single-stage power turbine is connected through a twostage planetary reduction gearbox to a flanged propeller
shaft. For the purpose of engine RPM/power ratio
familiarization and computation, compressor turbine
rotational speed (RPM) will be referred to as N1, and the
propeller speed N2 as actual indicated tachometer RPM.
The accessory drive at the aft end of the engine provides
power to drive the fuel pumps fuel control, the oil pumps,
the starter/generator, and the tachometer transmitter.
2-14 Change 7
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. The oil
tank, filler cap and dip stick are an integral part of the
compressor air inlet and the accessory section.
2-19. Engine Compartment Cooling.
The forward engine compartment including the
accessory section is cooled by air entering around the
exhaust stub cutouts, the gap between the propeller
spinner and forward cowling, and exhausting through
louvers located behind the exhaust stubs.
The
accessory section is cooled by air from the plenum
chamber which is exhausted through a flush vent on the
left side of the cowl.
2-20. Air Induction Systems - General.
Each engine receives ram air ducted from an air
scoop located within the lower section of the forward
nacelle. Each oil cooler uses ram air secured by a
TM 55-1510-215-10
(Figure 2-10 sheet 1 of 2)
(Figure 2-10 sheet 2 of 2)
1.
2.
3.
4.
5.
6.
7.
Propeller governor
Torque pressure transmitter
Auto ignition pressure switch
Auto feather pressure switch
Center fireseal
Rear fireseal
Oil filler cap and dipstick
8.
9.
10.
11.
12.
13.
Oil-to-fuel heat exchanger
Starter/generator
Ignition regulator box
Air intake screen
Oil scavenge tubs
Propeller overspeed governor
Figure 2-10. T74-CP-700 Engine (sheet 1 of 2)
separate air scoop attached to the lower section of the
nacelle. Special components of the engine induction
system protects the power plant from icing and foreign
object damage.
2-21. Foreign Object Damage Control.
The engine has an integral air inlet screen designed
to obstruct objects large enough to damage the
compressor.
2-22. Inlet Air Separator System.
a. An inlet air system is designed to prevent the
engines from ingesting dust and foreign matter when the
aircraft lands on unimproved runways and reverse
propeller thrust is used. This system is for use during
landing operation only. It is comprised of plumbing
which routes engine bleed air through a control valve
into the bypass air duct mounted aft of the engine oil
cooler. When opened, bleed air-flow causes an increase
in the velocity of the air passing through the engine air
intake. The increased air velocity prevents dust particles
present in the air from making the sharp turn up into the
engine plenum chamber, forcing them to follow the air
stream out through the bypass duct. The control valve
mounted on top of the bypass duct, is normally closed
blocking the flow of bleed air. Both engine particle
separator systems are controlled by one INLET AIR
SEPARATOR SWITCH ON THE LEFT SUBPANEL
(fig. 2-7). A yellow PARTICLE SEP indicator on the
2-15
TM 55-1510-215-10
14.
15.
16.
17.
18.
19.
Fuel control unit
Fuel control unit control rod
Bleed air line
Fuel manifold
Exhaust
Power turbine governor
20.
21.
22.
23.
24.
25.
Reduction gearbox housing
Oil pressure tube
Ignitor plug
Engine mount
Oil filter cover
Accessory gearbox housing
Figure 2-10. T74-CP-700 Engine (sheet 2 of 2)
annunciator panel (fig. 2-22) illuminates when the bleed
air control valves are in the open position.
CAUTION
Monitor inlet turbine temperature
during ground roll with inlet air
separator ON and propellers in
reverse pitch. Do not allow ITT to
exceed engine limits. If excessive
ITT temperatures occur, reduce
power.
b. During the DESCENT-ARRIVAL check (chap.
8), the INLET AIR SEPARATOR switch may be placed
to AUTO as required. Extend the inertial separator antiice vanes by pulling the ENG ICE VANE handles (fig.
2-7) out. During the AFTER LANDING check, turn the
INLET AIR SEPARATOR switch off, and reposition the
engine ice vane handles as required.
2-16
NOTE
Propeller reversing can be used
below 40 knots, if required, with the
inlet
air
separators
operating.
Propeller blade erosion may result
but engine dust and foreign matter
ingestion will be minimized.
2-23. Power Plant Ice Protection Systems.
a. Inertial Separators.
An inertial separation
system is built into each engine air inlet to prevent
moisture particles from entering the engine inlet plenum
under freezing conditions. This is done by introducing 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. During normal operation, a movable vane is
raised out of the direct ram airstream. For cold weather
(+ 5°C or below) operation in visible moisture, it should
be lowered into the airstream. The anti-ice vanes are
TM 55-1510-215-10
operated by individual T-handle, push-pull controls,
located below the left subpanel.
The controls are
placarded LEFT ENG ICE VANE PULL TO EXT, PUSH
TO RET, RIGHT ENG ICE VANE, PULL TO EXT, PUSH
TO RET (fig. 2-7). Vane position during operation is
indicated by the position of the T-handles, and by a
slight decrease in torque with the engine ice protection
controls extended. The vanes should be either fully
retracted or fully extended; there are no intermediate
positions.
made by movement of the power and condition levers for
a specific engine. Engine shutdown is accomplished by
moving the appropriate condition lever to the full aft,
FUEL CUT-OFF position, which shuts off the fuel supply.
b. Engine Lip Boot Heat. The engine air inlet lip
boots are electrically heated to prevent the formation of
ice and consequent distortion of the airflow. The boots
are operated by the two 5-ampere circuit breaker
switches on the pilot's subpanel placarded: ENG LIP
BOOT, LEFT-RIGHT. The circuit is connected through
the left landing gear safety switch and is therefore
operable only during flight. The circuit is protected by
two 25-ampere circuit breakers, placarded LIP ANTIICE, LH-RH located on the copilot's circuit breaker panel
(fig. 2-19). During flight, when icing conditions are
anticipated, position both ENG LIP BOOT heat switches
ON (up), (fig. 2-7). Continue use as required and shut
off when icing conditions are no longer present or
anticipated.
Moving the power levers into reverse
range without the engines running
may result in damage to the reverse
linkage mechanisms.
2-25. Power Levers.
CAUTION
Two power levers are located on the control
pedestal (fig. 2-6). The left power lever incorporates a
go-around button. 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
c. Fuel Control Heat.
Each fuel control's
temperature compensating line is protected against ice
by electrically heated jackets. Power is supplied to each
fuel control air line heater by two switches, placarded
FUEL CONTROL HEAT LEFT AND RIGHT on the pilot's
subpanel. Fuel control heat should be turned on before
all flight operations.
speed governor in the fuel control unit. Power is
increased when N1 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 below 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.
2-24. Engine Fuel Control System.
2-26. Condition Levers.
a. Description. The fuel control system consists of
an engine-driven primary (high pressure) fuel pump, an
engine driven boost pump, a fuel control unit, and a fuel
nozzle manifold. An automatic fuel dump valve and two
drain valves are provided to drain residual fuel from the
engine after shutdown or after a discontinued start.
Two condition levers are located on the control
pedestal (fig. 2-6). Each lever starts or stops the fuel
supply, and controls the idle speed for its engine. The
levers have three placaded 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 50% to 53% (at
sea level) minimum N1 and HIGH IDLE position sets the
b. Fuel Control Unit. One fuel control unit is on the
accessory case of each engine. This unit is a hydromechanical 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. N1 speed is controlled by varying
the amount of fuel injected into the combustion chamber
through the fuel nozzles. All fuel control operations are
rate to attain 70% to 73% N1, minimum N1. The power
lever for the corresponding engine can select N1 from
the respective idle setting to maximum power. An
increase in low idle N1 will be experienced at higher field
elevation.
Change 7 2-17
TM 55-1510-215-10
2-27. Friction Lock Knobs.
Four friction lock knobs are located on the control
pedestal to adjust friction drag for the engine and
propeller control levers. One knob is below the propeller
levers, one below the condition levers, and two under
the power levers. When a knob is rotated clockwise,
friction restraint is increased opposing movement of the
affected lever as set by the pilot. Counterclockwise
rotation of a knob will decrease friction drag thus
permitting free and easy lever movement.
Two
FRICTION LOCK placards are located on the pedestal
adjacent to the knobs (fig. 2-6).
switch placarded TEST SWITCH FIRE DETECTION OFF, 1, 2, 3, on the control pedestal, is provided to test
the engine fire detection system (fig. 2-6).
Before
checkout, battery power must be on and the FIRE
DETECTOR circuit breaker must be in. Switch position
1, checks area forward of the air intake of each nacelle,
including circuits to the cockpit alarm and indication
devices. Switch position 2, checks the circuits for the
upper accessory compartment of each nacelle. Switch
position 3, checks the circuits for the lower accessory
compartment of each nacelle. Each numbered switch
position will initiate the cockpit alarm and indications
previously described.
2-28. Engine Fire Detection System.
2-29. Oil Supply System.
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 audio and 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.
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 non-congealing external oil cooler keeps the
engine oil temperature within the operating limits. The
capacity of each engine oil tank is 9.2 quarts and
includes an expansion space of 0.72 quarts. The total
system capacity for each engine, which includes the oil
tank, oil cooler, lines, etc., is 14 quarts - of which 6
quarts are usable. The oil level is indicated by a dipstick
attached to the oil filler cap. For oil grade, specification
and servicing points, refer to Section XII, Servicing.
a. 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
alarms. 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 DETECTOR circuit breaker.
b. Warning of internal nacelle fire is provided as
follows: A warning horn sounds in the cockpit;
simultaneously the red MASTER WARNING lights on
the instrument panel start flashing. These alarms are
accompanied by the continuous illumination of a red
FIRE L ENG or FIRE R ENG light also on the
annunciator panel (fig. 2-22). Fire detector circuits are
protected by a single 3-ampere circuit breaker,
placarded FIRE DETECTOR, located on the right
subpanel (fig. 2-7).
c. An erroneous indication of engine fire may 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, close the cowling and/or
change the aircraft heading away from the light source.
d. Fire detection system test switch.
2-18
One rotary
b. The oil system of each engine is coupled into 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.
2-30. Engine Chip Detection System.
A magnetic chip detector is installed in the bottom of
each engine reduction 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 an
annunciator panel yellow light placarded L or R CHIP
DETECT (fig. 2-22) and the MASTER CAUTION
TM 55-1510-215-10
lights (fig. 2-22). Chip detector circuits are protected by
a single 5-ampere circuit breaker, placarded CHIP
DETECTOR, on the right subpanel (fig. 2-7).
2-31. Engine Ignition System.
a. The basic ignition system consists of a current
regulator unit, two igniter plugs, two shielded ignition
cables,
pilot-controlled
IGNITION
&
ENG
START/STARTER ONLY switches. Activation of an
IGNITION & ENG START switch will cause coils in the
respective igniter plugs to heat, 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
on the left subpanel, will initiate starter motoring and
ignition in the IGNITION & ENG START position, or will
motor the engine in the STARTER ONLY position (fig.
2-7). The switches are placarded LEFT AND RIGHT to
designate the appropriate engine. The IGNITION &
ENG START 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 20-ampere circuit breakers on the copilot's circuit
breaker panel, placarded IGNITER - LH and RH, protect
ignition circuits. Two 7.5-ampere circuit breakers on the
right subpanel, placarded START CONT - LH and RH,
protect starter control circuits (fig. 2-7).
2-32. Autoignition System.
If "armed", the autoignition 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 autoignition control switch,
a green press-to-test light and a yellow indicator on the
annunciator panel. Autoignition is accomplished by
energizing the two igniter heating elements in each
engine.
a. Autoignition Switches. Two switches placarded
ENG AUTOIGNITION, LEFT and RIGHT, with positions
ARM and OFF, are located on the left subpanel
(fig. 2-7). ARM position initiates a readiness mode for
the autoignition system of the corresponding engine.
OFF position disarms the system. Each switch is
protected by a corresponding START CONT, LH or RH
circuit breaker of 7.5-amperes on the right subpanel
(fig. 2-7).
b. Autoignition Lights. Two green press-protest
lights are positioned below the autoignition switches (fig.
2-7). Each light, although not placarded, is associated
only with the switch directly above it and will illuminate
when that switch is placed in the ARM position.
Illumination of a light indicates that the autoignition
system is in a "ready condition".
If an ARMED
autoignition system changes from "ready condition" to an
"operating condition" energizing the two igniter elements
in an engine, the green light will extinguish and a
corresponding yellow annunciator panel light will
illuminate. The annunciator panel light is placarded L or
R IGN ON and indicates that the igniters are energized.
The autoignition system is triggered from a "ready
condition" to an "operating condition" when engine
torque drops below 12 PSI (350 to 450 ft-lb torque).
Therefore, when an autoignition system is ARMED, the
igniters will be energized continuously during the time
when an engine is operating at a level below 12 PSI
(350 to 450 ft-lb torque). The autoignition lights are
protected by 7.5-ampere START CONT - LH and RH
circuit breakers on the right subpanel (fig. 2-7).
c. Ignition-On Indicators. The ignition on state is
confirmed by an illuminated yellow function indicator on
the annunciator panel placarded L, or R IGN ON
(fig. 2-22).
2-33. Engine Starter-Generators.
One starter-generator is mounted on each engine
accessory drive section. Each is able to function either
as a starter or as a generator. In starter function, a
stabilized minimum of 22 volts DC is required to power
rotation. In generator function, each unit is capable of
250 amperes DC output. When the starting function is
selected, the starter generator receives power through
the respective 7.5-ampere START CONT circuit
breakers on the right subpanel from either the aircraft
battery or an external power source.
When the
generating function is selected, the starter-generator
provides electrical power to its respective bus. For
additional description of the starter generator system,
refer to Section IX.
2-34. Engine Instruments.
Instruments which display engine conditions or state
of operation are discussed in the following paragraphs.
Engine instruments are located horizontally in the
Change 7 2-19
TM 55-1510-215-10
upper-center section and vertically on the pilot's side of
the instrument panel.
a. Interstage Turbine Temperature Indicators. Two
interstage turbine temperature (ITT) gages on the
instrument panel are calibrated in degrees celsius units
(fig. 2-22). Each gage is connected to ten thermocouple
probes located in the hot gases between the turbine
wheels. The ITT 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 indicate torque applied to the propeller
shafts of the respective engines (fig. 2-21). Each gage
shows torque by foot-pound measure using 100 pound
graduations and is actuated by an electrical signal from
a pressure sensing system located in the respective
propeller reduction gear case.
Torquemeters are
protected separately by a 1-ampere circuit breaker,
placarded TORQUEMETER LH, RH, on the copilot's
circuit breaker panel (fig. 2-19).
c. Turbine Tachometers. Two tachometers on the
instrument panel register compressor turbine RPM (N1)
for the respective engine (fig. 2-22). 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.
2-20 Change 5
d. Oil Pressure Indicators. Two gages on the
instrument panel register oil pressure in PSI as taken
from the delivery side of the main oil pressure pump
(fig. 2-22). Each gage is connected to a pressure
transmitter installed on the respective engine. Both
instruments are protected by a single, 1-ampere circuit
breaker, placarded OIL PRESS, on the copilot's circuit
breaker panel (fig. 2-19).
e. Oil Temperature Indicators. One oil temperature
gage on the instrument panel is provided for each
engine (fig. 2-22). These instruments are connected to a
thermal sensor unit attached to the respective engines.
Each gage registers oil temperature in °C as it leaves
the delivery side of the oil pressure pump. Both gages
are protected by a single 5-ampere circuit breaker,
placarded OIL TEMP IND, on the right subpanel
(fig. 2-7).
f. Fuel Flow Indicators.
Two gages on the
instrument panel register the rate of flow for consumed
fuel as measured by sensing units coupled into the fuel
supply lines of the respective engines (fig. 2-22). The
fuel flow indicators are calibrated in increments of
hundreds of pounds per hour.
Both circuits are
protected by a single, 1-ampere circuit breaker,
placarded FUEL FLOW, on the copilot's circuit breaker
panel (fig. 2-19).
TM 55-1510-215-10
Section IV. FUEL SYSTEM
2-35. Fuel Supply System.
The engine fuel supply system (fig. 2-11) consists of
two identical systems sharing a common fuel
management panel and fuel crossfeed manifold. Each
fuel system consists of four interconnected wing tanks, a
nacelle tank, an engine driven boost pump mounted on
each engine, an auxiliary fuel pump located within the
nacelle tank, a fuel transfer pump located within the
inboard wing tank, a fuel heater (engine oil-to-fuel heat
exchanger unit), a tank vent system, a tank vent heating
system and interconnecting wiring and plumbing.
a. Engine Driven Boost Pumps.
CAUTION
Engine operation using only the
engine-driven
primary
(high
pressure) fuel pump without auxiliary
fuel pump or engine-driven boost
pump fuel pressure is limited to 10
cumulative hours. This condition is
indicated by either R or L FUEL FAIL
lights. All time in this category shall
be entered on DA Form 2408-13 for
the
attention
of
maintenance
personnel.
A gear-driven boost pump, mounted on each engine
supplies fuel under pressure to the inlet of the enginedriven primary high-pressure pump. Either the enginedriven boost pump or auxiliary fuel pump is capable of
supplying sufficient pressure to the engine-driven
primary high pressure pump and thus maintain normal
engine operation.
NOTE
Boost pump failure is indicated by
steady illumination of either the L
FUEL FAIL or R FUEL FAIL light on
the annunciator panel, and the
simultaneous
flashing
of
both
MASTER CAUTION lights. Refer to
chapter 9.
b. Auxiliary
Fuel
Pumps.
A
electrically-operated auxiliary fuel pump, located within
each nacelle tank, serves as a backup unit for the
engine-driven boost pump. The auxiliary pumps are
switched off during normal system operations, except for
takeoff, landing or crossfeed. An auxiliary fuel pump
must be operated during crossfeed to pump fuel from
one system to the other. The auxiliary fuel pumps are
protected by two 10-ampere circuit breakers placarded
AUX PUMP, located on the fuel system circuit breaker
panel (fig. 2-12).
c. Fuel Transfer Pumps.
A submerged,
electrically-operated fuel transfer pump is located in
each inboard wing tank. Fuel level in the nacelle tank is
automatically maintained by gravity feed from the wing
tanks. However, approximately 28 gallons (182 pounds)
in each of the inboard wing tanks will not gravity feed.
This fuel is moved by the fuel transfer pumps. These
transfer pumps are energized and de-energized by a fuel
quantity sensor located in each nacelle tank (when the
transfer pump switches are in the ON position).
Whenever the nacelle tank quantity drops to 8 gallons
(52 pounds), the transfer pump is automatically
energized and begins transfer of fuel from the wing tanks
to the nacelle tank. Fuel transfer continues until the fuel
quantity in the nacelle tank reaches 24 gallons (156
pounds). The fuel transfer pumps are protected by two
5-ampere circuit breakers placarded TRANSFER PUMP,
located on the fuel system circuit breaker panel
(fig. 2-12).
d. Fuel Gaging System. Four fuel quantity gages
are mounted on the fuel management panel (fig. 2-12).
All gages are calibrated in quarters of total tank capacity.
Two gages indicate quantity of each fuel system: one
gage indicates fuel quantity of the nacelle tank and the
other indicates fuel quantity of the four interconnected
wing tanks. Refer to table 2-1 for usable fuel, capacity
and weight. The fuel gages are protected by two 5ampere circuit breakers placarded QTY IND, located on
the fuel system circuit breaker panel (fig. 2-12).
e. Fuel Management Panel (fig. 2-12). The fuel
management panel is located on the cockpit sidewall, on
the left side of the pilot. It contains the fuel gages,
auxiliary fuel pump switches, transfer pump switches,
transfer test light, crossfeed valve switch, firewall shutoff
valve switches, and nine circuit breakers protecting the
fuel system.
submerged,
2-21
TM 55-1510-215-10
NOTE
•
RIGHT SYSTEM IS IDENTICAL TO LEFT
EXCEPT THE LATTER CONTAINS THE
CROSSFEED VALVE, AND THERE IS A
THERMAL RELIEF VALVE AND LINE
FROM THE CROSSFEED LINE IN THE
RIGHT FUEL SYSTEM.
•
TOTAL USABLE FUEL 370 GALLONS
Figure 2-11. Fuel System Schematic
2-22
TM 55-1510-215-10
Table 2-1. Fuel Quantity Data
TANKS
NUMBER
USABLE FUEL
**WEIGHT/POUNDS
Wing tanks
4
128
832.0
Nacelle tank
1
57
370.5
Wing tanks
4
128
832.0
Nacelle tank
1
57
370.5
10
370
2405.0
LEFT ENGINE
RIGHT ENGINE
*Totals
* Unusable fuel quantity and weight (3.6 gallons, 24 pounds not included in totals).
** Fuel weight is based on standard day conditions at 6.5 pounds per U.S. gallon. Total fuel system capacity is 373.6
gallons.
1.
2.
3.
4.
5.
6.
7.
Left wing tanks (4) fuel quantity gage
Left auxiliary fuel pump switch
Left transfer pump switch
Crossfeed valve switch
Left or right transfer test switch
Right transfer pump switch
Right auxiliary fuel pump switch
8.
9.
10.
11.
12.
13.
14.
Right wing tanks (4) fuel quantity gage
Right nacelle tank fuel quantity gage
Left nacelle tank fuel quantity gage
Right firewall shutoff valve switch
Switch guard
Fuel system circuit breaker panel
Left firewall shutoff valve switch
Figure 2-12. Fuel Management Panel
2-23
TM 55-1510-215-10
(1) Auxiliary fuel pump switches.
Two
switches, placarded AUX PUMP ON and OFF, located
on the fuel management panel (fig. 2-12) control a
submerged fuel pump located in the corresponding
nacelle tank. During normal aircraft operation both
switches are OFF (except during takeoff, landing or
crossfeed, refer to chapter 9) so long as the enginedriven boost pumps function. The loss of fuel pressure,
due to failure of an engine driven boast pump will initiate
two flashing MASTER CAUTION lights on the instrument
panel and will illuminate the yellow L FUEL FAIL or R
FUEL FAIL on the caution annunciator panel (fig. 2-22).
Turning ON the AUX FUEL PUMP will extinguish the
FUEL FAIL lights. The MASTER CAUTION lights must
be manually turned off.
(2) Fuel transfer pump switches.
Two
switches on the fuel management panel (fig. 2-12),
placarded TRANSFER PUMP, ON and OFF control
arming of the fuel transfer pumps in the normal mode.
During normal operation both switches are on, which
allows the pump to be automatically turned off and on by
a quantity sensor located in each nacelle tank. If either
transfer pump fails to operate when switched ON and
triggered to function by its quantity sensor, the fault
condition is indicated by two flashing MASTER
CAUTION lights on the instrument panel and a steadily
illuminated yellow FUEL XFR light on the caution
annunciator panel (fig. 2-22).
(3) Fuel transfer test switch.
A switch,
placarded TRANSFER TEST (fig. 2-12) on the fuel
management panel provides a means of checking the
operation of either fuel transfer system. This switch is a
three-position toggle type, spring-loaded to the OFF
(center) position. When positioned to either L (left) or R
(right) the switch applies power to the selected transfer
pump by bypassing the normal automatic circuit. If the
nacelle tank is full, the selected transfer pump will be
energized momentarily, which is enough to establish the
operating status of that transfer system, indicated by the
momentary flash of a yellow L FUEL XFR or R FUEL
XFR indicator light on the caution annunciator panel
(fig. 2-22).
(4) Fuel transfer indicator lights. When all
usable fuel has been transferred from a wing tank
system, a sensing switch detects the pressure drop in
the fuel transfer line. After 30 seconds the affected
transfer pump is shut off. This will illuminate the flashing
MASTER CAUTION lights (fig. 2-22) and the appropriate
yellow L or R FUEL XFR indicator light on the caution
2-24
annunciator panel (fig. 2-22).
NOTE
The L or R FUEL XFR light will also
serve as an operation indicator for
the designated transfer pump. If the
light illuminates, and the respective
wing tank gage does not show
empty, the transfer pump has
stopped transferring fuel into the
nacelle tank.
A positive transfer
pump test (as accomplished in
chapter 8 during ENGINE RUNUP)
without fuel transfer indicates a
nacelle tank fuel quantity sensor
failure. Holding the TRANSFER TEST
switch to ON will override the fuel
quantity sensors and allow transfer
of the remaining fuel. The L or R
FUEL XFR light may be extinguished
by turning the TRANSFER PUMP
switch to off.
(5) Fuel crossfeed switch. The fuel crossfeed
valve (fig. 2-13) is controlled by a two position switch
(fig. 2-12), located on the fuel management panel,
placarded OPEN, and CLOSED. Under normal flight
conditions the switch is left in the CLOSED position. For
crossfeed system operation, refer to chapter 9.
Crossfeed operation is indicated by the illumination of
the yellow FUEL CROSSFEED indicator light on the
caution annunciator panel (fig. 2-22), when the switch is
placed in the OPEN position. The crossfeed valve is
protected by a 5-ampere circuit breaker placarded
CROSSFEED VALVE located on the fuel system circuit
breaker panel (fig. 2-22).
(6) Fuel crossfeed indicator light. Illumination
of the yellow FUEL CROSSFEED indicator light on the
caution annunciator panel (fig. 2-22) indicates that the
electrically operated crossfeed valve is open.
NOTE
The fuel crossfeed light may remain
illuminated if the crossfeed valve
closes due to malfunction. The same
circuit that opens the crossfeed valve
also illuminates the fuel crossfeed
light.
TM 55-1510-215-10
NOTE
DIAGRAM
SHOWS
TYPICAL
FUEL
CROSSFEED SITUATION WITH LEFT WING
FUEL SYSTEM SUPPLYING BOTH ENGINES
(ALL
AUX.
AND
TRANSFER
PUMPS
OPERABLE).
Figure 2-13. Crossfeed Fuel Flow
Change 7 2-25
TM 55-1510-215-10
(7) Firewall shutoff valves.
CAUTION
Do not use the fuel firewall shutoff
valve to shut down an engine, except
in an emergency. The engine-driven
high pressure 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.
Two guarded switches, placarded FIREWALL
SHUT-OFF VALVE on the fuel management panel, (fig.
2-12), are provided to give the pilot electrical fuel shutoff
capability at each engine firewall. Each switch is a twoposition unit controlling the corresponding firewall shutoff
valve. OPEN position opens the firewall shutoff valve
admitting fuel to the engine. In the CLOSED position
fuel flow to the affected engine is cut off, thereby
isolating the fuel supply from that engine, although the
isolated fuel may be supplied to the opposite engine by
crossfeed. A hinged red-colored metal guard engages
each firewall valve switch toggle when the switch is in
the OPEN position. This guard prevents inadvertent
movement of the switch to the CLOSED position. The
guard must be manually disengaged from the switch
toggle to move the switch to the CLOSED position. The
firewall shutoff valves are protected by two 5-ampere
circuit breakers placarded FIREWALL VALVE, located
on the fuel system circuit breaker panel (fig. 2-12).
f. Fuel Tank Sump Drains. The fuel system tanks
and interconnecting lines may be drained of moisture
condensate and sediment by means of 10 drains (plus
two for the ferry system when installed), located at the
system low points on the nacelle tanks, wing tanks and
at fuel filter drain at the inertia separator air bypass duct.
g. Fuel Vent System. Three fuel vents are located
under each wing. One serves both the nacelle and wing
tanks, and is protected against icing conditions by an
electric heating element. (Refer to paragraph 2-57 for
fuel system anti-icing information.)
h. Thermal Pressure Relief Value.
Volume
expansion in the fuel system is relieved by a thermal
pressure relief valve. Normally, thermal expansion
occurs only during hot weather while the aircraft is static
on the ground.
i. 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
2-26 Change 10
might contain. The temperature of the delivered fuel is
thermostatically regulated to remain between + 21°C
and + 32°C.
2-36. Fuel Management
CAUTION
During normal fuel system operation
all usable fuel of each system can be
used. However, if a transfer pump
becomes inoperative, approximately
28 gallons of wing fuel from the
respective system cannot be used
because gravity feed will stop with
approximately 28 gallons of fuel
remaining within the respective wing
tanks (fig. 2-14).
a. Fuel Transfer System. Fuel used from nacelle
tanks is replenished by gravity feed alone, until the level
within the tank is depleted to 8 gallons (52 pounds) (fig.
2-14). At the 8 gallon (52 pound) level, the low switch
position on the level sensor float is actuated which
causes fuel transfer to start. When the fuel quantity
rises to 24 gallons (156 pounds), transfer action is cutoff
by the second switch position on the float. Unless the
pilot uses manual transfer control after the first transfer
cycle, all subsequent fuel transfer will maintain quantity
within the nacelle tanks at a level between 8-24 gallons
(52-156 pounds), until all fuel is used from the wing
tanks.
NOTE
During normal fuel system operation
(not the TRANSFER TEST mode), a
fuel transfer pump will be activated
when approximately 8 gallons remain
in the nacelle tank. If that transfer
pump should become inoperative, 28
gallons (182 pounds) of fuel in the
inboard wing tank will not be usable
since it cannot gravity feed (fig. 2-13).
With this situation, only 8 gallons (52
pounds) or less of fuel will be
available
for
continued
flight
(approximately 10 minutes flying
time).
Crossfeed fuel (fig. 2-13),
however, may be used for continued
engine
operation
utilizing
the
operative fuel transfer pump and the
auxiliary fuel boost pump from the
opposite engine's fuel system. Do
not operate with the crossfeed in the
OPEN mode with both auxiliary fuel
pumps operating.
Fuel may be
inadvertently crossfeed from either
fuel system due to normal variances
in pump pressure.
TM 55-1510-215-10
NOTE
•
DIAGRAM SHOWS TYPICAL GRAVITY FEED FUEL
FLOW FOR THE LEFT SIDE.
LEFT SYSTEM IS
IDENTICAL TO RIGHT EXCEPT THE FORMER
CONTAINS THE CROSSFEED VALVE, PARALLELED BY
A THERMAL RELIEF VALVE, AND THERE IS A
THERMAL RELIEF VALVE AND LINE FROM THE CROSS
LINE IN THE RIGHT FUEL SYSTEM.
•
EACH WING SYSTEM WILL GRAVITY FEED ONLY TO
IT’S RESPECTIVE ENGINE, I.E. LEFT OR RIGHT. FUEL
WILL NOT GRAVITATE THROUGH THE CROSSFEED
SYSTEM.
•
THE ENGINE DRIVEN PRIMARY (HIGH PRESSURE)
FUEL PUMP IS LIMITED TO 10 CUMULATIVE HOURS OF
OPERATION THROUGHOUT IT’S TBO PERIOD WITHOUT
AUXILIARY FUEL PUMP OR ENGINE-DRIVEN BOOST
PUMP FUEL PRESSURE.
Figure 2-14. Gravity Fuel Flow.
2-27
TM 55-1510-215-10
(1) Transfer pump manual operation. While
sufficient fuel remains in the wing tanks, the pilot may
use manual control to refill the nacelle tanks to capacity.
Refilling must be conducted with the TRANSFER PUMP
switch on, and by placing TRANSFER TEST switch at L
or R position (whichever applies). If the TRANSFER
TEST switch is held at the L or R setting momentarily,
then released, the initiated fuel transfer will be shut off
by either the 24 or 57 gallon (156 or 370.5 pound) float
switch position, whichever the rising fuel level
encounters first. If the TRANSFER TEST switch is
manually retained in the ON position fuel transfer will
continue until the switch is released. In this instance, the
transfer pump shutoff circuits on the 24 and 57 gallon
(156 and 370.5 pound) float switches are bypassed, and
if enough fuel is present in the wing tanks, the nacelle
tank may be overfilled, resulting in excess fuel flowing
through the vent lines back into the wing tanks. A safe
continuous transfer loop is thus established but is not a
recommended procedure. Unless a nacelle tank is full to
capacity whenever a TRANSFER PUMP switch is turned
ON, after an OFF period, the affected pump will start an
immediate transfer cycle. Transfer will continue until fuel
within the receiving tank rises enough to actuate a floatswitching position which was not already exceeded by
the fuel level when the TRANSFER PUMP switch was
placed ON. If fuel quantity before the start of transfer
was between 8-24 gallons (52-156 pound), transfer
power will be cut of at the 24 gallon (156 pound) floatswitching level. However, if quantity within a nacelle
tank is above the 24 gallon (156 pound) switching level,
when the TRANSFER PUMP switch is turned ON, a
transfer will initiate, continuing to fill the tank until the top
float-switching position of 57 gallons (370.5 pounds) is
reached.
(2) Top nacelle fuel quantity sensor failure
(continuously full nacelle tank). If fuel transfer into a
nacelle tank should not be terminated by actuation of the
top level quantity sensor, over-fill fuel will flow back into
the wing tanks through vent lines. Operation would
stabilize into a continuous fuel transfer loop.
A
continuously full nacelle tank, monitored on the gage,
would indicate this condition and can be corrected by
using the following procedure:
1. Reset the appropriate TRANSFER
2-28 Change 10
PUMP switch to OFF.
2. When the gage of the affected tank
indicates a quantity below 24 gallons (156 pound), reset
the TRANSFER PUMP switch to ON. This will reset the
automatic level control between the limits of 24-57
gallons (156-370 pounds).
(3) Transfer pump operational check. When
engines are running, transfer pump operation must be
checked by a method which does not depend on sound.
Another factor is that transfer may be already underway
when the pilot decides to conduct the transfer check. He
cannot hear pumps running, and if transfer is in
progress, the associated FUEL XFR annunciator light
will not flash, due to pressure within the transfer line.
For fuel transfer pump operational check procedure,
refer to chapter 8.
b. Operation with failed engine-driven boost pump
or auxiliary pump. Two pumps in each fuel system
provide inlet head pressure to the engine-driven primary
high-pressure fuel pump and if crossfeed is used, a third
pump, the auxiliary fuel pump from the opposite system,
will supply the required pressure. 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
engine-driven primary high-pressure fuel pump, but only
for a 10-hour period due to the limitation on the primary
pump. The total time that the engine-driven primary
high-pressure pump is operated without fuel being
supplied under pressure from the engine-driven boost
pump or an auxiliary pump shall be entered on DA Form
2408-13.
2-37. Ferry Fuel System.
CAUTION
Do not use AVGAS in the ferry fuel
system. It has not been tested using
AVGAS. Flow rates will be less, and
consequently ferry fuel pump may
not be able to keep up with engine
demand.
TM 55-1510-215-10
a. Description. The ferry fuel system (fig. 2-15)
provides additional fuel for ferry missions. The system
consists of two removable fuel tanks, a removable ferry
fuel control panel assembly, a permanently installed vent
system and permanently installed interconnecting
plumbing and wiring.
A 15-ampere circuit breaker
located on the lower wall behind the copilot's seat
placarded FERRY SYSTEM POWER protects the ferry
fuel system.
(1) Ferry fuel tanks.
Two removable
rectangular aluminum fuel tanks of 120 gallons (780
pounds) capacity each are bolted to the seat tracks in
the cabin for ferry missions. A water sump and drain
valve is provided at the lower aft end of each tank.
(a) Ferry tank fuel gages. One float
actuated dial type fuel gage is provided on the forward
end of the top side of each ferry fuel tank.
(b) Ferry fuel tank shutoff valves. A
shutoff valve is located on the fuel line between each
ferry fuel tank and the ferry fuel management panel.
(2) Ferry fuel control panel assembly. This
unit is removed from the aircraft when the ferry fuel
system is not required. It mounts to the seat tracks
across the aisle directly aft of the pilot's and copilot's
seats. It consists of two fuel switching valves, two
circuit-breaker switches, a fuel filter, an electric fuel
pump, and a manual fuel pump.
(a) Manual fuel pump. If the electric
fuel pump fails to operate, fuel may be pumped from the
ferry fuel system tanks into the wing tanks by means of a
manual fuel pump, located on the copilot's side of the
fuel control panel assembly. When operated at 34
strokes per minute, the manual fuel pump will transfer 48
gallon/hour (315 pounds/hour) of fuel to the left wing and
48 gallon/hour (315 pounds/hour) of fuel to the right wing
simultaneously (a total of 97 gallons/hour (630 pounds)).
The manual fuel pump will transfer 80 gallons/hour (520
pounds/hour) of fuel to one wing only. To maintain an
equal fuel flow with engine consumption set at 200
pounds/hour, 22 strokes/minute are required.
(b) Ferry fuel selector vale.
A four
position valve located on the ferry fuel control panel,
placarded OFF, LEFT WING TANK, BOTH, RIGHT
WING TANK, controls the flow of fuel from the ferry
tanks.
(c) Electric fuel pump and switch. Fuel
is normally pumped from the ferry tanks by an electric
fuel pump, located in the ferry system control panel
assembly. It is controlled by a 7.5-ampere circuitbreaker switch, located on the ferry system fuel control
panel, placarded FUEL PUMP, OFF, ON. The electric
fuel pump will transfer 46 gallons/hour (300
pounds/hour) of fuel to the left wing and 46 gallons/hour
(300 pounds/hour) of fuel to the right wing
simultaneously (a total of 92 gallons (600 pounds/hour)).
The electric fuel pump will transfer 77 gallons/hour (500
pounds/hour) of fuel to one wing only.
(d) Ferry system fuel vent heater
switch. The ferry system fuel vent heater is controlled by
a 7.5-ampere circuit-breaker switch located on the ferry
system fuel control panel, placarded VENT HEAT, OFF,
ON.
(e) Manual fuel pump control valve. A
two position valve located on the ferry fuel control panel
placarded FUEL SELECTOR - OFF, ON controls the
flow of fuel through the ferry fuel system. Fuel cannot be
pumped into the wing tanks by either the electric or
manual pump unless this valve is in the ON position.
(f) Ferry system fuel filter. Fuel from
the ferry tanks is pumped through a fuel filter, located on
the copilot side of the fuel system control panel
assembly, placarded FUEL FILTER.
(3) Ferry fuel tank vent system. The ferry fuel
tanks are connected to a heated fuel vent located on the
underside of the aft fuselage. Secondary tank vents are
installed on the top of each tank.
(4) Ferry system wheel well fuel shutoff
valves. A fuel shutoff valve is installed on the ferry
system fuel line in the forward section of each main
wheel well. These valves must be open and secured
with safety wire when the ferry system is being used.
NOTE
A safety of flight release is required
for takeoff above maximum takeoff
weight.
b. Normal Operation.
WARNING
Smoking in the aircraft is prohibited
when the ferry system is installed.
2-29
TM 55-1510-215-10
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
Figure 2-15. Ferry Fuel System
2-30
Tank sump drains
Tank vent lines
Tank fuel gages
Filler caps
Tank shutoff valves
Ferry fuel management pane
Power cable
Manual fuel pump handle
Wing tank selector valve
Fuel pump switch
Vent heat switch
Fuel pump control valve
TM 55-1510-215-10
WARNING
Continuing to pump fuel from the
ferry tanks into the wing tanks after
they are full will result in fuel being
dumped overboard through the wing
fuel tank vent system.
12. Ferry fuel tanks - Fill in accordance
with fuel handling precautions and servicing instructions
in Section XII, Servicing, Parking and Mooring.
13. Ferry system
connectors - Check for leaks.
fuel
lines
and
14. Left ferry fuel tank shutoff valve Off.
(1) Preflight functional test.
NOTE
15. Right wing tank fuel gage - Note
indicated fuel quantity.
The aircraft's wing tanks should be
nearly empty before beginning the
ferry system preflight functional test.
This will provide tank space for fuel
transferred from the ferry tanks.
16. Right ferry tank fuel gage - Note
indicated fuel level. Remove right ferry tank filler cap
and visually check that indicated fuel quantity
corresponds to gage indication.
1. Ferry fuel tank shutoff valves (two,
located between ferry tank outlet and ferry fuel control
panel) - On.
17. Manual fuel pump control valve FUEL ON.
NOTE
2. Ferry fuel selector valve - OFF.
The manual fuel pump control valve
must be in the FUEL ON position to
transfer fuel whether using the
manual or electric fuel pump.
3. Manual fuel pump control valve OFF.
4. Heated fuel vent switch - OFF.
5. Ferry system electric fuel pump
switch - OFF.
6. Wing fuel drains - Closed.
7. Ferry tank water sumps (two) Closed.
8. Ferry system wheel-well fuel shutoff
valves (two) - Check open and safetied.
9. Aircraft
switches (all) - Off.
electrical
equipment
10. GPU - Connect.
11. Battery switch - ON.
WARNING
The battery switch may be left on to
monitor fuel quantity gages and ferry
fuel system electric pump, but must
not be moved during fueling
operations.
18. Ferry fuel selector valve - RIGHT
WING TANK.
19. Ferry system electric fuel pump ON. Transfer approximately 30 gallons of fuel (roughly
one quarter of right ferry tank capacity) from the right
ferry tank to the right wing tank system. Transfer of 30
gallons of fuel should take approximately 40 minutes.
20. Right ferry system fuel lines and
connectors - Check for leaks while fuel is flowing.
21. Ferry system electric fuel pump OFF.
22. Ferry fuel selector valve - OFF.
23. Right wing tank fuel gage - Check
for increased indicated fuel quantity of approximately
one quarter tank.
24. Right ferry tank fuel gage - Note
indicated fuel level. Remove right ferry tank filler cap
and visually check that gage reading approximately
corresponds with quantity of fuel that has actually been
transferred.
2-31
TM 55-1510-215-10
25. Right ferry fuel tank shutoff valve -
39. Ferry fuel selector valve - BOTH.
Off.
26. Left ferry fuel tank shutoff valve On.
27. Ferry fuel selector valve - LEFT
40. Ferry system
ON.
Transfer approximately
(approximately one quarter of the
tanks). Transfer of 60 gallons
approximately 40 minutes.
electric fuel pump 60 gallons of fuel
capacity of both ferry
of fuel should take
WING TANK.
41. Ferry system electric fuel pump 28. Left wing tank fuel gage - Note
indicated fuel quantity.
OFF.
42. Ferry fuel selector valve - OFF.
29. Left ferry tank fuel gage - Note
indicated fuel level. Remove left ferry tank filler cap and
visually check that indicated fuel quantity corresponds to
gage indication.
43. Ferry fuel tank shutoff valves (2) On.
44. Manual fuel pump control valve -
30. Ferry system electric fuel pump
switch - ON. Transfer approximately 30 gallons of fuel
(approximately one quarter of left ferry tank capacity)
from the left ferry tank to the left wing tank system.
Transfer of 30 gallons of fuel should take approximately
40 minutes.
OFF.
45. Ferry fuel system vent heat switch ON. Check that temperature of heated vent increases
then turn switch OFF.
46. Battery switch - OFF.
31. Left ferry fuel system fuel lines and
connectors - Check for leaks while fuel is flowing.
32. Ferry system electric fuel pump
switch - OFF.
47. GPU - Disconnect as required.
48. Ferry fuel tanks - Fill in accordance
with fuel handling precautions and servicing instructions
in Section XII, Servicing Parking and Mooring.
33. Left wing tank fuel gage - Check for
increased indicated fuel quantity of approximately one
quarter tank.
NOTE
Settling time for jet fuels is one hour
per foot of tank depth. Allow the fuel
to settle for the prescribed period of
time before any fuel samples are
taken.
34. Left ferry tank fuel gage - Note
indicated fuel level. Remove left ferry tank filler cap and
visually check that indicated fuel quantity corresponds to
gage indication.
35. Manual fuel pump - Operate at
approximately 34 strokes per minute.
Transfer
approximately 10 gallons of fuel in to left wing tank
system. Visually check fuel level inside left ferry tank
before and after manual fuel pump check to insure that
fuel actually has been transferred.
36. Wing tank fuel gages (2) - Note
indicated fuel quantity.
37. Ferry tank fuel gages (2) - Note
indicated fuel quantity.
38. Right ferry fuel tank shutoff valve ON.
2-32
49. Fuel sample - Take from each ferry
tank drain.
50. Ferry fuel tank fuel and caps Check fuel level visually. Check seal is installed, cap is
tight and properly installed.
(2)
Before takeoff. Heated fuel vent switch -
ON.
(3) During flight.
1. Wing fuel gages - Monitor (until wing
tanks are 1/2 to 3/4 full).
TM 55-1510-215-10
2. Ferry system fuel selector valve LEFT WING TANK, BOTH or RIGHT WING TANK as
required.
3. Manual fuel pump selector valve ON.
4. Ferry system electric fuel pump
switch - ON.
2-38. Approved Fuels.
system. It has not been tested using
AVGAS. Flow rates will be less, and
consequently ferry fuel pump may
not keep up with engine demand.
The aircraft may use JP-4 or JP-5 in any ratio
Civilian jet fuel may not contain an anti ice/fungicide and
may thicken when temperature drop below -40F. In the
event aviation kerosene is not available, aviation
gasoline is an approved emergency fuel. Refer to
Section XII.
CAUTION
NOTE
The use of aviation gasoline is time
limited to 150 hours of operation
during any Time-Between-Overhaul
(TBO) period.
It may be in any
quantity with aviation kerosene.
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)
shall be used. If more than 10% of
the total fuel on board is AVGAS, the
total operating time shall be entered
in DA Form 2408-13.
CAUTION
Do not use AVGAS in the ferry fuel
Section V. FLIGHT CONTROLS
2-39. Description.
2-41. Rudder Pedals.
The aircraft's primary flight control system consists
of rudder, elevator and aileron control surfaces. These
surfaces are manually operated from the cockpit through
mechanical linkage using control wheels for the ailerons
and elevators, and adjustable rudder/brake pedals for
the rudder. Trim control for the rudder, elevator and
ailerons is accomplished through a manually actuated
cable-drum system for each set of control surfaces.
Aircraft rudder control and nose wheel steering is
accomplished by actuation of the rudder pedals from
either pilot's or copilot's station (fig. 2-5). A fore and aft
position adjustment of the pedals is provided through an
adjustment control lever on each pedal. For toe brake
coverage refer to paragraph 2-8.
2-40. Control Wheels.
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 consisting of two pins and an
elongated U-shaped strap interconnected by a chain (fig.
2-16). Installation of the control locks is accomplished
by inserting the strap over the aligned engine 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 wheels. The rudder pedals are held in the
Elevator and aileron control surfaces are
operated by manually actuating either the pilot's or
copilot's control wheel. Electric switches are installed in
the outboard grip of each wheel to operate the elevator
trim tabs, to disengage the autopilot/yaw damp, control
wheel steering, and press-to-talk microphone switch.
These control wheels (fig. 2-16) are installed on each
side of the instrument subpanel. An electric digital clock
is installed in the center of each wheel.
2-42. Flight Controls Lock.
Change 5 2-33
TM 55-1510-215-10
1.
2.
3.
4.
Clock
Control wheel steering switch
Electric trim switches
Microphone switch
5. Autopilot/yaw damp/trim switch
6. Engine controls lock bar
7. Rudder lock pin
AP011649
Figure 2-16. Control Wheels and Control Locks
neutral position by the largest of the two pins, which is
installed horizontally through the pilot's rudder pedals.
Removal sequence is a reverse of the installation
procedure.
2-43. Trim Tabs.
Trim tabs are provided for all flight control surfaces.
These tabs are manually actuated, and are mechanically
controlled by a cable-drum and jack-screw actuator
system. Elevator and aileron trim tabs incorporate antiservo action, i.e., as the elevators or ailerons are
displaced from the neutral position, the trim tab moves in
the same direction as the applied control surface, thus
increasing the effective control surface area and the
manual force required to apply it. This action increases
control pressure. The rudder trim tab is adjustable left or
right as required and maintains an "as adjusted" position
throughout the full range of rudder deflection.
control pedestal. The dual element thumb switch is
moved forward for trimming nose down, and aft for nose
up. When released, the switch returns to the center (off)
position. Any activation of the trim system through the
copilot's trim switch will be over ridden by activation of
the pilot's switch. A preflight check of the dual element
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 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 system. The
manual trim control wheel and the electric trim system
cannot be used simultaneously.
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. 2-6). The amount of
elevator tab deflection, in degrees from a neutral setting,
is indicated by a position arrow.
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. 2-6). The amount of aileron tab deflection,
from a neutral setting, as indicated 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.
b. Electric Elevator Trim. The electric elevator trim
system is controlled by dual element thumb switches on
the control wheels, a trim disconnect switch on each
control wheel, and a circuit breaker located on the
d. Rudder Trim Tab Control. The rudder trim tab
control knob, placarded RUDDER TAB - LEFT,
RIGHT on the control pedestal, controls adjustment
of the rudder trim tab (fig. 2-6). The amount of rudder
2-34 Change 10
TM 55-1510-215-10
tab deflection, in degrees from arrow.
2-44. 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, either up or down, is indicated in
percent of travel by a flap position indicator on the center
of the control pedestal. Full flap extension and retraction
time is approximately 11 seconds. The flap control
switch is also located on the control pedestal. No
emergency wing flap actuation system is provided. The
circuit is protected by a 20-ampere push-pull circuit
breaker, placarded FLAP MOTOR POWER, located on
the copilot’s circuit breaker panel (fig. 2-19).
a. Wing Flap Control Switch. Flap operation is
controlled by a three position switch with a flap-shaped
handle on the control pedestal (fig. 2-6). The handle of
this switch is discarded: 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 - 35%, 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 IND, located on the right
subpanel (fig. 2-7). Flap positions between UP and
APPROACH cannot be selected. For intermediate flap
positions between APPROACH 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.
b. Wing Flap Position Indicator. Flap position in
percent of travel from 0 percent (UP) to 100 percent
(DOWN), is shown on an indicator, placarded FLAPS on
the control pedestal (fig. 2-6). The approach and full
down or extended flap position is 15 and 43 degrees,
respectively. The flap position indicator is protected by a
5-ampere circuit breaker, placarded FLAP IND, located
on the right subpanel (fig. 2-7).
Section VI. PROPELLERS
2-45. Description.
A three-bladed aluminum propeller is installed on
each engine, These propellers are hydraulically
controlled, constant-speed full-feathering and reversible.
Each propeller is controlled by engine oil acting through
an engine-driven propeller governor.
Feathering is
accomplished by the feathering springs assisted by
centrifugal force applied to the blade shank
counterweights. Governor boosted engine oil pressure
moves the propeller blades to the high RPM (low pitch)
hydraulic stop and into reverse pitch.
Low pitch
propeller position is determined by a mechanically
monitored hydraulic stop. A back-up system, referred to
as the secondary low pitch stop, protects against
propeller reversing in the event of failure of the primary
hydraulic low pitch stop system. In the event of loss of
oil pressure, the propeller blades will go to the feathered
position.
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, if armed, will
sense loss of torque oil pressure and will feather an
unpowered propeller. Feathering springs will feather the
propeller when it is not turning.
a. Automatic Feathering. Automatic feathering can
occur only when the PROP AUTOFEATHER switch is in
the ARM position, both power levers are above 88% to
92% N1 and the torque value of one engine drops below
160 to 290 ft-lbs. The autofeather system has a crossinterlocking safety feature designed into the control
circuit to prevent automatic feathering of both propellers.
Before a propeller feathers automatically, the interlock
disarms, the autofeather circuit of the opposite propeller.
After autofeathering has occurred for one propeller the
opposite propeller can be feathered only by the manual
control.
2-46. Feathering Provisions.
b. Propeller Autofeather Switch.
The aircraft is equipped with both manual and
automatic propeller feathering. Manual feathering is
Change 7 2-35
TM 55-1510-215-10
Autofeathering is controlled by a PROP AUTHFEATHER
switch on the subpanel (fig. 2-7). The three-position
switch is placarded ARM, OFF and TEST, and is springloaded from TEST to OFF. The ARM position is used
only during takeoff and landing. At ARM, if an engine
loses power above 88% to 92% N1, two torque-sensing
switches of the affected engine are actuated by loss of
torque pressure.
Switch actuation applies current
through an autofeather relay to a corresponding dump
valve, causing the release of oil pressure which held an
established pitch angle on the blades of the affected
propeller Following the release of oil pressure, feathering
movement is accomplished by the leathering springs
assisted by centrifugal force applied to the blade shank
counterweights.
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. Refer to chapter 8.
c. Autofeather Lights. Two amber lights on the
instrument panel, placarded AUTOFEATHER LEFT and
RIGHT, when illuminated indicate that the autofeather
system is armed (fig.
2-22).
Both lights will be
extinguished if other propeller has been autofeathered or
if the system is disarmed by retarding a power lever.
Both lights may be dimmed by rotating the cover.
Autofeather circuits are protected by one 5-ampere
circuit breaker placarded PROP FEATHER, located on
the right subpanel (fig. 2-7).
2-47. Propeller Governors.
Each propeller system utilizes three governors, one
“primary" and two '”backup", to control propeller RPM.
Each propeller lever establishes RPM for the respective
propeller by altering the setting of a primary governor
attached to the engine gear reduction housing. It is the
primary governor which controls RPM through the entire
range.
Should a primary governor malfunction
(exceeding 2200 RPM) an overspeed governor cuts-in
(2248 to 2328 RPM), dumping oil from the propeller to
prevent RPM from exceeding safe limits. A solenoid
actuated by the PROP GOV TEST switch enables the
overspeed governor to be reset for test purposes (1980
to 2060 RPM). If a propeller should stick or move too
slowly during a transient condition, the corresponding
governor would be unable to prevent an overspeed
condition. To provide for this contingency, the engine
power turbine governor acts as a fuel topping governor.
Thus, when the propeller RPM reaches 2332 RPM, this
governor limits the fuel flow into the engine thereby
reducing the power during the propeller.
During
2-36 Change 7
propeller operation in the reverse range, the engine
power turbine governor will automatically be reset to
allow a maximum of 2040 RPM.
2-48. Propeller Governor Test Switches.
Two, three-position switches on the left subpanel,
are provided for operational test of the propeller systems
(fig. 2-7). The switches are placarded PROP GOV
TEST, LEFT, RIGHT (above) and SECONDARY IDLE
STOP TEST (below). Each switch is a double unit,
controlling two different test circuits for the
corresponding propeller. In up position, the switches are
used to test the function of the corresponding overspeed
governor. In the down position, the switches are used to
test function of the corresponding secondary low pitch
stop. Each switch is spring-loaded to the OFF (center)
position. Refer to chapter 8, for steps of both test
procedures. Propeller test circuits are protected by one
5-ampere circuit breaker, placarded PROP GOV IDLE
STOP, located on the fight subpanel (fig. 2-7).
2-49. Propeller Levers.
Two propeller levers on the control pedestal,
placarded PROP, are used to regulate propeller speeds
(fig. 2-6). Each lever controls a primary governor, which
acts to regulate propeller speeds within the normal
operating range. The full levers forward position is
placarded TAKEOFF, LANDING AND REVERSE, and
also HIGH RPM. Full levers aft position is placarded
FEATHER. When a lever is placed at HIGH RPM, the
propeller may attain a static RPM of 2200, 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.
2-50. Propeller Reversing.
CAUTION
Moving the power levers into reverse
range without the engine running will
result in damage to the reverse
linkage mechanisms.
TM 55-1510-215-10
CAUTION
To prevent damage to reversing
linkage and an asymmetric 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 are
positioned at HIGH RPM (full forward), and the power
levers are lifted up to pass over an IDLE detent, then
pulled aft into REVERSE setting. In REVERSE position,
each power lever overrides the corresponding secondary
idle stop, allowing engine power for the beta and reverse
ranges. Power levers must be pulled back through
normal idle speed range before being positioned in
REVERSE.
a. Propeller Reverse Not-Ready Annunciator Light.
One yellow caution light, placarded REVS NOT READY,
on the caution annunciator panel alerts the pilot not to
reverse the propellers (fig. 2-21). 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 5-ampere circuit breaker,
placarded LDG GR CONTROL, located on the fight
subpanel (fig. 2-7).
b. Propeller Primary Pitch Annunciator Lights. Two
yellow lights placarded L PRI PITCH and R PRI PITCH
are located on the caution annunciator panel (fig. 2-22).
Illumination of a light indicates malfunction of the normal
propeller low-pitch stop, and that the secondary low pitch
functions for the affected propeller. Propeller reverse
pitch operation must not be attempted while either of
these lights are illuminated. Refer to chapter 9.
2-51. Propeller Tachometer.
Two tachometers on the instrument panel register
propeller speed in hundreds of RPM (fig. 2-22). Each
indicator is slaved to a tachometer generator unit
attached to the corresponding engine.
Section VII. UTILITY SYSTEMS
2-52. Defrosting System.
a. Description.
The defrosting 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 right subpanel (fig. 2-7), 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. Normal Operation
NOTE
For maximum windshield defrosting,
pull out the CABIN AIR, VENT AIR,
and DEFROST AIR controls and
position the HTR switch to MAN.
Regulate the temperature by opening
one or more of the cold air outlets.
c. Emergency Operation.
If the automatic
temperature control should fail to operate, the
temperature (of defrost air and cabin air) may be
controlled manual by manipulating the HTR control
switch between the OFF and MAN positions.
1. Vent blower operation - Check.
2-53. Surface Deicer System.
2. HTR switch - AUTO.
a. Description. Ice accumulation is removed from
each outboard wing leading edge, both horizontal
stabilizers, and the vertical stabilizer 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.
3. CABIN TEMP control - As required.
4. CABIN AIR, VENT AIR, and DEFROST
AIR controls - As required.
2-37
TM 55-1510-215-10
A pressure regulator protects the system from overinflation.
When the system is not in operation, a
distributor valve applies vacuum to the boots for holddown. When a solenoid in the distributor valve is
energized by the pneumatic deicer timer, or when either
the surface or antenna DEICE CYCLE switch is
positioned to MNL (manual), a servo valve changes the
inlet to the boots from vacuum to pressure which allows
the boots to inflate. When the solenoid valve is deenergized, the airflow through the valve is cut off. The
air then discharges out of the boots through an integral
check valve until the pressure reaches approximately 1
in Hg, at which time the boots are ported to vacuum and
the remaining air is evacuated. The boots are again
held down by vacuum.
(2) Normal operation of the surface deice
system is accomplished by use of the three-position
DEICE CYCLE switch, which has a momentary SGL
(single) position, and a momentary MNL, (manual)
position. The switch returns to the center, OFF position,
when the toggle is released. When switched to the SGL
position, the deicer boots automatically inflate for seven
to eight seconds, then deflate and return to the vacuum
hold down position. In MNL position, all boots inflate
and stay inflated while the switch is in this mode. When
the switch is released, all boots deflate. The manual
position is for use in case of timer failure. In either
switch position, the boots cannot be over-inflated.
(1) 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 single-engine operation.
Bleed air passes
through a 16 PSI regulator and then enters the deicer
and vacuum systems. Vacuum pressure is created by
the ejector and is proportional to pneumatic pressure
supplied by the deicer pressure regulator valve.
Regulated pressure is indicated on a gage, placarded
DE ICING PRESS, located on the control pedestal
(fig. 2-6).
a. Description.
Elecrothermal deicer boots are
cemented to each propeller blade to prevent ice
formation or to remove ice from the propellers. Each
thermal boot consists of one outboard and one inboard
heating element, and receives electrical power from the
deicer timer. This timer sends current to all propeller
thermal boots and prevents the deicers from overheating
by limiting the time each element is energized. Four
intervals of approximately 3 seconds each complete one
cycle. Current consumption is monitored by a PROP
AMMETER on the left subpanel (fig. 2-7). A 20-ampere
circuit breaker switch, placarded PROP, on the left
subpanel (fig. 2-7), controls the propeller electrothermal
deicer system.
(2) The operation of the surface deicer
system is controlled by a three-position switch located
on the left subpanel. The switch is placarded DEICE
SURF, SGL, MNL.
The surface deicer system is
protected by two 5-ampere circuit breakers, placarded
SURF DEICE located on the fight subpanel.
b. Normal 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 on the boots to accumulate before attempting
ice removal. Very thin ice may crack and cling to the
boots instead of shedding.
NOTE
Never cycle the system rapidly, since
this may cause the ice to accumulate
outside the contour of the inflated
boots and prevent ice removal.
2-38
2-54. Propeller Electrothermal Deicer System.
b. Normal Operation. Operation of the propeller
deicing system is controlled by the propeller heat switch,
placarded PROP, which controls two inboard and
outboard, heating elements in each propeller boot.
When ice formation becomes visible on the aircraft, or
when ice is expected, place this switch in the HEAT
position. The timer will then cycle power to the heating
elements. The timer successively delivers current to the
outer heaters on one propeller, the inner heaters on the
same propeller, the outer heaters on the opposite
propeller and the inner heaters on the same propeller.
The timer energizes each of these four phases in turn for
about 30 seconds and then repeats the cycle as long as
the control switch is on. When the timer shuts off, it
advances one cycle. Each cycle is 30 seconds in
duration, which makes a complete cycle lasting two
minutes. When the timer switches from one phase to
the next, the ammeter will register a momentary
deflection. These fluctuations inform the pilot that the
timer is switching properly.
TM 55-1510-215-10
NOTE
On aircraft equipped with electronic
timers, cycle switching occurs very
rapidly and may or may not be
detectable as a flicker of the ammeter
needle.
These timers cannot be
manually stepped through cycles by
alternately
energizing
and
deenergizing the system on/off switch.
Heating may begin at any phase in the cycle depending
on the timer position when the switch was turned off
from previous use. If ammeter readings are above or
below the operating limits, propeller unbalance may
occur when operating in icing conditions.
2-55. 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 a single
pilot tube attached to the underside of the left wing
leading edge, 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. Refer to Pitot and Stall Warning Heat System,
Section VIII.
b. Emergency Static Air Source.
A knob type
control valve located at the upper right corner of the
instrument panel permits the selection of an alternate
static air pressure source. It is placarded EMERGENCY
STATIC AIR SOURCE, NORM OFF.
The normal
operating position (NORMAL, OFF) supplies static air
pressure from the external pressure ports on the aft
fuselage.
When required, static pressure may be
obtained from the alternate source by rotating the control
knob counterclockwise. For airspeed calibration, when
using the emergency static air source, refer to chapter 7.
2-56. Pitot and Stall Warning Heat System.
CAUTION
Pitot or stall warning heat should not
be used for more than 15 minutes
while the aircraft is on the ground.
Overheating may damage the heating
elements.
a. Description. The pitot tube and stall warning
vane have electrical heating elements to prevent icing.
Each heating elements is controlled by a 5-ampere,
circuit-breaker type switch, located on the left subpanel,
placarded STALL WARN HEAT (fig. 2-7), and HEAT,
LH PlTOT (fig. 2-7).
b. Normal Operation. When the pitot heat switch is
in the LH (up) position, the heating element in the
exposed position of the pitot tube is energized. The
PITOT (down) position shuts the pitot heating element
off. The stall warning heat switch activates the healing
element in the stall warning vane when the switch is in
the STALL WARN (up) position. The HEAT (down)
position of the switch shuts off the vane heating element.
In the event of overload, the circuit breaker element of
either switch will disconnect the respective heating
circuit, tripping the switch toggle to a down position. To
reset power to a healing circuit, move the respective
toggle switch to the up position.
2-57. 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. One external fuel vent on each
wing serves both the nacelle and wing tanks and is
protected against icing by externally attached electric
heat elements, controlled by the 5-ampere circuit
breaker switch placarded HEAT-FUEL VENTS-LEFT,
and the 7.5-ampere circuit breaker switch placarded
HEAT-FUEL VENTS-RIGHT (fig. 2-7).
CAUTION
To prevent overheat damage to
electrically heated anti-ice jackets,
FUEL VENT HEAT switches should
not be turned ON unless cooling air
will soon pass over the jackets.
b. Normal Operation. For normal operation, HEAT
switches for the FUEL VENTS anti-ice circuits are turned
ON as required during the BEFORE TAKEOFF
procedures.
FUEL CONTROL HEAT switches are
turned ON during the STARTING ENGINES procedures.
Refer to chapter 8.
Change 6 2-39
TM 55-1510-215-10
2-58. Windshield Electrothermal Anti-Ice Systems.
a. Description. Both pilot and copilot windshields
are provided with an independent electrothermal anti-ice
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, a relay switch, and a control
switch.
Both ON-OFF control switches, placarded
WINDSHIELD ANTI-ICE -PILOT, COPILOT, are located
on the left subpanel (fig. 2-7). Each switch controls one
electrothermal windshield system. The circuits of each
system are protected by a respective 1/2-ampere circuit
breaker and a respective 25-ampere circuit breaker on
the right subpanel, placarded WINDSHIELD ANTI-ICE PILOT, COPILOT (fig. 2-7).
b. Normal Operation.
Each elecrothermal
windshield is activated by placing the corresponding
WINDSHIELD ANTI-ICE switch to the ON position. If
glass temperature is below 43°C, the electrothermal
controller will actuate a relay switch applying power to
the heating wires sandwiched within the glass. A
windshield will warm to a maximum of 43°C and then will
cycle off. It will recycle ON again when the glass
temperature drops 2.5°C below cutoff. Refer to chapter
8.
2-59. Oxygen System.
a. Description. The pilot, copilot, and passengers
will receive oxygen from a supply system which has a
capacity which ranges from 11 cu ft to 214 cu ft,
depending upon the combination of supply cylinders
which are coupled into provisions of the aircraft. Figure
2-17 shows the location of the cylinders, some of which
are in the left nose compartment below the avionics
equipment, and some may be located aft of the main
entrance door in the cabin.
All cylinders are
interconnected, using check valves, so that refilling of all
cylinders can be accomplished through a single filler
valve located in the cabin adjacent to the 11-cu ft
cylinder. Each cylinder has a pressure gage. The
oxygen system pressure gage, placarded OXYGEN
SUPPLY PRESSURE, is located aft of the fuel
management panel. This gage shows the amount of
pressure remaining in the oxygen supply cylinders.
Table 2-2 shows oxygen duration capacities of the
system for all crew and passenger combinations.
Supply lines from the cylinders are routed to two high
pressure, slow-release supply valves located below the
2-40
fuel management panel. Each oxygen supply valve has
a flow governing feature controlled by an integral
pressure differential sensor. When initially turned ON by
the pilot, only a small flow is allowed to meter into the
line. Flow restriction continues until the pressure levels
are equalized at which time the valve opens up
completely, permitting a full flow condition. The forward
valve (fig.
2-17), placarded COCKPIT OXYGEN,
controls high pressure flow to a pressure reducer located
on the forward cabin wall behind the instrument panel.
This reducer lowers pressure to a 300 to 400 PSI range
and routes oxygen to the regulator control panels for the
pilot and copilot. The pilot's and copilot's regulators are
of the diluter demand type.
That is, when in the
NORMAL OXYGEN mode the regulator mixes the
proper amount of oxygen for a given amount of air at
flight altitude. The aft slow-release valve, placarded
CABIN OXYGEN, controls supply to the cabin oxygen
regulators, which provides passenger-regulated 100%
oxygen to the passenger oxygen outlets. A plug-in type
valve at each of six stations above the seat positions
delivers oxygen for passenger use. The 100% oxygen
flows continuously to passengers, when the masks are
coupled into an oxygen outlet.
(1) Regulator control panels.
The pilot's
oxygen regulator control panel is located on the cockpit
sidewall below the fuel management panel.
The
copilot's oxygen regulator control panel is located on the
cockpit sidewall below the copilot's circuit breaker panel.
The cabin regulator panels are located on each side of
the cabin sidewall (fig. 2-17). Each panel contains a
blinker-type flow indicator, a 2000 PSI gage, a red
emergency
pressure
control
lever
placarded:
EMERGENCY-NORMAL-TEST MASK, a white diluter
control lever placarded 100% OXYGEN and NORMAL,
OXYGEN, and a green supply control lever placarded
ON/OFF. The supply control lever turns the individual
regulator ON and OFF. The diluter control lever gives
the selection of either normal or 100% oxygen but acts
to select only when the emergency pressure control
lever is placed in NORMAL position.
CAUTION
When not in use, the diluter control
lever should be left in the 100%
OXYGEN
position
to
prevent
regulator contamination.
TM 55-1510-215-10
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
Provisions for one 64 cu. ft. oxygen cylinder
Provisons for one 11 cu. ft. oxygen cylinder
Cylinder pressure gage
Pressure reducer
Cabin oxygen regulator
Cabin oxygen slow release valve
Pilot/copilot's oxygen regulator control panel (pilot's side)
Cockpit oxygen slow release valve
Provisions for two 64 cu. ft oxygen cylinder
Provisions for one 11 cu. ft .oxygen cylinder
Passenger oxygen outlet (typical 6 places)
Oxygen supply pressure gage
Figure 2-17. Oxygen Supply System
2-41
TM 55-1510-215-10
Table 2-2. Oxygen Duration (sheet 1 of 11)
CREW MEMBER DURATION IN HOURS
CABIN
ALTITUDE
FEET
ONE
MAN
CREW
TWO
MAN
CREW
TWO
MAN
CREW
PLUS
ONE
PASS
TWO
MAN
CREW
PLUS
TWO
PASS
TWO
MAN
CREW
PLUS
THREE
PASS
TWO
MAN
CREW
PLUS
FOUR
PASS
TWO
MAN
CREW
PLUS
FIVE
PASS
TWO
MAN
CREW
PLUS
SIX
PASS
NOTES:
GAGE PRESSURE - PSI, U-21G
11 CUBIC FOOT SYSTEM
1500
1200
900
2000
1800
600
300
10,000
0.6
2.2
0.6
1.9
0.5
1.6
0.4
1.3
0.3
1.0
0.2
0.6
0.1
0.3
15,000
0.8
2.2
0.7
1.9
0.6
1.6
0.5
1.3
0.3
1.0
0.2
0.6
0.1
0.3
20,000
1.0
1.8
0.9
1.6
0.8
1.3
0.6
1.1
0.5
0.8
0.3
0.5
0.1
0.3
10,000
0.3
1.1
0.3
1.0
0.2
0.8
0.2
0.6
0.1
0.5
0.1
0.3
0.0
0.2
15,000
0.4
1.1
0.3
1.0
0.3
0.8
0.2
0.6
0.2
0.5
0.1
0.3
0.1
0.2
20,000
0.5
0.9
0.5
0.8
0.4
0.7
0.3
0.5
0.2
0.4
0.2
0.3
0.1
0.1
10,000
0.3
0.8
0.3
0.7
0.2
0.6
0.2
0.5
0.1
0.4
0.1
0.2
0.0
0.1
15,000
0.3
0.8
0.3
0.7
0.3
0.6
0.2
0.5
0.1
0.3
0.1
0.2
0.0
0.1
20,000
0.4
0.7
0.4
0.6
0.3
0.5
0.3
0.4
0.2
0.3
0.1
0.2
0.1
0.1
10,000
0.3
0.6
0.2
0.6
0.2
0.5
0.2
0.4
0.1
0.3
0.1
0.2
0.0
0.1
15,000
0.3
0.6
0.3
0.6
0.2
0.5
0.2
0.4
0.1
0.3
0.1
0.2
0.0
0.1
20,000
0.4
0.5
0.3
0.5
0.3
0.4
0.2
0.3
0.2
0.2
0.1
0.2
0.1
0.1
10,000
0.2
0.5
0.2
0.5
0.2
0.4
0.1
0.3
0.1
0.2
0.1
0.2
0.0
0.1
15,000
0.3
0.5
0.2
0.5
0.2
0.4
0.2
0.3
0.1
0.2
0.1
0.2
0.0
0.1
20,000
0.3
0.4
0.3
0.4
0.2
0.3
0.2
0.3
0.1
0.2
0.1
0.1
0.0
0.1
10,000
0.2
0.5
0.2
0.4
0.2
0.3
0.1
0.3
0.1
0.2
0.1
0.1
0.0
0.1
15,000
0.2
0.4
0.2
0.4
0.2
0.3
0.1
0.3
0.1
0.2
0.1
0.1
0.0
0.1
20,000
0.3
0.4
0.3
0.3
0.2
0.3
0.2
0.2
0.1
0.2
0.1
0.1
0.0
0.1
10,000
0.2
0.4
0.2
0.4
0.2
0.3
0.1
0.2
0.1
0.2
0.1
0.1
0.0
0.1
15,000
0.2
0.4
0.2
0.3
0.2
0.3
0.1
0.2
0.1
0.2
0.1
0.1
0.0
0.1
20,000
0.3
0.3
0.2
0.3
0.2
0.2
0.2
0.2
0.1
0.1
0.1
0.1
0.0
0.0
10,000
0.2
0.4
0.2
0.3
0.1
0.3
0.1
0.2
0.1
0.2
0.1
0.1
0.0
0.1
15,000
0.2
0.3
0.2
0.3
0.2
0.2
0.1
0.2
0.1
0.1
0.1
0.1
0.0
0.0
20,000
0.2
0.3
0.2
0.3
0.2
0.2
0.1
0.2
0.1
0.1
0.1
0.1
0.0
0.0
1. Top figures indicate 100 percent oxygen mode.
2. Bottom figures indicate normal oxygen, diluter-demand mode.
3. For conservative time estimates, use next lower gage pressure.
AP 006280.1
2-42
TM 55-1510-215-10
Table 2-2. Oxygen Duration (sheet 2 of 11)
CREW MEMBER DURATION IN HOURS
CABIN
ALTITUDE
FEET
ONE
MAN
CREW
TWO
MAN
CREW
TWO
MAN
CREW
PLUS
ONE
PASS
TWO
MAN
CREW
PLUS
TWO
PASS
TWO
MAN
CREW
PLUS
THREE
PASS
TWO
MAN
CREW
PLUS
FOUR
PASS
TWO
MAN
CREW
PLUS
FIVE
PASS
TWO
MAN
CREW
PLUS
SIX
PASS
NOTES:
GAGE PRESSURE - PSI, U-21G
22 CUBIC FOOT SYSTEM
1500
1200
900
2000
1800
600
300
10,000
1.2
4.3
1.1
3.9
0.9
3.2
0.7
2.6
0.5
1.9
0.4
1.3
0.2
0.6
15,000
1.5
4.3
1.4
3.9
1.1
3.2
0.9
2.6
0.7
1.9
0.5
1.3
0.2
0.6
20,000
2.0
3.6
1.8
3.2
1.5
2.7
1.2
2.1
0.9
1.6
0.6
1.1
0.3
0.5
10,000
0.6
2.2
0.6
1.9
0.5
1.6
0.4
1.3
0.3
1.0
0.2
0.6
0.1
0.3
15,000
0.8
2.2
0.7
1.9
0.6
1.6
0.5
1.3
0.3
1.0
0.2
0.6
0.1
0.3
20,000
1.0
1.8
0.9
1.6
0.8
1.3
0.6
1.1
0.5
0.8
0.3
0.5
0.1
0.3
10,000
0.6
1.6
0.5
1.5
0.4
1.2
0.3
1.0
0.3
0.7
0.2
0.5
0.1
0.2
15,000
0.7
1.6
0.6
1.4
0.5
1.2
0.4
0.9
0.3
0.7
0.2
0.5
0.1
0.2
20,000
0.8
1.3
0.8
1.2
0.6
1.0
0.5
0.8
0.4
0.6
0.3
0.4
0.1
0.2
10,000
0.5
1.3
0.5
1.2
0.4
1.0
0.3
0.8
0.2
0.6
0.2
0.4
0.1
0.2
15,000
0.6
1.2
0.5
1.1
0.4
0.9
0.4
0.7
0.3
0.5
0.2
0.4
0.1
0.2
20,000
0.7
1.1
0.7
1.0
0.5
0.8
0.4
0.6
0.3
0.5
0.2
0.3
0.1
0.2
10,000
0.5
1.1
0.4
1.0
0.4
0.8
0.3
0.6
0.2
0.5
0.1
0.3
0.1
0.2
15,000
0.5
1.0
0.5
0.9
0.4
0.8
0.3
0.6
0.2
0.5
0.2
0.3
0.1
0.1
20,000
0.6
0.9
0.6
0.8
0.5
0.7
0.4
0.5
0.3
0.4
0.2
0.3
0.1
0.1
10,000
0.4
0.9
0.4
0.8
0.3
0.7
0.3
0.6
0.2
0.4
0.1
0.3
0.1
0.1
15,000
0.5
0.9
0.4
0.8
0.4
0.6
0.3
0.5
0.2
0.4
0.1
0.3
0.1
0.1
20,000
0.6
0.8
0.5
0.7
0.4
0.6
0.3
0.5
0.3
0.3
0.2
0.2
0.1
0.1
10,000
0.4
0.8
0.4
0.7
0.3
0.6
0.2
0.5
0.2
0.4
0.1
0.2
0.1
0.1
15,000
0.5
0.7
0.4
0.7
0.3
0.6
0.3
0.4
0.2
0.3
0.1
0.2
0.1
0.1
20,000
0.5
0.7
0.5
0.6
0.4
0.5
0.3
0.4
0.2
0.3
0.2
0.2
0.1
0.1
10,000
0.4
0.7
0.4
0.6
0.3
0.5
0.2
0.4
0.2
0.3
0.1
0.2
0.1
0.1
15,000
0.4
0.7
0.4
0.6
0.3
0.5
0.3
0.4
0.2
0.3
0.1
0.2
0.1
0.1
20,000
0.5
0.6
0.4
0.5
0.4
0.4
0.3
0.3
0.2
0.3
0.1
0.2
0.1
0.1
1. Top figures indicate 100 percent oxygen mode.
2. Bottom figures indicate normal oxygen, diluter-demand mode.
3. For conservative time estimates, use next lower gage pressure.
AP 006280.2
2-43
TM 55-1510-215-10
Table 2-2. Oxygen Duration (sheet 3 of 11)
CREW MEMBER DURATION IN HOURS
CABIN
ALTITUDE
FEET
ONE
MAN
CREW
TWO
MAN
CREW
TWO
MAN
CREW
PLUS
ONE
PASS
TWO
MAN
CREW
PLUS
TWO
PASS
TWO
CREW
PLUS
THREE
PASS
TWO
MAN
CREW
PLUS
FOUR
PASS
TWO
MAN
CREW
PLUS
FIVE
PASS
TWO
MAN
CREW
PLUS
SIX
PASS
NOTES:
GAGE PRESSURE - PSI, U-21G
64 CUBIC FOOT SYSTEM
1500
1200
900
2000
1800
600
300
10,000
3.6
12.5
3.2
11.3
2.7
9.4
2.1
7.5
1.6
5.6
1.1
3.7
0.5
1.8
15,000
4.4
12.5
4.0
11.3
3.3
9.4
2.6
7.5
2.0
5.6
1.3
3.7
0.6
1.8
20,000
5.9
10.4
5.3
9.3
4.4
7.8
3.5
6.2
2.6
4.6
1.7
3.1
0.9
1.5
10,000
1.8
6.3
1.6
5.6
1.3
4.7
1.1
3.7
0.8
2.8
0.5
1.9
0.3
0.9
15,000
2.2
6.3
2.0
5.6
1.7
4.7
1.3
3.7
1.0
2.8
0.7
1.9
0.3
0.9
20,000
2.9
5.2
2.7
4.7
2.2
3.9
1.8
3.1
1.3
2.3
0.9
1.5
0.4
0.8
10,000
1.6
4.7
1.5
4.2
1.2
3.5
1.0
2.8
0.7
2.1
0.5
1.4
0.2
0.7
15,000
2.0
4.6
1.8
4.1
1.5
3.4
1.2
2.7
0.9
2.0
0.6
1.3
0.3
0.7
20,000
2.5
3.9
2.2
3.5
1.8
2.9
1.5
2.3
1.1
1.7
0.7
1.1
0.4
0.6
10,000
1.5
3.8
1.4
3.4
1.1
2.8
0.9
2.2
0.7
1.7
0.4
1.1
0.2
0.5
15,000
1.7
3.6
1.6
3.2
1.3
2.7
1.0
2.1
0.8
1.6
0.5
1.1
0.3
0.5
20,000
2.1
3.1
1.9
2.8
1.6
2.3
1.3
1.8
0.9
1.4
0.6
0.9
0.3
0.4
10,000
1.4
3.1
1.3
2.8
1.0
2.3
0.8
1.9
0.6
1.4
0.4
0.9
0.2
0.5
15,000
1.6
2.9
1.4
2.7
1.2
2.2
0.9
1.8
0.7
1.3
0.5
0.9
0.2
0.4
20,000
1.9
2.6
1.7
2.3
1.4
1.9
1.1
1.5
0.8
1.1
0.6
0.8
0.3
0.4
10,000
1.3
2.7
1.2
2.4
1.0
2.0
0.8
1.6
0.6
1.2
0.4
0.8
0.2
0.4
15,000
1.4
2.5
1.3
2.3
1.1
1.9
0.9
1.5
0.6
1.1
0.4
0.7
0.2
0.4
20,000
1.7
2.2
1.5
2.0
1.2
1.6
1.0
1.3
0.7
1.0
0.5
0.6
0.2
0.3
10,000
1.2
2.3
1.1
2.1
0.9
1.0
0.7
1.4
0.5
1.0
0.4
0.7
0.2
0.3
15,000
1.3
2.2
1.2
2.0
1.0
1.6
0.8
1.3
0.6
1.0
0.4
0.6
0.2
0.3
20,000
1.5
1.9
1.3
1.7
1.1
1.4
0.9
1.1
0.7
0.9
0.4
0.6
0.2
0.3
10,000
1.1
2.1
1.0
1.9
0.9
1.6
0.7
1.2
0.5
0.9
0.3
0.6
0.2
0.3
15,000
1.2
1.9
1.1
1.7
0.9
1.4
0.7
1.2
0.6
0.9
0.4
0.6
0.2
0.3
20,000
1.4
1.7
1.2
1.5
1.0
1.3
0.8
1.0
0.6
0.8
0.4
0.5
0.2
0.2
1. Top figures indicate 100 percent oxygen mode.
2. Bottom figures indicate normal oxygen, diluter-demand mode.
3. For conservative time estimates, use next lower gage pressure.
AP 006280.3
2-44
TM 55-1510-215-10
Table 2-2. Oxygen Duration (sheet 4 of 11)
CREW MEMBER DURATION IN HOURS
CABIN
ALTITUDE
FEET
ONE
MAN
CREW
TWO
MAN
CREW
TWO
MAN
CREW
PLUS
ONE
PASS
TWO
MAN
CREW
PLUS
TWO
PASS
TWO
CREW
PLUS
THREE
PASS
TWO
MAN
CREW
PLUS
FOUR
PASS
TWO
MAN
CREW
PLUS
FIVE
PASS
TWO
MAN
CREW
PLUS
SIX
PASS
NOTES:
GAGE PRESSURE - PSI, U-21G
75 CUBIC FOOT SYSTEM
1500
1200
900
2000
1800
600
300
10,000
4.2
14.7
3.8
13.2
3.1
11.0
2.5
8.8
1.9
6.6
1.2
4.3
0.6
2.1
15,000
5.2
14.7
4.7
13.2
3.9
11.0
3.1
8.8
2.3
6.6
1.5
4.3
0.8
2.1
20,000
6.9
12.1
6.2
10.9
5.2
9.1
4.1
7.3
3.1
5.4
2.0
3.6
1.0
1.8
10,000
2.1
7.3
1.9
6.6
1.6
5.5
1.3
4.4
0.9
3.3
0.6
2.2
0.3
1.1
15,000
2.6
7.3
2.3
6.6
1.9
5.5
1.5
4.4
1.2
3.3
0.8
2.2
0.4
1.1
20,000
3.5
6.1
3.1
5.5
2.6
4.5
2.1
3.6
1.5
2.7
1.0
1.8
0.5
0.9
10,000
1.9
5.5
1.7
4.9
1.4
4.1
1.1
3.3
0.9
2.5
0.6
1.6
0.3
0.8
15,000
2.3
5.3
2.1
4.8
1.7
4.0
1.4
3.2
1.0
2.4
0.7
1.6
0.3
0.8
20,000
2.9
4.5
2.6
4.1
2.2
3.4
1.7
2.7
1.3
2.0
0.9
1.3
0.4
0.7
10,000
1.8
4.4
1.6
4.0
1.3
3.3
1.1
2.6
0.8
2.0
0.5
1.3
0.3
0.6
15,000
2.0
4.2
1.8
3.8
1.5
3.1
1.2
2.5
0.9
1.9
0.6
1.2
0.3
0.6
20,000
2.5
3.6
2.2
3.2
1.9
2.7
1.5
2.2
1.1
1.6
0.7
1.1
0.4
0.5
10,000
1.6
3.7
1.5
3.3
1.2
2.7
1.0
2.2
0.7
1.6
0.5
1.1
0.2
0.5
15,000
1.9
3.5
1.7
3.1
1.4
2.6
1.1
2.1
0.8
1.5
0.5
1.0
0.3
0.5
20,000
2.2
3.0
2.0
2.7
1.6
2.2
1.3
1.8
1.0
1.3
0.6
0.9
0.3
0.4
10,000
1.5
3.1
1.4
2.8
1.1
2.4
0.9
1.9
0.7
1.4
0.4
0.9
0.2
0.5
15,000
1.7
2.9
1.5
2.6
1.3
2.2
1.0
1.8
0.8
1.3
0.5
0.9
0.2
0.4
20,000
1.9
2.6
1.7
2.3
1.5
1.9
1.2
1.5
0.9
1.1
0.6
0.8
0.3
0.4
10,000
1.4
2.8
1.3
2.5
1.1
2.1
0.8
1.6
0.6
1.2
0.4
0.8
0.2
0.4
15,000
1.6
2.6
1.4
2.3
1.2
1.9
0.9
1.5
0.7
1.1
0.5
0.8
0.2
0.4
20,000
1.8
2.2
1.6
2.0
1.3
1.7
1.0
1.3
0.8
1.0
0.5
0.7
0.3
0.3
10,000
1.3
2.4
1.2
2.2
1.0
1.8
0.8
1.5
0.6
1.1
0.4
0.7
0.2
0.4
15,000
1.4
2.3
1.3
2.0
1.1
1.7
0.9
1.3
0.6
1.0
0.4
0.7
0.2
0.3
20,000
1.6
2.0
1.4
1.8
1.2
1.5
1.0
1.2
0.7
0.9
0.5
0.6
0.2
0.3
1. Top figures indicate “100 percent oxygen” mode.
2. Bottom figures indicate” normal oxygen”, diluter-demand mode.
3. For conservative time estimates, use next lower gage pressure.
AP 006280.4
2-45
TM 55-1510-215-10
Table 2-2. Oxygen Duration (sheet 5 of 11)
CREW MEMBER DURATION IN HOURS
CABIN
ALTITUDE
FEET
ONE
MAN
CREW
TWO
MAN
CREW
TWO
MAN
CREW
PLUS
ONE
PASS
TWO
MAN
CREW
PLUS
TWO
PASS
TWO
CREW
PLUS
THREE
PASS
TWO
MAN
CREW
PLUS
FOUR
PASS
TWO
MAN
CREW
PLUS
FIVE
PASS
TWO
MAN
CREW
PLUS
SIX
PASS
NOTES:
GAGE PRESSURE - PSI, U-21G
86 CUBIC FOOT SYSTEM
1500
1200
900
2000
1800
600
300
10,000
4.8
16.8
4.3
15.1
3.6
12.6
2.9
10.1
2.1
7.5
1.4
5.0
0.7
2.5
15,000
5.9
16.8
5.3
15.1
4.4
12.6
3.6
10.1
2.7
7.5
1.8
5.0
0.9
2.5
20,000
7.9
13.9
7.1
12.5
5.9
10.4
4.7
8.3
3.5
6.2
2.3
4.1
1.2
2.0
10,000
2.4
8.4
2.2
7.6
1.8
6.3
1.4
5.0
1.1
3.8
0.7
2.5
0.4
1.2
15,000
3.0
8.4
2.7
7.6
2.2
6.3
1.8
5.0
1.3
3.8
0.9
2.5
0.4
1.2
20,000
4.0
7.0
3.6
6.3
3.0
5.2
2.4
4.2
1.8
3.1
1.2
2.1
0.6
1.0
10,000
2.2
6.3
2.0
5.7
1.6
4.7
1.3
3.8
1.0
2.8
0.7
1.9
0.3
0.9
15,000
2.6
6.1
2.4
5.5
2.0
4.6
1.6
3.7
1.2
2.7
0.8
1.8
0.4
0.9
20,000
3.3
5.2
3.0
4.7
2.5
3.9
2.0
3.1
1.5
2.3
1.0
1.5
0.5
0.8
10,000
2.0
5.0
1.8
4.5
1.5
3.8
1.2
3.0
0.9
2.3
0.6
1.5
0.3
0.7
15,000
2.3
4.8
2.1
4.3
1.8
3.6
1.4
2.9
1.0
2.1
0.7
1.4
0.3
0.7
20,000
2.9
4.1
2.6
3.7
2.1
3.1
1.7
2.5
1.3
1.9
0.8
1.2
0.4
0.6
10,000
1.9
4.2
1.7
3.8
1.4
3.1
1.1
2.5
0.8
1.9
0.6
1.2
0.3
0.6
15,000
2.1
4.0
1.9
3.6
1.6
3.0
1.3
2.4
1.0
1.8
0.6
1.2
0.3
0.6
20,000
2.5
3.4
2.3
3.1
1.9
2.6
1.5
2.1
1.1
1.5
0.7
1.0
0.4
0.5
10,000
1.7
3.6
1.6
3.2
1.3
2.7
1.0
2.2
0.8
1.6
0.5
1.1
0.3
0.5
15,000
1.9
3.4
1.7
3.0
1.5
2.5
1.2
2.0
0.9
1.5
0.6
1.0
0.3
0.5
20,000
2.2
2.9
2.0
2.6
1.7
2.2
1.3
1.8
1.0
1.3
0.7
0.9
0.3
0.4
10,000
1.6
3.2
1.5
2.8
1.2
2.4
1.0
1.9
0.7
1.4
0.5
0.9
0.2
0.5
15,000
1.8
2.9
1.6
2.6
1.3
2.2
1.1
1.7
0.8
1.3
0.5
0.9
0.3
0.4
20,000
2.0
2.6
1.8
2.3
1.5
1.9
1.2
1.5
0.9
1.2
0.6
0.8
0.3
0.4
10,000
1.5
2.8
1.4
2.5
1.1
2.1
0.9
1.7
0.7
1.3
0.5
0.8
0.2
0.4
15,000
1.7
2.6
1.5
2.3
1.2
1.9
1.0
1.5
0.7
1.2
0.5
0.8
0.2
0.4
20,000
1.8
2.3
1.6
2.1
1.4
1.7
1.1
1.4
0.8
1.0
0.5
0.7
0.3
0.3
1. Top figures indicate 100 percent oxygen mode.
2. Bottom figures indicate normal oxygen, diluter-demand mode.
3. For conservative time estimates, use next lower gage pressure.
AP 006280.5
2-46
TM 55-1510-215-10
Table 2-2. Oxygen Duration (sheet 6 of 11)
CREW MEMBER DURATION IN HOURS
CABIN
ALTITUDE
FEET
ONE
MAN
CREW
TWO
MAN
CREW
TWO
MAN
CREW
PLUS
ONE
PASS
TWO
MAN
CREW
PLUS
TWO
PASS
TWO
CREW
PLUS
THREE
PASS
TWO
MAN
CREW
PLUS
FOUR
PASS
TWO
MAN
CREW
PLUS
FIVE
PASS
TWO
MAN
CREW
PLUS
SIX
PASS
NOTES:
GAGE PRESSURE - PSI, U-21G
128 CUBIC FOOT SYSTEM
1500
1200
900
2000
1800
600
300
10,000
7.2
25.0
6.4
22.5
5.4
18.8
4.3
15.0
3.2
11.2
2.1
7.4
1.0
3.6
15,000
8.8
25.0
8.0
22.5
6.6
18.8
5.3
15.0
4.0
11.2
2.6
7.2
1.3
3.6
20,000
11.8
20.7
10.6
18.6
8.8
15.5
7.0
12.4
5.3
9.3
3.5
6.1
1.7
3.0
10,000
3.6
12.5
3.2
11.3
2.7
9.4
2.1
7.5
1.6
5.6
1.1
3.7
0.5
1.8
15,000
4.4
12.5
4.0
11.3
3.3
9.4
2.6
7.5
2.0
5.6
1.3
3.7
0.6
1.8
20,000
5.9
10.4
5.3
9.3
4.4
7.8
3.5
6.2
2.6
4.6
1.7
3.1
0.9
1.5
10,000
3.3
9.4
2.9
8.4
2.4
7.0
2.0
5.6
1.5
4.2
1.0
2.8
0.5
1.4
15,000
3.9
9.1
3.5
8.2
2.9
6.8
2.3
5.4
1.7
4.1
1.2
2.7
0.6
1.3
20,000
4.9
7.7
4.4
6.9
3.7
5.8
3.0
4.6
2.2
3.5
1.5
2.3
0.7
1.1
10,000
3.0
7.5
2.7
6.8
2.3
5.6
1.8
4.5
1.3
3.4
0.9
2.2
0.4
1.1
15,000
3.5
7.2
3.1
6.4
2.6
5.4
2.1
4.3
1.6
3.2
1.0
2.1
0.5
1.0
20,000
4.2
6.2
3.8
5.5
3.2
4.6
2.5
3.7
1.9
2.8
1.3
1.8
0.6
0.9
10,000
2.8
6.3
2.5
5.6
2.1
4.7
1.7
3.7
1.2
2.8
0.8
1.9
0.4
0.9
15,000
3.2
5.9
2.8
5.3
2.4
4.4
1.9
3.5
1.4
2.6
0.9
1.7
0.5
0.9
20,000
3.7
5.1
3.3
4.6
2.8
3.8
2.2
3.1
1.7
2.3
1.1
1.5
0.5
0.7
10,000
2.6
5.4
2.3
4.8
1.9
4.0
1.5
3.2
1.2
2.4
0.8
1.6
0.4
0.8
15,000
2.9
5.0
2.6
4.5
2.2
3.8
1.7
3.0
1.3
2.2
0.9
1.5
0.4
0.7
20,000
3.3
4.4
3.0
3.9
2.5
3.3
2.0
2.6
1.5
2.0
1.0
1.3
0.5
0.6
10,000
2.4
4.7
2.2
4.2
1.8
3.5
1.4
2.8
1.1
2.1
0.7
1.4
0.4
0.7
15,000
2.7
4.4
2.4
3.9
2.0
3.3
1.6
2.6
1.2
1.9
0.8
1.3
0.4
0.6
20,000
3.0
3.8
2.7
3.4
2.2
2.9
1.8
2.3
1.3
1.7
0.9
1.1
0.4
0.6
10,000
2.3
4.2
2.0
3.8
1.7
3.1
1.4
2.5
1.0
1.9
0.7
1.2
0.3
0.6
15,000
2.5
3.9
2.2
3.5
1.8
2.9
1.5
2.3
1.1
1.7
0.7
1.1
0.4
0.6
20,000
2.7
3.4
2.4
3.1
2.0
2.5
1.6
2.0
1.2
1.5
0.8
1.0
0.4
0.5
1. Top figures indicate 100 percent oxygen mode.
2. Bottom figures indicate normal oxygen, diluter-demand mode.
3. For conservative time estimates, use next lower gage pressure.
AP 006280.6
2-47
TM 55-1510-215-10
Table 2-2. Oxygen Duration (sheet 7 of 11)
CREW MEMBER DURATION IN HOURS
CABIN
ALTITUDE
FEET
ONE
MAN
CREW
TWO
MAN
CREW
TWO
MAN
CREW
PLUS
ONE
PASS
TWO
MAN
CREW
PLUS
TWO
PASS
TWO
CREW
PLUS
THREE
PASS
TWO
MAN
CREW
PLUS
FOUR
PASS
TWO
MAN
CREW
PLUS
FIVE
PASS
TWO
MAN
CREW
PLUS
SIX
PASS
NOTES:
GAGE PRESSURE - PSI, U-21G
139 CUBIC FOOT SYSTEM
1500
1200
900
2000
1800
600
300
10,000
7.8
27.2
7.0
24.5
5.8
20.4
4.6
16.3
3.5
12.2
2.3
8.1
1.1
4.0
15,000
9.6
27.2
8.6
24.5
7.2
20.4
5.7
16.3
4.3
12.2
2.8
8.1
1.4
4.0
20,000
12.8
22.5
11.5
20.2
9.6
16.9
7.7
13.5
5.7
10.1
3.8
6.7
1.9
3.3
10,000
3.9
13.6
3.5
12.2
2.9
10.2
2.3
8.1
1.7
6.1
1.2
4.0
0.6
2.0
15,000
4.8
13.6
4.3
12.2
3.6
10.2
2.9
8.1
2.1
6.1
1.4
4.0
0.7
2.0
20,000
6.4
11.3
5.8
10.1
4.8
8.4
3.8
6.7
2.9
5.0
1.9
3.3
0.9
1.6
10,000
3.5
10.2
3.2
9.2
2.7
7.6
2.1
6.1
1.6
4.6
1.1
3.0
0.5
1.5
15,000
4.2
9.9
3.8
8.9
3.2
7.4
2.5
5.9
1.9
4.4
1.3
2.9
0.6
1.4
20,000
5.4
8.4
4.8
7.5
4.0
6.3
3.2
5.0
2.4
3.8
1.6
2.5
0.8
1.2
10,000
3.3
8.2
2.9
7.3
2.4
6.1
2.0
4.9
1.5
3.6
1.0
2.4
0.5
1.2
15,000
3.8
7.8
3.4
7.0
2.8
5.8
2.3
4.6
1.7
3.5
1.1
2.3
0.6
1.1
20,000
4.6
6.7
4.1
6.0
3.5
5.0
2.8
4.0
2.1
3.0
1.4
2.0
0.7
1.0
10,000
3.0
6.8
2.7
6.1
2.3
5.1
1.8
4.1
1.4
3.0
0.9
2.0
0.4
1.0
15,000
3.4
6.4
3.1
5.8
2.6
4.8
2.1
3.8
1.5
2.9
1.0
0.9
0.5
0.9
20,000
4.0
5.6
3.6
5.0
3.0
4.2
2.4
3.3
1.8
2.5
1.2
1.6
0.6
0.8
10,000
2.8
5.8
2.5
5.2
2.1
4.4
1.7
3.5
1.3
2.6
0.8
1.7
0.4
0.8
15,000
3.1
5.4
2.8
4.9
2.3
4.1
1.9
3.3
1.4
2.4
0.9
1.6
0.5
0.8
20,000
3.6
4.8
3.2
4.3
2.7
3.6
2.2
2.8
1.6
2.1
1.1
1.4
0.5
0.7
10,000
2.6
5.1
2.4
4.6
2.0
3.8
1.6
3.0
1.2
2.3
0.8
1.5
0.4
0.7
15,000
2.9
4.7
2.6
4.3
2.2
3.5
1.7
2.8
1.3
2.1
0.9
1.4
0.4
0.7
20,000
3.2
4.2
2.9
3.7
2.4
3.1
1.9
2.5
1.5
1.9
1.0
1.2
0.5
0.6
10,000
2.5
4.5
2.2
4.1
1.9
3.4
1.5
2.7
1.1
2.0
0.7
1.3
0.4
0.7
15,000
2.7
4.2
2.4
3.8
2.0
3.1
1.6
2.5
1.2
1.9
0.8
1.2
0.4
0.6
20,000
3.0
3.7
2.7
3.3
2.2
2.8
1.8
2.2
1.3
1.7
0.9
1.1
0.4
0.5
1. Top figures indicate 100 percent oxygen mode.
2. Bottom figures indicate normal oxygen, diluter-demand mode.
3. For conservative time estimates, use next lower gage pressure.
AP 006280.7
2-48
TM 55-1510-215-10
Table 2-2. Oxygen Duration (sheet 8 of 11)
CREW MEMBER DURATION IN HOURS
CABIN
ALTITUDE
FEET
ONE
MAN
CREW
TWO
MAN
CREW
TWO
MAN
CREW
PLUS
ONE
PASS
TWO
MAN
CREW
PLUS
TWO
PASS
TWO
CREW
PLUS
THREE
PASS
TWO
MAN
CREW
PLUS
FOUR
PASS
TWO
MAN
CREW
PLUS
FIVE
PASS
TWO
MAN
CREW
PLUS
SIX
PASS
NOTES:
GAGE PRESSURE - PSI, U-21G
150 CUBIC FOOT SYSTEM
1500
1200
900
2000
1800
600
300
10,000
8.4
29.3
7.5
26.4
6.3
22.0
5.0
17.5
3.7
13.1
2.5
8.7
1.2
4.3
15,000
10.4
29.3
9.3
26.4
7.8
22.0
6.2
17.5
4.6
13.1
3.1
8.7
1.5
4.3
20,000
13.8
24.3
12.4
21.8
10.3
18.2
8.3
14.5
6.2
10.9
4.1
7.2
2.0
3.5
10,000
4.2
14.7
3.8
13.2
3.1
11.0
2.5
8.8
1.9
6.6
1.2
4.3
0.6
2.1
15,000
5.2
14.7
4.7
13.2
3.9
11.0
3.1
8.8
2.3
6.6
1.5
4.3
0.8
2.1
20,000
6.9
12.1
6.2
10.9
5.2
9.1
4.1
7.3
3.1
5.4
2.0
3.6
1.0
1.8
10,000
3.8
11.0
3.4
9.9
2.9
8.2
2.3
6.6
1.7
4.9
1.1
3.3
0.6
1.6
15,000
4.6
10.7
4.1
9.6
3.4
8.0
2.7
6.4
2.0
4.8
1.4
3.2
0.7
1.6
20,000
5.8
9.1
5.2
8.1
4.3
6.8
3.5
5.4
2.6
4.0
1.7
2.7
0.8
1.3
10,000
3.5
8.8
3.2
7.9
2.6
6.6
2.1
5.3
1.6
3.9
1.0
2.6
0.5
1.3
15,000
4.1
8.4
3.7
7.5
3.1
6.3
2.4
5.0
1.8
3.7
1.2
2.5
0.6
1.2
20,000
5.0
7.2
4.5
6.5
3.7
5.4
3.0
4.3
2.2
3.2
1.5
2.1
0.7
1.1
10,000
3.3
7.3
2.9
6.6
2.4
5.5
1.9
4.4
1.5
3.3
1.0
2.2
0.5
1.1
15,000
3.7
6.9
3.3
6.2
2.8
5.2
2.2
4.1
1.7
3.1
1.1
2.0
0.5
1.0
20,000
4.4
6.0
3.9
5.4
3.3
4.5
2.6
3.6
2.0
2.7
1.3
1.8
0.6
0.9
10,000
3.0
6.3
2.7
5.7
2.3
4.7
1.8
3.8
1.4
2.8
0.9
1.9
0.4
0.9
15,000
3.4
5.9
3.0
5.3
2.5
4.4
2.0
3.5
1.5
2.6
1.0
1.7
0.5
0.9
20,000
3.9
5.1
3.5
4.6
2.9
3.8
2.3
3.1
1.7
2.3
1.2
1.5
0.6
0.7
10,000
2.8
5.5
2.6
4.9
2.1
4.1
1.7
3.3
1.3
2.5
0.8
1.6
0.4
0.8
15,000
3.1
5.1
2.8
4.6
2.3
3.8
1.9
3.1
1.4
2.3
0.9
1.5
0.5
0.7
20,000
3.5
4.5
3.2
4.0
2.6
3.4
2.1
2.7
1.6
2.0
1.0
1.3
0.5
0.7
10,000
2.7
4.9
2.4
4.4
2.0
3.7
1.6
2.9
1.2
2.2
0.8
1.4
0.4
0.7
15,000
2.9
4.5
2.6
4.1
2.2
3.4
1.7
2.7
1.3
2.0
0.9
1.3
0.4
0.7
20,000
3.2
4.0
2.9
3.6
2.4
3.0
1.9
2.4
1.4
1.8
0.9
1.2
0.5
0.6
1. Top figures indicate 100 percent oxygen mode.
2. Bottom figures indicate normal oxygen, diluter-demand mode.
3. For conservative time estimates, use next lower gage pressure.
AP 006280.8
2-49
TM 55-1510-215-10
Table 2-2. Oxygen Duration (sheet 9 of 11)
CREW MEMBER DURATION IN HOURS
CABIN
ALTITUDE
FEET
ONE
MAN
CREW
TWO
MAN
CREW
TWO
MAN
CREW
PLUS
ONE
PASS
TWO
MAN
CREW
PLUS
TWO
PASS
TWO
CREW
PLUS
THREE
PASS
TWO
MAN
CREW
PLUS
FOUR
PASS
TWO
MAN
CREW
PLUS
FIVE
PASS
TWO
MAN
CREW
PLUS
SIX
PASS
NOTES:
GAGE PRESSURE - PSI, U-21G
192 CUBIC FOOT SYSTEM
1500
1200
900
2000
1800
600
300
10,000
10.7
37.6
9.7
33.8
8.0
28.1
6.4
22.5
4.8
16.8
3.2
11.1
1.6
5.5
15,000
13.3
37.6
11.9
33.8
9.9
28.1
7.9
22.5
5.9
16.8
3.9
11.1
1.9
5.5
20,000
17.7
31.1
15.9
28.0
13.2
23.3
10.6
18.6
7.9
13.9
5.2
9.2
2.6
4.5
10,000
5.4
18.8
4.8
16.9
4.0
14.1
3.2
11.2
2.4
8.4
1.6
5.6
0.8
2.7
15,000
6.6
18.8
6.0
16.9
5.0
14.1
4.0
11.2
3.0
8.4
2.0
5.6
1.0
2.7
20,000
8.8
15.5
8.0
14.0
6.6
11.6
5.3
9.3
4.0
7.0
2.6
4.6
1.3
2.3
10,000
4.9
14.1
4.4
12.7
3.7
10.5
2.9
8.4
2.2
6.3
1.5
4.2
0.7
2.1
15,000
5.9
13.7
5.3
12.3
4.4
10.2
3.5
8.2
2.6
6.1
1.7
4.0
0.9
2.0
20,000
7.4
11.6
6.7
10.4
5.5
8.7
4.4
6.9
3.3
5.2
2.2
3.4
1.1
1.7
10,000
4.5
11.3
4.1
10.1
3.4
8.4
2.7
6.7
2.0
5.0
1.3
3.3
0.7
1.6
15,000
5.2
10.7
4.7
9.7
3.9
8.0
3.1
6.4
2.3
4.8
1.6
3.2
0.8
1.6
20,000
6.4
9.2
5.7
8.3
4.8
6.9
3.8
5.5
2.8
4.1
1.9
2.7
0.9
1.3
10,000
4.2
9.4
3.8
8.4
3.1
7.0
2.5
5.6
1.9
4.2
1.2
2.8
0.6
1.4
15,000
4.7
8.8
4.3
8.0
3.6
6.6
2.8
5.3
2.1
4.0
1.4
2.6
0.7
1.3
20,000
5.6
7.7
5.0
6.9
4.2
5.7
3.3
4.6
2.5
3.4
1.7
2.3
0.8
1.1
10,000
3.9
8.0
3.5
7.2
2.9
6.0
2.3
4.8
1.7
3.6
1.2
2.4
0.6
1.2
15,000
4.3
7.5
3.9
6.8
3.2
5.6
2.6
4.5
1.9
3.4
1.3
2.2
0.6
1.1
20,000
5.0
6.6
4.5
5.9
3.7
4.9
3.0
3.9
2.2
2.9
1.5
1.9
0.7
1.0
10,000
3.6
7.0
3.3
6.3
2.7
5.3
2.2
4.2
1.6
3.1
1.1
2.1
0.5
1.0
15,000
4.0
6.5
3.6
5.9
3.0
4.9
2.4
3.9
1.8
2.9
1.2
1.9
0.6
1.0
20,000
4.5
5.7
4.0
5.2
3.4
4.3
2.7
3.4
2.0
2.6
1.3
1.7
0.7
0.8
10,000
3.4
6.3
3.1
5.6
2.6
4.7
2.0
3.7
1.5
2.8
1.0
1.9
0.5
0.9
15,000
3.7
5.8
3.3
5.2
2.8
4.3
2.2
3.5
1.7
2.6
1.1
1.7
0.5
0.8
20,000
4.1
5.1
3.7
4.6
3.1
3.8
2.4
3.0
1.8
2.3
1.2
1.5
0.6
0.7
1. Top figures indicate 100 percent oxyg en mode.
2. Bottom figures indicate normal oxygen, diluter-demand mode.
3. For conservative time estimates, use next lower gage pressure.
AP 006280.9
2-50
TM 55-1510-215-10
Table 2-2. Oxygen Duration (sheet 10 of 11)
CREW MEMBER DURATION IN HOURS
CABIN
ALTITUDE
FEET
ONE
MAN
CREW
TWO
MAN
CREW
TWO
MAN
CREW
PLUS
ONE
PASS
TWO
MAN
CREW
PLUS
TWO
PASS
TWO
CREW
PLUS
THREE
PASS
TWO
MAN
CREW
PLUS
FOUR
PASS
TWO
MAN
CREW
PLUS
FIVE
PASS
TWO
MAN
CREW
PLUS
SIX
PASS
NOTES:
GAGE PRESSURE - PSI, U-21G
203 CUBIC FOOT SYSTEM
1500
1200
900
2000
1800
600
300
10,000
11.3
39.7
10.2
35.7
8.5
29.7
6.8
23.7
5.1
17.8
3.4
11.8
1.7
5.8
15,000
14.0
39.7
12.6
35.7
10.5
29.7
8.4
23.7
6.3
17.8
4.2
11.8
2.0
5.8
20,000
18.7
32.9
16.8
29.6
14.0
24.6
11.2
19.7
8.4
14.7
5.5
9.7
2.7
4.8
10,000
5.7
19.9
5.1
17.9
4.2
14.9
3.4
11.9
2.5
8.9
1.7
5.9
0.8
2.9
15,000
7.0
19.9
6.3
17.9
5.2
14.9
4.2
11.9
3.1
8.9
2.1
5.9
1.0
2.9
20,000
9.3
16.4
8.4
14.8
7.0
12.3
5.6
9.8
4.2
7.3
2.8
4.9
1.4
2.4
10,000
5.2
14.9
4.7
13.4
3.9
11.2
3.1
8.9
2.3
6.7
1.5
4.4
0.8
2.2
15,000
6.2
14.4
5.6
13.0
4.6
10.8
3.7
8.6
2.8
6.5
1.8
4.3
0.9
2.1
20,000
7.8
12.3
7.0
11.0
5.9
9.2
4.7
7.3
2.3
5.5
2.3
3.6
1.1
1.8
10,000
4.8
11.9
4.3
10.7
3.6
8.9
2.8
7.1
2.1
5.3
1.4
3.5
0.7
1.7
15,000
5.5
11.3
5.0
10.2
4.1
8.5
3.3
6.8
2.5
5.1
1.6
3.4
0.8
1.7
20,000
6.7
9.8
6.1
8.8
5.0
7.3
4.0
5.8
3.0
4.4
2.0
2.9
1.0
1.4
10,000
4.4
9.9
4.0
8.9
3.3
7.4
2.6
5.9
2.0
4.4
1.3
2.9
0.6
1.4
15,000
5.0
9.3
4.5
8.4
3.8
7.0
3.0
5.6
2.2
4.2
1.5
2.8
0.7
1.4
20,000
5.9
8.1
5.3
7.3
4.4
6.1
3.5
4.9
2.6
3.6
1.8
2.4
0.9
1.2
10,000
4.1
8.5
3.7
7.7
3.1
6.4
2.5
5.1
1.8
3.8
1.2
2.5
0.6
1.2
15,000
4.6
7.9
4.1
7.1
3.4
5.9
2.7
4.7
2.0
3.6
1.4
2.4
0.7
1.2
20,000
5.3
6.9
4.7
6.2
3.9
5.2
3.1
4.2
2.4
3.1
1.6
2.1
0.8
1.0
10,000
3.8
7.4
3.5
6.7
2.9
5.6
2.3
4.5
1.7
3.3
1.1
2.2
0.6
1.1
15,000
4.2
6.9
3.8
6.2
3.2
5.2
2.5
4.1
1.9
3.1
1.3
2.0
0.6
1.0
20,000
4.7
6.1
4.3
5.5
3.6
4.5
2.8
3.6
2.1
2.7
1.4
1.8
0.7
0.9
10,000
3.6
6.6
3.2
6.0
2.7
5.0
2.2
4.0
1.6
3.0
1.1
2.0
0.5
1.0
15,000
3.9
6.1
3.5
5.5
2.9
4.6
2.3
3.7
1.7
2.7
1.2
1.8
0.6
0.9
20,000
4.3
5.4
3.9
4.8
3.2
4.0
2.6
3.2
1.9
2.4
1.3
1.6
0.6
0.8
1. Top figures indicate 100 percent oxygen mode.
2. Bottom figures indicate normal oxygen, diluter-demand mode.
3. For conservative time estimates, use next lower gage pressure.
AP 006280.10
2-51
TM 55-1510-215-10
Table 2-2. Oxygen Duration (sheet 11 of 11)
CREW MEMBER DURATION IN HOURS
CABIN
ALTITUDE
FEET
ONE
MAN
CREW
TWO
MAN
CREW
TWO
MAN
CREW
PLUS
ONE
PASS
TWO
MAN
CREW
PLUS
TWO
PASS
TWO
CREW
PLUS
THREE
PASS
TWO
MAN
CREW
PLUS
FOUR
PASS
TWO
MAN
CREW
PLUS
FIVE
PASS
TWO
MAN
CREW
PLUS
SIX
PASS
NOTES:
GAGE PRESSURE - PSI, U-21G
214 CUBIC FOOT SYSTEM
1500
1200
900
2000
1800
600
300
10,000
12.0
41.9
10.8
37.7
9.0
31.3
7.2
25.0
5.3
18.7
3.5
12.4
1.7
6.1
15,000
14.8
41.9
13.3
37.7
11.1
31.3
8.8
25.0
6.6
18.7
4.4
12.4
2.2
6.1
20,000
19.7
34.6
17.7
31.2
14.8
25.9
11.8
20.7
8.8
15.5
5.8
10.3
2.9
5.0
10,000
6.0
20.9
5.4
18.8
4.5
15.7
3.6
12.5
2.7
9.4
1.8
6.2
0.9
3.0
15,000
7.4
20.9
6.6
18.8
5.5
15.7
4.4
12.5
3.3
9.4
2.2
6.2
1.1
3.0
20,000
9.9
17.3
8.9
15.6
7.4
13.0
5.9
10.4
4.4
7.7
2.9
5.1
1.4
2.5
10,000
5.5
15.7
4.9
14.1
4.1
11.8
3.3
9.4
2.4
7.0
1.6
4.7
0.8
2.3
15,000
6.5
15.2
5.9
13.7
4.9
11.4
3.9
9.1
2.9
6.8
1.9
4.5
1.0
2.2
20,000
8.2
12.9
7.4
11.6
6.2
9.7
4.9
7.7
3.7
5.8
2.4
3.8
1.2
1.9
10,000
5.0
12.6
4.5
11.3
3.8
9.4
3.0
7.5
2.2
5.6
1.5
3.7
0.7
1.8
15,000
5.8
12.0
5.3
10.8
4.4
9.0
3.5
7.2
2.6
5.3
1.7
3.5
0.9
1.7
20,000
7.1
10.3
6.4
9.3
5.3
7.7
4.2
6.2
3.2
4.6
2.1
3.1
1.0
1.5
10,000
4.7
10.5
4.2
9.4
3.5
7.8
2.8
6.3
2.1
4.7
1.4
3.1
0.7
1.5
15,000
5.3
9.9
4.8
8.9
4.0
7.4
3.2
5.9
2.4
4.4
1.6
2.9
0.8
1.4
20,000
6.2
8.6
5.6
7.7
4.7
6.4
3.7
5.1
2.8
3.8
1.8
2.5
0.9
1.2
10,000
4.3
9.0
3.9
8.1
3.2
6.7
2.6
5.4
1.9
4.0
1.3
2.7
0.6
1.3
15,000
4.8
8.4
4.3
7.5
3.6
6.3
2.9
5.0
2.2
3.7
1.4
2.5
0.7
1.2
20,000
5.5
7.3
5.0
6.6
4.2
5.5
3.3
4.4
2.5
3.3
1.6
2.2
0.8
1.1
10,000
4.1
7.9
3.6
7.1
3.0
5.9
2.4
4.7
1.8
3.5
1.2
2.3
0.6
1.1
15,000
4.4
7.3
4.0
6.5
3.3
5.5
2.7
4.4
2.0
3.3
1.3
2.2
0.6
1.1
20,000
5.0
6.4
4.5
5.8
3.7
4.8
3.0
3.8
2.2
2.9
1.5
1.9
0.7
0.9
10,000
3.8
7.0
3.4
6.3
2.8
5.2
2.3
4.2
1.7
3.1
1.1
2.1
0.6
1.0
15,000
4.1
6.4
3.7
5.8
3.1
4.8
2.5
3.9
1.8
2.9
1.2
1.9
0.6
0.9
20,000
4.6
5.7
4.1
5.1
3.4
4.3
2.7
3.4
2.0
2.5
1.3
1.7
0.7
0.8
1. Top figures indicate 100 percent oxygen mode.
2. Bottom figures indicate normal oxygen, diluter-demand mode.
3. For conservative time estimates, use next lower gage pressure.
AP 006280.11
2-52
TM 55-1510-215-10
(a) The emergency pressure control
lever has three positions, 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 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.
(b) The 2000 PSI oxygen pressure
gage provided on the oxygen control panels should at no
time indicate over 400 PSI. If the pressure exceeds 400
PSI, a malfunction of the pressure reducer is indicated.
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.
NOTE
Check to insure that the OXYGEN
SUPPLY PRESSURE gage registers
adequate pressure before each flight.
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 an altitude which does
not require the use of oxygen.
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 diluter control lever to 100%
OXYGEN and continue breathing
undiluted oxygen until the danger is
past.
(2) 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 CREW OXYGEN shut-off
valve is turned on, then the individual places the supply
control lever on the regulator control panel to the ON
position, dons and adjusts his mask, selects TEST
MASK position, and checks for leaks.
NOTE
A loss of oxygen will occur if the
CABIN OXYGEN valve is OPEN, and a
passenger
oxygen
mask
is
connected to any of the outlets.
b. Normal Operation. For normal operation of the
oxygen system, the pilot rotates the cockpit or cabin
oxygen shut-off valve (fig. 2-17) to OPEN, then each
individual places the supply lever (green) on his
regulator control panel to the ON position, the diluter
lever (white) to the NORMAL OXYGEN position, and the
emergency pressure lever (red) to the NORMAL
position. To obtain oxygen for the passengers, the pilot
rotates the CABIN OXYGEN valve to OPEN causing
100% oxygen to flow continuously into the distribution
lines which supply passenger masks coupled into the
service outlets.
NOTE
The shutoff valves of high pressure
oxygen systems should be opened
fully to prevent the possibility of
leakage.
2-53
TM 55-1510-215-10
c. Emergency Operation.
For emergency
operation, the affected crew member selects the
EMERGENCY position of the emergency pressure
control lever on his oxygen regulator control panel. This
selection provides 100% oxygen, at a positive pressure,
regardless of the position of the diluter control lever on
his panel. The cabin oxygen system is not equipped
with oxygen regulator control panels.
Oxygen flow
control to the passenger oxygen outlets is maintained by
the pilot who actuates the cabin oxygen system slow
release valve.
2-60. Windshield Wipers.
a. Description.
Two
electrically
operated
windshield wipers are provided for use at takeoff and
landing speed (fig. 2-5). Operation at cruise speed may
result in damage to the wiper operating mechanism. A
rotary switch (fig. 2-21), placarded WINDSHIELD
WIPER, located on the overhead control panel, selects
mode of windshield wiper operation. An information
placard below 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 spring-loaded 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 settings for wiper operation.
The windshield wiper circuit is protected by one 10ampere circuit breaker, placarded WSHLD WIPER,
located on the right subpanel (fig. 2-7).
CAUTION
Do not operate windshield wipers on
dry glass. Such action can damage
the linkage as well as scratch the
windshield glass.
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 potation
will stop the windshield wipers, without returning them to
the normal inactive position.
2-54
2-61. Stall Warning System.
Approach to a stall is indicated by a steady tone of a
warning horn located behind the right subpanel. A small
metal vane located on the left wing leading edge, is
moved by any change in airflow over the leading edge of
the wing.
When the airspeed decreases to
approximately 5 to 10 knots above stall speed,
movement of the vane actuates switch which completes
a DC electrical circuit to the start warning horn. Since
the vane is affected by the same aerodynamic forces
that result in the stall, the system functions regardless of
the type of stall or configuration of landing gear and wing
flaps, the only variation in performance being the margin
of airspeed at which the warning occurs. To prevent ice
formations on the stall warning vane an electrically
operated heating element is installed, which is controlled
by a 5-ampere STALL WARN circuit breaker switch
located on the left subpanel. The stall warning circuit is
protected by a 5-ampere circuit breaker, placarded
STALL WARN HORN, on the right subpanel (fig. 2-7)
and a 5-ampere circuit breaker placarded STALL
WARNING CKT BKR on a bracket below the battery box
in the right center wing section (fig. 2-17).
2-62. Seating Provisions.
a. Troop Transport. The troop transport versions
will accommodate ten combat equipped troops in the
cabin with five non-adjustable, bench-type seat units.
One rode of each seat unit attaches to wall structure of
the fuselage, and the other side is supported by foldable
legs which insert into slotted tracks mounted on the floor.
All seat units are removable, however, they may also be
folded up and strapped close to the compartment walls
using straps attached for that purpose.
Six seat
positions are located along the right wall and four along
the left wall.
All face the center aisle.
Combat
equipment and personnel packs may be stowed beneath
the seats. Seat belts are installed at each seat position.
b. Air Ambulance.
When configured for air
ambulance use the cabin has three seating positions to
accommodate medical attendants and ambulatory
patients, and provisions to install three litters. The three
forward seat positions along the right side are utilized.
Litter suspension mountings are provided for two litters
on the aft-right side, and one litter on the left-forward
side.
c. Staff Transport.
When configured as staff
transport the cabin has two forward-facing chair
TM 55-1510-215-10
seats and two aft-facing chair seats attached to floor
railings on each side of the center aisle.
d. Cargo Transport. Troop seat unit can be folded
up to prevent conflict with the cargo mission, or may be
remove, if required.
2-63. Cigarette Lighters and Ash Trays.
One electrical push-in type cigarette lighter is
located on the pilot's control pedestal (fig. 2-6). A pushpull type ash tray is built into the inboard arm rest of the
pilots and copilot's seats and is also included on the
back side of each staff personnel chair. One push-pull
type ash tray is installed on the fuselage wall adjacent to
each of the four passenger seats in the cabin. The
circuit is protected by a 5-ampere circuit breaker,
placarded CIGARETTE LIGHTER, on the right subpanel
(fig. 2-7).
2-64. Relief Tubes.
Two relief tubes are provided, one under the pilot's
seat and the other located aft of the main entrance door
in the aft cabin.
2-65. Rear View Mirror.
A single rear-view type mirror is externally mounted
below the pilot's side window to enable surveillance
along the left fuselage and aft tail surfaces (fig. 2-5).
Sighting angle of the mirror is adjustable from within the
cockpit using a rotary control knob located aft of the fuel
management panel (fig. 2-5). When not in use, the
mirror may be adjusted to streamline against the
fuselage exterior for minimum drag.
2-66. Sun Visors.
CAUTION
When adjusting the sun visors, grasp
only by the top metal attachment to
avoid damage to the plastic shield.
Two sun visors are provided for the pilot and copilot
respectively (fig.
2-5).
Each visor is manually
adjustable. 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.
2-66A. Battery Vent Anti-Icing.
a. Description.
The battery box ram air vent,
located on the top of the right wing center section,
utilizes anti-icing protection provided by an externally
attached spirally-wound electrical element. Power is
supplied to the element by the 7.5 amp circuit breaker
switch placarded HEAT--FUEL VENTS-RIGHT located
on the pilot's subpanel (fig. 2-7).
CAUTION
To prevent overheat damage to the
electrically heated anti-ice jackets,
the HEAT-FUEL VENTS-RIGHT switch
should not be turned ON unless
cooling air will soon pass over the
jackets.
b. Normal Operation. For normal operation the
HEAT - FUEL VENTS-RIGHT switch is turned OFF
during BEFORE STARTING ENGINES procedure, and
turned ON as required the BEFORE TAKEOFF
procedure. Refer to Chapter 8.
Section VIII. HEATING, VENTILATION, COOLING, AND ENVIRONMENTAL, CONTROL UNIT
2-67. Heating and Ventilation Systems.
a. Description. The heating and ventilation system
operates utilizing fresh air from outside of the aircraft.
On the ground, the outside air enters the system cold air
plenum through the ventilation louvers in the left door of
the nose avionics compartment. Two ventilation air
blowers operate only while the aircraft is on the ground,
forcing air from the nose avionics compartment into the
cold air plenum. When the aircraft becomes airborne, a
switch on the left main landing gear strut turns the
blowers off. Part of the ram air bypasses the hearer and
is ducted to the nose avionics compartment, and to the
cold air outlets in both the cockpit and cabin. The
remainder of the air is ducted into the heater. After the
air is heated, it is ducted to three warm air outlets in the
cabin, and to the two warm air and defroster outlets in
the cockpit. The cockpit's warm air outlets are located at
floor level. Stale air is vented through the exhaust air
plenum installed in the cabin ceiling.
(1) Vent blower and switch.
A switch,
placarded VENT BLWR on the right subpanel, controls
activation of the blowers for the ventilation system (fig.
2-7). VENT BLWR position, activates the ventilation
blowers for cooling air when the aircraft is on the ground.
When the aircraft becomes airborne, a switch on the left
main gear turns the ventilation blower circuits off.
Circuits are protected by a 35-ampere circuit breaker
placarded VENT BLOWER located on the copilot's
circuit breaker panel (fig. 2-18).
Change 10 2-54A/(2-54B blank)
TM 55-1510-215-10
(2) Fresh air outlets. Fresh air is ducted
directly from the fresh air plenum to individual outlets in
the cockpit and cabin positioned in the ceiling above the
seats. Airflow volume is controlled at each outlet. The
direction of airflow is controlled by moving the outlet in it
spherical socket.
(3) Stale air exhaust system. Stale air is
exhausted from the cabin area though the exhaust
plenum in the cabin ceiling and is ducted to two vents
located in the cabin on each side of the fuselage.
(4) Push-pull cockpit ventilation control. Two
push-pull type air inlet controls, placarded VENT AIR PUSH ON, are located below the pilot and copilot
subpanels to manually regulate cockpit ventilation.
When pushed IN, either control will cause outside airflow
from outlets above the respective rudder pedals As a
control is pulled out, there is a corresponding decrease
in the amount of air flow.
(5) Push-pull cabin ventilation control.
A
push-pull type control, placarded CABIN AIR, located on
the fight subpanel (fig. 2-7, manually controls ventilation
of the cabin. Airflow is at maximum when the control is
pushed in. As the control is pulled out there is a
corresponding decrease in the amount of airflow.
(6) Heater combustion unit. The heater is a
combustion unit that uses the same fuels as the engines.
The heater is located in the lower right m-d of the nose
avionics compartment. Air for combustion is ducted from
the nose avionics compartment to a combustion blower
which forces air into the combustion chamber. Taking
air from the nose avionics compartment helps to cool the
compartment by absorbing heat from installed electronic
equipment. The combustion blower operates on the
ground whenever electrical power is applied.
(7) Heater fuel pump. A heater fuel pump in
the left wing outboard of the fuselage, forward of the
main wing spar, operates whenever the heater control
switch is ON. Fuel for the heater is obtained from the
left nacelle tank (fig. 2-11). The cabin heater will
continue to operate until all fuel is consumed from the
left nacelle tank.
(8) Pressure differential switch and duct
temperature thermal switch. When either combustion
blower air flow or vent blower airflow is insufficient, a
differential pressure switch prevents heater operation.
Also, to prevent heat damage to equipment located
adjacent to a heater outlet, a duct temperature thermal
switch allows the ventilation blower to continue operation
on the ground
after heater shutdown to purge heat from the system.
This switch will shut off the ventilation blower when the
duct temperature is reduced to 52°C.
(9) Heater cycling switch. A cycling switch
located in the hot air plenum disconnects power to the
heater fuel solenoid valve at 107°C. If this automatic
function should fail, temperature within the plenum will
ultimately exceed 149°C and cause a 7.5-ampere fuse to
blow, shutting down the heater. The 7.5-ampere fuse,
which is in series with the HTR control switch and the
TEMP CONTROL circuit breaker, cannot be replaced in
flight due to its location.
(10) Heater control switch. A switch placarded
HEATER OFF AUTO, located on the right subpanel,
controls cabin heater operation. Either mode, AUTO or
MAN, will activate the combustion air blower, the
ventilation blower (if aircraft is on the ground), the heater
fuel pump, open the fuel solenoid valve; and deliver
power to the igniter for combustion. AUTO position of
the heater control switch activates the cabin heater
system and couples it with a temperature-regulation
circuit, which maintains cabin heat between 18°C and
29°C, as established by the temperature control
thermostat. MAN position activates the cabin heater
system but cuts out the temperature-regulation circuit
allowing the heater to operate continuously until limit
switches within the hot air plenum cuts it off at either
107°C or 149°C.
(11) Temperature control circuit breakers. The
combustion blower, igniter fuel pump, heater control box,
and all sensing and regulating circuits are protected by a
single 2-ampere circuit breaker, placarded TEMP
CONTROL, located on the copilot's circuit breaker and
fuse panel.
(12) Cabin temperature rheostat. A rheostat,
placarded CABIN TEMP, located on the right subpanel,
controls cabin temperature between 18°C and 29°C in
the AUTO mode of cabin heater operation.
(13) Cabin heat out indicator light.
A red
press-to-test light, placarded CABIN HEAT OUT located
on the copilot's instrument panel, illuminates if the cabin
heater is inoperative (heater control switch ON or at
AUTO and cabin temperature lower than 18°C).
b. Normal Operation - Cabin Heat.
NOTE
Operation with GPU is the same as
operation with aircraft power.
Change 10 2-55
TM 55-1510-215-10
1. Vent blower operation - Check
2. VENT BLOWER switch - ON.
2. HTR switch - AUTO
3. Cold air outlets - Adjust as required.
3. CABIN TEMP control - As required.
4. CABIN AIR and VENT AIR controls Position as required.
4. CABIN AIR, DEFROST AIR control - As
required.
NOTE
c. Emergency Operation - Cabin Heat.
If the
automatic temperature control should fail to operate, the
temperature may be controlled manually by manipulating
the HTR control switch between the OFF and MAN
positions.
d. Normal Operation - Ventilation System.
1. HTR switch - OFF.
With the heater control switch off,
cold air will enter the cockpit and
cabin through the warm air outlets.
NOTE
The two ventilation air blowers
operate
in
nose
avionics
compartment only when the aircraft
is on the ground with landing gear
shock struts compressed by aircraft
weight.
Section IX. ELECTRICAL POWER SUPPLY AND DISTRIBUTION SYSTEM
2-68. Description.
This aircraft employs both direct current (DC) and
alternating current (AC) electrical power (figs. 2-18,
2-19). The DC electrical supply forms 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 transfer and
auxiliary fuel pumps, heater blower, ventilation blower,
lights and electronic equipment. AC power is obtained
from Dc power through inverters. The three sources of
DC power consist of one 24 volt, 34 ampere/hour battery
and two 250 ampere starter-generators. The starter
generators are controlled by generator control units.
The output of each generator passes through a cable to
the respective generator bus. Other busses distribute
power to aircraft DC loads, and derive power from the
generator buses.
When the generator buses are
coupled, the generators may be paralleled to balance
the Dc loads between the two units.
When the
generating systems are not coupled, if no fault exists,
only generator paralleling and load balance between the
two units is lost. In this circumstance, all aircraft DC
power requirements continue to be supplied, from one or
the other generator source, but not from both. Most DC
distribution buses are connected to both generator
buses but have isolation diodes to prevent power
crossfeed between the generator buses is uncoupled.
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.
2-56 Change 10
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.
2-69. DC Power Supply.
One nickel-cadmium battery furnishes DC power when
the engines are not operating. This is a 24-volt, 34ampere/hour battery, located in the right wing center
section, and accessible through a panel on the top of the
wing. DC power is produced by two engine-driven 28volt 250-ampere starter-generators.
Controls and
indicators associated with the DC supply system are
located on the left subpanel, and consist of a single
battery switch (BAT), two generator switches (GEN 1
AND GEN 2), and single MASTER SWITCH and two
volt-loadmeters.
(Refer to Section VII, UTILITY
SYSTEMS for battery vent system anti-icing.)
a. Battery Switch.
A switch, placard BAT, is
located on the left subpanel under the MASTER
SWITCH (fig. 2-7). The BAT 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 placed down,
the BAT switch is forced OFF.
b. Generator Switches. Two switches, placarded
GEN 1 and GEN 2, are located on the left subpanel
under the MASTER SWITCH (fig. 2-7).
TM 55-1510-215-10
Figure 2-18. Typical Circuit Breaker and Fuse Panels
Change 10 2-57/(2-58 blank)
TM 55-1510-215-10
(Figure 2-19 Sheet 1 of 3)
(Figure 2-19 Sheet 2 of 3)
(Figure 2-19 Sheet 3 of 3)
Figure 2-19. Typical Electrical System Schematic (sheet 1 of 3)
Change 5 2-59
TM 55-1510-215-10
AVIONICS BUS
VHF #1 & #2
NAV #1 & #2
ADF
UHF
XPDR #1
ENC ALT
HF
XPDR-IFF
GPA AS POWER
AUDIO #1 & #2 (PHONE)
TELEPHONE
MARKER BEACON
WX RADAR
FM
DME
RMI #1 & #2
AUDIO #1 & #2 (SPEAKERS)
CABIN AUDIO (PHONE)
RADAR ALT
ALT ALERT
DIRECTIONAL GYRO #2
AUTOPILOT
AVIONICS FANS
ELEC TRIM
HDG FLAG
SUBPANEL BUS NO. 1
TAXI LIGHT
NAV LIGHTS
STALL WARN HEAT
LH PITOT HEAT
L FUEL CONTROL HEAT
L ENG LIP BOOT
L FUEL VENT HEAT
LANDING GEAR CONTROL
SURFACE DEICE
ENG INST LIGHTS
FLIGHT INST LIGHTS
CABIN LIGHTS
FIRE DETECTOR
WINDSHIELD WIPER
STALL WARN HORN
GROUND FAULT
CIGARETTE LIGHTER
R START CONTROL
LANDING GEAR WARN HORN
LANDING GEAR INDICATOR
ANNUNCIATOR PANEL
ANTENNA DEICE
PLT WINDSHIELD ANTI-ICE POWER
PLT WINDSHIELD ANTI-ICE CONTROL
SUBPANEL BUS NO. 2
BEACON LIGHTS
ICE LIGHTS
R FUEL CONTROL HEAT
PROP HEAT
R ENG LIP BOOT
R FUEL VENT HEAT
BATTERY VENT HEAT
INST INDIRECT LIGHTS
SUBPANEL & PEDESTAL LIGHTS
OVERHEAD & FUEL PNL LIGHTS
PROP GOV IDLE STOP
TURN & SLIP IND
FLAP IND
OIL TEMP IND
L START CONTROL
PROP FEATHER
CHIP DETECTOR
INLET AIR SEP
COPILOT WINDSHIELD ANTI-ICE POWER
COPILOT WINDSHIELD ANTI-ICE CONTROL
FUEL PANEL BUS
L FIREWALL VALVE
L AUX PUMP
L TRANSFER
R AUX PUMP
R FIREWALL VALVE
FUEL QTY SYS
CROSSFEED VALVE
R FUEL TRANSFER
CIRCUIT BREAKER PANEL BUS NO. 2
FWD VENT BLOWER
RH IGNITER
RH LIP ANTI-ICE
RH LDG LIGHT
CIRCUIT BREAKER PANEL BUS NO. 1
FLAP MOTOR
LH LDG LIGHT
LH LIP ANTI-ICE
LH IGNITER
TEMP CONT
COMB AIR BLOWER
MAPCO UNIT CONT
AP0058894.2
Figure 2-19. Electrical System Schematic (sheet 2 of 3)
2-60 Change 9
TM 55-1510-215-10
Figure 2- 19. Typical Electrical System Schematic (sheet 3 of 3)
Change 5 2-61
TM 55-1510-215-10
The toggle switches control electrical power from the
designated generator to paralleling circuits and the bus
distribution system. Switch positions RESET, ON and
OFF are placarded on the MASTER SWITCH. RESET
is up (spring-loaded back to ON), ON is center, and OFF
is down. 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.
may also illuminate for short intervals after landing gear
and/or flap operation.
If the caution light should
illuminate during normal steady-state 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 OFF and may be turned
back ON only for gear and flap extension and approach
to landing.
c. Master Switch. All electrical current may be shut
off using the MASTER SWITCH bar (fig. 2-7) which
extends above the battery and generator switches. The
MASTER SWITCH bar is raised when a battery or
generator switch is turned on. Placed down, the bar
forces each switch to the OFF position.
Battery may be usable after a 15 to 20
minute cool down period.
d. Volt-Loadmeters.
Two meters on the left
subpanel display voltage readings and show the
electrical load from left and fight generating systems
(fig. 2-7). 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
member normally shows output electrical lad from the
respective generator, unless the push-button switch is
pressed to obtain bus voltage reading.
Current
consumption is indicated as a percentage of total output
amperage capscrew for the generating system
monitored.
e. Battery Monitor.
Nickel-cadmium battery
overheating will cause the battery charge current to
increase. 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, if installed, 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 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
2-62 Change 10
NOTE
f. Generator Out Warning Lights. Two annunciator
panel fault lights inform the pilot when either generator is
not delivering current to the aircraft DC bus system.
These lights are placarded L GEN OUT and R GEN
OUT (fig. 2-22). Two flashing MASTER WARNING
lights and illumination of either fault light indicates that
either the identified generator has failed or voltage is
insufficient to keep it connected to the bus distribution
system.
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 (fig. 2-23). 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. The holding coil circuit of the relay
is energized by the external power source. The BAT
switch must be in the ON position to connect external
DC power to the aircraft circuits.
2-70. AC Power Supply.
AC power for the aircraft is supplied by two inverters,
which obtain operating current from the DC power
system.
The inverters are operated by a switch
placarded AIRCRAFT INVERTER 1 OFF 2, located on
the pilot's subpanel. Only one inverter can be operated
at a time. The second inverter remains as a spare to be
used in case of AC power failure. The single-phase
inverters provide 115 VAC power to the vertical and
directional gyros. The inverters also provide 26 VAC
power to the following engine instruments: fuel flow, oil
pressure, torquemeter; and avionics systems: HDG,
RMI, compass. ADI, yaw damp.
TM 55-1510-215-10
a. AC Power Warning Lights.
Two flashing
MASTER CAUTION lights and the illumination of an
annunciator caution light INV 1, INV 2, (fig. 2-22)
indicate an inverter failure.
b. Inverter Control Switch.
One three-position
toggle switch on the left subpanel controls the selection
of aircraft inverter source. The switch is placarded
AIRCRAFT INVERTER, 1, OFF, and 2, (fig.
2-7).
Neither inverter is a preferred unit, allowing the pilot free
choice to select either unit. If an inverter fails in service,
power may be restored by selecting the alternate unit.
The aircraft inverter switch placarded AIRCRAFT
INVERTER 1, OFF, and 2 provides a choice between
two 2500 volt-ampere units for single-phase power.
NOTE
When inverters are switched, the
magnetic compass will be affected.
c. Inverter Control Circuit Breakers.
Inverter
control relay circuits are protected by two 1-ampere
circuit breakers placarded INVERTER NO 1 and NO 2
CONTROL, RELAYS, located on the copilot’s circuit
breaker panel (fig. 2-18).
the cable. This will isolate the faulted cable from power
input and will also terminate power crossfeed between
the two generating systems. Generator paralleling and
balanced lead sharing between the two units is lost.
Both generators will remain in operation and all DC
loads will continue to receive power. Each generator will
supply power only to the distribution buses and other
loads which connect to its respective generator bus. In
this circumstance, battery power will be applied only to
the right generator bus. The shorted tie-cable will be
indicated to the pilot only by divergent readings on the
two volt-loadmeters: which will display the unbalanced
loads due to loss of paralleling. To restore connection of
the bus hie-able to the generator buses, and recouple
the two generating systems, cycle the GND FAULT
circuit breaker on the copilot's circuit breaker panel (fig.
2-18). If the fault has been cleared, bus-tie will be
restored. If a fault persists, the ground fault system will
maintain isolation of the tie-cable from the power input,
both generators will continue to operate, all DC loads will
be supplied, and only paralleling and power crossfeed
will be lost.
2-72. Bus Overload Protection System.
CAUTION
2-71. Ground Fault Protection System.
a. A ground fault detection and isolation system is
integrated with the output cable from each generator to
the respective generator bus, and with the tie-cable
which links the two generator buses together. A short in
either generator-to-generator-bus cable will cause the
ground fault system to open the respective line contactor
relay connecting the cable to the generator bus, and to
stop power production from the affected generator. Both
the pilot's and copilot's MASTER WARNING lights will
stare to flash, and the L GEN OUT or R GEN OUT
annunciator warning light will illuminate (fig. 2-22). To
restore a generator to service, move the respective GEN
switch to RESET position, then back to ON. If the
ground fault has cleared, the generator will be operative,
if not, the generator will not reset. If the generator will
not reset, all aircraft DC loads will be supplied from the
remaining generator and the aircraft battery.
b. A short in the wiring which links the two
generator buses together will cause the ground fault
system to open the line contactor relay at each end of
Do not reset the GND FAULT circuit
breaker more than one time. If it trips
after a reset no attempt should be
made to reset the affected generator
using the GEN 1 switch.
A bus overload and protection system is integrated
with the respective generator buses and also with the
battery-generator bus. A temperature-actuated sensor
detects the passage of excessive current through any of
the buses designated, and will isolate the affected bus
from current input and from connection with the rest of
the system. Loss of the left generator bus is indicated
by the flashing MASTER WARNING lights, illumination
of the L, GEN OUT annunciator light, the GND FAULT
circuit breaker trapped, and by a no load reading on the
left volt-loadmeter. In this case, all aircraft DC loads are
supplied by the fight generator, except for those loads
which only receive power from the left generator. To
restore operation of the left generator and reconnect it to
the respective generator bus, reset the GND FAULT
circuit breaker, then
Change 5 2-63
TM 55-1510-215-10
To restore operation of the left generator and reconnect
it to the respective generator bus, reset the GND FAULT
circuit breaker, then position the GEN 1 switch to
RESET, then back to ON.
CAUTION
If the BAT RELAY circuit breaker
trips again, after a reset attempt, do
not attempt to reset the GND FAULT
and BUS OVERLOAD circuit beakers.
CAUTION
Do not attempt to reset the BUS
OVERLOAD circuit breaker more than
one time. If it trips again, after a
reset attempt, no attempt should be
made to reset the affected generator
using the GEN 2 switch.
a. When right generator operation is terminated,
and the right generator bus is disconnected, the
indication will be as follows: both MASTER WARNING
lights will flesh, the R GEN OUT annunciator light will
illuminate, the BUS OVERLOAD circuit breaker on the
copilot's circuit breaker panel will trip, and a no load
reading will appear on the fight volt-loadmeter. To
restore operation of the fight generator, and reconnect it
to the generator bus, reset the BUS OVERLOAD circuit
breaker, then position the GEN 2 switch to RESET, then
back to ON.
b. When a current overload is sensed on the
battery-generator bus, the battery will be isolated from
the bus, and the tie-cable which links the two generator
buses will be disconnected.
Both generators will
continue to operate supplying all loads, although
paralleling will be lost, and battery power will be lost to
all circuits. Fault warning will consist of divergent load
readings on the volt-loadmeters, and the BAT RELAY,
BUS OVERLOAD and GND FAULT circuit breakers will
be tripped (fig. 2-19). To reset the disconnected bus,
reset the BAT RELAY, BUS OVERLOAD and GND
FAULT circuit breakers.
NOTE
In no event will a single bus overload
cause the loss of more than one
power source.
Section X. LIGHTING
2-73. Exterior Lighting.
a. Description. Exterior lighting (fig. 2-20) consists
of conventional navigation lights on the tail section and
on each wing tip, an anti-collision light (STROBE
BEACON) on the vertical stabilizer and the underside of
the fuselage center section, a single taxi light mounted
on the nose landing gear strut, dual landing lights flush
mounted in each wing leading edge near the wing tips,
and two ice lights, one light flush mounted in each
nacelle, positioned to illuminate along the outboard wing
leading edge. All exterior lights except the STROBE
BEACONS are controlled by individual two-position
circuit breaker switches on the left subpanel.
b. Exterior Lighting Circuit Breakers and Switches.
The navigation lights are controlled by a 5-ampere circuit
breaker switch, placarded NAV, on the left subpanel (fig.
2-7). The taxi light is controlled by a 15-ampere circuit
breaker switch, placarded TAXI, on the left subpanel (fig.
2-7). The landing lights for each wing are controlled by
individual circuit breaker switches, placarded LANDING
2-64
LIGHTS, on the left subpanel (fig. 2-7). Both landing
lights have redundant circuit protection. Each light
switch has an internal 20-ampere circuit breaker element
which is in series with a respective 20-ampere circuit
breaker, placarded LANDING LIGHTS-LH-RH, located
on the copilot’s circuit breaker and fuse panel (fig. 2-19).
The STROBE BEACON lights are protected by a 15ampere push-pull circuit breaker placarded BEACON, on
the left subpanel (fig. 2-7). The STROBE BEACON
control and select swatches are located on the control
pedestal placarded STROBE BEACON (fig. 2-6). The
two ice light are controlled by a 5-ampere circuit breaker
switch, placarded ICE, on the left subpanel (fig. 2-19).
Placing any of the control switches ON completes the
circuit to the associated light. In the event of circuit
overload, the circuit breaker portion of the affected
switch will disconnect the circuit from power and the
switch toggle arm will drop back to the OFF potation.
Repositioning a switch toggle to the ON position will
reset the circuit breaker. In the event the beacon circuit
is overloaded, the push-pull circuit breaker will pop out.
To reset the circuit breaker, push in.
TM 55-1510-215-10
1. Wing Tip navigation light (typical both wings)
2. Anti-collision lights (strobe beacon)
3. Tail navigation light
4. Dual landing lights (typical both wings)
5. Ice light (typical both nacelles)
6. Taxi light
Figure 2-20. Exterior Lightning
2-74. Interior Lighting.
Instrument panel lighting is provided by individual
post or eyebrow type red lights installed at the top edge
of each instrument. In addition, instrument indirect
lighting is provided by nine white lamps and nine red
lamps located in the glareshield overhang along the top
edge of the instrument panel. Additional cockpit lighting
is furnished by two white flood lights flush mounted in
the overhead light control panel, and a cockpit utility light
mounted on the window ledge on each side of the
cockpit. Two overhead white flood lights are installed in
the cabin area. All aircraft interior lighting is controlled
by either rheostat, press-to-light, or toggle type switches
located on the overhead control panel (fig. 2-21) in the
cockpit. All switches are appropriately placarded as to
their identity and function position.
a. Master Panel Lights Switch. A MASTER PANEL
LIGHTS switch is located at the center of the overhead
control panel (fig. 2-21). This switch exercises the ON
or OFF control for all lighting circuits, except flood lights,
installed on the panel. It is in series with the individual
lighting control switches and must be ON to enable the
individual switches to function. This switch does not
interfere with the brightness settings, which must be set
on the individual rheostat switches, but it relieves the
pilot of a need to shut off each individual light switch
before leaving the aircraft. Instead, the pilot may leave
the brightness settings as established by placing the
MASTER PANEL LIGHTS switch in the OFF position.
Subsequently, when the MASTER PANEL LIGHTS
switch is placed ON, each individual light will again
illuminate at its previous brightness level.
b. Interior Lighting Circuit Breakers.
Interior
lighting circuits are protected by eight 5-ampere circuit
breakers on the right subpanel under the group
placarded LIGHTS (fig. 2-7). The circuit breaker INST
INDIRECT protects the red and white lights on the
instrument panel glareshield. The circuit breaker PED &
SUB PANEL protects both subpanel lighting circuits and
the control pedestal lights. The FLT INST circuit breaker
protects the pilot and copilot instrument lights, the clock
light, the pilot and copilot oxygen regulator lights, the
magnetic compass light, and the free air temperature
gage light. The OVHD & FUEL PNL circuit breaker
protects the copilot's circuit breaker panel lights, the
edge lights on the fuel control panel, the oxygen
pressure gage light, and the edge lights and flood lights
of the overhead control panel. The ENG INST circuit
breaker protects the instrument panel post lights and the
2-65
TM 55-1510-215-10
Figure 2-21. Overhead Control Panel
radio lights. The CABIN circuit breaker protects the aft
cabin and utility lights.
c. Interior Lighting Controls. Two white flood lights
are integrally mounted on the bottom part of the
overhead control panel (fig. 2-21). These flood lights are
controlled by a rheostat switch, placarded OVERHEAD
FLOOD LIGHTS, located on the overhead control panel.
One cockpit utility light assembly is mounted adjacent to
the inertia reel release handle on each side of the
cockpit. The multipurpose light assembly is designed to
serve either as a red or white map light, or as a red or
white flood light.
The different lighting effects are
controlled by depressing the red button on the top of the
light assembly and turning the lens assembly to the
desired position. The light is controlled by a rheostat
and push-button switch at the aft end of the light
assembly. The rheostat is used during normal operation
of the light and is placarded OFF, DIM, and BRIGHT.
When depressed, the push-button switch provides
momentary full light output regardless of the rheostat
position. A swivel-hinge allows the light to be positioned
as desired. The light assembly can be removed from the
base assembly to provide more mobility in the use of the
light. The overhead panel flood lights are protected by a
5-ampere circuit breaker, placarded OVHD & FUEL PNL
on the right subpanel (fig. 2-7). The cockpit utility light
assembly is protected by a 5-ampere circuit breaker,
placarded CABIN, on the right subpanel (fig. 2-7).
d. Free Air Temperature Light.
The free air
temperature light is mounted overhead in the cockpit
adjacent to the free air temperature gage. The light is
operated by the push-button switch, placarded FREE
AIR TEMP-PRESS TO LIGHT (fig. 2-21), to illuminate
2-66
the face of the gage. The light is protected by a 5-mpere
circuit breaker on the right subpanel placarded FLT INST
(fig. 2-7). The free air temperature light is similar in
construction to the instrument post lights.
e. Magnetic Compass Light.
The magnetic
compass light, above the face of the compass, makes
the compass card readable. This light is controlled by a
rheostat-switch,
placarded
PILOT
INSTRUMENT
LIGHTS - BRT, OFF, located on the overhead control
panel (fig. 2-21). This light is protected by a 5-ampere
circuit breaker, placarded FLT INST, on the right
subpanel (fig. 2-7).
f. Subpanels and Copilot's Circuit Breaker Panel
Edge Lighting.
The plastic cover panels on the
subpanels and the copilot's circuit breaker panel are
illuminated by edge lights. The edge-light socket is
recessed into the plastic panel. The cap assembly,
which contains a red filter, screws down against the
plastic panel. The edge lights on the subpanels are
controlled by a rheostat-switch, placarded SUBPANEL &
CONSOLE LIGHTS, on the overhead control panel
(fig. 2-21). Edge light circuits receive power from a 5ampere circuit breaker, placarded FED & SUB PANEL,
located on the right subpanel (fig. 2-7).
g. Normal Operation of Interior Lights. Normal
operation will consist of placing the MASTER PANEL
LIGHTS switch ON, then at the pilot's choice turning
individual lights ON or OFF or readjusting the
established brightness level at which they illuminate
when the master panel lights switch is positioned ON.
The passengers individually control the lights at their
respective seats or station.
TM 55-1510-215-10
Section XI. FLIGHT INSTRUMENTS
2-75. Airspeed Indicators.
Two airspeed indicators are installed separately on the
pilot and copilot sides of the instrument panel (fig. 2-22).
The indicator dials are calibrated in knots from 40 to 260.
2-76. Turn-And-Bank Indicators.
Two turn-and-bank indicators are installed, one each
on the pilot and copilot sides of the instrument panel
(fig.
2-22).
These indicators are gyroscopically
operated. The pilot's unit is operated by DC power and
is protected by a 5-ampere circuit breaker, placarded
T&B IND, located on the right subpanel (fig. 2-7). The
copilot's indicator is driven by the aircraft vacuum
system.
2-77. Pilot's Encoding Altimeter.
The pilot's altimeter is located on the upper left side
of the instrument panel (fig. 2-22). Altitude is displayed
on the pilot's altimeter by a 10,000 foot counter a 1,000
foot counter, a 100 foot counter, and a single needle
pointer. Below an altitude of 10,000 feet, black and
white cross-hatching will appear on the 10,000 foot
counter. Altitude below sea level is indicated by a wavy
blue and white line in the 10,000 foot counter. A
barometric pressure setting knob is provided to insert the
desired altimeter setting. The altimeter contains dual
baroscales which permits barometric pressure setting in
both millibars and inches of mercury (Hg). A red and
white striped warning flag will cover the digital altitude
indicator if DC power is lost. The encoding altimeter is
protected by a 1-ampere circuit breaker, placarded ENC
ALT, located on the right subpanel (fig. 2-7).
(1) Encoding Altimeter. The altimeter is a
self-contained unit which consists of a sensitive pressure
altimeter combined with an altitude encoder.
The
display indicates and the encoder transmits,
simultaneously, pressure altitude reporting. 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.
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 left portion of the instrument face, indicating
that the altitude encoder is inoperative, and that the
system is not reporting altitude to ground stations.
(2) Barometric Altimeter.
The barometric
altimeter, displays altitude in hundreds of feet by a long
pointer, in thousands of feet by a second pointer, and in
ten-thousands of feet by a disc mounted pointer. A
barometric pressure setting knob is provided to insert the
desired altimeter setting in inches Hg. This unit requires
no electrical power for operation.
An alternate unit may be installed which is similar in
appearance to the encoding altimeter. This unit requires
DC power for proper operation. It does not contain the
CODE OFF flag, and does not have altitude encoding
capability.
2-79. Vertical Velocity Indicators.
Two vertical velocity indicators are installed
separately on the pilot and copilot sides of the
instrument panel (fig. 2-22). 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.
2-78. Copilot's Altimeter.
2-80. Pilot's Fight Director Indicator.
a. Copilot's Altimeter. The copilot's altimeter is
located on the upper right side of the instrument panel
(fig. 2-22). Either one of the following altimeters may be
installed.
The pilot's flight director indicator (FDI) (fig. 3-10)
combines the attitude sphere display with computed
steering information to provide the commands required to
intercept and maintain a desired flight path. All guidance
Change 5 2-67
TM 55-1510-215-10
NOTE
INSTRUMENT RANGE MARKINGS
ARE CONTAINED IN CHAPTER 5,
SECTION II, SYSTEMS LIMITS,
AS APPLICABLE.
(Figure 2-22 Sheet 1 of 2)
(Figure 2-22 Sheet 2 of 2)
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
Turn-and-bank indicator (electric)
Vertical gyro fast erect switch
Airspeed indicators
Flight director indicator
Flight director mode annunciator panel
Encoding altimeter
Propeller autofeather arm annunciators
Turbine tachometers
Torquemeters
ITT gages
Propeller tachometers
DME transfer switch
13.
14.
15.
15A.
15B.
15C.
16.
17.
18.
19.
20.
Avionics master switch
NAV 1 control panel
COMM 1 control panel
GPAAS volume control
GPAAS VOICE OFF, VA FAIL switch indicator
GPAAS circuit breaker
IFF caution annunciator
Instrument suction gage
COMM 2 control panel
Emergency static air control
NAV 2 control panel
AP011804.1
Figure 2-22. Typical Instrument Panel (sheet 1 of 2)
2-68 Change 10
TM 55-1510-215-10
21.
22.
23.
24.
25.
26.
27.
28.
29.
30.
Marker beacon receiver
Attitude indicator
Copilot's horizontal situation indicator
Altimeter
Vertical speed indicators
Turn-and-bank indicator (air)
Slave control units
Audio control panels
DME indicator
Radio magnetic indicators
31.
32.
33.
34.
35.
36.
37.
38.
39.
40.
Radar altimeter
Weather radar
ADF control panel
Caution annunciator pane
Transponder control panel
Oil pressure gages
Oil temperature gages
Fuel flow gages
Altitude select controller
Pilot's horizontal situation indicator
AP011804.2
Figure 2-22. Typical Instrument Panel (sheet 2 of 2)
Change 9 2-69
TM 55-1510-215-10
commands are depicted through the use of a command
V-bar arrangement. Warning flags are provided to
indicate invalid attitude or computed command displays.
Any warning flag in view indicates that that portion of
information is unreliable. An inclinometer, located on the
bottom of the indicator, provides the pilot a conventional
display of aircraft slip or skid.
a. Fast Erect Switch. A switch placarded V/G
FAST ERECT, located on the pilot's side of the
instrument panel, provides for fast erection of the FDI
attitude sphere.
2-80A. Copilot's Attitude Director Indicator.
The copilot's attitude director indicator (ADI) (fig. 3-11)
displays aircraft attitude as a conventional pneumatic
operated attitude gyro.
Attitude displayed is in
relationship to an artificial horizon. An indicating plane in
front of the face of the instrument represents the aircraft
and a movable horizontal bar behind the indicating plane
represents the horizon. The symbolic aircraft may by
adjusted vertically by means of a knob located on the
indicator, to correct for variations in level flight attitude at
different airspeeds. This unit is designed to operate
through all attitudes and need not be caged for any
maneuver. The ADI is equipped with a decision height
(DH) annunciator, which works in conjunction with the
radar altimeter.
2-81. Free Air Temperature Gage.
The free air temperature gage, centered above the
windshield (fig. 2-5), indicates the outside air temperature
in degrees celsius.
2-82. Magnetic Compass.
The magnetic compass is located on the top of the
windshield divider. A compass correction chart indicating
deviation is located adjacent to the magnetic compass.
2-84. Miscellaneous Instruments.
a. Annunciator Panels. Two annunciator panels
are installed. One in a WARNING panel with red fault
identification lights, and the other is a CAUTION panel
with yellow identification lights. The WARNING panel is
mounted in the center of the glare shield above the
instrument panel (fig. 2-22) and the CAUTION panel is
located in the bottom center of the instrument panel
(fig. 2-22). 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. 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. Both panels employ an automatic
brightness control circuit to adjust brightness of the
legend indicators to assure good readability in the
presence of other cockpit lighting. Automatic brightness
circuits will function only if the following switching
combination is established: OVERHEAD FLOOD
LIGHTS - OFF, MASTER PANEL LIGHTS - ON, and
GEN 1 and GEN 2 switches ON. If any one of the
switching positions stated are changed, all lights of both
annunciator panels will illuminate at maximum bulb
brightness. The bulbs of all annunciator panel lights are
tested by pressing the momentary push button switch
which is located on the right side of the warning
annunciator panel.
(1) Master warning light (red).
A master
warning light is provided for both the pilot and copilot,
and is located at the top center of their respective
instrument panels (fig. 2-22). Any time a warning light
illuminates, the master warning light will flash, and will
continue to flash until the illuminated warning light
condition is corrected and/or the master warning light is
pressed to reset the flashing circuit. If a new condition
occurs, the flashing light will be reactivated, and the
applicable annunciator panel light will illuminate.
2-83. Radio Magnetic Indicators (RMI).
Identical radio magnetic indicators are installed for
the pilot and copilot (fig. 3-6). Each RMI provides
aircraft heading and radio bearing information to-from a
VOR, ADF facility or RNAV waypoint. A selector switch
on the RMI provides for selection of either NAV or ADF
as the bearing information to be displayed by either
needle. Slave control units located on the pilot and
copilot sides of the instrument panel (fig. 2-22) controls
the RMI gyro slaving circuit. Refer to chapter 3.
2-70 Change 10
(2) Master caution light (yellow). A master
caution light is provided for both the pilot and copilot and
is located adjacent to the master warning light near the
top center of the pilot's and copilot's respective
instrument panel area (fig. 2-22). Whenever a caution
light illuminates, the master caution will flash, and will
continue to flash until the illuminated caution light
condition is corrected and/or the master caution light is
pressed to reset the flashing circuit. If a new condition
occurs, the flashing light will be reactivated and the
TM 55-1510-215-10
appropriate annunciator panel lights will illuminate.
(3) Caution legend light and switch.
One
green light and a switch are located above the caution
annunciator panel (fig. 2-22). The switch is spring
loaded to the center OFF position and is placarded ON,
OFF, CAUTION LEGEND. The green light is located
below the words CAUTION LEGEND. If the pilot tires of
seeing long existent conditions illuminated on the
caution panel, he may momentarily position the switch
OFF, to extinguish all lights on the caution panel, and
cause the green light to illuminate as the only indication
that these conditions still exist. If the pilot chooses, he
may momentarily press the switch to ON, to extinguish
the green light and cause redisplay of each existing
condition on the caution panel. If a new monitored event
occurs while the green light is illuminated and the
caution panel is darkened, the signal of the new
condition will automatically extinguish the green light and
redisplay all existing events on the caution panel.
Circuits of the annunciator panels are protected by a
5-ampere circuit breaker, placarded ANN PANEL,
located on the right subpanel (fig. 2-7).
b. Clocks. A digital clock/timer is located in the
center of each control wheel (fig. 2-16). Each digital
clock is protected by a 1-ampere circuit breaker, located
in the DC junction box. The DC junction box is located
in the nose compartment of the aircraft.
(1) Digital clock operating procedures. The
MODE button is pressed to select the desired operation.
The mode annunciator is displayed above the word
TIMER or above the word CLOCK. The position of the
annunciator indicates the function mode, either timer or
clock, except for the 24 hour clock operation.
1. Press the MODE button to position
the annunciator above the word
CLOCK.
Date Setting:
2. Press the SET button one time. The
month's digit will flash. Press the
DT/AV button to advance the
month's digit to the desired month.
Press the SET button one time and
the day's digit will flash. Press the
DT/AV button to advance to the
desired day. The date is now set.
Time Setting.
NOTE
It is always necessary to go through
the date step before setting the time.
If the date is correct, simply press the
SET button four times to cause the
hour's digit to flash, then proceed
with the following instructions.
NOTE
Twelve or twenty-four hour time is
displayed depending upon installation.
In the 24 hour clock operation only,
the clock annunciator on the display
does not appear.
Clock or timer
function is indicated only by the timer
annunciator on the display.
3. Press the SET button two times to
cause the hours digit to flash. Then
press the DT/AV button to advance
the hour digit. Press the SET button
one time to cause the minutes digit
to flash. Then press the DT/AV
button to advance the minute's digit.
The minute should be set to the next
minute to come up on the time
standard being used for the correct
time. The SET button is pressed
once more to hold the time
displayed, press the DT/AV button
to start the clock function at the
exact second. If the minutes are not
advanced when they are flashing in
the set mode, pressing the SET
button will return the clock to the
normal time keeping mode without
altering the minutes timing. This
feature is useful when changing
time zones when only the hours are
to be changed.
(2) Digital clock operation. Normal operation
of the clock will be indicated by activity of the colon
indicator. The colon dots will blink off for one second
every ten seconds to indicate correct clock functioning.
(3) Date operation.
With
selected, pressing the DT/AV button
activate the date mode to the display.
date will be visible for 1.5 seconds.
the clock mode
momentarily will
The month and
The display will
Change 10 2-70A
TM 55-1510-215-10
then automatically return to the clock display mode.
During the display of the date, the clock annunciator will
be blanked from the display. If, when in the clock mode,
the DT/AV button is pressed continuously for two
seconds, the display will return from the date to the time
function. In this time mode the colon activity is altered
and may read continuously or be blanked from the
display. To return to the normal time mode with its colon
activity indication, press the DT/AV button again for two
seconds and the correct time mode will be restored. The
calendar function will automatically advance the date
correctly according to the four year perpetual calendar.
One day must be added manually on February 29 of
every leap year.
The date advances correctly at
midnight each day. In the 12 hour time mode, AM or PM
indication is not provided. It is possible, in rare cases
such as initial setup after a battery change, for the date
to change at 12 noon instead of 12 midnight. If this is
observed just add 12 hours to the time displayed using
the set methods previously described. A simple test can
be made to determine that the clock is reading AM or
PM correctly. Set in 11:59 on the clock and note the
date. Activate the clock and wait one minute. If the date
remains the same after the clock times to 12:00, the
clock is reading noon and can now be set to the correct
AM or PM time.
(4) Timer. The timer may be used effectively
as a cumulative trip timer for single or multileg flights.
The timer is started from zero at the beginning of a flight.
The time into the flight is continuously available by
selecting the timer mode. Returning to the clock mode
will not disturb the operation of the timer. At the end of
the first flight leg, the ST/SP button is pressed while in
the timer mode to hold the total flight time accumulated
so far. This will be retained in memory and will be
available for view any time the timer mode is selected
even when on the ground and shut down. At the
beginning of the next flight leg the previously
accumulated flight time may be zeroed or maintained.
Pressing the ST/SP button will again start the counter,
either adding to the previous flight time or from zero, as
desired. If the flight timer feature is not used, the timer
may be used for any short term timing function such as
ground and speed checks, timed turns or timed
approaches. For ground speed checks, the timer is
started at the beginning check point and then stopped at
the final check point. The elapsed time is now held on
the display for ground speed calculation. For timed
approaches, the timer is previously set to zero. At the
final approach fix, the ST/SP button is activated and the
time shown on the display will be elapsed time into the
approach.
2-70B Change 10
NOTE
The flight time up to approach may
be recorded just before the time is
set to zero for the approach. Total
flight time, less approach time, will
then be available.
(5) Timer operation.
1. Press the MODE button to position
the annunciator above the word
TIMER.
2. Press the RST button so that the
time will read zero.
3. To start the timer counting, press
the ST/SP button one time. The
timer will initially count in minutes
and seconds and the colon will blink
off for one-tenth second each
second. After 59 minutes and 59
seconds, the timer will change to
count in hours and minutes up to a
maximum of 23 hours and 59
minutes.
During the hours and
minutes count, the colon will
indicate counting activity by blinking
off one second each ten seconds.
The timer count may be stopped
and held at a particular time by
pressing the ST/SP button. To add
to the existing count, press the
ST/SP button again and the timer
will continue counting up from the
previous total. To start a new count,
press the RST button and zero the
display. Press the ST/SP button to
start a new count.
(6) Display test. Pressing both the SET and
DT/AV buttons at the same time will result in a display
test function.
All display segments, colon, and
annunciators will be activated. If the display test is done
from the clock mode it will return to the time set mode. If
the display test is done from timer mode, the display will
return to the timer mode in the set condition.
c. Other Instruments.
The instrument pressure gage is located on the
instrument panel (fig. 2-22).
TM 55-1510-215-10
Section XII. SERVICING, PARKING, AND MOORING
2-85. General.
The following paragraphs include the procedures
necessary to service the aircraft, except lubrication
(fig. 2-23). The lubrication requirements of the aircraft are
covered in TM 55-1510-209-23. Refer to tables 2-3, 2-4,
2-5, and 2-6, for identification of fuel, oil, etc. used to
service the aircraft. The servicing instructions provide
procedures and precautions necessary to service the
aircraft.
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.
2-86. Servicing Fuel System.
The fuel system is made up of a 57 gallon
(370.5 pounds) nacelle tank and wing tank fuel cells of
128 gallons (832 pounds) capacity, making a total fuel
capacity of 370 gallons (2,405 pounds).
NOTE
Service the fuel tanks after each
flight to keep the bladder type cells
from drying out.
CAUTION
Proper procedures for handling JP-4
and JP-5 fuel cannot be over
stressed. Clean, fresh fuel must be
used and the entrance of water into
the fuel storage or aircraft fuel
system must be kept to a minimum.
CAUTION
NOTE
At least thirty minutes after servicing
the tanks, open the sump drains and
fuel strainer drains and allow a small
amount of fuel to drain from each of
these points.
NOTE
Settling time for AVGAS is 15
minutes per foot of tank depth and
one hour per foot depth for jet (JP)
fuels. Allow the fuel to settle for the
prescribed period before any fuel
samples are taken.
a. Fuel Handling Precautions. While handling fuel
it is well to remember that even though the aircraft uses
JP-4 (primary) and JP-5 (alternate) as its principal fuel, it
may be operated on aviation gasoline as an emergency
fuel.
When conditions permit, the aircraft
should 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.
NOTE
The use of aviation gasoline as an
emergency fuel is permitted for a
period of not more than 150 hours
during time between engine overhauls
(TBO). A mixture of more than 10%
aviation gasoline shall be entered on
DA Form 2408-13. The lowest octane
aviation
gasoline
available
is
recommended to avoid build up of
deposits on turbine blades.
1. Shut off all electrical equipment on the
aircraft, including radar and radar equipment.
The
master switch may be left on, but it must 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.
2-71
TM 55-1510-215-10
1. Fuel tank filler caps (typical left and right)
Fuel: Spec. MIL-F-5624 (JP-3, JP-4, or JP-5)
2. Wheel brake fluid reservoir
Hydraulic fluid: Spec. MIL-H-5606
3. Battery
4. Oxygen supply cylinders
Breathing oxygen: Spec. MIL-O-27210
5. Oxygen filler valve
Breathing oxygen: Spec. MIL-O-27210
6. Engine oil filler caps (typical, left and right)
Engine oil: Spec. MIL-L-7808, synthetic base, -40°F (-40°C) and below
Engine oil: Spec. MIL-L-23699, synthetic base, above -40°F (-40°C)
7. Fire extinguishers, hand type CF3Br
Extinguishing agent: Spec. MIL-E-52031
8. DC external power receptacle (24 volt)
Figure 2-23. Servicing Locations
2-72
TM 55-1510-215-10
Table 2-3. Fuels, Lubricants, Specifications, and Capacities
SYSTEM
SPECIFICATION
CAPACITY
Fuel
MIL-T-5624 (JP-4)
373.6 U.S. Gals.
Engine oil
MIL-L-7808 (See Notes 1, 3, and 4)
MIL-L-23699 (See Notes 2 and 3)
14 U.S. Quads
per engine
Hydraulic brake system
MIL-H-5606
1 U.S. Pint
Oxygen system
MIL-O-27210
203 Cubic feet
NOTE 1:
MIL-L-7808 oil used in the engine oil system is specified for operation in ambient
temperatures below -40°C (-40°F). This oil may also be used when MIL-L-23699 oil is not
available.
NOTE 2:
CAUTION
Under no circumstances shall MIL-L-23699 oil be used at ambient temperatures
below -40°C (-40°F). MIL-L-23699 oil used in engine oil system is authorized and
directed for use in ambient temperatures above -40°C (-40°F).
CAUTION
NOTE 3:
Do not mix MIL-L-23699 oil with MIL-L-7808 oil except in case of emergency. If it
becomes necessary to mix the oils, the applicable system shall be flushed within
six hours and filled with the proper oil.
CAUTION
NOTE 4:
Lubrication oil made to MIL-L-7808 by Shell Oil Company under their part number
307, qualification number 7D-1 shall NOT be used in U-21 series engines.
2. Refuel aircraft as soon as possible after
landing.
3. Keep fuel servicing nozzles free of snow,
water, and mud at all times.
4. Carefully remove snow, water, and ice
from the aircraft fuel filler cap area before removing the
fuel filler cap. Remove only one aircraft tank filler cap at
any one time, and replace each one immediately after
the servicing operation is completed.
5. Wipe all frost from fuel filler necks before
WARNING
Prior to opening the fuel tank filler,
the hose nozzle static ground wire
must be attached to the grounding
lugs that are located adjacent to filler
openings.
servicing.
6. Drain water from fuel tanks, filter cases,
and pumps at least 30 minutes after each servicing, at
least 30 minutes after each removal from heated shelter,
and before each flight. Preheat, when required, to
insure free fuel drainage.
2-73
TM 55-1510-215-10
Table 2-4. Approved Oils
SPECIFICATION
PRODUCT
MIL-L-23699
(5 centistokes)
Esso Turbo Oil 2380
Exxon International Co.
New York, NY
Toronto, Ontario,
Canada
Exxon Turbo Oil 2380
Exxon Co.
Houston, Texas
Aero Shell Turbine Oil 500
Shell Oil Co.
Houston, Texas
Toronto, Ontario,
Canada
Stauffer Jet II
Stauffer Co.
Westport, CT
Castrol 5000
United Kingdom
7. Insofar as
serviced aircraft outside.
possible,
leave
only
fully
8. Avoid dragging the fueling hose where it
can damage the soft, flexible surface of the deicer boots.
9. Insure that the aircraft and components
being serviced are securely coupled to a low resistance
ground. Connect the static ground cable and the fueler
to a grounding stake. When engaged in the fueling
operation, discharge the static electricity accumulated by
the body and clothing by touching the ground cable or
stake before each operation.
VENDOR
15. Wear only nonsparking shoes near aircraft
or fueling equipment, as shoes with nailed soles or metal
heel plates can be a source of sparks.
b. Fuel Handling Precautions For Extreme Weather
Conditions. When fueling is conducted on ice or on
sandy or desert terrain, or when a satisfactory ground
cannot be secured, additional precautions must be taken
to avoid the buildup of static electricity.
Refer to
TM 55-410. Since draining the static charges is not
possible without an adequate ground, reliance must be
placed on equalizing the potentials in order to prevent a
dangerous sparking discharge as in the following
bonding procedure:
10. Observe NO SMOKING precautions.
11. Prior to transferring the fuel, insure that
the hose is grounded to the aircraft.
2. Connect grounding cable to some
convenient unpainted metal part of aircraft, excluding
propellers or antennas.
12. Wash off spilled fuel immediately.
13. Handle the fuel hose and
cautiously to avoid damaging the wing skin.
nozzle
14. Do not conduct fueling operations within
100 feet of energized airborne radar equipment or within
300 feet of energized ground radar equipment
installations.
2-74
1. Connect ground cable from fuel servicing
unit to satisfactory ground, such as a grounding post.
3. Connect bonding cable from aircraft to fuel
servicing unit. A conductive-type fuel hose is not a
satisfactory method of bonding.
4. Connect bonding cable from fuel nozzle to
aircraft before fuel tank cover is opened.
TM 55-1510-215-10
Table 2-5. Approved Fuels
SOURCE
PRIMARY OR
STANDARD FUEL
ALTERNATE FUEL
US Military Fuel
NATO Code No.
JP-4 (MIL-T-5624)
F-40 (Wide Cut Type)
JP-5 (MIL-T-5624)
F-44 (High Flash Type)
COMMERCIAL FUEL
(ASTM-D-1655)
American Oil Co.
Atlantic Richfield
Richfield Div.
B. P. Trading
Caltex Petroleum Corp.
Cities Service Co.
Continental Oil Co.
Gulf Oil
EXXON Co. USA
Mobil Oil
Philips Petroleum
Shell Oil
Sinclair
Standard Oil Co.
Chevron
Texaco
Union Oil
FOREIGN FUEL
Belgium
Canada
Denmark
France
Germany (West)
Greece
Italy
Netherlands
Norway
Portugal
Turkey
United Kingdom
(Britain)
JET B
American JP-4
Arcojet B
JET A
American Type A
Arcojet A
Richfield A
B.P.A.T.G.
Caltex Jet B
Conoco JP-4
Gulf Jet B
EXXON Turbo Fuel B
Mobil Jet B
Philjet JP-4
Aeroshell JP-4
Chevron B
Texaco Avjet B
Union JP-4
NATO F-40
BA-PF2B
3GP-22F
JP-4 MIL-T-5624
Air 3407A
VTL-9130-006
JP-4 MIL-T-5624
AA-M-C-1421
JP-4 MIL-T-5624
JP-4 MIL-T-5624
JP-4 MIL-T-5624
JP-4 MIL-T-5624
D. Eng RD 2454
CITGO A
Conoco Jet-50
Gulf Jet A
EXXON A
Mobil Jet A
Philjet A-50
Aeroshell 640
Superjet A
Jet A Kerosene
Chevron A-50
Avjet A
76 Turbine Fuel
JET A-1
NATO F-34
Arcojet A-1
Richfield A-1
B.P.A.T.K.
Caltex Jet A-1
Conoco Jet-60
Gulf Jet A-1
EXXON A-1
Mobil Jet A-1
Aeroshell 650
Superjet A-1
Jet A-1 Kerosene
Chevron A-1
Avjet A-1
NATO F-44
3-6P-24e
UTL-9130-007/UTL-9130-010
AMC-143
D. Eng RD 2493
D. Eng RD 2498
NOTE:
Anti-icing and Biocidal Additive for Commercial Turbine Engine Fuel - The fuel system icing inhibitor shall
conform to MIL-1-27686. The additive provides anti-icing protection and also functions as a biocide to kill
microbial growths in aircraft fuel systems. Icing inhibitor conforming to MIL-1-27686 (PRIST) shall be added to
commercial fuel not containing an icing inhibitor during refueling operations, regardless of ambient
temperatures. Refueling operations shall be accomplished in accordance with accepted commercial procedures
Refer to paragraph 2-86. (c)
Change 7 2-75
TM 55-1510-215-10
Table 2-6. Standard, Alternate, and Emergency Fuels
ENGINE
ARMY STANDARD
FUEL
T74-CP-700
MIL-T- 5624
Grade JP-4
ALTERNATE
FUEL
MIL-T-5624
Grade JP-5
EMERGENCY FUEL
TYPE
*MAX.
HOURS
MIL-G-5572
Any AV Gas
150
*Maximum operating hours with indicated fuel between engine overhauls (TBO).
5. When disconnecting, reverse the order of
steps 1 through 4.
d. Filling Fuel Tanks. Fill tanks as follows:
1. Attach bonding cables to aircraft.
c. Blending Anti-icing Additives to Fuel.
2. Attach bonding cable from hose nozzle to
ground socket adjacent to fuel tank being filled.
CAUTION
3. Always fill nacelle tank before wing tank.
Insure that the additive is directed
into the flowing fuel stream; start
additive flow after fuel flow starts,
and stop before fuel flow stops. Do
not allow concentrated additive to
contact coated interior of fuel cells or
aircraft painted surfaces. Use not
less than 20 fl oz of additive per 260
gallons of fuel or more than 20 fl oz
of additive per 104 gallons of fuel.
The following procedure is to be used when blending
anti-icing additive (which must conform to specification
MIL-I-27686) with the fuel as the aircraft is being refueled:
1. Using "HI-FLO PRIST" blender (Model PHF204), remove cap containing the tube and clip assembly.
2. Attach pistol grip on collar.
3. Press tube into button.
4. Clip tube end to fuel nozzle.
5. Pull trigger fi rmly to ensure full flow, then
lock in place.
6. Start
begins. (Refueling
gallons per minute.
minute may be used
2-76
flow of additive when refueling
should be at a rate of 30 to 60
A rate of less than 30 gallons per
when topping off tanks.)
4. Open applicable fuel tank filler cap.
5. Fill fuel tank with fuel.
6. Secure applicable fuel tank filler cap.
CAUTION
Make sure fuel filler cap latch tab is
pointed aft.
7. Disconnect bonding cables from aircraft.
e. Filling Ferry Fuel System Tanks.
Fill ferry
system tanks in accordance with procedures and safety
precautions which are used in filling the aircraft's regular
fuel tanks. Observe placard on ferry system tanks which
reads FERRY TANK CAPACITY 120 U.S. GALLONS.
FILL TO BOTTOM OF TAB ON FILLER NECK. USE
AVIATION KEROSENE OR SEE PILOT'S OPERATING
MANUAL FOR ALTERNATES.
f. Draining Moisture From Fuel System.
To
remove moisture and sediment from the fuel system, 10
fuel drains are installed. The locations are as follows:
one outboard of nacelle, one in the main gear wheel well
(low point of fuel system), one just ahead of wheel well
(to drain boost pumps), one inboard of nacelle (to drain
TM 55-1510-215-10
transfer pumps), and one at the inertial separator air
bypass duct (to drain fuel filter).
g. Fuel Types. Approved fuel types are as follows:
(1) Army standard fuels. These are the Armydesignated primary fuels adopted for worldwide use, and
are the only fuels available in the Army supply system.
(2) Alternate fuels. These are fuels which can
be used continuously when Army standard fuel is not
available, without reduction of power output. Power
setting adjustments and increased maintenance may be
required when an alternate fuel is used.
(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.
However, 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.
2-87. Servicing Oil System.
(3) Emergency fuel. These are fuels which
can be used if Army standard and alternate fuels are not
available. Their use is subject to a specific time limit.
CAUTION
h. Use of Fuel. Fuel is used as follows:
(1) Fuel limitations.
There is no special
limitation on the use of Army standard fuel, but certain
limitations are imposed when Alternate or Emergency
fuels are used. For the purpose of recording, fuel
mixtures shall be identified as to the major component of
the mixture, except when the mixture contains leaded
gasoline. A fuel mixture which contains over 10 percent
leaded gasoline shall be recorded as all leaded gasoline.
The use of any fuels other than standard will be entered
in the FAULTS/REMARKS column of DA Form 2408-13,
Aircraft Maintenance and Inspection Record, noting the
type of fuel, additives, and duration of operation.
(2) 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.
(3) 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.
(4) Fuel specifications.
Fuels having the
same NATO code number are interchangeable. Jet
fuels 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 refueled with NATO F-40
Do not mix MIL-L-23699 oil with MILL-7808 oil except in case of
emergency. If it becomes necessary
to mix the oils, the applicable system
shall be flushed within six hours and
filled with the proper oil.
CAUTION
Make sure oil filler cap latch tab is
pointed aft.
An integral oil tank occupies the cavity formed
between the accessory gearbox housing and the
compressor inlet case on the T74-CP-700 engine. The
tanks have a total oil capacity of 2.3 gallons and features
a calibrated oil dipstick and an oil drain plug. If engine
has been stationary for more than 12 hours, carry out a
motoring run (no ignition) before checking oil level.
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.
1. Open the upper rear cowling to gain access to
the oil filler cap and dipstick.
2-77
TM 55-1510-215-10
2. Remove oil filler cap.
3. Insert a clean funnel, with a 100 micron screen
incorporated, into the filler neck.
4. Replenish with oil to within 1 quart below MAX
mark on dipstick.
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 again frost.
NOTE
5. Check oil filler cap for damaged preformed
packing, general condition and locking.
6. Secure oil filler cap.
7. Check for and remedy any oil leaks.
2-88. Servicing Hydraulic Brake System Reservoir.
1. Gain access to brake hydraulic system reservoir
(fig. 2-23).
NOTE
The hydraulic brake system reservoir
for the aircraft is located on the right
side of the bulkhead in the nose
avionics compartment.
2. Remove brake reservoir cap and fill reservoir to
washer on dipstick with hydraulic fluid.
3. Install brake reservoir cap.
2-89. Inflating Tires.
Inflate tires as follows:
1. Inflate nose wheel tire to a pressure between 50
and 55 PSI. (For very soft field takeoffs, nose wheel tire
pressure should be between 30 and 32 PSI.)
2. Inflate main wheel tires to a pressure between
49 and 55 PSI for 12 ply tires.
2-90. Anti-icing, Deicing and Defrosting Protection.
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
2-78
Do not apply anti-icing, deicing and
defrosting fluid on 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.
2-91. Anti-icing, Deicing and Defrosting Treatment.
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.
1. Remove frost or ice accumulations from aircraft
surfaces by spraying with diluted anti-icing, deicing, and
defrosting fluid mixed in accordance with table 2-7.
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.
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 is
expensive and slow.
TM 55-1510-215-10
Table 2-7. Recommended Fluid Dilution Chart
AMBIENT
TEMPERATURE
(°F)
PERCENT
DEFROSTING
FLUID BY VOLUME
PERCENT WATER
BY
VOLUME
+ 30° and above
+ 20°
+10°
0°
-10°
-20°
-30°
20
30
40
45
50
55
60
80
70
60
55
50
45
40
NOTES:
FREEZING POINT OF
MIXTURE (°F)
(APPROXIMATE)
+10°
0°
-15°
-25°
-35°
-45°
-55°
1. Use anti-icing and deicing fluid (MIL-A-8243).
2. Heat mixture to a temperature of 180° to 200°F (82° to 93°C)
4. If tires are frozen to ground, use undiluted
defrosting fluid to melt ice around tire. Move aircraft as
soon as tires are free.
2-92. Application of External Power.
2-93. Servicing Oxygen System.
The oxygen system furnishes breathing oxygen to
the pilot, copilot and observer. Refer to Section VII for
location of oxygen cylinders.
a. Oxygen System Safety Precautions.
CAUTION
WARNING
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. The
GPU shall not exceed 28 VDC.
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. A GPU capable of delivering a continuos load
of 1000 amperes is required.
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 must
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.
Change 7 2-79
TM 55-1510-215-10
b. Replenishing Oxygen System.
1. Remove protective cap on oxygen cylinder
filler valve.
2. Attach oxygen hose from oxygen servicing
unit to filler valve.
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:
WARNING
If the oxygen cylinder pressure is
below 50 PSI, do not attempt to
service system. Make an entry on DA
Form 2408-13.
3. Insure that supply cylinder shutoff valves
on the aircraft are open.
4. Fill system slowly to prevent overheating
by adjusting recharging rate with pressure regulating
valve on oxygen servicing unit.
5. Close pressure regulating valve on
oxygen servicing unit when pressure gage on oxygen
cylinder indicates the pressure obtained from the oxygen
system servicing chart (fig. 2-24).
6. Disconnect oxygen hose from oxygen
servicing unit and filler valve.
7. Install protective cap on oxygen filler
valve.
2-94. Ground Handling.
Ground handling covers all the essential information
concerning movement and handling 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 standards and recognized ground
handling procedures.
Carelessness or insufficient
knowledge of the aircraft or equipment being handled
can be fatal. The applicable technical manuals and
pertinent directives should be studied for familiarization
with the aircraft, its components, and the ground
2-80
1. Keep intake air ducts free of loose articles
such as rags, tools, etc.
2. Stay clear of exhaust outlet areas (fig. 8-2).
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 properly
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. Keep personnel clear of exhaust danger
area.
12. 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:
TM 55-1510-215-10
OXYGEN SYSTEM SERVICING PRESSURE
Figure 2-24. Oxygen System Servicing Pressure
2-81
TM 55-1510-215-10
(1) Taxiing.
with chapter 8.
Taxiing shall be in accordance
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 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.
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 (fig. 2-25). In the event towing lines are
necessary, use towing lugs on the main landing gears.
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
2-82
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 aircraft are
covered by the various phases of the ground handling
procedures included in this section of general ground
handling instructions. (Refer to TM 55-1500-204-25/1.)
2-95. 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 shocking the wheels to the
more elaborate mooring procedures described in
paragraph 2-103. The proper steps for securing the
aircraft must be based on the time the aircraft will be left
unattended, the aircraft weight, 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 pilot's toe 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.
NOTE
Parking brake cannot be set by using
copilots brake pedals.
3. Release pilot's toe brake pedals.
TM 55-1510-215-10
Figure 2-25. Parking, Covers, Ground Handling and Towing Equipment
2-83
TM 55-1510-215-10
CAUTION
Do not set parking brakes when the
brakes are hot during freezing
ambient temperatures. Allow brakes
to cool before setting parking brakes.
1. Use mooring cables of 1/4 inch aircraft
cable and clamp (clip-wire rope), chain or rope 3/8 inch
or over. 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.
4. To release the parking brakes push in on
the parking brake handle.
CAUTION
b. The control lock (fig. 2-16) 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:
Do not use slip knots. Use bowline
knots to secure aircraft to mooring
stakes.
1. With engine and propeller control levers in
secure position, slide lock onto control pedestal to
prevent operation of levers.
2. Install elevator and aileron lock pin
vertically through pilot's control column to lock control
wheel.
3. Install rudder lock pin through right hand
pilot's rudder pedal; then neutralize and lock pedals
together.
4. Reverse steps 1 through 3 above to
remove control lock. Store control lock.
2-96. Installation of Protective Covers.
While in transit, the crew will insure that the aircraft
is protected during inclement weather. If protective
covers (fig. 2-25) are not on board, steps will be taken to
procure them from the airfield maintenance facility.
2-97. 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. 2-26)
are provided beneath the wing, nose, tail and on each
main 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.
2-84
2. One-piece wheel chocks or wood blocks
may be used to chock the main landing gear wheels.
They must be equipped with rope or wood cleats to
retain them against the wheels.
NOTE
In ice or snow conditions, collapsible
ice grip wheel chocks should be
used. However, sandbags may be
used if collapsible chocks are not
available or if parking or mooring the
aircraft on steel mats.
b. Mooring Procedures for High Winds.
If an
aircraft is to remain securely moored during high velocity
winds, it is necessary to use the proper size and type of
wheel chock. Since the factor of weight is significant in
determining
adequate
mooring
provisions,
the
approximate weight must be known if the aircraft is to be
properly secured. During emergencies, knowledge of
this information is very useful in selecting the aircraft that
should be tied down first, as a heavy aircraft will better
stand high winds than an empty aircraft.
CAUTION
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.
TM 55-1510-215-10
Figure 2-26. Mooring the Aircraft
2-85
TM 55-1510-215-10
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
(wood) together with rope 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.
5. Accomplish aircraft tiedown by utilizing
mooring points (fig. 2-26). Make tiedown with 1/4 inch
aircraft cable, using two wire rope clips, or bolts, and a
chain 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. When
anchor kits are not available, use metal stakes or deadman type anchors, providing they can successfully
sustain a minimum pull of 3000 pounds.
6. In event nose position tiedown is
considered to be of doubtful security due to existing soil
condition, drive additional anchor rods at nose tiedown
2-86
position. Place padded work stand or other suitable
support under the aft fuselage tiedown position and
secure.
7. Place control surfaces in locked position
and trim tab controls in neutral position. Place wing
flaps in up position.
8. 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. 2-25).
9. Secure propellers to prevent windmilling
(fig. 2-25).
10. Disconnect battery.
NOTE
Where
typhoon
or
hurricane
conditions exist, it is well to
remember that the storm appears to
pass two times, each time with a
different wind direction. This will
necessitate turning the aircraft after
the first passing.
11. 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.
12. After high winds, inspect aircraft for visible
signs of structural damage and for evidence of damage
from flying objects.
Service nose shock strut and
reconnect battery.
TM 55-1510-215-10
CHAPTER 3
AVIONICS
SECTION I. GENERAL
3-1. Introduction.
This
chapter
covers
avionics
equipment
configuration installed in the aircraft. It includes a brief
description of the avionics equipment, its technical
characteristics, capabilities, and location.
3-2. Avionics Equipment Configuration.
The avionics configuration of the aircraft is
comprised of three groups of electronic equipment. The
communication equipment group consists of the radio
telephone (if installed), interphone, FM liaison (if
installed), UHF command, VHF command, and HF
command (if installed) systems.
The navigation
equipment 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, 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. A ground
proximity altitude advisory system is also provided.
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. 2-22) on the pilot's side of
the instrument panel. Individual system circuit breakers
and associated avionics busses are shown in figure 2-18
and figure 2-19. With the AVIONICS MASTER power
switch in the ON position, the avionics power relay is
deenergized and power is applied through two 50ampere AVIONICS NO. 1 and 2 circuit breakers to the
individual avionics circuit breakers on the circuit breaker
panel (fig. 2-18). In the OFF 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 to the avionics equipment.
NOTE
The avionics master switch and radar
should not normally be turned on,
until the generators are on. This will
help protect the solid-state circuitry.
b. AC Power.
AC power for the avionics
equipment is provided by two separately selected
inverters. The inverters supply 115-volt and 26 volt
single-phase AC power. The inverters are selected by a
switch located on the left subpanel, placarded
INVERTER LEFT OFF RIGHT. Either inverter may be
selected for use.
SECTION II. COMMUNICATIONS
3-4. Description.
The communication equipment group consists of
radio telephone (if installed), interphone, FM liaison (if
installed), UHF command, VHF command, and HF
command (if installed) systems.
copilot's seats: and a pushbutton switch located in each
headset microphone jack (pigtail).
b. Controls and Functions.
(1) INPH-XMT MIC switch.
or selected facility.
Keys interphone
3-5. Microphone Switches.
a. Description. Three microphone switches are
provided for the pilot and copilot: A bi-level microphone
switch placarded INPH XMT MIC, located on the pilot's
and copilot's control wheels; a floor switch placarded
MIC located on the floor in front of the pilot's and
(a) INPH.
When depressed to first
detent, keys interpone regardless of audio panel selector
switch position.
(a) XMIT MIC. When depressed fully,
keys facility indicated on audio panel selector switch.
Change 9 3-1
TM 55-1510-215-10
3-6. Interphone System.
affects intercom volume only. It does not affect volume
of other audio inputs selected with the pushbuttons.
a. Description. Identical audio control panels are
provided for the pilot and copilot. The controls and
switches on each panel provide for audio reception of
interphone, communication, navigation audio signals,
and a choice of transmission on VHF, UHF, and HF (if
installed). Cabin speakers and crew headphones are
utilized for audio reception. Aircraft power for speaker
and headphone isolation amplifiers is derived from
separate sources to provide a high degree of audio
integrity. The power is routed through 3-ampere circuit
breakers placarded SPKR 1 and SPKR 2, and 1-ampere
circuit breakers placarded PHONE 1 and PHONE 2,
located on the lower right sub panel.
b. Audio Panel (KMA 24H).
The controls and
switches on the KMA 24H audio control panel (fig. 3-1)
provide for selection and volume control of interphone,
communication, and navigation audio signals.
c. Controls and Functions.
(1) INT VOL control. Volume of the intercom
input is controlled by the INT VOL control. This control
1.
2.
3.
4.
(2) Speaker audio select buttons. Selects
desired audio inputs to be heard on the speakers.
(3) MIC SELECT switch. The MIC SELECT
switch performs several functions. It routes microphone
audio and keying to the appropriate system and switches
the speaker amplifier output to the appropriate speaker.
In the 1, 2, 3, or 4 positions, microphone audio and
keying are routed to the appropriate transceiver and the
speaker amplifier output is connected to the cockpit
speakers. In the PA position, microphone audio is
routed to the speaker amplifier. The speaker amplifier
output is connected to the passenger address speakers.
In the INT position, microphone audio is routed to the
intercom output. In INT, 1, 2, 3, or 4 positions, any audio
output selected on the top row of pushbuttons will be
heard through the cabin speakers. Keying causes the
output audio to be muted and the sidetone audio to be
heard.
communicate
headphones.
(a) INT.
Permits the flight crew to
with
cabin
occupants
through
INT VOL switch
SPEAKER audio select buttons
MIC SELECT switch
PHONE audio select buttons
AP011655
Figure 3-1. Audio Control Panel
3-2 Change 5
TM 55-1510-215-10
(b) #1. Selects COMM 1 position.
(c) #2. Selects COMM 2 position.
(d) #3. Selects UHF position.
(e) #4. Selects HF radio position (if
installed).
(f) PA.
Permits the flight crew to
address aft cabin occupants over the passenger address
speakers.
(4) Headphone audio select buttons. Selects
audio input to heard on the headphones.
d. Operating Procedures.
1.
2.
3.
4.
AVIONICS MASTER switch - ON.
MIC SELECT switch - As required.
Audio select button - As required.
INT VOL - Adjust as required if INT is
selected.
3-7. FM Liaison Set (AN/ARC-131) (If installed).
The FM liaison set (fig. 3-2) is a radio transceiver
which provides two-way frequency modulated (FM)
communications in the 30.00 to 75.95 MHz range for a
distance of approximately 50 miles. The FM liaison set
is used for voice communications and FM homing. The
audio output is applied to the respective audio control
panel where it is made available to the headsets. Power
for the FM liaison set is fed through 50-ampere
AVIONICS NO. 1 and AVIONICS NO. 2 circuit breakers,
located on the copilot's circuit breaker panel. The unit is
protected by a 10-ampere circuit breaker, placarded FM,
located on the copilot's subpanel.
a. Controls and Functions.
(1) Megahertz selector. Tunes transceiver in
10-MHz, as indicated by the first digit of the frequency
indicator.
(2) Frequency indicator.
to which transceiver is tuned.
Indicates frequency
(3) Megahertz selector. Tunes transceiver in
1-MHz, as indicated by the second digit of the frequency
indicator.
(4) Kilohertz selector. Tunes transceiver in
100-kHz, as indicated by the third frequency indicator.
1.
2.
3.
4.
5.
6.
7.
8.
Megahertz (10) selector
Frequency indicator
Megahertz (1) selector
Kilohertz (100) selector
Kilohertz (10) selector
Mode selector
VOL control
SQUELCH switch
AV 094957
Figure 3-2. FM Liaison Set (AN/ARC-131)
Change 5 3-3
TM 55-1510-215-10
(5) Kilohertz selector. Tunes transceiver in
50-kHz, as indicated by the fourth frequency indicator.
(6) Mode selector.
Determines operating
mode.
(a) OFF. Turns set off.
(b) T/R.
Provides for transceiver
operation of frequency displayed on frequency indicator.
(c) RETRAN.
Not
used
in
(3) PLAIN indicator.
Illuminates when the
PLAIN/CIPHER switch is in the PLAIN position.
(4) PLAIN/CIPHER switch. Controls unciphered
or ciphered communications.
(a) PLAIN.
Permits
communications on the FM liaison set.
unciphered
(b) CIPHER.
Permits
communications on the FM liaison set.
ciphered
this
installation.
(5) RE-X/REG
switch.
Controls
retransmission, or cipher/normal communications.
(d) HOME. Provides for operation in the
homing mode. May also be operated as a transceiver on
channels indicated on frequency indicator.
(7) VOL control. Adjusts volume.
(a) RE-X.
Permits retransmission of
ciphered communications at a distant location.
(b) REG.
plain communications.
Permits normal cipher or
(8) SQUELCH switch. Controls squelch.
NOTE
(a) DIS.
Disables receiver squelch
circuit.
(b) CARR.
Activates receiver squelch
Do not place the ZEROIZE switch in
the ON position unless a crash or
capture is imminent.
circuits.
(c) TONE. Not used in this installation.
b. Operating Procedures. Refer to Voice Security
System operating procedures.
3-8. Voice Security System (TSEC/KY-28) (If installed).
a. Description. The voice security system is used in
conjunction with the FM liaison set to provide secure
(ciphered), two-way voice communications (fig. 3-3).
Power for the voice security system is fed through the 50ampere AVIONICS NO. 1 and AVIONICS NO. 2 circuit
breakers on the copilot's circuit breaker panel. The voice
security system is protected by a 10-ampere circuit
breaker, placarded FM, located on the copilot's subpanel.
(6) ZEROIZE switch.
Normally in OFF
position.
Placed in ON position during emergency
situations to neutralize and make inoperative the
associated cipher equipment.
(7) CIPHER indicator. Illuminates when the
PLAIN/CIPHER switch is in the CIPHER position.
c. Operating Procedures for FM Liaison Set and
Voice Security System.
NOTE
Disregard the operating procedures
involving
the
voice
security
(CIPHONY) control-indicator if this
unit is not installed.
b. Controls and Functions.
NOTE
(1) POWER ON switch. Turns set on or off.
NOTE
Switch must be in the ON position for
FM liaison operations in either the
plain or cipher mode.
(2) POWER ON indicator. Illuminates when
the POWER ON switch is placed in the up (on) position.
3-4 Change 5
The
audio
automatically
receiver audio
whenever the
transmitter is
position.
discriminators
will
interrupt
the
FM
at a given position
UHF, VHF, or HF
keyed from that
(1) Turn-on procedure. POWER ON switch
(CIPHONY control-indicator) - ON.
TM 55-1510-215-10
1.
2.
3.
4.
5.
6.
7.
POWER ON switch
POWER ON indicator
PLAIN indicator
PLAIN-CIPHER switch
RE-X-REG switch
ZEROIZE switch
CIPHER indicator
AV 094958
Figure 3-3. Voice Security Control-Indicator
NOTE
2. PLAIN/CIPHER switch (CIPHONY
control-indicator) - PLAIN.
The POWER ON switch must be in
the ON position regardless of the
mode of operation, whenever the
voice security (CIPHONY) KY-28 is
installed in the aircraft.
3. Microphone switch - Press.
(4) Transmitter
(CIPHER).
(2) Receiver Operating Procedure.
1. SQUELCH switch (FM
control panel) -As required.
operating
procedure
1. Transmitter-interphone
selector
(audio control panel) - No. 1 position.
COMM
2. Mode selector (FM COMM control
panel) - T/R or PTT.
3. Frequency selectors (FM COMM
control panel) - As required.
(3) Transmitter operating procedure (PLAIN).
1. Transmitter-indicator selector (audio
control panel) - No. 1 position.
2. PLAIN/CIPHER switch (CIPHONY
control-indicator) - CIPHER.
3. RE-X REG switch (CIPHONY
control-indicator) - As required (REX position only if distant station is
using retransmitting equipment).
4. Microphone
switch
Press
momentarily (interrupted tone from
voice security unit should no longer
be heard).
Change 5 3-5
TM 55-1510-215-10
NOTE
3-9. UHF Command Set (AN/ARC-51BX) (If installed).
No traffic will be passed if the
interrupted tone is still heard from
pressing
and
releasing
the
microphone switch.
a. Description. The UHF command set (fig. 3-4)
provides two-way amplitude modulated (AM) voice
communication within the frequency range of 225./0 to
399.95 MHz for a distance range of approximately 50
miles line-of-sight. Channel selection is spaced at 50
kHz intervals.
Additionally, a separate receiver is
incorporated to provide monitoring capability for the UHF
guard frequency (243.0 MHz). The audio output of the
UHF set is applied to the audio control panel where it is
made available to the headsets. Power for the UHF
command set is fed through the 50-ampere AVIONICS
NO. 1 and AVIONICS NO. 2 circuit breakers on the
copilot's circuit breaker panel. The set is protected by a
15-ampere circuit breaker, placarded UHF, located on
the copilot's subpanel.
5. Microphone switch - Press (do not
talk). Wait until beep is heard, then
speak into microphone.
(5) Homing operating procedure.
NOTE
Accuracy of the FM liaison set has
been verified on the following
frequencies only: 32.00, 34.50, 38.90,
and 39.50 MHz.
1. Mode selector (FM COMM control
panel) - HOME.
2. Frequency selectors (FM COMM
control panel) - Select.
3. VOL control (FM COMM control
panel) - As required.
4. Pilot or copilot's COURSE IND
switch (instrument panel) - HOME.
NOTE
FM homing steering signals can be
displayed only on one course
indicator-selector) at a time.
5. Course deviation indicator (pilot or
copilot's course indicator-selector) Read lateral deviation from horning
course.
6. Glideslope
indicator
(pilot
or
copilot's course indicator-selector) Read relative signal strength.
Indicator will deflect towards the
bottom for a weak signal and move
toward the center position as the
signal strength increases.
b. Controls and Functions.
(1) Preset channel indicator. Indicates the
preset channel in use when the mode selector is in the
PRESET CHAN position.
(2) SQ DISABLE switch.
Turns squelch
circuit on or off.
(3) VOL. Control. Adjusts volume.
(4) Function switch. Selects type of operation
and turns set off.
(5) Manual frequency selector. Selects the
operating frequency when the mode selector is in the
MAN position.
(6) Frequency.
Indicates the operating
frequency of the UHF receiver and transmitter when the
mode selector is in the MAN position.
(7) Mode selector.
Determines operation
mode.
(a) PRESET CHAN. Permits selection
of one of 20 preset channels.
(b) MAN. Permits frequency selection
by means of manual frequency selector controls.
(c) GD XMIT. Automatically tunes to
the guard channel frequency.
(8) PRESET CHAN selector.
Selects the
desired preset channel when the mode selector is in the
PRESET CHAN position.
(6) Shutdown procedure.
1. Mode selector (FM COMM control
panel) - OFF.
2. POWER ON switch (CIPHONY
control-indicator) - OFF.
3-6 Change 5
c. UHF Command Set Operation. The function
switch must be ON for the following procedures.
(1) Receive operating procedure.
1. Mode selector - As required.
TM 55-1510-215-10
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
Preset Channel indicator
SQ DISABLE switch
VOL control
Function switch
Decimal MC selector
MC indicator
One MC selector
Ten MC selector
Tuning mode switch
PRESET CHAN selector
AV 095264
Figure 3-4. UHF Command Set (AN/ARC-51BX)
2. Frequency
Select
required
frequency using either preset
channel control or manual frequency
selectors.
NOTE
The preset CHAN selector and
MANUAL frequency selectors are
inoperative when the mode selector
is set to GD XMIT position.
(2) Transmit operating procedures.
1. Transmitter-interphone
selector
(audio control panel) - No.2 position.
2. Microphone switch - Press.
(3) Shutdown procedure.
1. Function switch - OFF.
d. UHF Command Set Emergency Operation.
3. Volume - Adjust.
NOTE
NOTE
To adjust volume when audio is not
being received, turn squelch disable
switch ON, adjust volume for
comfortable noise level, then turn
squelch disable switch OFF.
4. Squelch - As Required.
Transmission
on
emergency
frequency (guard channel) shall be
restricted to emergencies only. An
emergency frequency of 121.500 MHz
is also available on the VHF (KY-196)
COMM-1 and/or COMM-2 radio sets.
1. Transmitter - interphone selector
(audio control panel) - No. 2
position.
Change 5 3-7
TM 55-1510-215-10
2. PLAIN/CIPHER switch (CIPHONY
control-indicator) - PLAIN.
3. Mode selector (UHF control panel) GD XMIT.
4. Microphone switch - Press.
3-10. VHF Command Sets (KY 196).
position and will increase volume by turning clockwise.
This knob also overrides the automatic squelch by
pulling it out for audio test.
The radios are protected by two 7 1/2-ampere circuit
breakers placarded COMM VHF 1 VHF 2, located on the
copilot' subpanel (fig. 2-7).
b. Controls and Functions.
a. Description. Identical VHF command sets are
provided for the pilot and copilot, COMM 1 and COMM 2
respectively (fig. 3-5). The KY 196 transceiver is a
communications transceiver designed to provide twoway voice communication within the frequency range of
118.000 MHz to 135. 975 MHz, in 25 kHz increments.
Active and standby frequencies may be transferred at
anytime, by pressing a frequency transfer button. The
frequency is tuned by the selector on the right side of the
unit. It may be tuned in ascending or descending order,
by rotating the selector clockwise or counterclockwise.
The outer larger selector knob is used to change the
MHz portion of the frequency while the smaller knob
changes the kHz portion. The 5 kHz units will not be
indicated in the readout display.
(1) Active frequency display.
frequency being used.
Displays the
(2) Transmit symbol. A letter T will appear
between the USE and STANDBY windows when the
transceiver is in the transmit mode of operation.
(3) Standby frequency display.
frequency selected for standby.
(4) Frequency selector
frequencies in MHz and kHz.
(a) MHz
knob.
Displays the
knobs.
Selects
Selects
MHz
frequencies.
(b) kHz knob. Selects kHz frequencies.
The OFF-PULL TEST control knob turns power off to
the unit by rotating counterclockwise to the detent
1.
2.
3.
4.
5.
6.
Active frequency display
Transmit symbol
Standby frequency display
Frequency selector knobs
OFF-PULL TEST switch
Frequency transfer button
AP011858
Figure 3-5. VHF Control Panel
3-8 Change 5
TM 55-1510-215-10
(5) OFF-PULL TEST switch. When turned
clockwise. turns set on. When pulled out, provides for
manual squelch control.
(6) Frequency transfer button.
active and standby frequencies.
display. the remote KAC 952 power amplifier/antenna
coupler and KTR 953 receiver/exciter. The system
operates on any 0.1 kHz frequency between 2.000 and
29.999.9 kHz.
Transfers
c. Mode of Operation.
(1) Frequency mode. In the frequency mode
of operation frequencies are entered into the STBY
window of the display and may then be transferred into
the USE window by pressing the frequency transfer
button. To select frequencies the MHz portion of the
frequency displayed in the STBY window is incremented
or decremented in 1 MHz steps by rotating the MHz
knob either clockwise or counterclockwise.
MHz
frequencies will roll over or roll under at each band edge
(118 or 135 MHz). The kHz portion of the frequency
displayed in the STBY window may be incremented or
decremented by rotating the 25 kHz knob either
clockwise or counterclockwise. Frequency selection is in
50 kHz steps when the 25 kHz knob is pushed in and is
in 25 kHz steps when the 25 kHz knob is pulled out.
Frequencies roll over from 95 or 97 to 00 and roll under
from 00 to 95 or 97 according to the position of the 25
kHz knob. Frequencies are transferred from STBY to
USE window and vice versa by depressing the
frequency transfer button.
d. Operating Procedures.
1. AVIONICS MASTER switch -ON.
2. OFF-PULL TEST switch -Turn
clockwise out of detent.
The
frequencies displayed will be those
last used.
3. Audio select button (audio control
panel) - Select as required.
4. OFF-PULL TEST switch -Pull out
and rotate until the desired audio
level is obtained. Push the knob
back in to activate the automatic
squelch.
5. Active frequency - Insure desired
active frequency is displayed in USE
window.
6. Microphone switch - Press to
second detent.
e. Shutdown Procedure. Turn OFF-PULL TEST
switch counterclockwise to the OFF position.
3-11. HF Communication Set (KHF 950).
a. Description. The KHF 950 HF system consists
of three units; the pedestal mounted KCU 951 control
With the capability to preset and store 99
frequencies for selection during flight, the system also
allows for selection of other frequencies manual (direct
tuning). or reprogramming of any preset frequency. The
system automatically matches the antenna by keying the
microphone. Power to the system is routed through a
25-ampere circuit breaker. placarded HF, located on the
copilot's circuit breaker panel.
b. Controls and Functions.
(1) FREQ
display.
Displays
frequency
selected.
(2) Mode display. Displays LSB, AM or USB
mode of operation as selected.
(3) CHANNEL display. Displays from 1 to 99
channels as selected.
(4) Light sensor.
The light sensor is a
photocell which automatically adjusts brightness of the
display.
(5) MODE switch. The mode switch is a
momentary pushbutton switch that selects LSB, AM or
USB modes of operation.
(6) FREQ/CHAN switch. Transfers the HF
system from a direct frequency operation to a
channelized form of operation.
(7) PGM (program) recessed switch. Enables
channelized data to be modified. The PGM message
will be displayed whenever this switch is depressed.
(a) Program. The program mode must
be used for setting or changing any of the 99 preset
frequencies. Each of the 99 channels may be preset to
receive and transmit on separate frequencies (semiduplex). receive only. or transmit and receive on the
same frequency (simplex). The operating mode (LSB,
USB or AM) must be the same for both receive and
transmit and can also be preset.
(8) Frequency/channel selector. This selector
consists of two concentric knobs that control the channel
and frequency digits. plus the lateral position of the
cursor.
(a) Frequency control. The outer knob
becomes a cursor (flashing digit) control with the
FREQ/CHAN switch in the FREQ position. The flashing
digit is then increased/decreased with the inner knob.
(b) Channel control. The outer knob is
not functional when the FREQ/CHAN switch is in the
CHAN position. The inner knob will provide channel
control from 1 through 99, displayed at the right end of
the display window.
Change 5 3-9
TM 55-1510-215-10
(9) STO (store) recessed switch.
Stores
displayed data when programming preset channels.
(10) OFF-VOLUME. Applies power to the unit
and controls the audio output level.
(11) SQUELCH.
Provides variable squelch
threshold control. Squelch is set by rotating the knob
clockwise until background noise is heard then
counterclockwise until the noise is just barely audible or
absent. Squelch operation on HF is not as predictable
as is VHF.
Readjustment may be required more
frequent.
(12) CLARIFIER. Provides 250 Hz of local
oscillator adjustment, thus providing better SSB
reception when a ground station is slightly off frequency.
When receiving an unclear transmission, pull out the
knob and rotate in either direction to give a natural voice
quality. Then push the knob in, the clarifier control
should be left in the inner position.
c. Methods of Operation (Frequency Selection).
The HF system has two methods of operation The first
method is called direct tuning (frequency agile). The
second method is a channelized operation in which
desired operating frequencies are preset, stored and
referenced to a channel number.
(1) Direct tuning (simplex only). Each digit of
the frequency may be selected instead of dialing up or
down to a frequency. The larger concentric knob is used
to select the digit to be changed. This digit will flash
when selected. Rotation of the knob moves the flashing
cursor in the direction of rotation. After the digit to be
changed is flashing, the smaller concentric knob is used
to select the numeral desired. This process is repeated
until the new frequency has been selected. The flashing
cursor may then be stowed by moving it to the extreme
left or right of the display and then one more click. This
stows the cursor behind the display until needed again.
The cursor may be recalled by turning the concentric
knob one click left or right.
(2) Channel programming. The long distance
propagation of HF signals depends on such factors as
atmospheric conditions, conditions in the ionosphere, the
time of day, and the frequency being used. Therefore,
whenever possible, the 99 preset channels should be
chosen so that communications with each of several
stations along the route is possible on three or more
frequencies spaced out well across the HF band. Then
if there is some difficulty in communicating with a station
on one frequency, other frequencies that station is
guarding may be tried without having to set up a
frequency digit by digit.
3-10 Change 5
There are three ways to set up a channel: Receive
only, simplex, and semi-duplex. To gain access to
channelized operation, depress FREQ/CHAN button. To
utilize the existing programmed channels (i.e.
no
programming required) use the small control knob to
select the desired channel number. Then momentarily
key the microphone to tune the antenna coupler. If
channel programming is required, it is necessary to
activate the program mode as follows.
With the
FREQ/CHAN button in (CHAN), use a pencil or other
pointed object to push the PGM button in. The button is
an alternate action switch: push-on, push-off. The letters
PGM will appear in the lower part of the display window
and the system will remain in the program mode until the
PGM button is pressed again.
(a) Receive only.
1. Stow the cursor if a frequency digit
is flashing.
2. Select the channel to be preset.
3. Set the desired operating mode
(LSB,USB,or AM).
4. Set the desired frequency (refer to
direct tuning).
5. Push and release STO button once.
NOTE
T will flash in the display window,
however a receive only frequency is
being set. The flashing T should be
ignored.
NOTE
If another channel is to be set, the
cursor must be stowed before a new
channel can be selected. Use the
smaller concentric knob to select the
channel and repeat the steps for
selecting a new frequency.
6. To return to an operating mode,
push the PGM button.
(b) Simplex. Setting a channel up for
simplex operation (receive and transmit on the same
frequency).
1. FREQ/CHAN button in (cursor
stowed).
2. PGM button in (PGM displayed).
3. Select channel to be preset.
TM 55-1510-215-10
4. Set mode (LSB,USB or AM).
5. Set desired frequency (refer to
direct tuning).
6. Push and release STO button twice.
NOTE
The first press of the STO button
stores frequency in the receive
position and the second press stores
the same frequency in the transmit
position.
The second push also
stores the cursor.
1. AVIONICS MASTER switch - ON.
2. OFF-VOLUME
switch
Turn
clockwise out OFF position. Adjust
volume desired.
3. Squelch - Adjust as desired.
4. Clarifier - Adjusts desired.
5. Mode of operation - Select.
6. Frequency/channel
Set/select
program as desired.
e. Shutdown Procedures.
NOTE
1. OFF-VOLUME switch-OFF.
2. AVIONICS MASTER switch -OFF.
If another channel is to be reset, use
the smaller concentric knob to select
the channel and repeat the steps for
selecting a new frequency.
The
cursor was automatically stowed. To
return to one of the operating modes,
push the PGM button again.
3-12. Radio Telephone (KT 96)(Optional).
(c) Semi-duplex. Setting a channel for
semi duplex (transmit on one frequency and receive on
another)
1. Select channel to be preset.
2. Set desired frequency
direct tuning).
d. Operating Procedures.
(refer
to
a. Description. The radio telephone is installed
directly behind the copilot's seat. Its 10 watts of UHF
transmitting power provides full line-of-sight range.
Range depends on the aircraft’s altitude. A range of 110
miles is typical at an altitude of 1000 feet. The unit
incorporates 12 two-way channels, plus one signalling
channel on which the telephone listens for any incoming
calls. The radio telephone rings incoming calls through
the cabin speakers. If the call is not answered after forty
five seconds, the radio telephone stops ringing and goes
back to listening on the signalling channel. Power is
provided to the unit through a 5-ampere circuit breaker,
placarded TELEPHONE, located on the copilot’s circuit
breaker panel.
3. Set mode (LSB,USB, or AM).
b. Controls and Functions.
4. Push STO button once.
5. Set transmit frequency.
6. Push STO button again.
If another channel is to be reset, use the smaller
concentric knob to select the channel and repeat the
steps.
7. To return to operating mode, push
the PGM button.
NOTE
The mode for each channel (LSB,
USB AM) is stored along with the
frequency If the mode is changed, the
system will receive and transmit in
the mode selected for transmit.
(1) Volume control knob. The volume (VOL)
control knob allows for adjustment audio.
(2) Indicator light. The red indicator light will
illuminate when the handset is removed from the cradle,
and the CHAN SEL button is pressed.
(3) Hang up button. Pressing the HANG UP
button disconnects any incoming calls. If you are unable
to answer a incoming call, press the HANG UP button.
This however, will require the caller to replace his call.
(4) Channel display window.
Displays a
channel number as selected by the channel selector
knob. Channel numbers are derived from a map of
ground-air telephone facilities.
Change 10 3-11
TM 55-1510-215-10
Corporation. Although, is easier to use the ground
station map for reference, you can locate a receivable
station by rotating the channel selector knob listening to
each channel until a dial tone is heard.
(5) Channel selector knob.
The channel
selector knob is used to position the desired channel
number in the channel display window.
1. Handset - Remove from cradle.
2. CHAN SET button -Press.
3. Channel selector knob -Rotate to select
desired channel number in channel display window and
listen for a dial tone.
NOTE
(6) Cradle. The cradle is used to store the
handset. It incorporates a spring loading feature, which
keeps the handset firmly in place. Placing the handset
in the cradle will disconnect calls.
(7) Handset.
The handset
transmitting and receiving calls.
provides
for
(8) Press-to-talk button. For proper operation
the press-to-talk button must be pressed while talking
into be handset Release the button while listening
(receiving). The press-to-talk button is also used to
summon the ground station operator when placing a
radio telephone call.
If a ground service station map is not
available, rotate through the channel
numbers until a dial tone is received.
Press-to-talk button - Press momentarily to summon
operator. When the operator responds, place call giving
your city of registry, your phone number and the number
you wish call.
(2) To answer a call.
1. Handset - Remove from cradle.
2. Press-to-talk
answer call
(3) To end a call.
c. Operating Procedures.
(1) To make a call. First check your ground
service station map to determine the channel number of
the ground station within range, then proceed as follows:
button
-Press
and
1. Handset - Place in cradle, or press
HANG UP button.
Section III. NAVIGATION
3-13. Description.
The overall navigation equipment 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 under instrument
meteorological conditions (IMC).
The navigation
configuration includes equipment for determining
altitude, attitude, position, destination range and bearing,
heading reference, groundspeed, and drift angle.
3-13A. Emergency ocator Transmitter (ELT).
a. Description. An ELT, if installed, is provided to
assist in locating the aircraft and crew in the event an
emergency landing is necessitated.
The output
frequency is 121.5 and 243 MHz simultaneously. Range
approximately line-of sight. The transmitter unit has a 3position toggle switch, placarded AUTO-OFF-ON,
located on one end the case.
The transmitter is
accessible though the lower tail section access door.
b. Controls and Functions.
(1) AUTO. Arms set to be actuated by impact
switch (NORMAL mode).
(2) OFF. Turns set off.
(3) ON. Manually activates set. (fig. 3-5A).
3-14. Radio Magnetic Indicators (RMI).
a. Description. Identical KNI 582 RMI indicators
are installed for the pilot and copilot (fig. 3-6). Each
RMI provides aircraft heading and radio bearing
information to or from a VOR, ADF facility or RNAV
waypoint.
Selector switches on the RMI provides
selection of either VOR or ADF as the information be
displayed on either or both needles. The single needle
points to either a tuned NAV 1 VOR or ADF while the
double points to either a tuned NAV 2 VOR or ADF
navaid. Both RMI's are protected by 1-ampere circuit
breakers located on the DC J-BOX in the nose avionics
compartment.
b. Controls and Functions.
(1) Warning flag. Indicates loss of heading
signal, or that bearing information is unreliable.
3-12 Change 10
TM 55-1510-215-10
Figure 3-5A. Emergency Locator Transmitter (ELT).
Change 10 3 - 12.1/(3-12.2 Blank).
TM 55-1510-215-10
1.
2.
3.
4.
5.
6.
7.
HDG warning flag
Compass card
Heading index
Double needle pointer
Single needle pointer
Double needle switch
Single needle switch
AP011857
Figure 3-6. Radio Magnetic Indicator (RMI)
(2) Compass card. Gyro stabilized to indicate
aircraft heading and bearing information.
(3) Heading index.
aircraft heading.
Reference point for
(4) Double needle point. Indicates ADF or
NAV 2 VOR bearing as selected by double needle
switch.
(5) Single needle pointer. Indicates ADF or
NAV 1 VOR bearing as selected by single needle switch.
(6) Double needle switch. Selects desired
signal to be displayed on double needle pointer.
(a) ADF position. Selects ADF bearing
3-15. Gyromagnetic Compass Systems.
a. Description.
The aircraft incorporates two
gyromagnetic compass systems. The pilot's compass
system consists of the following components: KSG 105
slaved directional gyro, KMT 112 magnetic flux valve
and a KA 51B slave control (fig. 3-7) (located on left
side of pilot's instrument panel). This system provides
primary heading data to the pilot's HSI (KPI 553) and
copilot's RMI (KNI 582).
The copilot's compass system consists of the
following components: KG 102 directional gyro, K 112
magnetic flux valve and a KA 51B slave control
(fig. 3-7) (located on right side of copilot's instrument
panel). This system provides primary heading data to
the copilot's HSI (KI 525A) and pilot's RMI.
information.
(b) NAV position. Selects VOR bearing
information.
(7) Single needle switch.
Selects desired
signal to be displayed on single needle pointer.
(a) ADF position. Selects ADF bearing
information.
(b) NAV position. Selects VR or RNAV
For heading reference, two separate modes of
operation (FREE/SLAVED) are provided, as selected
utilizing the KA 51B slave control. In areas where
magnetic references are reliable, the systems should be
operated in the SLAVE mode.
In this mode, the
directional gyro is slaved to the magnetic flux valve
which supplies magnetic reference for correction of the
apparent drift of the gyro. In FREE mode, the systems
are operated as a free gyro. In this mode, latitude
corrections are manually introduced using the CCW
(counterclockwise) or CW (clockwise) switch.
b. Controls and Functions.
(1) Slave meter.
Indicates difference
between sensed heading and displayed heading. A
positive deflection indicates a clockwise error of the
compass card.
Change 10 3-13
TM 55-1510-215-10
1.
2.
3.
Slave meter
SLAVE/FREE gyro switch
CCW/CW switch
Figure 3-7. Gyromagnetic Compass Slave Control
(2) SLAVE/FREE gyro switch. When placed
in SLAVE position, the system is in the slaved gryo
mode. When placed in FREE position, the system is in
the free gryo mode.
(3) CCW/CW switch. Placing the switch in
the CCW position, when the system is in the gyro mode,
causes the compass card to rotate counterclockwise.
When the switch is placed in the CW position, the
compass card rotates in the clockwise direction.
3-16. Pilot's Horizontal Situation Indicator.
a. Description.
The pilot's horizontal situation
indicator (HSI) (fig. 3-8) combines numerous displays to
provide a presentation of the aircraft position. The
indicator displays the aircraft position relative to a VOR,
localizer, glide slope beam, RNAV waypoint, TACAN,
and selected heading, with respect to magnetic north.
The rotating heading dial is driven by the pilot's compass
system. Any warning flag in view indicates that portion
of the HSI display is unreliable.
In addition, the
horizontal situation indicator incorporates a DME readout
display. Readout brightness is automatically controlled,
with respect to cockpit ambient lighting, by a photo-cell.
b. Controls and Functions.
(1)
DME readout display. Provides
a digital display of DME information.
3-14 Change 10
(a) Distance. Distance to VORTAC to
RNAV waypoint is displayed on the left portion of the
display, indicated by the legend NM.
Distance is
indicated to the nearest tenth of a nautical mile, from 0 to
99.9 nautical miles, and to the nearest nautical mile,
from 100 to 389 nautical miles.
(b) Groundspeed.
Goundspeed is
shown by the middle portion of the display, indicated by
the legend KT It is indicated to the nearest knot from 0 to
99 knots. DME ground speed is only accurate when
flying a direct course to or from the VORTAC or RNAV
waypoint.
(c) Time-to-station.
Time-to-station
(TTS) is displayed by the right portion of the display,
indicated by the legend MIN. It is displayed to the
nearest minute, from 0 to 99 minutes, with 99 indicated
for any longer time-to-station. Time-to-station is only
accurate for a course directly to or from a VORTAC or
RNAV waypoint.
(d) Radar altitude.
Radar altitude is
shown as dashed lines on the middle display between
2500 and 1000 feet, and numerically to the nearest 10
feet from 990 to 0 feet. The appearance of the letters AL
on the right side of the display, and the blanking of KT
and MIN, indicate radar altitude information is being
displayed.
(e) DME controlling frequency source.
The digital display indicates the source of the frequency
information which is controlling the DME. Between the
left and middle displays, a 1 is displayed when the DME
XFER switch is set to NAV 1, likewise 2 when NAV 2 is
TM 55-1510-215-10
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
DME readout display
Lubber line
Compass flag
Course pointer
Rotating heading dial
To-From pointer
VOR/ADF switch
Heading knob
Heading bug
Course knob
Symbolic aircraft
Glideslope deviation pointer/scale
Course deviation bar and scale
NAV flag
Photocell
Bearing pointer
AP011855
Figure 3-8. Pilot's Horizontal Situation Indicator
selected. Switching to HOLD will store the channeling
information from the NAV frequency selector which was
previously selected. No frequency selector thereafter
affects the DME, and any frequency selector may then
be set to another frequency. The DME will remain on
the frequency being stored until NAV 1 or NAV 2 is
selected. 1H or H2 on the display indicates a HOLD
setting, as well as the NAV frequency selector whose
setting was held.
(f) RNV display.
The indicator will
display RNV when the displayed distance, groundspeed,
and time-to-station are derived from the area navigation
system. This RNV legend will flash when HOLD mode
has been selected. The display is blanked when RNV is
flashing.
(g) Photocell.
The
photocell
automatically controls readout brightness with respect to
cockpit ambient lighting.
the heading bug will maintain its position. The difference
between the bug and the fore lubber line represents the
heading select error applied to the flight director
computer.
(j) Compass .flag. When the compass
flag is in view, compass information is not valid.
(k) Course deviation bar and scale.
The yellow course deviation bar represents the
centerline of the selected VOR, RNAV, or localizer
course. This deviation is depicted as either linear or
angular deflection of the course deviation bar.
Angular deviation is displayed in the VOR or TACAN
modes. One dot deflection in the VOR or TACAN mode
equals 2-degree off course, whereas full scale or 5 dots
equals 10-degrees off course.
When a localizer
frequency is selected. one dot deflection equals 1/2degree deviation, whereas full scale or 5 dots equals 2
1/2-degree deviation off course.
(h) Lubber line. Fixed heading mark.
(i) Heading bug. The orange heading
bug is positioned on the rotating heading dial, by the
heading knob, to select and display a preselected
compass heading. Once set to the desired heading,
Linear deflection is displayed in the VOR RNV, VOR
RNV APR, TACAN RNV and TACAN RNV APR modes.
When using linear deflection, (excluding APR mode) one
dot deflection equals one nautical mile off course and
five dots equal five nautical miles off course. This
Change 5 3-15
TM 55-1510-215-10
deviation is true regardless of distance from the station,
therefore depicting linear instead of angular deviation.
The aircraft symbol pictorially shows actual aircraft
position in relation to this selected course.
(I) VOR/ADF switch. Selects NAV 1
VOR or ADF bearing information to be displayed by the
green VOR/ADF bearing bug.
deviation, and selected headings with respect to
magnetic north.
Deviation displacement is always
angular. Any warning flag in view indicates that portion
of the HSI display is unreliable. The rotating heading
dial is driven by the copilot's compass system.
b. Controls and Functions
(1) Lubber line. Fixed heading mark.
(m) Heading knob. Positions the orange
heading bug.
(n) Symbolic aircraft.
The fixed
miniature aircraft symbol corresponds to the longitudinal
axis of the aircraft and lubber line mark. The symbol
shows aircraft position and heading with respect to the
rotating heading dial, and aircraft position in relation to a
radio course.
(o) Course knob. Positions the yellow
course pointer.
(p) Glide slope deviation pointer/scale.
The glide slope pointer displays glide slope deviation.
The pointer is in view only when tuned to a localizer
frequency. If the aircraft is below glide slope path, the
pointer is displayed upward on the scale. Each dot on
the scale represents approximately 0.4 degrees
displacement.
(q) To-from pointer.
The to-from
pointers aligned on the course pointer, are located 180
degrees apart. One always points in the direction of the
station, along the selected VOR, TACAN, or RNAV
waypoint radial.
(r) Rotating heading dial.
Displays
gyro stabilized magnetic compass information. The
heading dial rotates with the aircraft throughout 360
degrees. The rotating heading dial is driven by the
pilot's compass system.
(s) Navigation flag.
When the
navigation flag is in view, the NAV portion of the system
is not valid.
(t) Course pointer. The yellow course
pointer is positioned on the heading dial by the course
knob, to a magnetic bearing that coincides with the
selected course being flown. The course pointer is also
positioned by RNAV, or flight director system modes of
operation. The course pointer rotates with the heading
dial to provide a continuous readout of course error to
the computer.
3-17. Copilot's Horizontal Situation Indicator.
a. Description. The copilot's horizontal situation
indicator (HSI) (fig. 3-9) operates independently from
the pilot's. It combines numerous displays to provide a
presentation of the aircraft position. The indicator
displays aircraft displacement relative to VOR or
localizer course, VOR/ localizer deviation, glide slope
3-16 Change 5
(2) HDG warning flag.
reliable heading information.
Indicates loss of
(3) Course pointer. The yellow course pointer
is positioned on the heading dial by the course knob to
select a magnetic bearing that coincides with the desired
VOR radial or localizer course.
(4) To-from pointer. The to-from pointers,
aligned on the course pointer, are located 180 degrees
apart. One always points in the direction of the station.
(5) Heading select knob.
orange heading bug.
Positions the
(6) Rotating heading dial.
Displays gyro
stabilized magnetic compass information. The heading
dial rotates with the aircraft throughout 360 . The
copilot's compass system provides power to drive the
rotating heading dial.
(7) Course select knob. Positions the yellow
course pointer.
(8) Course deviation bar and scale. The
yellow course deviation bar represents the centerline of
selected VOR or localizer courses. The symbolic aircraft
pictorially shows the actual aircraft position in relation to
this selected course. In VOR operation, each dot
represents 2-degree deviation from the centerline,
whereas 5 dots represent 10-degree deviation off
course. In ILS operation, each dot represents 1/2degree deviation from the centerline, whereas 5 dots
represent 2 1/2-degree deviation off course.
(9) Symbolic aircraft. The fixed miniature
aircraft symbol corresponds to the longitudinal axis of
the aircraft and lubber line markings. The symbol shows
the aircraft position and heading with respect to a radio
course and the rotating heading dial.
(10) Glide slope pointer and scale. The dual
glide slope displays glide slope deviation. The pointer is
in view only when tuned to and receiving a localizer
frequency. If the aircraft is below glide slope, the pointer
is displayed upward on the scale. Each increment on
the scale indicates an approximate 0.4 degree deviation
from the glide slope.
TM 55-1510-215-10
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
Lubber line
HDG warning flag
Course pointer
To-From pointer
Heading select knob
Rotating heading dial
Symbolic aircraft
Course select knob
Course deviation bar and scale
Glideslope pointer and scale
Heading bug
NAV warning flag
AP011854
Figure 3-9. Copilot's Horizontal Situation Indicator
(11) Heading bug. The orange heading bug is
positioned on the rotating heading dial by the heading
knob, and displays preselected compass heading. The
bug rotates with the heading dial.
(12) NAV warning flag. Indicates loss of NAV
2.
3-18. Course Deviation Indicator (KI 204) (Optional).
a. Description.
The course deviation indicator
when interfaced with the NAV 2 receiver (KN 53).
provides rectangular display of VOR/localizer, and
glideslope deviation.
The indicator incorporates
separate and independent to/from flags, and warning
flags for both VOR/localizer, and glideslope. The course
deviation indicator is protected through a 2-ampere
circuit breaker, placarded NAV 2, located on the lower
right subpanel.
b. Controls and Functions.
(1) Course deviation pointer.
The course
deviation pointer displays the aircraft's position relative
to the omni (VOR) course, or localizer (LOC) course.
When tracking a VOR course, each dot represents 1degree of deviation. When on a LOC course. each dot
represents 1/2-degree of deviation.
(2) VOR/localizer warning flag.
The NAV
warning flag will display whenever the VOR or localizer
frequency is unreliable.
(3) To-from flags. They operate only with a
usable VOR or localizer signal. The white TO/FR flags
indicate whether the selected course is to or from a
station. When receiving a reliable localizer signal. the
TO flag will be in view.
(4) Omni-bearing course card. The card is
rotated by the OBS knob for selection of a desired
VOR/localizer course.
(5) Omni-bearing selector (OBS) knob. The
OBS knob is used to set a desired VOR/ localizer course
under the course index.
(6) Glideslope deviation pointer.
The
glideslope deviation pointer displays the aircraft’s
position relative to the glideslope. Each dot represents
1/10-degree deviation from the glideslope. The pointer
is offset 15-degrees for correct viewing.
(7) Glideslope (GS) warning flag.
The
glideslope warning flag will display whenever the
glideslope frequency is unreliable.
Change 5 3-17
TM 55-1510-215-10
3-19. Pilot's Flight Director Indicator.
a. Description. The pilot's flight director indicator
(FDI) (fig. 3-10) combines the attitude sphere display
with computed steering information to provide the
commands required to intercept and maintain a desired
flight path.
All guidance commands are depicted
through the use of a command bar arrangement.
Warning flags are provided to indicate invalid attitude or
computed command displays. Any warning flag in view
indicates that portion of information is unreliable.
(5) Command bar.
The three-dimensional
command bar displays the computed steering
commands to intercept and maintain a desired flight
path. The bar moves up or down to present pitch
commands and rotates clockwise or counterclockwise
for roll commands. The command bar will bias out of
view and ATTITUDE/COMPUTER flags will display,
whenever the FD mode is not selected and/or when
flight director internal power is inadequate.
(6) Computer flag. When in view, indicates
the system is receiving unreliable flight computer data.
b. Controls and Functions.
(1) Roll attitude scale. Displays actual roll
attitude through a movable pointer and fixed reference
marks.
(2) Roll attitude pointer. Rotates with attitude
sphere, indicating roll attitude.
(3) Decision height annunciator. Illuminates
upon reaching altitude selected on the altitude
preselect/alerter.
(4) Attitude sphere. Moves with respect to
the symbolic aircraft reference to display actual pitch and
roll attitude.
Pitch attitude marks are in 5-degree
increments on a blue and brown sphere.
(7) Inclinometer.
Provides the pilot a
conventional display of aircraft slip or skid. Used as an
aid to coordinate maneuvers.
(8) Attitude flag. When in view, indicates the
system is receiving unreliable vertical gyro data.
(9) Attitude horizon test switch. When the
test switch is pressed, the ATTITUDE and COMPASS
flags will appear and the pictorial horizon will display a
climbing right turn.
(10) Symbolic aircraft. Serves as a stationary
symbol of the aircraft. Aircraft pitch and roll attitudes
1.
2.
3.
4.
5.
6.
7.
8.
RNAV annunciator
Roll attitude scale
Decision height annunciator
Attitude sphere
Command bar
Inclinometer
Attitude horizon test switch
Symbolic aircraft
AP011853
Figure 3-10. Pilot's Flight Director Indicator
3-18 Change 5
TM 55-1510-215-10
are displayed by relationship between the fixed miniature
aircraft and movable sphere. The symbolic aircraft is
flown to align the command bar to the aircraft symbol, in
order to satisfy commands of the selected flight director
mode.
(11) RNAV annunciator.
Illuminates when
navigation is being controlled by the flight director, in
NAV or APPR mode.
3-20. Copilot's Attitude Director Indicator.
a. Description.
The copilot's attitude director
indicator (ADI) (fig. 3-11) displays aircraft attitude as a
conventional pneumatically operated attitude gyro.
Attitude displayed is in relationship to an artificial
horizon. The instrument is equipped with a DH (decision
height) annunciator, which works in conjunction with the
radar altimeter. Pneumatic power to drive the gyro is
obtained from the aircraft pneumatic system.
b. Controls and Functions.
(1) Roll attitude index. Displays aircraft roll
attitude, as read against the roll attitude scale.
(3) Attitude sphere. Moves with respect to the
symbolic aircraft reference to display actual pitch and roll
attitude. Pitch up and down marks are in 5-degree
increments divided by an artificial horizon.
(4) Symbolic aircraft. Serves as a stationary
symbol of the aircraft. Aircraft pitch and roll attitudes are
displayed by the relationship between the fixed miniature
aircraft and the movable sphere.
(5) Symbolic
aircraft
alignment
knob.
Provides manual positioning of the symbolic aircraft for
pitch attitude alignment.
(6) Decision
height
(DH)
annunciator.
Illuminates when the aircraft descends below a selected
decision height, as set on the radio altimeter indicator.
3-21. Turn and Bank Indicators.
a. Description. Two gyroscopically operated turn
and bank indicators are installed separately on the pilot's
and copilot's sides of the instrument panel. The pilot's
unit is operated by DC power and is protected by a 5ampere circuit breaker placarded T&B located on the
copilot's subpanel.
(2) Roll attitude scale. Displays actual roll
attitude from 0, 10, 20, 30, 60, and 90 degrees.
1.
2.
3.
4.
5.
6.
Roll attitude index
Roll attitude scale
Attitude sphere
Symbolic aircraft
Symbolic aircraft alignment knob
Decision height annunciator
AP011852
Figure 3-11. Copilot's Attitude Director Indicator
Change 7 3-19
TM 55-1510-215-10
The copilot's turn and bank is pneumatically operated.
when the aircraft is at or below the selected decision
height.
b. Controls and Functions.
(1) Turn rate indicator. Indicates direction
and rate of turn. A two minute turn rate is indicated
when the turn rate indicator is deflected one needle
width to the left or right of the index.
(2) Index. A reference mark for alignment of
the turn rate indicator.
(3) Inclinometer.
acceleration (slip skid) of aircraft.
Indicates
lateral
(2) Altitude pointer.
point to the existing altitude.
(3) Decision height bug. The decision height
bug is set to the desired decision height, by the decision
height set knob.
(4) Altitude scale. The altitude scale is shown
in one hundred feet increments, from 0 to 2500 feet
AGL.
(5) Decision height set knob. The decision
height set knob is used to set the decision height bug.
3-22. Radar Altimeter Indicator.
a. Description. The KI 250 radar altimeter indicator
(fig. 3-12) displays actual altitude of the aircraft from
2500 feet above ground level (AGL) to touchdown. The
indicator is protected by a 1-ampere circuit breaker,
placarded RADAR ALT, located on the copilot's circuit
breaker panel.
b. Controls and Functions.
(1) Decision height (DH) annunciator.
decision height (DH) annunciator will illuminate
3-23. Altitude Select Controller.
a. Description.
The altitude select controller
(fig. 3-13) provides a means of selecting and displaying
the desired altitude reference for altitude alerting and
altitude capture. It is protected through a 1-ampere
circuit breaker, placarded ALT ALERT, located on the
copilot's circuit breaker panel.
The
1.
2.
3.
4.
5.
Decision height annunciator
Altitude pointer
Decision height bug
Altitude scale
Decision height set knob
AP011851
Figure 3-12. Radar Altimeter Indicator
3-20 Change 7
The altitude pointer will
TM 55-1510-215-10
1.
2.
3.
4.
Altitude ARM switch
Altitude ARM annunciator
Selected altitude display
Altitude select knob
AP011850
Figure 3-13. Altitude Select Controller
b. Controls and Functions.
(1) Altitude ARM switch. Pressing the ARM
(altitude arm) switch arms the altitude select controller.
ARM annunciation will display in the display window.
The ALT ARM annunciator, located in the flight director
mode annunciator panel on the pilot's instrument panel
(fig. 2-22), will be illuminated.
Within approximately 1000 feet of the preselected
altitude, ALERT will be displayed in the display window
and will remain displayed until reaching approximately
300 feet of the selected altitude.
(2) Selected altitude display.
altitude selected with altitude select knob.
Displays
(3) Altitude select knob.
(a) Inner knob. Used to select 100 foot
increments of desired altitude
(b) Outer knob. Used to select 1000
foot increments of desired altitude.
(4) Altitude alert annunciator. Will display
when the aircraft is within approximately 1000 feet, and
will extinguish when reaching/deviating approximately
300 feet of the selected display.
(5) Altitude arm annunciator.
The ARM
annunciator will display when altitude ARM has been
selected.
3-24. NAV 1 Receiver (KNS 81).
a. Descriptions. The KNS 81 is a self-contained
navigation system consisting of a 200-channel
VOR/LOC receiver, a 40-channel glideslope receiver,
and a digital RNAV computer with preselection storage
capability and display of 10 NAV frequencies (fig. 3-14).
The receiver interfaces with the pilot's TACAN/DME,
displaying bearing information on the HSI and RMI. It
receives and interprets VHF omnidirectional radio range
(VOR) and localizer (LOC) signals in the frequency
range of 108.00 to 117.95 MHz; glideslope signals in the
frequency range of 329.15 to 335.00 MHz: and marker
beacon signals to 75 MHz.
The system incorporates three TACAN modes
(TAC, TAC RNV, TAC RNV APR), three VOR modes
(VOR, VOR RNV, VOR RNV APR). and an ILS mode.
System flexibility is maintained with CHK and RAD
buttons. The CHK button, when pressed, permits a
momentary display of the TAC/VOR radial and
DME distance information in the display window.
Change 5 3-21
TM 55-1510-215-10
The RAD button when pressed permits displaying the
radial (in place of ground speed) from the VORTAC or
waypoint in the DME indicator.
In addition, the KNS 81 tunes the KTU 709
DME/TACAN system to 252 (126X and 126Y) TACAN
channels. The system is protected through a 3-ampere
circuit breaker, placarded NAV 1, located on the copilot's
subpanel. DME is protected through a 2-ampere circuit
breaker, placarded DME, located on the copilot's circuit
breaker panel.
b. Controls and Functions.
(1) System
status
displays.
Mode
annunciation displays the mode (VOR, VOR/RNV, VOR/
RNV/APR, TAC, TAC/RNV, TAC/RNV/APR) that is
active.
NOTE
When an ILS frequency is selected in
a
VOR
mode,
the
system
automatically goes into the ILS mode
regardless of the VOR mode
annunciated.
(2) WPT (waypoint) display.
Shows the
number of the waypoint data storage bin that is being
displayed in the display window. When the displayed
waypoint number is different from the active waypoint,
the WPT will blink.
(3) Caret display (> <).
Indicates the
waypoint parameter (frequency or channel, radial, or
distance) that the data input knobs are controlling. The
caret display cycles from FRQ to RAD to DST and back
to FRQ as the data button is pressed.
(4) FRQ (frequency) display.
Shows the
frequency entered into the system for VOR, VORTAC, or
localizer stations with the data input knobs. When in a
TACAN mode, the channel number is displayed. When
selecting a frequency or channel, the carets must be at
the FRQ display. When the system is first turned on, the
information being displayed prior to the last shutdown
will be displayed again.
(5) RAD (radial) display. Shows the VORTAC
or TACAN station radial on which the waypoint is
located.
NOTE
In the VOR or TAC mode, this display
will show dashes to indicate that
RNAV radial information is irrelevant.
The value is selected by the data input knobs when
the caret is at the RAD display. When the CHK button is
3-22 Change 5
pressed, this area will display the radial from the
VORTAC or TACAN station to the aircraft.
(6) DST (distance) display.
Shows the
distance the waypoint is offset from the VORTAC or
TACAN station.
NOTE
In the VOR or TAC mode, this display
will show dashes to indicate that
RNAV
distance
information
is
irrelevant.
Its value is selected by the data input knobs when
the caret is at the DST display. This area will display the
distance from the VORTAC or TACAN station to the
aircraft when the CHK button is pressed.
(7) Data input knobs. Two concentric knobs
used for selection of VOR/LOC frequency or TACAN
channel, waypoint radial, and waypoint distance. DME
and TACAN are channeled with selection of the paired
VOR frequency or TACAN channel.
The internal
glideslope is channeled with selection of the paired LOC
frequency. Turn these knobs to enter desired data into
the RNAV computer as follows:
(a) Frequency/channel selection. In a
VOR mode, the large knob controls the 10 and 1 MHz
digit in the display. The small knob controls the 0.1 and
0.05 MHz digit and selects FRQ in 0.05 MHz steps in
either the in or out position.
In a TACAN mode, the large knob controls the 10
and 100 digits in the display. The small knob controls
the 1 digit in the in position and x or y selection in the out
position. If the unit is displaying an illegal channel, i.e.,
00 or 127 through 129, the FRQ will flash.
(b) Radial selection. The large knob
changes the 10-degree and 100-degree digit of the
display. The small knob changes the 1-degree digit in
the inner position and the 0.1-degree digit in the outer
position.
(c) Distance selection. The large knob
changes the 100 and 10 nautical mile digit of the display.
The small knob changes the 1 nautical mile digit in the in
(inner) position and the 0.1 nautical mile digit in the out
(outer) position.
(8) DATA button.
The momentary push
button placarded, DATA, moves the caret (> <) display
from FRQ to RAD to DST and back to FRQ, providing a
visual reference of which waypoint parameter the data
input knobs are addressing.
(9) Power OFF PULL IDENT volume control.
TM 55-1510-215-10
1.
2.
3.
4.
5.
6.
7.
8.
9.
System status displays
Data input knobs
DATA button
Power OFF PULL IDENT volume control
CHK (check) button
RAD (radial) button
RTN (return) button
USE button
Waypoint select knob
AP011849
Figure 3-14. NAV 1 Receiver
Turn clockwise for power on and volume increase. Pull
out for VOR/LOC IDENT tone. In TACAN modes. this
audio is muted and the station is identified by selecting
DME on the audio panel.
(10) CHK (check) button.
Pressing this
momentary push-button displays the aircraft's position
from a VORTAC or TACAN by replacing the RAD and
DST waypoint parameters normally displayed with the
radial and distance from the VORTAC or TACAN station.
(11) RAD (radial) button. This two-position
button is normally operated in the outer position. When
depressed (inner position), the remote digital DME
indicator will display the radial you are on (from the
active waypoint, VORTAC, or TACAN station) instead of
groundspeed and time-to station. This feature becomes
inoperative in DME hold mode or when the DME is tuned
by another NAV receiver.
(12) RTN (return) button. A momentary push
button that switches the display to the active waypoint,
VORTAC, TACAN, or ILS frequency. This button is
used after entering new data for other waypoints, and
you desire returning to navigation data presently in use.
(13) USE button. After calling up a previously
entered RNAV waypoint, VORTAC, TACAN, or ILS
frequency with the waypoint select knob, depressing this
momentary push button will put in use that waypoint.
VOR. VORTAC. TACAN. or ILS frequency.
(14) Waypoint select knob. Selects any one of
ten data storage bins in which VOR/LOC frequencies or
TACAN channels, with or without waypoint coordinates,
have been inserted. Each storage bin can be called up
as required from 0 through 9, or 9 through 0.
(15) Mode select knob. This knob selects one
of the six available modes of operation. Modes are:
VOR (direct to or from a VOR or VORTAC station with
angular course width deviation), VOR/ RNAV (direct to
waypoint with linear crosstrack deviation ±5 nautical
mile). VOR/RNV/APR (direct to waypoint with linear
crosstrack deviation ±1.25 nautical mile), TAC (direct to
or from VORTAC or TACAN station with angular course
width deviation), TAC/RNV (direct to waypoint with linear
crosstrack deviation ±5 nautical mile). TAC/RNV/APR
(direct to waypoint with linear crosstrack deviation ±1.25
nautical mile).
Change 5 3-23
TM 55-1510-215-10
When an ILS frequency is entered, with the system
in a VOR mode, the system will automatically go to the
ILS mode. When the ILS frequency is removed, the
system will revert back to the VOR mode it was in
previously. Some TACAN channels correspond to ILS
frequencies. Selecting a TACAN mode will allow full use
of these channels without the system reverting to ILS
mode.
c. Operating Procedures.
1. AVIONICS MASTER switch - ON.
2. OFF PULL IDENT switch - Turn
clockwise out of detent.
Adjust
volume as required.
3. Mode of operation - Select as
required.
4. Waypoint -Select as required.
5. Tacan channel -Verify, if in a TAC
mode.
6. Frequency - Verify, if in a VOR,
VORTAC, or ILS mode.
7. Station ident -Check as follows:
a. TACAN - Select DME audio.
b. VOR - Pull out OFF PULL
IDENT switch. Push in after
identification has been made.
d. Shutdown Procedures.
1. OFF PULL IDENT switch - Turn
counterclockwise into the off
position.
2. AVIONICS MASTER switch - OFF.
3-25. NAV 2 Receiver (KN 53).
a. Description.
The KN 53 NAV 2 receiver
(fig. 3-15) is capable of being tuned to all 200 VOR/
LOC frequencies from 108.00 to 117.95 MHz, and
receive all 40 glideslope channels.
Navigation
information is displayed on the copilot's HSI. The
receiver allows for selecting and storing two frequencies,
USE and STBY. These frequencies are displayed at all
times. Both frequencies are provided with storage,
which protects them from loss during power
interruptions. Power for the receiver is routed through a
2-ampere circuit breaker placarded NAV 2. located on
the copilot's subpanel.
b. Controls and Functions.
(1) USE display. The active frequency tuned
to. This frequency is first tuned into the STBY position,
then transferred to the USE position by pressing the
frequency transfer button.
3-24 Change 5
(2) STBY display. A standby frequency (if
selected) will display in the STBY window.
(3) Frequency selector knobs. The outer
knob tunes the MHz portion of the frequency. The
smaller knob tunes the kHz portion in 50 kHz steps. By
rotating the frequency selector knobs either clockwise or
counterclockwise, the desired operating frequency is
programmed into the standby (STBY) display window. A
clockwise rotation will increase the displayed frequency
number, while a counterclockwise rotation will decrease it.
(4) OFF PULL ID knob. Rotating the OFF
PULL ID knob clockwise from the detented off position
applies power to the unit. No warm up time is required.
Further rotation of the knob increases NAV signal
volume. NAV voice is heard when the knob is pushed
in. When the knob is pulled out, the identification (ID)
signal, and voice may be heard.
(5) Frequency transfer button.
When
pressed, the frequency transfer button will transfer the
displayed standby frequency into the USE position and
moves the displayed USE frequency to the STBY
(standby) position.
c. Operating Procedure.
1. AVIONICS MASTER switch - ON.
2. OFF PULL ID knob - Turn knob
clockwise out of off (detent).
Increase volume by further rotation
of knob.
3. Frequencies - Select.
Insure
desired active frequency is in USE
display.
4. Identify station - As required.
d. Shutdown Procedure.
1. OFF PULL ID knob - Turn
counterclockwise to off (detent).
2. AVIONICS MASTER switch - OFF.
3-26. ADF Radio (KR 87).
a. Description. The KR 87 automatic directional
finder is a digitally tuned solid state receiver which
provides bearing information to stations in the 200 kHz
to 1799 kHz frequency band, along with audio reception
(fig. 3-16). The unit displays the active frequency in the
left display window. The right window will display either
the stand-by frequency (which can be transferred to the
active window), or a flight timer, or programmable
elapsed timer. The flight timer will keep track of
TM 55-1510-215-10
1.
2.
3.
4.
5.
USE frequency display
STBY frequency display
Frequency selector knobs
OFF PULL ID knob
Frequency transfer button
AP011848
Figure 3-15. NAV 2 Receiver
the total flight time, while the independent programmable
elapsed timer can be reset to count up from zero or be
preset to a value and count down to zero. The system is
protected by a 1-ampere circuit breaker, placarded ADF,
located on the right subpanel. The antenna located on
the lower side of the aircraft (fig. 2-1). contains both
loop and sense antennas, preamplifiers, and modulators,
which combine the antenna signals into a single RF
signal which is output to the receiver via a triaxial cable.
b. Operating Modes. The automatic directional
finder (ADF) has two operational modes. In the ANT
(antenna) mode, the loop antenna is disabled. and the
unit acts as a receiver, allowing audio reception through
the speaker or headphones. The indicator needle as
selected on the RMI's will park at a 90-degree relative
position and the ANT message on the left side of the
display window will display. This mode will provide a
slightly clearer audio reception, and is used for station
identification. In various parts of the world, some L/LM
stations use an interrupted carrier for identification
purposes. A beat frequency oscillator (BFO) function is
provided to permit these stations to be more easily
identified. Pushing the BFO switch (fig. 3-16) will cause
a 1000 Hz tone to be heard whenever there is a radio
carrier signal present at the selected frequency. It will
also light the BFO message in the center of the display.
With the ADF button depressed. the unit is placed
into the ADF mode and the loop antenna is enabled.
The ADF message on the left side of the window display
will display and the indicator needle as selected on the
RMI's will point to the relative bearing of the selected
station. In order to tell if there is a sufficient signal for
navigational purposes, place the unit back in the ANT
mode, parking the indicator needle as selected on the
RMI's at 90-degrees. When the unit is then switched to
the ADF mode, the needle should slew to the station
bearing in a positive manner.
without excessive
sluggishness, wavering, or reversals.
c. Controls and Functions.
(1) Mode display.
the unit is in.
Displays operating mode
(2) USE display.
the system is tuned to.
Displays frequency which
(3) FRQ display. Illuminates when a standby
frequency has been inserted for display.
(4) STBY/TIMER display. Displays standby
frequency, flight time or elapsed time.
Change 5 3-25
TM 55-1510-215-10
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
Mode display
USE frequency display
STBY/TIMER display
Frequency select knobs
OFF/VOL control switch
SET/RST button
FLT/ET button
FRQ button
BFO button
ADF button
AP011847
Figure 3-16. ADF Radio
The standby frequency is placed in memory when either
FLT (fight time) or ET (elapsed time) mode is selected.
(5) FLT/ET display. Flight time or elapsed
time are displayed and annunciated alternatively by
depressing the FLT/ET button.
(6) Frequency, select
active/standby frequencies.
knobs.
Selects
(a) Active frequency.
The active
frequency is displayed in the left portion of the window
display. This frequency may be changed with the
concentric knobs when either time mode (FLT or ET) is
being displayed. To set the 10 digit, push the small knob
in and rotate it. Clockwise rotation will increment the
digit. The digit will roll over at 9 to 0 and roll under
(when turning the knob counterclockwise) at 0 to 9. With
the small knob pulled out, the 1's digit may be set. It's
operation is the same as for the 10's digit.
Turning the large knob changes the 100's digit and
the 1000's digit. The 100's digit carries to the 1000's
digit from 9 to 10 and borrows from 10 to 9. The two
digits roll over from 17 to 02 and under from 02 to 17,
thus limiting the frequencies to the range of 200 kHz to
1799 kHz.
3-26 Change 5
(b) Standby frequency. The standby
frequency is displayed in the right portion of the display
window when the FRQ message is displayed. In this
case, the standby frequency may be changed with the
frequency control knobs as desired. If the standby
frequency is not being displayed, it may be called to the
window by pressing the FRQ button. Pressing this
button when the standby frequency is displayed causes
the current standby and active frequencies to be
exchanged.
(7) OFF/VOL control switch. Rotating this
control switch clockwise out of detent turns the unit on.
Further rotation of the control switch increases volume.
(8) SET/RST button. This button controls the
elapsed time. Time is reset back to zero each time the
button is pressed. The elapsed times has a countdown
mode which is set by depressing the SET/RST button for
2 seconds or until the ET display begins to flash. With
the ET flashing, the selector knob will set any time up to
59 minutes, 59 seconds. To start the elapsed time count
down, press the SET/RST button and the timer will start
counting down. At zero the timer will begin to flash for
15 seconds and then go to a solid display.
TM 55-1510-215-10
While the FLT or ET are displayed, the in use frequency
may be changed with the frequency control knobs.
(9) Timers.
(a) FLT/ET button. If elapsed time (ET)
is currently displayed pressing the FLT/ET button will
cause the flight timer (FT) to be displayed. Pressing this
button again will exchange the two timers in the display.
If the standby frequency is displayed. pressing the
FLT/ET button will cause the timer which was last
displayed to reappear in the window.
NOTE
When power is first applied the flight
timer is displayed.
(b) Flight timer. The flight timer is
displayed in the right portion of the display window when
the FLT message is displayed. The timer receives
power through the landing gear squat switch. The
counter will count up to 59 hours, 59 seconds. It begins
counting upon liftoff and will quit counting upon landing.
The elapsed timer may be reset back to 0 by pressing
the SET/RST button or turning the unit off.
(c) Elapsed timer. This counter has two
modes: count up and count down. When power is
applied it is in the count up mode starting at 0. The timer
will count up to 59 hours, 59 seconds, displaying
minutes and seconds until one hour has elapsed, then
displaying hours and minutes. When in the count up
mode. the timer may be reset to 0 by pressing the
SET/RST button.
NOTE
Pressing the reset button will reset
the elapsed timer regardless of what
is currently being displayed.
To enter the count down mode, the SET/RST button
is held depressed for approximately 2 seconds until the
ET message begins to flash. While the ET message is
flashing the timer is in the ET set mode. In this mode a
number up to 59 minutes, 59 seconds may be preset
into the elapsed timer with the frequency select knobs.
With the smaller knob pressed in, the 10's of seconds
digit may be changed; it will roll over from 5 to 0 and
under 0 to 5. With the knob pulled out, the 1's of
seconds digit may be changed. It rolls over for 9 to 0
and under from 0 to 9. The larger knob modifies the
minutes. It rolls over from 59 to 0 and under from 0 to
59. The timer will remain in the ET set mode (ET
message flashing) for 15 seconds after a number is set
in or until the SET/RST. FLT/FT or FRQ button is
pressed. The preset number will remain unchanged until
the SET/RST button is pressed. When the SET/RST
button is pressed after a number has been preset. the
elapsed timer will start counting down. When the
elapsed timer is counting down. pressing the SET/RST
button again will have no effect unless it is held for
approximately 2 seconds. This will cause the timer to
stop and enter the set mode (ET message flashing).
When the timer reaches 0 it changes to the count up
mode and continues up from 0. The elapsed time will
flash for 15 seconds, then annunciate a steady display.
(10) FRQ button. The selected frequency is
put into the active window by pressing the FRQ button.
The standby and active frequencies will be exchanged.
(11) BFO button. When the BFO button is
depressed, the BFO mode is activated and BFO will
display in the display window. In various parts of the
world, some L/LM stations use an interrupted carrier for
identification purposes. When in the BFO mode these
stations are more easily identified. Pushing the BFO
button will cause a 1000 Hz tone to be heard whenever
there is a radio carrier signal present at the selected
frequency.
(12) ADF button. When the ADF button is
depressed. the unit is placed into the ADF mode with
the loop antenna enabled. An ADF message on the left
portion of the window will display. Indicator needles. as
selected on the RMI's, will point to the relative bearing of
the selected station.
d. Operating Procedures.
1. AVIONICS MASTER switch - ON.
2. OFF VOL control switch - Turn
clockwise out of detent.
Adjust
volume as required.
NOTE
An audio muting feature causes the
audio output to be muted unless the
receiver is locked onto a valid
station.
This reduces interstation
noise and aids in identifying
navigable stations.
3. ADF test - Select ANT mode. Verify
bearing pointers park in the 90degree position.
4. Operating mode - Select as desired.
5. Frequencies - Select.
Insure
desired
active
frequency
is
displayed.
6. FLT/ET
switch - Operate as
required.
Change 5 3-27
TM 55-1510-215-10
e. Shutdown Procedures.
1. OFF
VOL
switch
Turn
counterclockwise to off (detent).
2. AVIONICS MASTER switch - OFF.
system are slipped to assure the pilot can overpower the
autopilot at all times. The system is protected by a 10ampere circuit breaker placarded AP/FD, and a 1ampere circuit breaker placarded YAW, both located on
the copilot's circuit breaker panel (fig. 2-18).
b. Modes of Operation.
3-27. Flight Control System (KFC 250).
a. Description.
The flight control system is
functionally divided into four parts: sense, compute,
display, and control. All sensor information (pitch and
roll reference. slaved compass, RNAV/VOR/ LOC/GS,
DME, marker receiver, and air data) is fed into a flight
computer. The flight computer computes pitch and roll
commands. These commands are routed to the pilot's
flight director indicator (fig. 3-10), where they are
displayed on the command V-bar as visual steering
commands. These steering commands are also fed to
the autopilot computation circuits contained in the flight
computer, where the steering commands and aircraft
yaw rate information are combined to generate the
aileron, elevator trim, and rudder drive commands for the
autopilot.
Each servo used in the system incorporates a
clutch, which allows for manually overriding the controls.
During the autopilot preflight check, all clutches in the
(1) Flight director (FD). The flight director
mode is activated by depressing the FD button on the
mode controller (fig. 3-17). FLT DIR will display in the
mode annunciator panel located on the pilot's instrument
panel (fig. 2-22). The FDI command V-bar will appear,
providing the pilot with steering commands, to maintain
wings level and pitch attitude that existed at the time of
flight director engagement. If pitch or roll attitudes are
changed, recycling the FD button will synchronize the
command V-bar to the new position.
If a change in the commanded pitch attitude is
desired, the control wheel steering (CWS) button,
installed on both (pilot and copilot) control wheels, allows
the pilot or copilot to manually synchronize the command
V-bar. The vertical trim switch on the mode controller
may also be used to adjust the selected pitch attitude up
or down at approximately 1-degree/per second.
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
Figure 3-17. Mode Controller
3-28 Change 5
Vertical trim switch
Heading (HDG) button
Flight director (FD) button
Altitude (ALT) button
Autopilot switch
Yaw damp switch
Yaw damp annunciator
Autopilot test switch
ELEC TRIM circuit breaker
TRIM TEST switch
ROLL TEST switch
PITCH TEST switch
Approach (APPR) button
Back course (BC) button
Navigation (NAV) button
TM 55-1510-215-10
NOTE
The flight director (FD) mode must be
activated before the autopilot can be
engaged.
(2) Autopilot engagement.
With the HDG SEL mode in operation, subsequent
changes made in the heading bug position will
immediately cause the command V-bar to call for a turn
to the new heading.
The HDL SEL mode is cancelled when NAV or
APPR coupling occurs, or whenever the FD or HDG
mode buttons are depressed.
CAUTION
CAUTION
Prior to autopilot engagement, insure
that command V-bar commands are
satisfied.
An invalid heading source (compass
flag in view) will automatically
disengage the autopilot. The autopilot
may be reengaged; however, only the
vertical modes will be useable.
The autopilot is engaged by moving the AP switch
on the mode controller to the ON position. FLT DIR,
AUTOPILOT will display in the mode annunciator panel
(fig. 2-22). With autopilot selected, yaw damp (YD)
mode is automatically engaged.
The autopilot, together with the yaw damp, provides
three-axis stabilization, automatic turn coordination, and
automatic elevator trim, as well as automatic response to
all selected flight director commands.
CAUTION
When the autopilot is engaged,
manual application of a force to the
pitch axis of the control wheel for a
period of 3 seconds or more will
result in the autotrim system
operating in the direction to create an
opposing force. If the autopilot is
disengaged under these conditions,
the pilot may be required to exert
control forces in excess of 50 pounds
to maintain the desired aircraft
attitude. This force will have to be
maintained until the aircraft is
manually retrimmed.
(3) Heading select/preselect mode (HDG
SEL). Prior to selecting the HDG SEL mode, position
the heading bug on the pilot's HSI. Depress the HDG
button on the mode controller, activating the HDG SEL
mode. HDG SEL will display on the mode annunciator
panel (fig. 2-22) and a computed, visually displayed
bank command will show on the FDI. The command Vbar on the FDI will deflect in the direction of the shortest
turn to satisfy the commanded turn to the preselected
heading. The aircraft may be manually banked to
realign the command V-bar and satisfy the command or,
if the autopilot is engaged, the aircraft will automatically
bank, turn to, roll out, and hold the preselected heading.
As the aircraft approaches the selected heading, the
command V-bar will command a rollout to wings level.
(4) Yaw damp (YD) mode. The yaw damp
controller is located beside the mode controller
(fig. 3-17). Yaw axis is automatically engaged when the
autopilot is engaged.
Disengagement of YD is
accomplished by using the alternate action switch
located on the yaw damp controller. YD may be
engaged alone or with any flight director mode. Yaw
damp engage status is indicated by a annunciator,
placarded ON, located on the YD controller.
(5) Navigation (NAV ARM and NAV CPLD)
mode. The NAV mode provides visual commands on
the FDI. and deviation guidance on the pilot's HSI to
intercept and track a VOR or RNAV course.
When the NAV button on the mode controller is
depressed. NAV/ARM will be displayed on the mode
annunciator panel and the automatic capture circuit is
armed. HDG SEL is retained until capture occurs. The
VOR/RNAV course-capture point is variable to prevent
an overshoot. It depends on the angle of intercept and
rate that course deviation is changing. Upon capture. a
bank command will be displayed on the FDI; HDG (if on)
will be cancelled and NAV/CPLD will be displayed on the
mode annunciator panel. The pilot can manually bank
the aircraft to satisfy the command display which will call
for a rollout to level flight when on course centerline to
track the course. Crosswind compensation is provided
in the track state.
If the NAV mode is selected with the aircraft level
±4-degrees of bank and within three dots of course
deviation, NAV/ARM will be bypassed and NAV/CPLD
will engage directly. If the autopilot is engaged, the
aircraft will bank to satisfy the command display and
rollout on course automatically.
Upon station (or waypoint) passage. an outbound
course other than the inbound reciprocal can be selected
by resetting the NAV course arrow on the HSI. This will
Change 5 3-29
TM 55-1510-215-10
cause an immediate command V-bar deflection on the
FDI directing a turn to the new course.
The NAV mode is cancelled by depressing the NAV
button. by selecting HDG (when in NAV coupled) or
APPR modes, or deselecting FD.
CAUTION
The NAV mode of operation will
continue to provide aircraft control
without a valid VOR/LOC signal (NAV
flag in view).
(6) Approach (APPR/ARM, APPR/CPLD and
GS CPLD) mode. The APPR mode provides visual roll
and pitch commands on the FDI command V-bar to
capture and track precision ILS beams. or non-precision
VOR radials. Lateral and vertical deviation is monitored
on the pilot's HSI.
The automatic APPR capture function will be
immediately armed. APPR/ARM will be displayed on the
mode annunciator panel. In APPR/ARM mode, prior to
capture, the heading select mode is retained.
The LOC beam or VOR capture point will vary,
depending on angle of intercept and rate of change of
deviation indication. Upon capture, a bank command
will be introduced on the FDI, the existing heading mode
will be cancelled and APPR/ CPLD will be displayed on
the mode annunciator panel. The pilot may manually
bank the aircraft to satisfy the command display, which
will command a rollout to level flight when the aircraft is
on course. Automatic crosswind compensation will
provide precise tracking. VOR/LOC deviation is shown
on the pilot's HSI. Actual crab angle will be indicated by
offset of the course arrow from the lubber line.
If the autopilot is engaged during operation in the
APPR mode, automatic steering response will follow the
command display on the FDI.
The glideslope mode is armed for automatic capture
if LOC front course capture has occurred. Automatic
glideslope capture occurs as the aircraft passes through
the glide path from above or below. Upon interception of
the glideslope, capture occurs and GS CPLD is
displayed on the mode annunciator panel. A capture
pitch command is displayed by the command V-bar.
The pilot or autopilot (if engaged) controls the aircraft to
satisfy the command V-bar. Upon GS capture, the ALT
HOLD mode (if active) is cancelled. However, ALT
HOLD may be manually reselected to maintain altitude
upon reaching MDA if visual contact is not established.
APPR/CPLD mod is cancelled by selection of HDG.
NAV, or go-around modes or by deselecting FD or
APPR.
3-30 Change 5
CAUTION
The APPR mode of operation will
continue to provide aircraft control
without a valid VOR/LOC signal (NAV
flag in view).
(7) Back course (BACK CRS) mode.
Whenever a LOC or ILS frequency is selected, the BC
mode may be activated by depressing the BC button on
the mode controller, after selecting APPR. When in BC
mode and localizer capture occurs, the system will turn
and track outbound on the front course or inbound on
the back course. The BC mode reverses the LOC
deviation signal and course datum to permit the FDI
steering command display to operate on a steer-to rather
than a steer-from basis on the reverse course. BACK
CRS will be displayed on the mode annunciator panel.
(8) Go-around mode. The go-around mode is
primarily designed to assist in establishing the proper
pitch attitude under missed approach conditions.
However. it can also be used to establish a proper climb
attitude during takeoff. The go-around switch is located
on the left power lever. Depression of the go-around
switch cancels all flight director modes and, if the
autopilot is engaged, disengages the autopilot. A wingslevel and 8-degree pitch-up command is displayed by
the FDl. GO AROUND will be displayed on the mode
annunciator panel.
Go-around is cancelled by use of vertical trim,
altitude hold mode, control wheel steering or by turning
off the flight director.
(9) Altitude select (ALT ARM) mode. This
mode allows for selection of an altitude, and upon
approaching that altitude obtain an automatic visual pitch
command on the FDI, to capture and hold this preselected altitude. As the aircraft reaches the selected
altitude, ALT HOLD will automatically engage. ALT
HOLD will display on the mode annunciator panel and
ALT ARM display will disappear. If the autopilot is
engaged the system will automatically capture and hold
the selected altitude.
ALT ARM is disengaged by depressing the ALT
ARM button, by engaging ALT HOLD, by GS capture, or
deselecting FD.
(10) Altitude hold (ALT HOLD) mode. This
mode will cause a computed visual pitch command on
the FDI command bars to hold the aircraft at the
pressure altitude existing at the time the mode was
activated. The mode is activated either automatically by
the ALT ARM function, or manually by depressing the
ALT button on the mode controller. If the autopilot is
engaged, it will automatically hold the aircraft at that
altitude.
TM 55-1510-215-10
The vertical trim switch is used to adjust the selected
altitude up or down at a constant rate of approximately
600 fpm without disengaging the mode.
because of back force generated by
the elevator downsprings or pilot
induced forces.
The ALT HOLD mode is cancelled by automatic
Glideslope capture, selection of ALT ARM/GO-AROUND
modes, or deselecting FD.
WARNING
(11) Control wheel steering (CWS). When the
autopilot is engaged, CWS provides the pilot with the
capability for manual maneuvering of the aircraft without
the need to disengage and reengage the autopilot, or
reselect any modes of operation. CWS is engaged by
continuous pressure on the CWS button, located on
either the pilot's or copilot's control wheel. Operation of
the CWS button causes immediate release of autopilot
servos allowing the pilot to assume manual control.
Upon release of the CWS button, the autopilot will
reassume control of the aircraft to the original lateral,
and existing vertical commands.
If the AUTOPILOT circuit breaker is
pulled,
the
red
TRIM
failure
annunciator light on the autopilot
annunciator panel will be disabled
and only the aural alert will sound if
an electric trim malfunction occurs.
If the aural alert sounds pull the
ELECT TRIM circuit breaker and
accomplish inflight trimming with the
manual trim wheel.
1. AVIONICS MASTER switch - ON.
2. Flight director switch - Press on.
Since all engaged modes remain coupled during
operation of the CWS switch, their annunciator displays
will continue to show on the mode annunciator panel.
CAUTION
Prior to autopilot engagement insure
that command V-bar commands are
satisfied.
c. Operating Procedures.
WARNING
Insure that the autopilot has been
disengaged and check that the
aircraft manual trim indicator is set to
the takeoff position before takeoff.
Operating the autopilot on the ground
may cause the autotrim to run
3. Autopilot switch (AP) - ON.
4. Autopilot modes - As required.
d. Shutdown Procedures.
1. Flight director switch - Disengage.
2. AVIONICS MASTER switch - OFF.
Section IV. TRANSPONDER AND RADAR
3-28. Transponder (KT 76A).
a. Description. The transponder, (fig. 3-18) is an
identification, position tracking , altitude reporting, and
emergency tracking device. The unit receives, decodes,
and responds to interrogations by search radar. The
range of the set is normally limited to line-of-sight. The
transponder is protected by a 3-ampere circuit breaker,
placarded XPDR 1, located on the right subpanel. The
associated antenna is shown in figure 2-1.
b. Controls and Functions.
(1) Function selector switch.
Provides for
selection of OFF, SBY, ON, ALT, and TST positions.
(a) OFF position. Removes power from
the unit.
(b) SBY position. The unit should be
placed in SBY after engine start for warm up. It takes
approximately 47 seconds for the transponder to warm
up and become operational.
(c) ON position. When the transponder
is in the ON position the unit is operating in mode A.
Mode A provides for normal operation without altitude
reporting.
(d) ALT
position.
When
the
transponder is in the ALT position the unit is operating in
mode C (altitude reporting). Mode C provides for normal
operation along with altitude reporting. The aircraft
altitude is automatically reported to the ground controller
in increments of 100 feet from minus 1000 feet up to
63,000 feet.
(e) TST position. When the function
selector switch is held in the TST position the reply
annuniator should illuminate and remain illuminated until
Change 5 3-31
TM 55-1510-215-10
1.
2.
3.
4.
5.
Function selector switch
Reply annunciator
Code windows
Control knobs
Ident button
AP011845
Figure 3-18. Transponder (KT 76A)
When the function selector switch is held in the TST
position the reply annunciator should illuminate and
remain illuminated until the selector switch is placed in
another position. This provides for a integral test of the
unit.
(2) Reply annunciator.
During normal
transponder operation, a flashing annunciator is an
indication of a transmitted reply. An interrogation will
normally be at 10-15 second intervals Flashes within this
interval may be from noise, a second or third
interrogator, or from side lobes (from interrogators
without side lobe suppression). When the IDENT button
is depressed the reply annunciator will glow steady as
an indication of the ident function.
(3) Code windows.
Displays transponder
reply codes as selected with the control knobs.
(4) Control knobs. The control knobs are
used to select the desired transponder reply codes.
Attention should be paid to the code selected. The
selected code should be in accordance with instructions
for IFR flight or rules applicable to transponder utilization
for VFR flight. Unless required, avoid selecting 7700,
7600, or 7500 codes. These codes are for emergencies,
loss of communications, and hijacking respectively.
3-32 Change 5
(5) IDENT button. The IDENT button feature
is used at the request of the traffic controller The IDENT
button is depressed momentarily and then released. A
memory holds the IDENT reply for an interval to assure
the proper reply for at least one radar sweep. This
memory also turns the reply lamp on steady as an
indication of the ident function.
c. Operating Procedures.
1. AVIONICS MASTER switch - ON.
2. Transponder reply code - Set as
required.
3. Function selector switch - SBY
(allow time for warm up).
4. Function selector switch - ON or
ALT prior to takeoff.
5. Ident button - Depress as required.
d. Shutdown Procedure.
1. Function selector switch - OFF.
2. AVIONICS MASTER switch - OFF.
TM 55-1510-215-10
3-29. Transponder Set (AN/APX-72) (If installed).
a. Description. The transponder set (fig. 3-19) is
an identification, position tracking, and emergency
tracking device. This set receives, decodes. and
responds to interrogations by search radar. It operates
in conjunction with a TS-1843/APX inflight test set which
provides a self test feature, and a KIT-1A/TSEC
computer which provides mode 4 feature. Power for the
system is fed through two 50-ampere circuit breakers,
placarded AVIONICS 1 and AVIONICS 2, located on the
copilot's circuit breaker panel. The system is protected
by a 10-ampere circuit breaker, placarded XPDR 2,
located on the right subpanel.
b. Controls and Functions.
(1) CODE control. Selects mode 4 code of
the day, A or B.
(a) HOLD position. Prevents zeroizing
when power is removed from the set or MASTER control
is turned OFF.
(b) A position. Selects keyed in code A.
(c) B position. Selects keyed in code B.
(2) REPLY light. Indicates valid mode 4
interrogations and replies when MODE 4 AUDIO LIGHT
switch is in AUDIO or LIGHT positions.
(3) TEST light. Indicates the proper response
has been generated when the M-1, M-2, M-3/A, and M-C
switches are placed in TEST position. Also illuminates
when RAD TEST-MON switch is in MON position and
replies are made to M-1, M-2, or M-3/A interrogations.
(4) MASTER control.
selecting the following:
Provides a means of
(a) OFF position. Turns set off.
(b) STBY position. Places set in warmup (standby) condition.
(c) LOW position. Places set at low
sensitivity.
(d) NORM position. Operates set at
normal sensitivity.
(e) EMER
position.
Transmits
emergency reply.
(5) M-1, M-2, M-3/A and M-C switches.
Provides a means of selecting the following:
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
CODE control
REPLY light
TEST light
MASTER control
M3/A switch
M-C switch
RAD TEST-MON switch
IDENT-MIC switch
MODE 3.A code selector
MODE 1 code selectors
MODE 4 ON-OUT switch
MODE 4 AUTO-LIGHT switch
M-1 switch
M-2 switch
Figure 3-19. Transponder Control Panel (AN/APX-72)
Change 5 3-33
TM 55-1510-215-10
(a) Up (on) position. Permits set to
reply in the selected mode.
(b) OUT position. Disables replies.
(c) TEST position. Permits self test in
the selected mode when test set TS-1843A/APX is
installed. The transponder set can also reply to ground
interrogations in the selected mode while being tested.
(d) M-1 and M-2 switches. Response
to military identification interrogations.
(e) M-3/A switch. Response to civilian
identification interrogations.
(f) M-C switch. Response to altitude
reporting interrogations.
(6) RAD TEST MON switch.
means of selecting the following:
Provides a
(a) RAD
TEST.
Enables
an
appropriately equipped transponder to reply to TEST
mode interrogations.
(b) MON position. Turns on circuits in
the transponder test set to monitor for proper replies to
M-1, M-2 or M-3/A interrogations.
(7) IDENT MIC switch. Provides a means of
selecting the following:
(a) IDENT
position.
Activates
identification feature.
(b) MIC position. Inoperative in this
installation.
(8) MODE 1 and MODE 3/A code selector.
Selects the desired reply codes for modes 1 and 3A.
NOTE
MODE 2 code selector is located on
the transponder receiver-transmitter
and should be set prior to flight when
required.
(9) MODE 4 IFF ON OUT switch. Provides a
means of selecting the following:
(a) ON position. Permits transponder
set to decode a mode 4 interrogation.
(b) OUT position.
Disables mode 4
decoding.
(10) AUDIO OUT LIGHT switch. Provides a
means of selecting the following:
(a) AUDIO position. Permits aural and
reply light monitoring of valid mode 4 interrogations and
replies.
(b) OUT position.
Disables mode 4
decoding.
3-34 Change 5
(c) LIGHT
REPLY light monitoring.
position.
Permits
only
c. IFF Caution Light. The function of the IFF
caution light (fig. 2-22), placarded IFF CAUTION, is as
follows:
(1) IFF. Indicates that the transponder set
has failed to reply to a valid mode 4 interrogation. It also
illuminates when mode 4 codes have been zeroized.
d. Transponder Set Operation. If mode 4 operation
is required, perform the following procedures after the
landing gear has been retracted.
NOTE
MODE 2 code selectors are located
on
the
transponder
receivertransmitter and should be set prior to
flight when required.
NOTE
To prevent zeroizing the mode 4
function of the transponder, when the
landing gear is down and the struts
compressed, place the CODE control
momentarily in HOLD position before
either transponder or aircraft power
is turned off. CODE HOLD condition
will only be removed when the struts
are extended and the MASTER
control is not in OFF position.
Modes
procedure.
1. MASTE R control -STBY (allow 2
minute warmup).
1, 2, 3/A, and/or mode 4 operating
2. MASTER CONTROL - LOW or
NORM. Set for required receiver
sensitivity.
3. 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.
4. MODE 1 code selectors - Set (if
applicable).
5. MODE 3/A code selectors - Set (if
applicable).
6. CODE control - Set (as required).
TM 55-1510-215-10
7. RAD TEST
desired).
MON
switch
-
MON
(if
8. TEST light - Monitor to determine when
transponder set is replying to a SIF
interrogation.
9. MODE 4 AUDIO LIGHT switch - Set (as
required to monitor mode 4 interrogations
and replies).
10. AUDIO and/or observe REPLY light Listen
and/or
observe
(mode
4
interrogations and replies.
11. IFF CAUTION light (instrument panel
fig. 2-22) - Monitor for an indication (light
will illuminate if transponder fails to reply
to a mode 4 interrogation or when the
mode 4 codes have been zeroized).
NOTE
If IFF CAUTION light illuminates,
check the position of the MODE 4 ON
OUT switch immediately.
If the
switch is in the ON position,
establish
contact
with
the
interrogating
station
on
a
communications set and explain the
situation.
The following self-tests will not
prevent transponder set replies to
external interrogations.
Test only
those modes that are being used.
1. M-1, M-2, M-3/A, M-C switches
-Momentarily set switches to test
position, one at a time.
2. TEST light -Observe for indication
as each code switch is held in TEST
position. If TEST light fails to light,
recheck applicable control settings.
If the light does not illuminate after
the control settings have been
rechecked, either the transponder or
the self-test feature is faulty.
identification
NOTE
The transponder set can make
identification-position replies while
2. IDENT MIC switch momentarily to IDENT,
directed.
Press
when
NOTE
Holding
circuits
within
the
transponder receiver-transmitter will
transmit
identification-position
signals
for
approximately
30
seconds. This is normally sufficient
time for ground control to identify the
aircraft's position.
During this 30
second period, it is normal procedure
to acknowledge via the aircraft
communications
set,
that
identification-position signals arc
being generated.
NOTE
NOTE
set
1. Modes 1, 2, and/or 3/A -Operating.
(3) Shutdown procedure.
(1) To
test transponder set operation in
modes 1, 2, C and/or 3/A.
(2) Transponder
operating procedure.
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.
position
To prevent zeroizing the MODE 4
function of the transponder. when
the landing gear is down and the
struts compressed, place the CODE
control
momentarily
in
HOLD
position before either transponder or
aircraft power is turned off. CODE
HOLD condition will only be removed
when the struts are extended and the
MASTER control is not in OFF
position.
1. CODE control - HOLD (when
required to hold MODE 4 code).
2. MASTER control - OFF.
e. Emergency Operation.
NOTE
The
MASTER
switch
on
the
transponder control panel shall be
placed in the EMER position only
under emergency conditions.
1. Modes 1, 2, and/or 3/A -Operating.
2. Master switch -EMER.
Change 5 3-35
TM 55-1510-215-10
3-30. Weather Radar Set (AN/APN-215(V) 1).
a. Description. The weather radar set (fig. 3-20)
provides a visual presentation of the general sky area of
approximately 120 degrees 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, such as thunderstorms and
hailstorms, in terms of distance and azimuth with respect
to the aircraft. The radar is capable of ground mapping
operations. Radar antenna stabilization signals are
supplied by the vertical gyro. The antenna is a flat-plate,
nose mounted unit (fig. 2-1). Protection is provided by a
5-ampere circuit breaker, placarded RADAR, located on
the right subpanel (fig. 2-7).
b. Controls and Functions.
(1) GAIN control.
Used to adjust radar
receiver gain in the MAP mode only.
(2) STAB OFF switch. Push on/push off type
switch. Used to control antenna stabilization signals.
(3) RANGE switches. Momentary action type
switches.
When pressed, clears the screen and
increases or decreases the range depending on switch
pressed.
(4) TILT control. Varies the elevation angle of
radar antenna 15 degrees up or down from horizontal
attitude of aircraft.
(5) 60° switch. Push on/push off type switch.
When activated, reduces antenna scan from 120
degrees to 60 degrees.
(6) TRACK switches. Momentary action type
switches. When activated, a yellow track line extended
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.
(7) HOLD switch.
Push on/push off type
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.
(8) Function switch. Controls operation of the
radar set.
1.
2.
3.
4.
5.
6.
7.
6.
9.
10.
11.
Figure 3-20. Weather Radar Control/Indicator (AN/APN-215(V))
3-36 Change 5
GAIN control/switch
STAB OFF switch
RANGE switches
TILT control
60° scan switch
TRACK switches
HOLD switch
Function switch
MODE select switches
NAV switch
BRT control
TM 55-1510-215-10
(a) OFF. Turns set off.
(b) STBY. Places set in standby mode.
This position also indicates a 90 second warmup delay
when first turned on.
(c) TEST.
Displays test pattern to
check for proper operation of the set. The transmitter is
disabled during this mode.
(d) ON. Places set in normal operation.
(9) 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 increase
or decrease depending on switch pressed. When either
top of bottom mode is reached, the opposite switch must
be pressed to further change the mode.
(1) Turn on procedure.
1. AVIONICS MASTER switch – ON.
2. Function switch - TEST or ON as
required (information will appear
after time delay period has elapsed).
(2) Initial adjustments operating procedure.
1. BRT control - As required.
2. MODE switches - Press and release
as required.
3. RANGE switches release as required.
Press
and
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.
(3) Test procedure.
1. Function switch - TEST.
(10) NAV switch. The words NO NAV will be
displayed in the lower left corner unless the indicator is
supplied with navigation data from other avionics not
covered by this manual.
2. RANGE switches - Press and
release as required to obtain 80 mile
display.
(11) BRT control.
brightness.
4. Screen - Observe for proper display.
The test display consists of two
green, two yellow, 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.
Used to adjust screen
c. Operating Procedures.
WARNING
Do not operate the weather radar set
while personnel or combustible
materials are within 18 feet of the
antenna flat-plate array. When the
weather radar set is operating, highpower radio-frequency energy is
emitted from the antenna flat-plate
array, which can have harmful effects
on the human body, and can ignite
combustible materials.
CAUTION
3. BRT control - As required.
(4) Weather observation operating procedure.
1. Function switch - ON.
2. MODE switches - Press and release
as required to select WX.
3. BRT control - As required.
Do not operate the weather radar set
in a confined space where the
nearest metal wall is 50 feet or less
from the antenna flat-plate array.
Scanning such surfaces may damage
receiver crystals.
4. TILT control - Adjust until weather
pattern is displayed. Include the
areas above and below the rainfall
areas to obtain a complete display.
Change 7 3-37
TM 55-1510-215-10
4. GAIN control - As required to
present usable display.
5. MODE switches - Press and release
to select WXA. Areas of intense
rainfall will appear as flashing red.
(6) Standby procedure.
6. TRACK switches - Press to move
track line through area of least
weather intensity.
Read relative
position, in degrees, in upper left
corner of screen.
1. Function switch - STBY.
(7) Shutdown procedure.
1. Function switch - OFF
2. AVIONICS MASTER switch - OFF.
NOTE
Refer to TM 11-5841-28-9-13 for both
weather
and
ground
mapping
functions of Radar Set AN/APN215(V).
(5) Ground mapping operating procedure.
1. Function switch - ON.
3-31. Pilot's Encoding Altimeter.
a. Description. The encoding altimeter (fig. 3-21)
provides the pilot with an indication of the aircraft's
altitude above sea level in addition to providing the
transponder with altitude information for use in mode C.
The encoding altimeter is protected by a 1-ampere
circuit breaker, placarded ENC ALT, located on the right
subpanel.
2. MODE switches - Press and release
as required to select MAP.
3. BRT control - As required.
b. Controls and Functions.
(1) Altitude alert annunciator.
The ALT
annunciator will illuminate when the aircraft is between
±300 feet and ±1000 feet of a preselected altitude.
When the aircraft is within ±300 feet of the selected
altitude
the
ALT
annunciator
will
extin-
1.
2.
3.
4.
5.
Altitude alert (ALT) annunciator
Altitude (drum) display
Barometric windows
Barometric correction knobs
Needle indicator
AP011844
Figure 3-21. Pilot's Encoding Altimeter
3-38 Change 7
TM 55-1510-215-10
guish. However, upon reaching the selected altitude, the
ALT annunciator will illuminate and stay illuminated for
two seconds, then extinguish. Upon deviating ±300 feet
from the selected altitude the ALT annunciator will again
illuminate and stay illuminated until the aircraft is ±1000
feet from the selected altitude, whereupon it will again
extinguish.
(2) Altitude (drum) display. The drum display
provides a full five figured digital readout of altitude in
increments of 100 feet. Black and white cross-hatching
is provided in place of the first digit of the counter to
command attention at altitudes below 10,000 feet.
Altitudes below sea level are indicated by a wavey blue
and white line which takes the place of the first digit of
the counter. In the event of power failure or a detection
of a malfunction by the failure monitor, a red and white
striped warning flag obscures the numbers on the digital
counter.
(3) Barometric windows .
Dual baroscales,
which permit barometric pressure settings in both
millibars and inches of mercury, display in the barometric
windows.
These settings are inputted with the
barometric correction knob, which simultaneously sets
millibars (MB) and inches (HG) of mercury.
(4) Barometric correction knob.
The
barometric correction knob is used to set the desired
altimeter setting in the display windows. The knob
simultaneously sets the millibars and inches of mercury.
(5) Needle indicator. The needle indicates
aircraft altitude in hundreds of feet with subdivisions at
twenty-foot intervals.
One revolution of the needle
equals 1000 foot of altitude change.
c. Operating Procedure.
1. Barometric Correction Knob - Set desired
altimeter setting in barometric window.
Note that needle indicator operates
properly.
2. Warning Flag - Check not vi sible.
NOTE
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 nullifies
use of the altimeter for IFR flight.
3-32.
Ground Proximity Altitude Advisory System
(GPAAS).
WARNING
The GPAAS will provide little, if any
warning for flight into abrupt vertical
terrain approaching a sheer wall if
there is little gradually rising terrain
before reaching the steep terrain.
The GPAAS will provide no warning
for stabilized descent into terrain
while the aircraft is in the landing
configuration, unless the aircraft is
following an operating electronic
glideslope or a correct minimum
descent altitude (decision height) has
been set on the radio altimeter
indicator.
a. Description.
The ground proximity altitude
advisory system (GPAAS) is provided to aid the flight
crew in terrain avoidance.
The GPAAS is a completely automatic system
(requiring no input from the crew) which continuously
monitors the aircraft's flight path at altitudes of between
100 and 2000 feet above ground level (AGL).
The GPAAS computer processes the data and,
when conditions warrant, selects the appropriate
digitized voice advisory/warning message from its
memory. This message is then announced over the
pilot's and copilot's audio systems. If the condition is not
corrected, the GPAAS will rearm, and will again
announce and repeat the warning if the condition recurs.
The GPAAS computer remains ready to announce a
different message during the intervals between
repetitions. All messages are disabled below 100 feet
AGL.
The GPAAS system receives 28 VDC power through
a 1-ampere circuit breaker placarded G.P.A.A.S.
POWER, located on the instrument panel.
(1) GPAAS switch-indicator lights. A switchindicator is located on the instrument panel. The upper
half of the switch-indicator (yellow) is placarded VOICE
OFF. The lower half is an indicator (red) only and is
placarded VA FAIL.
Depressing the upper (VOICE OFF) switch-indicator
disables the GPAAS voice advisory, and illuminates the
VOICE OFF indicator light.
The VA FAIL annunciator light (red) will illuminate
when the GPAAS fails.
Change 9 3-39
TM 55-1510-215-10
(2) GPAAS volume control.
A GPAAS
volume control placarded VOL, located on the
instrument panel, controls the audio volume of the
GPAAS advisory/warning messages down to a certain
minimum level.
(3) GPAAS Aural Warning Indications. The
following is a list of aural indications. Due to the
possibility of activating more than one condition at a
time, a warning priority has been established.
Figure 3-22. Ground Proximity Altitude Advisory System Controls and Indicators
3-40 Change 9
TM 55-1510-215-10
The highest priority message will be announced first. If
a higher priority item is received after a message is
started, voice annunciation of the higher priority
message shall be announced after a lower priority
message in progress at the end of the message
segment. It will not stop in the middle of a word. On
messages that are repeated three times at four second
intervals, the priority list will be scanned for higher
priority messages and will insert them in the interval
between the messages. The messages provided by the
system are listed in descending order of priority as
follows:
1. "Two thousand" at 2000 feet AGL.
2. "One thousand" at 1000 feet AGL.
3. "Nine hundred" at 900 feet AGL.
4. "Eight hundred" at 800 feet AGL.
5. "Seven hundred" at 700 feet AGL.
6. "Six hundred" at 600 feet AGL.
7. "Five hundred" at 500 feet AGL.
8. "Check gear" will be announced
immediately
after
500
foot
announcement if gear is not down.
9. "Four hundred" at 400 feet AGL.
10. "Check gear" will be announced
immediately
after
400
foot
announcement if gear is not down.
11. "Three hundred" at 300 feet AGL.
12. "Check gear" will be announced
immediately
after
300
foot
announcement if gear is not down.
13. "Two hundred" at 200 feet AGL.
14. "Check gear" will be announced
immediately
after
200
foot
announcement if gear is not down.
15. "One hundred" at 100 feet AGL.
16. "Check gear" will be announced
immediately
after
100
foot
announcement if gear is not down.
17. "Minimum, minimum" at decision
height.
18. "Localizer" at 1.3 to 1.5 dots either
side of center of beam. Will be
repeated three times at four second
intervals.
19. "Glideslope" at 1.3 to 1.5 dots above
or below center of beam.
Will be repeated three times at four
second intervals.
20. "Altitude, altitude" at excessive
deviation from altitude selected on
the altitude alerter.
21. "Check trim" when trim failure has
occurred. Will be repeated three
times at four second intervals.
22. "Autopilot" when
disconnected.
autopilot
has
The highest priority message will be announced first.
If a higher priority item is received after a message has
been started, voice annunciation of the higher priority
message shall immediately override the lower priority
message in progress at the end of the message
segment. It will not stop in the middle of a word. On
messages that are repeated three times at four second
intervals, the priority list will be scanned for higher
priority messages. If found, they will be inserted into the
interval between the messages.
b. Normal Operation.
(1) Turn-on procedure.
The GPAAS is
operable when the following conditions have been met:
1. Battery switch - ON.
2. Avionics master switch - On.
3. G.P.A.A.S. POWER circuit breaker
- SET.
4. RADIO ALTM circuit breaker - SET.
5. VA FAIL annunciator
Extinguished.
light
-
(2) GPAAS ground check.
1. GPAAS voice advisory VOL control
- Full clockwise.
2. VOICE OFF
Extinguished.
switch-indicator
-
3. Audio control panel - Set listening
audio level.
4. VA FAIL annunciator
Extinguished.
light
-
5. Radio altimeter DH SET control Set to 200 feet.
6. Radio altimeter TEST switch - Press
and hold. "Minimum, minimum" will
be annunciated once followed by
the illumination of the VA FAIL light.
Change 9 3-41
TM 55-1510-215-10
7. Radio altimeter
Release.
TEST
switch
-
c. GPAAS Modes of Operation.
The GPAAS
operates in the following modes of operation:
(1) Aural "TWO THOUSAND" advisory
(mode 1).
The aural advisory "TWO THOUSAND"
indicates that the aircraft is at a radio altitude of 2000
feet above ground level. This advisory is cancelled
when valid information from the radio altimeter is lost,
during climb, or whenever the aircraft is out of the
operating altitude range of the radio altimeter.
(2) Hundred foot increment aural altitude
advisories (mode 2).
The aural advisories "ONE
THOUSAND, NINE HUNDRED, EIGHT HUNDRED,
SEVEN HUNDRED, SIX HUNDRED, FIVE HUNDRED,
FOUR
HUNDRED,
THREE
HUNDRED,
TWO
HUNDRED, ONE HUNDRED" indicate that the aircraft is
at the associated radio altitude in feet above ground
level. This advisory is cancelled when valid information
from the radio altimeter is lost, during climb, or whenever
the aircraft is out of the operating altitude range of the
radio altimeter.
(3) Aural "LOCALIZER" advisory (mode 3).
The aural advisory "LOCALIZER" indicates that the
aircraft has deviated from the center of the localizer
beam in excess of 1.3 to 1.5 dots. The localizer advisory
is armed when a valid localizer signal is detected and
the aircraft is below 1000 feet above ground level. It will
be repeated no more that 3 times at 4 second intervals
unless the aircraft is returned to less than 1.3 to 1.5 dots
from the center of the localizer course. The localizer
advisory is disabled when a valid localizer signal has
been lost, during climb, below the decision height set on
the radio altimeter, or if the navigation receiver is not
tuned to a localizer frequency.
(4) Aural "CHECK GEAR" advisory (mode 4).
The aural "CHECK GEAR" advisory indicates that the
aircraft has descended to 500 feet AGL and the landing
gear is not down. This advisory is repeated once at 100
foot intervals down to 100 feet AGL.
3-42 Change 9
(5) Aural "GLIDESLOPE" advisor (mode 5).
The aural advisory "GLIDESLOPE" indicates that the
aircraft has exceeded 1.3 to 1.5 dots above or below the
center of the glideslope beam. The glideslope advisory
is armed when a valid glideslope signal is detected and
the aircraft is below 1000 feet AGL. It will be repeated
no more than three times at 4 second intervals unless
the aircraft is returned to less than 1.3 to 1.5 dots from
the center of the beam. The glideslope advisory is
disabled upon loss of a valid glideslope signal, during
climb, on a localizer back course, below the decision
height set on the radio altimeter or, if the navigation
receiver is not tuned to a localizer frequency. This
advisory is inhibited by the weight on wheels strut
switch.
(6) Aural advisory "MINIMUM, MINIMUM"
(mode 6). The aural advisory "MINIMUM, MINIMUM"
indicates that the aircraft at the radio altitude selected by
the crew with the radio altimeter indicator decision height
knob. This advisory is cancelled when valid information
from the radio altimeter is lost, during climb, whenever
the aircraft is above 1000 feet AGL, or whenever the
aircraft is out of the operating altitude range of the radio
altimeter.
(7) Aural "ALTITUDE, ALTITUDE" advisory
(mode 7). The aural advisory "ALTITUDE, ALTITUDE"
indicates the approach to a preselected altitude as the
aircraft reaches a point 1000 feet from the selected
altitude or, after reaching the selected altitude, when the
aircraft deviates more than 250 feet from the selected
altitude.
(8) Aural "CHECK TRIM, CHECK TRIM,
CHECK TRIM" advisory. The aural advisory "CHECK
TRIM, CHECK TRIM, CHECK TRIM" indicates that the
autopilot has had a trim failure.
(9) Aural "AUTOPILOT" advisory. The aural
advisory "AUTOPILOT" indicates that the autopilot has
disengaged.
d. Emergency procedures. If an emergency or
malfunction makes it necessary to disable the GPAAS,
pull the G.P.A.A.S. POWER circuit breaker located on
the instrument panel (GPAAS audio may be turned off
by depressing the VOICE OFF switch).
TM 55-1510-215-10
CHAPTER 4
MISSION EQUIPMENT
This aircraft is not equipped with mission equipment.
4-1/(4-2 blank)
TM 55-1510-215-10
CHAPTER 5
OPERATING LIMITS AND RESTRICTIONS
Section I. GENERAL
5-1. Purpose.
5-3. Exceeding Operational Limits.
This chapter identifies or refers to all important
operating limits and restrictions that shall be observed
during ground and flight operations.
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.
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.
5-4. Minimum Crew Requirements.
The minimum crew required for aircraft operation is
one pilot. Additional crewmembers as required will be
added at the discretion of the commander, in
accordance with pertinent Department of the Army
regulations.
Section II. SYSTEM LIMITS
5-5. Instrument Markings.
5-7. Instrument Glass Alignment Marks.
Several instruments display operating limitations
(fig. 5-1). The operating limitations are color coded on
the instrument faces. Color coding of each instrument is
explained in the illustration. The instrument illustration
also denotes, in bold type, the fuel grade upon which
limits are based.
Limitation
markings
consist
of
strips
of
semitransparent color tape which adhere to the glass
outside of an indicator dial. Each tape strip shall align to
increment marks on the dial face so correct operating
limits are portrayed. The pilot should occasionally verify
alignment of the glass to the dial face. For this purpose,
all engine instruments (except fuel flow meters) have
short, vertical white alignment marks extending from the
bottom part of the dial glass onto the fixed base of the
indicator. These slippage marks appear as a single
vertical line when limitation markings on the glass
properly align with reading increments on the dial face.
However, the slippage marks appear as separate radial
lines when a dial glass has rotated.
5-6. 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 on the dial faces of these instruments indicate
the limit above or below which continued operation is
likely to cause damage or shorten life. The GREEN
markings on instruments indicate the safe or normal
range of operation.
The YELLOW markings on
instruments 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
instruments indicate flap operating range.
5-8. Propeller Limitations.
Propeller limitations consist of RPM limits and
situation limits for the use of reverse pitch. The normal
propeller operating range (green arc) extends from 1800
to 2200 RPM, with a red line at 2200 RPM. However,
the actual governor controlled limits are 1750 to 2332
RPM.
5-1
TM 55-1510-215-10
FUEL GRADE JP-4
INTERSTAGE TURBINE TEMPERATURE
750°C MAXIMUM
TORQUE
1315 FT LB MAXIMUM
PROPELLER TACHOMETER
1800-2200 RPM NORMAL OPERATING RANGE
2200 RPM MAXIMUM
TURBINE TACHOMETER (N1 SPEED)
(Figure 5-1 Sheet 1 of 3)
(Figure 5-1 Sheet 2 of 3)
(Figure 5-1 Sheet 3 of 3)
101.5% MAXIMUM
G
R
AV 095183.1
Figure 5-1. Instrument Markings (Sheet 1 of 3)
5-2
TM 55-1510-215-10
OIL PRESSURE
40 PSI MINIMUM
65-85 PSI NORMAL OPERATING RANGE
OIL TEMPERATURE
0°C MINIMUM
99°C MAXIMUM USING ENGINE OIL SPEC. ML-L23699
*85°C MAXIMUM USING ENGINE OIL SPEC. MLL-7808
0-10°C CAUTIONARY
10-99°C NORMAL OPERATING RANGE USING
ENGINE OIL SPEC. MIL-L-23699
*10-85°C NORMAL OPERATING RANGE USING
ENGINE OIL SPEC. ML-L-7808
SUCTION
HG 4.5 - 5.2 NORMAL OPERATING RANGE
*FOR OPERATION BELOW -40°F (-40°C)
G
R
Y
AV 095183.2
Figure 5-1. Instrument Markings (Sheet 2 of 3)
Change 5 5-3
TM 55-1510-215-10
AIRSPEED INDICATOR
75-130 KNOTS NORMAL OPERATING RANGE
WITH FLAPS
92-208 KNOTS NORMAL OPERATING RANGE
208 KNOTS MAXIMUM
PROPELLER DEICER AMMETER
14-18 AMPERES NORMAL OPERATION
DE-ICING PRESSURE
12-20 PSI NORMAL OPERATING RANGE
20 PSI MAXIMUM
G
R
AV 095183.3
Figure 5-1. Instrument Markings (Sheet 3 of 3)
5-4
TM 55-1510-215-10
The lower limit of 1725 to 1775 RPM is controlled by the
primary governor. The upper limit of 2332 RPM is
maintained by the power turbine governor (should the
primary and overspeed governors fail). During reverse
pitch operation, the power turbine governor prevents
propeller speed from exceeding 2040 RPM. Refer to
chapter 2 for propeller governor details.
5-9A. Autopilot Limitations.
a. During autopilot operation, one pilot must be
seated at the controls with seat belt fastened.
b. Autopilot and yaw damper must be disengaged
during takeoff or landing.
c. The system is approved for
operation only (approach mode selected).
5-9. Starter Limitations.
The starters in this aircraft are limited to an operating
period of 40 seconds on, then 60 seconds OFF, for two
starter operations.
After two starter operations the
starter shall be operated for 40 seconds on, then 30
minutes OFF.
Category
1
d. Do not operate autopilot with flaps extended
beyond the approach position.
e. Maximum altitude for autopilot operation is
25,000 feet.
Section III. POWER LIMITS
5-10. Engine Limitations.
Operation of the T74-CP-700 engines is monitored
by instruments with the operating limits marked on the
face of the dial. Table 5-1 shows all operating conditions
and limits for the engine.
a. Engine operation using only the engine driven
primary (high pressure) fuel pump without auxiliary fuel
pump or engine-driven boost pump fuel pressure is
limited to 10 accumulative hours.
All time in this
category shall be entered on DA Form 2408-13 for the
attention of maintenance personnel.
b. Use of aviation gasoline is time-limited to 150
hours of operation during any Time-Between-Overhaul
(TBO) period. It may be used in any quantity with
primary or alternate fuel.
of 850°C is time-limited to two seconds duration when
accelerating engine.
b. During engine starting the temperatures and
time limits listed in the Engine Operating Limitations
Chart shall be observed (Table 5-1). When these limits
are exceeded, the incident shall be entered as an engine
discrepancy on DA Form 2408-13. It is particularly
important to record the amount and duration of
overtemperature.
c. Whenever the prescribed engine overspeed limit
or engine RPM operating limit is exceeded the incident
must be reported as an engine discrepancy on DA Form
2408-13.
It is particularly important to record the
maximum percent of RPM registered by the tachometer,
and the duration of overspeed.
5-12. Power Definitions For Engine Operation.
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.
5-11. Overtemperature and Overspeed Limitations.
a. Whenever the limiting temperatures listed in the
Engine Operating Limitations Chart are exceeded and
cannot be controlled by retarding the power levers and
the engine shall be shut down or a landing shall be
made as soon as possible. It should be noted that
maximum observed interstage turbine temperature (ITT)
The following definitions describe the engine power
ratings listed in the Engine Operating Limitations Chart,
Table 5-1.
a. Takeoff Power. The maximum power available
from an engine for takeoff, limited to periods of five
minutes duration.
b. Maximum Power.
The maximum power
available from an engine for use during an emergency
operation.
c. Normal Rated Climb Power. The maximum
power available from an engine for continuous normal
climb operations.
d. Normal Rated Power. The maximum power
available from an engine for continuous operation in
cruise (with lower ITT limit than normal rated climb
power).
Change 5 5-5
TM 55-1510-215-10
Table 5-1. Engine Operating Limitations
OPERATING LIMITS
POWER
RATING
TAKEOFF
5
MAXIMUM
NORMAL RATED
CLIMB
NORMAL RATED
HIGH IDLE (73% N1)
LOW IDLE (54% N1)
STARTING
ACCELERATION 7
MAX REVERSE
PROP FEATHER
N1%
h
N2 RPM
(PROP)
OIL
PRESS
PSIG
750
750
725
101.5
101.5
---
2200
2200
2200
705
--685 6
1090 (2 SEC)
850
750
-----
--------102.6
88
-----
2200
---------------
H
TORQUE
FT LB
H MAX
OBSERVED
ITT
5 MINUTES
CONTINUOUS
CONTINUOUS
1315
1315
1315Q
CONTINUOUS
CONTINUOUS
CONTINUOUS
--2 SECONDS
1 MINUTE
CONTINUOUS
2 SECONDS
1315Q
------1500
--525
1500
MAX
TIME
1
GENERATOR LOAD
MINIMUM ENGINE RPM
0 to .5 Load
50%
.5 to .75 Load
57%
.75 to .90 Load
59%
.90 to 1.0 Load
63%
2
3
4
OIL
TEMP °C
OIL
TEMP °C
65-85
65-85
65-85
10-99
10-99
0-99
10-85
10-85
0-85
65-85
--40 (MIN)
----65-85
-----
0-99
0-99
-40-99
-40 (MIN)
0-99
0-99
-----
0-85
0-85
-40-85
-40 (MIN)
0-85
0-85
-----
H
Each column is a separate limitation. The stated limits do not necessarily occur simultaneously.
h
The limit values within the N2 RPM (PROP) column are not propeller limitations. These values specify propeller
RPMs which correspond to stress limits of the engine power section.
Q
This is an engine gearbox torque limit which shall not be exceeded under steady state or continuous engine
operation.
1
For every 10° below -30°C ambient temperature,
reduce maximum allowable N1 by 2.2%.
4
85°C maximum when using engine oil spec.
MIL-L-7808.
2
Oil pressure below 65 PSIG is undesirable at power
settings above 75% N1. Flight may be completed at a
reduced power setting, but the cause of low oil
pressure should be corrected prior to next flight. Oil
pressure below 40 PSIG requires engine shutdown.
5
This power rating is intended for emergency use at
the discretion of the pilot.
6
High ITT may be decreased by reducing accessory
load and/or increasing N1 speed.
99°C maximum when using engine oil spec.
MIL-L-23699.
7
High generator loads at low N1 speeds may cause
the ITT acceleration temperature limit to be
exceeded. Observe the above generator limits.
3
APOO5761
5-6 Change 7
TM 55-1510-215-10
Figure 5-2. Takeoff Temperature Limitations
5-7
TM 55-1510-215-10
5-13. Ambient Temperature Takeoff Limitation.
A limitation based on pressure altitude and ambient
temperature prohibits aircraft takeoff under certain high
ambient temperature conditions.
The Takeoff
Temperature Limitations Chart determines this limitation
(fig. 5-2).
Section IV. LOADING LIMITS
5-14. Center-of-Gravity Limitations.
5-15. Weight Limitations.
The maximum designed gross weight is 9650
pounds for takeoff and 9168 pounds for landing.
Maximum ramp weight is 9705 pounds.
CAUTION
When oxygen bottles are not
installed in rear cabin, check aircraft
loading to avoid exceeding forward
CG limit. Flight operations involving
forward loadings while the oxygen
bottles are removed are critical, i.e.,
pilot and copilot only or forward
loaded cargo and/or passengers.
5-16. Floor Loading Limits.
The floor is stressed to support a maximum vertical
load of 200 pounds per square foot on the seat tracks.
Secondary supports should be used to distribute highly
condensed weights evenly over the cargo areas.
Center-of-gravity
limits
and
instructions
for
computation of the center of gravity are contained in
chapter 6.
Section V. AIRSPEED LIMITS
5-17. Airspeed Limitations.
Airspeed indicator readings contained in procedures,
text, and illustrations throughout the Operators 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). Refer to the
Airspeed Calibration Chart in chapter 7.
5-18. Maximum Allowable Airspeed Vmo.
NOTE
Altitude variations do not affect the
limits shown in the Flight Envelope
Chart (fig. 5-3).
5-20. Landing Gear Extension Speed.
The airspeed limit for extending the landing gear and
for flight with the landing gear extended is 156 KCAS
(154 KIAS). Above this speed, air loads may damage
the landing gear doors or their operating mechanisms.
The maximum allowable airspeed under all
conditions is 208 KCAS (208 KIAS) Vmo. Operation
above this speed may cause structural damage.
5-21. Landing Gear Retraction Speed.
5-19. Turbulence Penetration Speed.
The airspeed limit for retracting the landing gear is
130 KCAS (127 KIAS). If the landing gear is retracted
above this speed, air loads may damage the landing
gear operating mechanism.
The maximum safe penetration speed in severe
turbulence is 169 KCAS (168 KIAS).
5-8 Change 7
TM 55-1510-215-10
5-22. Wing Flap Extension Speeds.
5-24. Minimum Single-Engine Control (V mc ).
The airspeed limit for lowering the flaps to the
approach position (35%) is 174 KCAS (173 KIAS). The
airspeed limit for extending the flaps from the approach
position to the full down, or any intermediate position is
130 KCAS (127 KIAS). If indicated airspeeds exceed
these limitations the flaps or their operating mechanisms
may be damaged.
The placard Vmc airspeed of 92 knots is the
calibrated airspeed value for sea level standard day
conditions. For Vmc as a function of pressure altitude
and atmospheric temperature, refer to Chapter 7,
Section X.
5-23. Cockpit Vent/Storm Window Speed.
The maximum design maneuvering speed is 169
KCAS (168 KIAS).
No airspeed limitations are imposed on cockpit
vent/storm windows.
5-25. Maximum Design Maneuvering Speed Va .
Section VI. MANEUVERING LIMITS
5-26. Maneuvers.
1.68G's with wing flaps up or a positive load factor of
2.0G's, or negative G's with wing flaps down.
WARNING
b. The maximum design maneuvering speed is 169
KCAS (168 KIAS). For turbulent air penetration use an
airspeed of 169 KCAS (168 KIAS). Avoid over action on
power levers, turn off autopilot altitude hold, keep wings
level, maintain attitude and avoid use of trim. Do not
chase airspeed and altitude. Penetration should be at
an altitude which provides adequate maneuvering
margins when severe turbulence is encountered.
Operation beyond the structural
capabilities of the aircraft will result
in complete failure of one or more
airframe components.
a. The following maneuvers are prohibited.
5-27. Bank and Pitch Limit.
(1) Spins.
a. Bank limit is 60°.
(2) Aerobatics of any kind.
(3) Abrupt
(168 KIAS).
maneuvers
above
169
KCAS
b. Pitch limit is 30° above or below the horizon.
(4) Any maneuver which results in a positive
load factor of 3.70G's or a negative load factor of
Change 7 5-9
TM 55-1510-215-10
FLIGHT ENVELOPE
FLIGHT ENVELOPE
U-21G
T74-CP-700
NORMAL OPERATING RANGE
OPERATION WITH FLAPS DOWN
PROHIBITED OPERATION
Vf - DESIGN FLAP SPEED
NOTE
Va - DESIGN MANEUVERING SPEED
Vle - DESIGN GEAR EXTENDED SPEED
BELOW 30,000 FEET, ALTITUDE
VARIATIONS DO NOT AFFECT
THE LIMITS SHOWN
G
R
Vmo - MAX OPERATING LIMIT SPEED
AP 000578
Figure 5-3. Flight Envelope Chart
5-10
TM 55-1510-215-10
Section VII. ENVIRONMENTAL RESTRICTIONS
5-28. Altitude Limitations.
crewmember in compliance with AR95-1 for planned
flight above 10,000 feet.
The maximum altitude that the aircraft may be
operated at is 26,150 feet. The single engine service
ceiling is 12,000 feet.
5-29. Auto Ignition During Night Operation.
Auto ignition shall be used during night operations at
or above 14,000 feet.
5-30. Oxygen Requirements.
5-31. Flight Under IMC (Instrument Meteorological
Conditions).
This aircraft is approved for flight under instrument
conditions.
5-32. Wind Limitations.
The maximum demonstrated crosswind for landing is
25 knots. Refer to chapter 7 for wind limitations.
One oxygen mask and an adequate supply of
aviators breathing oxygen shall be provided each
Section VIII. OTHER LIMITATIONS
5-33. Passenger Seats.
The forward passenger seat on each or either side
of the aircraft may face aft, but only seats approved for
aft facing installation may be installed facing aft. When
an approved seat faces aft, the occupant shall not weigh
more than 170 pounds. The headrest and seat back,
when occupied, must be in the fully upright position for
takeoff and landing.
5-11/(5-12 blank)
TM 55-1510-215-10
CHAPTER 6
WEIGHT/BALANCE AND LOADING
SECTION I. GENERAL
6-1. Extent of Coverage.
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 basic weight and moment of the aircraft, any
combination of weight and balance can be computed.
6-2. Class.
Army Model U-21G is in Class 2.
Additional
directives governing weight and balance of Class 2
aircraft forms
6-3. Aircraft Compartment and Stations.
The aircraft is separated into two compartments
associated with loading of the aircraft.
These
compartments are the cockpit and the cabin. Figure 6-1
illustrates
the
general
description
of
aircraft
compartments.
SECTION II. WEIGHT AND BALANCE
6-4. Purpose.
The data to be inserted on weight and balance
charts and forms are applicable only to the individual
aircraft, the serial number of which appears on the title
page of the booklet entitled WEIGHT AND BALANCE
DATA supplied by the aircraft manufacturer and on the
various forms and charts which remain with the aircraft.
The charts and forms referred to in this chapter may
differ in nomenclature and arrangement from time to
time, but the principle on which they are based will not
change.
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. Deleted.
6-8. Deleted.
6-9. Deleted.
6-5. Charts and Forms.
The standard system of weight and balance control
require the use of several different charts and forms.
Within this chapter, the following are used:
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 (Transport).
6-6. Responsibility.
The aircraft manufacturer inserts all
identifying data on the title page of the booklet
WEIGHT AND BALANCE DATA and on the
charts and forms. All charts, including one
aircraft
entitled
various
sample
Change 8 6-1
TM 55-1510-215-10
1
NOSE AVIONICS COMPARTMENT
7
OXYGEN CYLINDER (11 CU. FT.)
15 STARTER-GENERATOR
2
PILOT - COPILOT SEATS
8
OXYGEN CYLINDER (64 CU. FT.)
16 OXYGEN CYLINDER
(64 CU. FT.)
3
BATTERY (R WING ONLY)
9
STROBE BEACON POWER SUPPLY
4
FWD PASSENGER SEATS (AFT FACING)
10
NACELLE TANK
5
CABIN TABLES
11
WING FUEL TANKS
6
AFT PASSENGER SEATS (FWD FACING)
12
WING FUEL TANKS
13
WING FUEL TANKS
14
WING FUEL TANKS
17 AFT AVIONICS SHELF
18 COCKPIT EMERGENCY
ENTRANCE/EXIT HATCH
19 REFRESHMENT CABINET
AND MAGAZINE RACK
AP 00525
Figure 6-1. Aircraft Compartment and Station
6-2
TM 55-1510-215-10
6-10. Chart C - Basic Weight and Balance Record,
DD Form 365-3.
6-11. Weight and Balance Clearance Form F, DD
Form 365-4 (Transport).
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
and moment/1000 entry is considered the current weight
and balance status of the basic aircraft.
Refer to TM 55-1500-342-23 for Form F 365-4
instructions. Refer to figures 6-4, 6-6, and 6-7 for
moments.
Change 8 6-3
TM 55-1510-215-10
Deleted.
Figure 6-2. Basic Weight and Balance Record
6-4 Change 8
TM 55-1510-215-10
Deleted.
Figure 6-3. Weight and Balance Form DD 365-4 Transport
All Data on Page 6-6 Deleted
Change 8 6-5
TM 55-1510-215-10
SECTION III. FUEL/OIL
6-12. Fuel Load Fuel loading imposes a restriction on
the amount of cargo which can be carried. The required
fuel must first be determined, then that weight subtracted
from the total weight of cargo and fuel. Cargo weights
up to and including the remaining allowable capacity can
be subtracted directly from the weight of cargo and fuel.
As the fuel load is increased, the cargo capacity is
reduced.
6-13. Fuel and Oil Data
a. Fuel Moment Chart (fig. 6-4). This chart shows
fuel moment/1000 given gallons and pounds of JP-4
fuel.
b. Oil Data. Total oil weight is included in the basic
weight of the aircraft. Servicing information is provided
in paragraph 2-87.
SECTION IV. PERSONNEL
6-14. Aircraft Personnel Cargo Features.
a. Cabin. The cabin extends from the back of the
cockpit to the aft cabin wall. This area, when completely
stripped of seating provisions, provides as 230.0 cubic
foot cargo space with maximum measurements of 155.0
inches long, 57.0 inches high and 55.0 inches wide.
Access is gained through the rectangular main entrance
door which measures 51.5 inches high and 26.5 inches
wide. In conjunction with the main entrance door, a
cargo door is provided to give an opening of 51.5 inches
high and 53.5 inches wide. The floor is designed to
withstand cargo loads of 200 pounds per square foot.
Refer to Section V to determine maximum cargo
capacity and load position. Payload, must be limited in
conjunction with fuel loading to stay within the design
gross weight limitations.
one litter on the left side of the cabin. One three-man
bench seat is provided at the forward right side of the
compartment for ambulatory patients or medical
personnel (fig. 6-6, sheet 2).
d. Staff Transport Features. Two versions of staff
transport cabin seating arrangements are available. The
first version has three forwardfacing chair seats secured
to the floor ratings on each side of the center aisle
(fig. 6-5, sheet 5). The second version has two forwardfacing chair seats and two aft-facing chair seats secured
to the floor railings on each side of the center aisle.
Also, foldout tables are attached to the compartment wall
between each pair of facing seats. A storage cabinet is
located along the right wall aft of the main entrance door
(fig. 6-5, sheet 4).
NOTE
b. Troop Cargo Features.
The troop transport
version is designed to carry 10 combat-equipped troops
on center-facing bench-type seats (fig. 6-5, sheet 1).
c. Staff Transportation Features.
The air
ambulance version is equipped for three litters and has
three seat positions. Two litters are to be placed one
above the other, on the right aft side of the cabin, and
The forward passenger seat on each
or either side of the aircraft may face
aft but only seats approved for an aftfacing installation may be installed
facing aft. When an approved seat
faces aft, the aft-facing occupant
shall weigh not more than 170
pounds.
Change 8 6-7
TM 55-1510-215-10
FUEL MOMENT CHART
FUEL MOMENT
U-21G
T74-CP-700
G
AP012885
Figure 6-4. Fuel Moment Chart
6-8 Change 8
TM 55-1510-215-10
TROOP TRANSPORT CONFIGURATION
(Figure
(Figure
(Figure
(Figure
(Figure
6-5
6-5
6-5
6-5
6-5
Sheet
Sheet
Sheet
Sheet
Sheet
1
2
3
4
5
of
of
of
of
of
5)
5)
5)
5)
5)
Figure 6-5. Personnel Loading (1 of 5)
6-9
TM 55-1510-215-10
AIR AMBULANCE VERSION
Figure 6-5. Personnel Loading (2 of 5)
6-10
TM 55-1510-215-10
CARGO VERSION
Figure 6-5. Personnel Loading (3 of 5)
6-11
TM 55-1510-215-10
COMMAND TRANSPORT VERSION
Figure 6-5. Personnel Loading (4 of 5)
6-12
TM 55-1510-215-10
STAFF TRANSPORT VERSION
Figure 6-5. Personnel Loading (5 of 5)
6-13
TM 55-1510-215-10
PERSONNEL MOMENTS
STAFF TRANSPORT
(Figure 6-6 Sheet 1 of 3)
(Figure 6-6 Sheet 2 of 3)
(Figure 6-6 Sheet 3 of 3)
Figure 6-6. Personnel Moments (1 of 3)
6-14
PERSONNEL MOMENT
U-21G
T74-CP-700
TM 55-1510-215-10
PERSONNEL MOMENTS
AMBULANCE CONFIGURATION
PERSONNEL MOMENTS
U-21G
T74-CP-700
Figure 6-6. Personnel Moments (2 of 3)
6-15
TM 55-1510-215-10
PERSONNEL MOMENTS
TROOP CONFIGURATION
Figure 6-6. Personnel Moments (3 of 3)
6-16
PERSONNEL MOMENT
U-21G
T74-CP-700
TM 55-1510-215-10
d. Air Ambulance Cargo Features.
The air
ambulance version is equipped for three litters and has
three seat positions. Two litters are to be placed one
above the other, on the right aft side of the cabin, and
one litter on the left side of the cabin. One three-man
bench seat is provided at the forward right side of the
compartment for ambulatory patients or medical
personnel (fig. 6-7, sheet 4).
(3) After patients have been strapped to the
litters, load them in the following sequence:
6-15. Personnel Loading and Unloading.
(4) Confer with medical attendants to
determine maximum altitude patients can endure (unless
respirators are to be used).
a. Troop Seat Installation. The center facing bench
seats are fastened to the sidewall and to the inboard
seat rail. See chapter 2 for a description of seat
installation.
b. Cabin Safety Belts and Harnesses. The troop
transport and air ambulance versions are not equipped
with cabin occupant restraints other than those built into
the litters. The staff transport versions are equipped with
lap belts attached to the seat tracks.
c. Comfort and Emergency Provisions. The entire
cabin is padded and there are air vents, lights, and
oxygen outlets in the ceiling. A relief tube is located just
aft of the main entrance door. A fire extinguisher is
located under the copilot's seat. For description and
operation of emergency exits see chapter 9.
d. Litters. The air ambulance version is readied by
removing or folding all troop bench seats, except one
three-man seat unit for medical attendants, and installing
one single litter bed and one double-deck litter bed unit
(3 litters) in the cabin. The following procedure is
recommended for loading litter patients in the Air
Ambulance Version:
(a) Top right side.
(b) Bottom right side.
(c) Left side.
(5) Inform medical attendants of any expected
weather hazards (turbulence, etc.).
(6) Complete normal pre-flight briefing for
medical attendants.
6-16. Personnel Load Computation.
When aircraft are operated at critical gross weights,
the exact weight of each individual occupant plus
equipment should be used. If weighing facilities are not
available, or if the tactical situation dictates otherwise,
loads shall be computed as follows:
a. Combat
individual.
Equipped
Soldiers:
240
lb.
per
b. Combat Equipped Paratroopers: 260 lb. per
individual.
c. Crew and Passengers with no Equipment:
complete weight according to each individual's estimate.
NOTE
(1) Confirm the presence of installation plates
for each litter bracket required (each litter requires two
plates and two brackets).
(2) Confirm the
straps and litter brackets.
presence
of
suspension
Personnel loading configurations
other than those shown in the
Personnel Loading Diagram (fig. 6-7)
shall be computed using Cargo
Moment Chart in Section II.
6-17
TM 55-1510-215-10
Section V. CARGO LOADING
6-17. Air Cargo Features.
The air cargo version lacks seating within the cabin.
Eighteen tiedown rings, which swivel within cuplike
depressions, are provided in the compartment flooring.
The cleared interior space accommodates varied cargo
arrangements within the 3000 pound total and 200
pounds per square foot cargo limitations (fig. 6-7, sheet
5). The main entrance door measures 51.5 inches high
and 26.5 inches wide. Loading work is facilitated by an
outward opening cargo door which mates with the main
entrance door.
Opening both doors provides an
entrance 51.5 inches high and 53.5 inches wide. The
cabin has fittings for three litters (fig. 6-7, sheet 4).
6-18. Aerial Delivery System.
WARNING
Procedures for aerial delivery of
personnel and cargo have not been
developed.
Paradrops with the
standard 15-foot static line cannot be
performed without the danger of
interference of the trailing static line
with the elevator control surface.
WARNING
The cargo door is a structural panel
and shall be closed for flight.
A static line is provided along the center of the
ceiling for equipment and personnel drop. The main
entrance door (air-stair) may be removed, on the ground,
for personnel or cargo drops.
6-19. Cargo Center-of-Gravity Planning.
CAUTION
If the aft-mounted 64 cubic feet
oxygen cylinder is removed and the
nosemounted 64 cubic feet oxygen
cylinder is installed in the aircraft,
ballast must be placed aft to hold the
aircraft center-of-gravity within limits.
6-18
The cargo loading (fig. 6-5) must be planned so that
the center-of-gravity of the loaded aircraft will fall within
the operating limits given in chapter 5. The allowable
cargo center-of-gravity range is determined by these
operating limits. All cargo capacities are based on the
aircraft operating weight with all passenger seats
removed. The weight and location of passengers and
crew members must be considered when determining
the cargo center-of-gravity location and cargo weight
capacity. The maximum gross weight of the aircraft is
9650 pounds for takeoff and 9168 pounds for landing.
The approximate total weight available for fuel and
cargo/personnel, for each configuration, is as follows:
a. Troop Transport Version: approximately 3950
pounds available for crew, fuel and cargo/personnel.
b. Staff Transport Version: approximately 3720
pounds available for crew, fuel and cargo/personnel.
c. Air Ambulance Version: approximately 3930
pounds available for crew, fuel and cargo/personnel.
d. Air Cargo Version: approximately 4110 pounds
available for crew, fuel and cargo/personnel.
6-20. Cargo Center-of-Gravity Computation.
Table 6-1 shows an example of cargo center-ofgravity planning.
6-21. Preparation of General Cargo.
Before loading cargo, loading personnel should
determine such data as weight, dimensions, center-ofgravity, and contact areas of the individual cargo items
for use in positioning the load. For final load position of
the aircraft see weight and balance computation in
Section II.
6-22. Preparation of Cabin for Loading.
Air cargo conversion is made by removing the troop
bench seats creating storage space within the cabin.
Cargo tie-down fittings are flush-mounted in the floor
structure.
TM 55-1510-215-10
CARGO MOMENT
CARGO MOMENT
U-21G
T74-CP-700
Figure 6-7. Cargo Moments
6-19
TM 55-1510-215-10
Table 6-1. Cargo Center-of-Gravity Location Planning Example
ITEM
WEIGHT
Aircraft basic weight from chart c
Add: Fuel unusable
Pilot
Operating weight
Cargo (2000 pounds)
Cargo
Cargo
Cargo
Fuel (296 gallons)
Nacelle fuel cell
Wing fuel cell
Takeoff weight
STATION
MOMENT
5421
24
200
5701
600
600
600
200
146.7
140
129
145.6
160
180
200
220
795316
3360
25800
830132
96000
108000
120000
44000
741
1183
9625
131
168
155.2
97000
199000
1494132
NOTES:
1.
The cargo center-of-gravity is located at station 184.0. To get the landing weight of the aircraft down to a
maximum of 9168 pounds, 457 pounds approximately (70 gallons) of fuel must be consumed.
2.
This chart is for planning purposes only. Final aircraft loading operations and weight and balance
computation must be checked for the particular aircraft. (See Section II).
6-23. Load Planning.
A thorough check by the pilot before each flight will
insure the best loading arrangement. The amount of
control surface deflection required to correct for forward
or aft CG conditions will restrict maneuverability of the
aircraft. The stability and controllability of the aircraft is
improved, particularly at low airspeeds, by loading as
close to the neutral position as possible. The degree of
load planning will vary with each operation, depending
on the amount and bulk of the load. The basic factors to
be considered in any loading situation are as follows:
Secondary supports should be used to distribute highly
condensed weights evenly over the cargo areas.
f. Cargo destination should be considered when
applicable. If part of the cargo is to be removed at an
intermediate stop, the cargo should be arranged
accordingly.
g. All cargo must be adequately secured to prevent
damage to the aircraft, other cargo, or the item itself.
6-24. Loading Procedure.
a. The location of the cargo must be planned so
that the center-of-gravity of the loaded aircraft will be
within the operating limits.
b. The total weight of the loaded aircraft must not
exceed the maximum allowable gross weight.
c. Cargo must be arranged to permit access to all
emergency equipment and exits during flight.
d. Bulk cargo must be properly arranged to prevent
damage to fragile items.
e. Floorboard and bulkhead structural capacity
must be considered in the loading of heavy or sharpedged containers and equipment.
6-20
Loading of cargo is accomplished through the main
cabin entrance and cargo doors.
Extreme caution
should be exercised to prevent damaging the wing flaps,
doors, floorboards, seat tracks, upholstery, etc.
Personnel shall observe NO STEP areas. Cargo should
have at least one secondary support on each side of the
compartment. The floor is stressed to support a vertical
load of 200 pounds per square foot.
6-25. Securing Loads.
Various aircraft maneuvers tend to move the cargo
vertically, sideways, forward, rearward, or in any
combination of directions.
For this reason all
TM 55-1510-215-10
cargo must 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.
6-26. Tiedown Devices.
18 recessed rings located in the floor are provided
for tiedown. Each has a full swiveling 2-1/2 inch ring
that is designed to take a 2000 pound load in any
direction. (fig. 6-7, sheet 5.)
6-27. Cargo Unloading.
Unloading of cargo will be accomplished through the
main entrance door and the cargo door. Extreme
caution must be exercised to prevent damaging the
wings, wing flap, floorboards, seat tracks, upholstery,
etc. Personnel shall observe NO STEP areas. After
completing the unloading operation, check the aircraft for
possible damage from loading, transporting, or
unloading the cargo; then replace the passenger seats.
Section VI. CENTER OF GRAVITY
6-28. Center-of-Gravity Limitations.
Center-of-gravity limits are expressed in ARM inches
which refers to a positive measurement from
the aircraft's reference datum. The forward CG limit is
144.7 ARM inches for all weights below 7400 pounds
and tapers to 153.2 ARM inches at 9650 pounds. The
aft CG limit is 160.4 ARM inches at all weights (fig. 6-8).
6-21
TM 55-1510-215-10
CENTER OF GRAVITY
LIMITATIONS
Figure 6-8. Center of Gravity
6-22
CENTER-OF-GRAVITY
LIMITATIONS
U-21G
T74-CP-700
TM 55-1510-215-10
CHAPTER 7
PERFORMANCE DATA
Section I. INTRODUCTION
7-1. Description.
The charts presented in this chapter are based on
and are consistent with the recommended operating
procedures and techniques set forth in other chapters of
this manual. The charts contain the performance data
necessary for preflight and in-flight mission planning.
Explanatory text applicable to each type of chart is
included to illustrate the use of the data presented.
and other parameters relating to that phase of flight are
presented. In addition to the presented data, the pilot's
judgement and experience will be necessary to
accurately obtain performance under a given set of
circumstances. The conditions for the data are listed
under the title of each chart. The effects of different
conditions are discussed in the text accompanying each
phase of performance.
Where practical, data are
presented at conservative conditions. However, NO
GENERAL CONSERVATISM HAS BEEN APPLIED.
7-2. Purpose.
a. The purpose of this chapter is to provide the
best available performance data for the U-21G aircraft.
Regular use of this information will enable the pilot to
receive maximum safe utilization from the aircraft.
Although maximum performance is not always required,
regular use of this chapter is recommended for the
following reasons:
(1) Knowledge of performance margin will
allow the pilot to make better decisions when
unexpected conditions or alternate missions are
encountered.
(2) Situations
requiring
performance will be more readily recognized.
maximum
(3) Familiarity with the data will allow
performance to be computed more easily and quickly.
(4) Experience will be. gained in accurately
estimating the effects of variables for which data are not
presented.
b. The information is primarily intended for mission
planning and is most useful when planning operations in
unfamiliar areas or at extreme conditions. The data may
also be used in flight, to establish unit or area standing
operating
procedures,
and
to
inform
ground
commanders of performance/risk tradeoffs.
7-3. General.
The data presented shall cover the maximum range
of conditions and performance that can reasonably be
expected. In each area of performance, the effects of
altitude, temperature, gross weight, and gross weight,
WARNING
Exceeding operating limits may
cause permanent damage to critical
components. Overlimit operation can
decrease
performance,
cause
immediate failure, or failure on a
subsequent flight.
7-4. Limits.
Applicable limits are shown on the charts as red
lines.
Performance generally deteriorates rapidly
beyond limits. If limits are exceeded, minimize the
amount and time. Enter the maximum value and time
above limits on DA Form 2408-13 so proper
maintenance action can be taken.
7-5. Chart Explanation.
A complete series of performance charts are
provided for U-21G aircraft in this manual. These charts
furnish the pilot with sufficient data to make an intelligent
and safe flight plan. The charts include data on takeoff,
climb, landing, and operating instructions for cruising
flight from maximum endurance to normal rated power.
No allowance has been made for navigational error,
formation flight, or other contingencies. Appropriate
allowances for these items should be dictated by local
regulations and should be accounted for when the fuel
available for cruise is determined. The charts are
arranged to give maximum facility of use in preflight and
in-flight planning. All charts are based on free air
temperature (FAT) conditions and pressure altitude.
7-1
TM 55-1510-215-10
7-6. Chapter 7 Index.
SECTION
TITLE
I
II
Introduction
Performance Planning
III
Crosswind-Takeoff or
Landing
Idle Fuel Flow
Torque Available for
Takeoff
Normal Takeoff
Normal Rotation/
Takeoff Airspeed
Acceleration Check
Accelerate-Stop Distance
Minimum
Single-Engine
Control Airspeed
Single-Engine Climb
Operation Envelope
Cruise Climb
Cruise
IV
V
VI
VII
VIII
IX
X
XI
XII
XIII
XIV
7-2
FIGURE
NUMBER
CHART SUBJECT
PAGE
NO.
N/A
7-1
7-2
7-3
7-4
7-5
7-6
7-7
N/A
Performance Planning Card
Temperature Conversion/Correction
Airspeed Calibration-Normal System
Airspeed Calibration-Emergency System
Altimeter Correction-Normal System
Altimeter Correction-Emergency System
Crosswind-Takeoff or Landing
7-1
7-6
7-8
7-11
7-12
7-13
7-14
7-17
7-8
7-9
Idle Fuel Flow
Torque Available for Takeoff
7-19
7-21
7-10
7-11
Takeoff-Normal
Normal Rotation/Takeoff Airspeed
7-23
7-25
7-12
7-13
Acceleration Check
Accelerate-Stop Distance
7-27
7-29
7-14
Minimum Single-Engine Control
Airspeed
7-31
7-15
7-16
7-17
7-18
7-19
7-20
7-21
7-22
7-23
7-24
7-25
7-26
7-27
7-28
7-29
7-30
7-31
7-32
7-33
7-34
7-35
7-36
7-37
7-38
7-39
7-40
7-41
7-42
7-43
7-44
7-45
7-46
Single-Engine Climb
Operation Envelope
Cruise Climb
Cruise Example
Cruise FAT -10°C, Sea Level
Cruise FAT 0°C, Sea Level
Cruise FAT 10°C, Sea Level
Cruise FAT 20°C, Sea Level
Cruise FAT 30°C, Sea Level
Cruise FAT 40°C, Sea Level
Cruise FAT 50°C, Sea Level
Cruise FAT -20°C, 4000 Ft
Cruise FAT -10°C, 4000 Ft
Cruise FAT 0°C, 4000 Ft
Cruise FAT 10°C, 4000 Ft
Cruise FAT 20°C, 4000 Ft
Cruise FAT 30°C, 4000 Ft
Cruise FAT 40°C, 4000 Ft
Cruise FAT -30°C, 8000 Ft
Cruise FAT -20°C, 8000 Ft
Cruise FAT -10°C, 8000 Ft
Cruise FAT 0°C, 8000 Ft
Cruise FAT 10°C, 8000 Ft
Cruise FAT 20°C, 8000 Ft
Cruise FAT 30°C, 8000 Ft
Cruise FAT -30°C, 12,000 Ft
Cruise FAT -20°C, 12,000 Ft
Cruise FAT -10°C, 12,000 Ft
Cruise FAT 0°C, 12,000 Ft
Cruise FAT 10°C, 12,000 Ft
Cruise FAT 20°C, 12,000 Ft
Cruise FAT -40°C, 16,000 Ft
7-33
7-35
7-37
7-39
7-42
7-44
7-46
7-48
7-50
7-52
7-54
7-56
7-58
7-60
7-62
7-64
7-66
7-68
7-70
7-72
7-74
7-76
7-78
7-80
7-81
7-82
7-84
7-86
7-88
7-89
7-90
7-91
TM 55-1510-215-10
XV
XVI
XVII
Climb/Descent
Approach Speed
Landing
7-47
7-48
7-49
7-50
7-51
7-52
7-53
7-54
7-55
7-56
7-57
7-58
7-59
7-60
7-61
7-62
7-63
7-64
7-65
7-66
7-67
7-68
7-69
7-70
7-71
7-72
7-73
Cruise FAT -30°C, 16,000 Ft
Cruise FAT -20°C, 16,000 Ft
Cruise FAT -10°C, 16,000 Ft
Cruise FAT 0°C, 16,000 Ft
Cruise FAT 10°C, 16,000 Ft
Cruise FAT 20°C, 16,000 Ft
Cruise FAT -50°C, 20,000 Ft
Cruise FAT -40°C, 20,000 Ft
Cruise FAT -30°C, 20,000 Ft
Cruise FAT -20°C, 20,000 Ft
Cruise FAT -10°C, 20,000 Ft
Cruise FAT 0°C, 20,000 Ft
Cruise FAT 10°C, 20,000 Ft
Cruise FAT -60°C, 24,000 Ft
Cruise FAT -50°C, 24,000 Ft
Cruise FAT -40°C, 24,000 Ft
Cruise FAT -30°C, 24,000 Ft
Cruise FAT -20°C, 24,000 Ft
Cruise FAT -10°C, 24,000 Ft
Cruise FAT -60°C, 25,000 Ft
Cruise FAT -50°C, 25,000 Ft
Cruise FAT -40°C, 25,000 Ft
Cruise FAT -30°C, 25,000 Ft
Cruise FAT -20°C, 25,000 Ft
Climb/Descent
Approach Speed
Landing
7-92
7-93
7-94
7-95
7-96
7-97
7-98
7-99
7-100
7-101
7-102
7-103
7-104
7-105
7-106
7-107
7-108
7-109
7-110
7-111
7-112
7-113
7-114
7-115
7-117
7-119
7-121
7-3
TM 55-1510-215-10
7-7. Color Coding.
The performance charts are color coded as follows:
a. Green: For example guidelines.
a. Flight Test Data. Data obtained by flight test of
the aircraft by experienced flight test personnel at
precise conditions using sensitive calibrated instruments.
b. Calculated Data. Data based on tests, but not
on flight test of the complete aircraft.
b. Red: Limit lines.
c. Yellow: Precautionary or time-limited operation.
c. Estimated Data. Data based on estimates using
aerodynamic theory or other means but not verified by
flight test.
7-8. Reading The Charts.
7-10. Specific Conditions.
The primary use of each chart is given in an
example and a green guideline is provided to help you
follow the route through the chart. The use of a straight
edge (ruler or page edge) and a hard fine point pencil is
recommended to avoid cumulative errors. The majority
of the charts provide a standard pattern for use as
follows: enter first variable on top left scale, move right to
the second variable, deflect down at right angles to the
third variable, deflect left at right angles to the fourth
variable, deflect down, etc. until the final variable is read
out at the final scale. In addition to the primary use,
other uses of each chart are explained in the text
accompanying each set of performance charts. Colored
registration blocks located on the bottom and top of each
chart are used to determine if slippage has occurred
during printing. If slippage has occurred, refer to chapter
5 for correct operating limits.
The data presented is accurate only for specific
conditions listed under the title of each chart. Variables
for which data is not presented, but which may affect
that phase of performance, are discussed in the text.
Where data is available or reasonable estimates can be
made, the amount that each variable affects
performance will be given.
7-11. General Conditions.
The following general conditions might have
deteriorating effects on the aircraft performance:
atmospheric humidity, fuel flow, rigging, pilot technique,
aircraft variation, engine variation, and instrument
variation.
7-12. Performance Discrepancies.
7-9. Data Basis.
The type of data used is indicated at the bottom of
each performance chart under DATA BASIS. The data
provided generally is based on one of three categories:
7-4
Regular use of this chapter will allow you to monitor
instrument and other aircraft systems for malfunction, by
comparing
actual
performance
with
planned
performance. Knowledge will also be gained concerning
the effects of variables for which data is not provided,
thereby increasing the accuracy of performance
predictions.
TM 55-1510-215-10
Section II. PERFORMANCE PLANNING
7-13. Performance Planning Card (PPC).
(9) LDG RUN.
This card (fig. 7-1), is provided to assist the pilot in
recording data applicable to the mission and may be
reproduced at the local level. This does not preclude the
use of locally developed cards. The PPC provides
readily available information for departure, climb, cruise,
arrival and prevailing conditions. Pertinent data required
to fill in the blanks on the PPC shall be computed from
the performance charts and tables contained in this
chapter, and from existing conditions at time of takeoff or
landing. The takeoff and landing data shall be computed
prior to takeoff, as a precaution against emergency
conditions which could develop after takeoff.
The
following blocks are provided on the front of the PPC for
entry of data as applicable:
a. Weather Data.
(1) PA (pressure altitude).
(2) FAT (free air temperature).
(10) Additional data as required.
7-14. Performance Planning Sequence.
The following information may be extracted from the
charts in this chapter.
NOTE
The pressure altitude may be
determined by setting the altimeter to
29.92 and reading the pressure
altitude, or by adding 100 feet to the
field elevation for each 0.1 in. Hg,
below 29.92, or by subtracting 100
feet from the field elevation for each
0.1 in. Hg. above 29.92.
a. Preflight Planning.
(3) WIND (speed and direction).
(4) RWY (runway heading, length and slope).
(1) Determine the following conditions for
each phase of the flight, as appropriate, before entering
the performance charts and enter the information on the
PPC.
b. Aircraft Data.
(a) Pressure altitude.
(1) T/O WT (takeoff weight).
(b) Free air temperature.
(2) LDG WT (landing weight).
(c) Wind.
c. Performance Data.
(d) Aircraft weight (both takeoff and
(1) T/O PWR (torque).
(2) T/O RUN
clearance (if required)).
(no
landing).
obstacle
or
obstacle
NOTE
(3) ACC-STOP (accelerate-stop distance).
(4) Vmc
airspeed).
(5) Vr-V lof
airspeed).
(minimum
single
(Rotation
engine
airspeed
and
control
liftoff
(6) Vx -V y (best angle of climb and best rate of
climb).
(7) Vyse (single engine best rate of climb).
Weight and balance blocks are
provided on the rear of the PPC and
should be utilized to determine exact
weight and loading conditions prior
to computing takeoff and landing
data.
Weight information may be
obtained from either Chart C or the
current Form 365F.
(e) Obstacles (if applicable).
(f) Ceiling and visibility.
(8) Vref (landing speed).
7-5
TM 55-1510-215-10
(Figure 7-1 Sheet 1 of 2)
(Figure 7-1 Sheet 2 of 2)
Figure 7-1. Performance Planning Card (Sheet 1 of 2)
7-6
TM 55-1510-215-10
Figure 7-1. Performance Planning Card (Sheet 2 of 2)
7-7
TM 55-1510-215-10
TEMPERATURE
CONVERSION/CORRECTION
TEMPERATURE CONVERSION
/CORRECTION
U-21G
T74-CP-700
Figure 7-2. Temperature Conversion Correction
7-8
TM 55-1510-215-10
(g) Instrument departure and approach
computed for landing immediately after takeoff in the
event of an emergency.
(h) Hazards.
(3) After the takeoff and landing data has
been logged, evaluate the performance/risk tradeoff of
the following:
procedures.
(2) Determine the following conditions from
the performance charts and enter the information on the
PPC blocks provided.
(a) T/O PWR -Takeoff power. Obtain
from TORQUE AVAILABLE FOR TAKEOFF chart.
(b) T/O RUN -Take off distance. Obtain
from TAKEOFF-NORMAL chart.
1. Crosswind conditions.
2. Takeoff, accelerate-stop and landing
distances.
3. Obstacle clearance.
b. Rear of PPC.
(c) ACC-STOP - Accelerate-stop distance.
Obtain from ACCELERATE-STOP DISTANCE chart. An
ACCELERATE-CHECK chart is provided for optional
use during critical length takeoffs.
(d) Vmc -Minimum single engine control
airspeed. Obtain from MINIMUM SINGLE ENGINE
CONTROL AIRSPEED (V mc ) chart.
(e) Vr-V lof -Rotation airspeed and liftoff
airspeed. Obtain from NORMAL ROTATION/TAKEOFF
AIRSPEED chart.
NOTE
For the purposes of this manual Vx
will be approximated by the Obstacle
Clearance Climb airspeed obtained
using the Normal Rotation/Takeoff
Airspeed Chart.
(f) Vx -V y - Best angle of climb airspeed
and best rate of climb airspeed.
Obtain Vx from
NORMAL ROTATION/TAKEOFF AIRSPEED chart
(obstacle clearance airspeed). Obtain Vy from the MAX
R/C LINE (two engine) using the appropriate CRUISE
chart.
(g) Vyse - Best single engine rate of
climb. Obtain from appropriate CRUISE chart.
(h) Vref - Landing approach airspeed.
Obtain from APPROACH SPEED chart. This airspeed
shall be computed for landing immediately after takeoff
in the event of an emergency.
(i) LDG RUN - Landing distance.
Obtain from Landing chart. This distance shall be
(1) Cruise data. Space is provided on the
rear of the PPC for such information as pressure altitude
(PA), free air temperature (FAT), wind speed and
direction (WIND), aircraft weight (WT), power required
(PWR), airspeed (KIAS and KTAS), fuel flow (FUEL
RATE) and single engine ceiling (SE CEILING). Power,
airspeed and fuel rate may be obtained from the Cruise
charts. Single engine ceiling may be obtained from the
Operation Envelope chart.
NOTE
BLOCK
SPEED
AND
FUEL
REQUIRED space is provided on the
rear of the PPC for use as required.
Performance charts are available for
climb
and
cruise
information.
Descent information may be obtained
by interpolating between Cruise
charts and the Climb/Descent chart.
The approach spaces may be utilized
to
compute
total
time/fuel/and
distance to actual landing.
(2) Landing data. Space is provided on the
rear of the PPC for entry of landing data.
This
information should be computed prior to landing and
reflect actual landing weight so the landing approach
speed (V ref) and landing distance will be correct. The
instrument approach in use and current weather
conditions may also be entered.
7-15. Performance Information.
Performance information obtained may make it
necessary to alter gross weight, airspeed, altitude, or
other variables in order to safely operate the aircraft. If
any of these variables are changed on one chart,
corresponding changes will be necessary on all other
charts where that information is used.
7-9
TM 55-1510-215-10
7-16. Airspeed Position Error Correction.
7-17. Altimeter Position Error Correction.
The relationship between indicated airspeed and
calibrated airspeed for various flap settings is shown for
the normal static air source.
The altitude corrections to be made to the altimeter
reading are shown for various altitudes and flap
positions for the normal static air source.
NOTE
Indicated airspeed assumes a zero
instrument error, also, no significant
change
in
airspeed
position
correction is apparent due to power
settings, altitude or landing gear
position.
All airspeed calibrations
were conducted in level flight and
may not be appropriate for stall.
7-10
NOTE
Indicated altitude assumes a zero
instrument error, also, no significant
change in position error correction is
apparent due to power settings or
landing gear position.
TM 55-1510-215-10
AIRSPEED CALIBRATION - NORMAL SYSTEM
AIRSPEED CALIBRATION
NORMAL SYSTEM
U-21G
T74-CP-700
Figure 7-3. Airspeed Calibration-Normal System
7-11
TM 55-1510-215-10
AIRSPEED CALIBRATION - EMERGENCY SYSTEM
AIRSPEED CALIBRATION
EMERGENCY SYSTEM
U-21G
T74-CP-700
Figure 7-4. Airspeed Calibration-Emergency System
7-12
TM 55-1510-215-10
ALTIMETER CORRECTION - NORMAL SYSTEM
ALTIMETER CORRECTION
NORMAL SYSTEM
U-21G
T74-CP-700
Figure 7-5. Altimeter Correction-Normal System
7-13
TM 55-1510-215-10
ALTIMETER CORRECTION - EMERGENCY SYSTEM
EXAMPLE
ALTIMETER CORRECTION
EMERGENCY SYSTEM
U-21G
T74-CP-700
Figure 7-6. Altimeter Correction-Emergency System
7-14
TM 55-1510-215-10
Section III. CROSSWIND - TAKEOFF OR LANDING
7-18. Description.
The Crosswind Chart (fig. 7-7) shows the crosswind
conditions under which a takeoff or landing is not
recommended.
7-19. Use of Chart.
Recommended rotation airspeed is obtained from
the Normal Rotation/Takeoff Airspeed Chart (fig. 7-11).
(7-15 blank)/7-16
Takeoff should not be attempted when the rotation
speed intercept point falls within the non-recommended
area.
Select a touchdown airspeed less than the
approach airspeed obtained from the Approach Speed
Chart (fig. 7-72) such that the touchdown airspeed falls
within the recommended area.
TM 55-1510-215-10
CROSSWIND - TAKEOFF OR LANDING
CROSSWIND TAKEOFF
OR LANDING
U-21G
T74-CP-700
Figure 7-7. Crosswind-Takeoff or Landing
7-17
TM 55-1510-215-10
Section IV. IDLE FUEL FLOW
7-20. Description.
7-22. Conditions.
The Idle Fuel Flow Chart (fig. 7-8) shows idle fuel
flow in pounds per hour at various altitudes and FAT.
All fuel flow data is based on JP-4 aviation fuel.
Variation in fuel temperature may change the fuel flow
values slightly.
7-21. Use of Chart.
NOTE
Fuel flow values should be doubled
for two engine operation.
7-18
TM 55-1510-215-10
IDLE FUEL FLOW
JP-4 FUEL
IDLE FUEL FLOW
U-21G
T74-CP-700
Figure 7-8. Idle Fuel Flow
7-19
TM 55-1510-215-10
Section V. TORQUE AVAILABLE FOR TAKEOFF
7-23. Description.
The Torque Available For Takeoff Chart (fig. 7-9)
shows the torque in foot-pounds available for takeoff.
The torque limits to observe for takeoff are also shown.
7-24. Use of chart.
Takeoff distances scheduled on the Takeoff-Normal
Chart (fig. 7-10) are based on power set according to
this chart.
7-20
7-25. Conditions.
a. Power levers - Advance to takeoff power.
b. Propeller speed - 2200 RPM.
TM 55-1510-215-10
TORQUE AVAILABLE FOR TAKEOFF
PROP SPEED = 2200 RPM
FUEL JP-4 AIRSPEED = 0 KNOTS
TORQUE AVAILABLE
FOR TAKEOFF
U-21G
T74-CP-700
EXAMPLE
WANTED
TORQUE AVAILABLE
FOR TAKEOFF
NOTE
1.
2.
3.
TORQUE INCREASES APPROXIMATELY 15 FT-LB FROM ZERO TO 70 KNOTS.
THIS CHART IS BASED ON AN ENGINE WITH MAXIMUM DETERIORATION OF
EXCESS TORQUE WHICH CAN BE PERMITTED WITHOUT EXCEEDING ENGINE
LIMITATIONS.
TAKEOFF PERFORMANCE WAS DEVELOPED UTILIZING THE TORQUE
VALVES SHOWN.
R
G
AP 000570
DATA BASIS: CALCULATED FROM ENGINE MODEL SPEC.
Figure 7-9. Torque Available for Takeoff
7-21
TM 55-1510-215-10
Section VI. NORMAL TAKEOFF
7-26. Description.
The Takeoff-Normal Chart (fig. 7-10), shows the
ground roll distance and the total distance required to
clear varying obstacle heights at a known free air
temperature, pressure altitude and weight.
normal takeoff distances in chapter 7.
Rotate at
recommended rotation speed (V r), establish a pitch
attitude that allows liftoff at recommended liftoff speed
(V lof ). When flight is assured, retract the gear and
establish proper initial climb attitude (obstacle clearance
climb speed (V x ) or (best rate of climb speed Vy )
dependent on existing conditions.
7-27. Use of Chart.
In order to achieve these distances, airspeeds
obtained from the Normal Rotation/Takeoff Airspeed
Chart (fig. 7-11), and power set according to the Torque
Available for Takeoff Chart (fig. 7-9), shall be used.
e. Wind - All data presented are based on calm
wind conditions. Since surface wind speed and direction
can not be accurately predicted, all takeoffs shall be
planned based on calm wind. Distance decreases
approximately 1% per knot headwind.
Distance
increases approximately 3% per knot tailwind.
7-28. Conditions.
a. Engine - Both engines operating with takeoff
power and 2200 RPM with torque set according to the
Torque Available for Takeoff Chart (fig. 7-9).
b. Flaps - UP.
c. Landing gear - UP (after liftoff).
d. Technique - Apply maximum allowable takeoff
power and accelerate on the runway. Takeoff power
must be applied prior to releasing brakes to obtain the
7-22
f. Runway - Runway conditions for this chart are
based on a dry, hard-surface, level runway. Conditions
other than these will vary aircraft takeoff distances.
Ground roll distance decreases approximately 5% per
1% downhill gradient. Ground roll distance increases
approximately 7% per 1% uphill gradient.
NOTE
Refer to the Single-Engine Climb
Chart (fig. 7-15), to determine if
adequate performance is available in
the event of engine failure during
takeoff.
TM 55-1510-215-10
TAKEOFF - NORMAL
CALM WINDS FLAPS 0 PERCENT POWER - TAKEOFF
LEVEL HARD SURFACE
TAKEOFF NORMAL
U-21G
T74-CP-700
Figure 7-10. Takeoff-Normal
7-23
TM 55-1510-215-10
Section VII. NORMAL ROTATION/TAKEOFF AIRSPEED
7-29. Description.
The Normal Rotation/Takeoff Airspeed Chart
(fig. 7-11), shows rotation, takeoff, and obstacle
clearance airspeeds for a known weight.
7-30. Use of Chart.
The rotation airspeed line shows recommended
rotation indicated airspeed for a given weight. The
takeoff airspeed line shows recommended takeoff
indicated airspeed for a given weight. The obstacle
clearance airspeed line shows the recommended
indicated airspeed for a given weight to use for
7-24
clearance of obstacles during the initial climb.
Performance scheduled on the Takeoff-Normal Chart,
(fig. 7-10), is based on use of these speeds.
7-31. Conditions - Flaps UP.
NOTE
These speeds have been selected in
order to provide adequate margins
above the flaps up stall speed and
single-engine
minimum
control
airspeed (V mc).
TM 55-1510-215-10
NORMAL ROTATION/TAKEOFF AIRSPEED
FLAPS 0 PERCENT
NORMAL ROTATION/
TAKEOFF AIRSPEED
U-21G
T74-CP-700
NOTE
OBSTACLE CLEARANCE AIRSPEED
LINE ALSO INDICATES THE BEST
TWO AND SINGLE-ENGINE ANGLE
OF CLIMB ENGINE SPEEDS.
Figure 7-11. Normal Rotation Takeoff Airspeed
7-25
TM 55-1510-215-10
Section VIII. ACCELERATION CHECK
7-32. Description.
The Acceleration Check chart (fig. 7-12), shows the
relationship between indicated airspeed and ground roll
distance during the takeoff run.
7-33. Use of Chart.
Required airspeeds at designated points along the
runway are obtained from this chart. This chart is used
in conjunction with the ground roll distance from the
Takeoff-Normal Chart (fig.
7-10), and the takeoff
airspeed from the Normal Rotation/Takeoff Airspeed
chart (fig. 7-11). If required indicated airspeed is not
obtained at a selected distance and the runway length is
critical, the takeoff should be aborted.
d. Technique - Apply takeoff power and accelerate
on the runway.
When the aircraft passes an
acceleration check point, insure that the required
airspeed has been achieved.
e. Runway - This chart is based on a dry, hard
surface runway. Conditions other than this may vary
aircraft operation. Adjust ground roll distance for runway
gradient before entering this chart. Refer to TakeoffNormal chart (fig. 7-10).
f. Wind - All data presented is based on calm wind
conditions. Since surface wind conditions cannot be
accurately predicted, all takeoffs shall be planned based
on calm wind.
NOTE
7-34. Conditions.
a. Engines - Both engines operating with takeoff
power and 2200 RPM according to Torque Available for
Takeoff Chart (fig. 7-9).
b. Flaps - UP.
c. Landing Gear - DN.
7-26
This chart should always be entered
with the takeoff airspeed appropriate
to the weight.
g. Acceleration Check - Is the distance from start of
takeoff ground roll to an acceleration check point
(runway distance marker, etc.)
TM 55-1510-215-10
ACCELERATION CHECK
POWER - TAKEOFF CALM WINDS
FLAPS 0 PERCENT LEVEL HARD SURFACE
ACCELERATION
CHECK
U-21G
T74-CP-700
Figure 7-12. Acceleration Check
7-27
TM 55-1510-215-10
Section IX. ACCELERATE - STOP DISTANCE
7-35. Description.
b. Flaps - Up.
The Accelerate Stop Distance Chart (fig. 7-13)
shows the distance required to accelerate to takeoff
airspeed then stop, at a known free air temperature,
pressure altitude, and weight.
c. Technique - Apply takeoff power prior to
releasing brakes. Should engine failure occur at or
before reaching takeoff airspeed, place both power
levers in the idle position and apply maximum braking.
7-36. Use of Chart.
d. Wind - All data presented are based on calm
wind conditions. Since surface wind speed and direction
can not be accurately predicted, all takeoffs shall be
planned based on calm wind. Accelerate stop distances
decrease approximately 1% per knot headwind and
increase approximately 3% per knot tailwind.
In order to achieve these distances, takeoff
airspeeds obtained from the Normal Rotation/Takeoff
Airspeed Chart (fig. 7-11), and power set according to
the Torque Available for Takeoff Chart (fig. 7-9), shall
be assumed.
7-37. Conditions.
a. Engines - Both engines operating with takeoff
power and 2200 RPM with torque set according to the
Torque Available for Takeoff Chart (fig. 7-9).
7-28
e. Runway - Runway conditions for this chart are
based on a dry, hard surface, level runway. Conditions
other than these will vary aircraft accelerate-stop
distances.
Accelerate-stop
distances
increase
approximately 4% per 1% downhill gradient. Acceleratestop distances decrease approximately 3% per 1% uphill
gradient.
TM 55-1510-215-10
ACCELERATE-STOP DISTANCE
HARD SURFACE RUNWAY WIND CALM FLAPS ZERO PERCENT
TAKEOFF POWER FOLLOWED BY IDLE POWER AND MAX BRAKING
ACCELERATE-STOP DISTANCE
U-21G
T74-CP-700
Figure 7-13. Accelerate-Stop Distance
7-29
TM 55-1510-215-10
Section X. MINIMUM SINGLE-ENGINE CONTROL AIRSPEED
7-38. Description.
7-40. Conditions.
The Minimum Single-Engine Control Airspeed
Chart (fig. 7-14), shows the minimum airspeed (V mc )
which will allow aircraft directional control during
single-engine operation for various air temperatures
and altitudes. Directional control of the aircraft
cannot be maintained for speeds below the rudder
limit line with the non-critical engine operating at
takeoff power.
a. Operating Engine - Operate with takeoff power
and 2200 RPM according to the Torque Available for
Takeoff Chart (fig. 7-9).
b. Inoperative Engine - Propeller FEATHER.
c. Flaps - UP.
d. Landing Gear - DN.
7-39. Use of Chart.
Flight at Vmc implies aircraft directional control
only and does not provide the pilot with singleengine rate-of-climb information.
7-30
TM 55-1510-215-10
MINIMUM SINGLE ENGINE
CONTROL AIRSPEED (Vmc)
POWER - TAKEOFF GEAR DOWN
FLAPS 0 PERCENT PROP FEATHERED
MINIMUM SINGLE ENGINE
CONTROL AIRSPEED
U-21G
T74-CP-700
NOTE
AT SOME WEIGHTS A STALL
CONDITION
CAN
OCCUR
AT
AIRSPEEDS HIGHER THAN THE
RUDDER LIMITED Vmc.
7-31
Figure 7-14. Minimum Single-Engine Control Airspeed
TM 55-1510-215-10
Section XI. SINGLE-ENGINE CLIMB
7-41. Description.
b. Inoperative Engine - Propeller FEATHER.
The Single-Engine Climb Chart (fig. 7-15),
shows the critical engine inoperative takeoff rate of
climb for varying air temperatures, altitudes, and
weights for the gear down and gear up
configurations.
For single-engine cruise
information refer to the Cruise Charts
(fig. 7-18 thru fig. 7-70), and for Vmc refer to the
Minimum Single Engine Control Speed Chart (fig.
7-14).
c. Flaps - UP.
7-42. Use of Chart.
This chart is to be used to determine single
engine performance in the event of engine failure
during the takeoff and initial climb segments.
Either weight for a desired rate-of-climb, or rate-ofclimb for a known weight can be determined.
Performance within the red zone(s) indicates the
lack of positive climb capability.
7-43. Conditions.
a.Operating Engine - Operate at takeoff power
and 2200 RPM according to the Torque Available
for Takeoff Chart (fig. 7-9).
7-32
d. Landing Gear - UP (after liftoff).
e. Technique - Achieve and maintain the obstacle
clearance airspeed appropriate to the weight. Refer to
Normal Rotation/Takeoff Airspeed Chart (fig. 7-11).
NOTE
Refer to Normal Rotation/Takeoff
Airspeed Chart (fig. 7-11). Gear
down rates of climb are based on
climb at the scheduled takeoff
airspeed. Gear up rates of climb are
based on climb at the scheduled
obstacle clearance airspeed.
TM 55-1510-215-10
SINGLE ENGINE CLIMB
TAKE OFF CONFIGURATION
FLAPS 0 PERCENT PROP FEATHERED
POWER - TAKEOFF
SINGLE-ENGINE
CLIMB
U-21G
T74-CP-700
Figure 7-15. Single-Engine Climb
7-33
TM 55-1510-215-10
Section XII. OPERATION ENVELOPE
7-44. Description.
7-45. Use of Chart.
The Operation Envelope Chart (fig. 7-16), shows
the maximum and minimum temperatures for which
operation is recommended; the maximum altitudes
possible for both two and single-engine operation; and
identifies the altitude and temperature combinations for
which Cruise Charts are available. The maximum and
minimum temperature lines are based on International
Standard Atmosphere (ISA) temperatures plus 37
degrees Celsius and minus 30 degrees Celsius,
respectively.
The airframe manufacturer cannot
recommend operation outside of this envelope.
This chart is used to determine if the proposed flight
is within the operational temperature/altitude envelope
and to identify the temperature and altitude combinations
which should be used for flight planning.
7-34 Change 8
7-46. Conditions.
Service ceilings are based on the aircraft in a clean
configuration.
TM 55-1510-215-10
OPERATION ENVELOPE
OPERATION ENVELOPE
U-21G
T74-CP-700
Figure 7-16. Operation Envelope
Change 8 7-35
TM 55-1510-215-10
Section XIII. CRUISE CLIMB
7-47. Description.
b. Climb Speed - 140 KIAS
The Cruise Climb Chart (fig. 7-17), shows the time,
fuel, and distance required to climb from sea level. For
climb at other than the conditions of this chart, refer to
Cruise Charts (fig. 7-18 thru fig. 7-70), and the Climb/
Descent Chart (fig. 7-71).
c. Flaps - UP.
7-48. Use of chart.
To determine the time, fuel and distance required to
climb from one altitude to another, subtract the time, fuel
and distance from sea level to the initial altitude as
shown.
7-49. Condition.
a. Power - Maximum cruise climb power at 2000
RPM.
7-36
d. Landing Gear - UP.
e. Atmosphere - Temperature lapse rate is
assumed to be linear from the initial altitude to the final
altitude. Distances to climb are based on calm wind.
NOTE
Time and fuel to climb are not
affected by wind.
TM 55-1510-215-10
CRUISE CLIMB
GEAR UP FLAPS 0 PERCENT CALM WIND
POWER - CRUISE CLIMB
CRUISE CLIMB
U-21G
T74-CP-700
Figure 7-17. Cruise Climb
7-37
TM 55-1510-215-10
Section XIV. CRUISE
7-50. Description.
NOTE
The Cruise Charts (fig. 7-18 thru fig. 7-70), show
the horsepower per engine for various airspeeds and
weights at selected combinations of altitude and
temperature. Cruise and climb fuel flows are shown
corresponding to engine power. Maximum performance
lines are shown. Torque variation with propeller speed
and engine horsepower are shown. Two engine and
single-engine data are presented on the chart. Each
individual chart is for a single pressure altitude and FAT.
7-51. Use of Chart.
f. Horsepower Per Engine.
Horsepower per
engine, although not readable in the cockpit is utilized in
conjunction with the Climb/Descent Chart (fig. 7-71), to
obtain rate of climb or descent.
7-52. Conditions.
a. Configuration - All cruise data are based on the
aircraft in a clean configuration.
The primary use of the Cruise Chart is shown in the
Cruise Example (fig. 7-18). Other uses for the chart are
described under "Other uses for Chart" (para. 7-53).
The Cruise charts are for use in determining power
settings to obtain a desired level of performance for a
known set of conditions of altitude, temperature, weight
and number of operating engines. Normally, sufficient
accuracy can be obtained by selecting the chart nearest
to the planned cruising altitude and temperature. If
greater accuracy is desired, interpolation between two or
more charts should be used. The power setting is
obtained by adjusting torque and propeller speed.
a. Maximum Range. The maximum range line
shows the airspeed which will provide the maximum
range for a given aircraft weight at the chart conditions.
b. Maximum
Endurance.
The
maximum
endurance line shows the airspeed which will provide
maximum endurance for a given aircraft weight at the
chart conditions.
c. Maximum Rate-of-Climb. The maximum rate of
climb line shows the airspeed which would provide
maximum rate of climb for a given aircraft weight if
maximum horsepower was used at the chart conditions.
d. Airspeed. The relationship between indicated
and true airspeed is shown for the specific chart
conditions of altitude and temperature.
Conversion
between the two airspeeds can be accomplished without
regard to other chart information.
e. Fuel Flow.
settings is shown.
Fuel flow values should be doubled
for two engine operation.
The fuel flow for cruise and climb
b. Fuel - Fuel flow is based on use of JP-4 aviation
fuel.
c. Operation - Areas not recommended for cruise
operation are shown in yellow.
7-53. Other uses for Chart.
a. Selected True Airspeed
Wanted
Indicated airspeed, horsepower per engine.
Fuel flow, and torque for a selected cruise
true airspeed.
Known
Pressure altitude
FAT
Weight
Prop RPM
Method
Enter chart for true airspeed
Move right
Read indicated airspeed
Move down from weight line
Read horsepower per engine
Move down
Read fuel flow
Move left from prop RPM
Read torque
b. Climb Conditions
Wanted
Maximum rate-of-climb airspeed.
Horsepower required for level flight at
MAX R/C airspeed.
7-38
TM 55-1510-215-10
CRUISE
PRESSURE ALTITUDE 12,000 FT EXAMPLE
FLAPS UP GEAR UP
CRUISE
U-21 G
T74-CP-700
Figure 7-18. Cruise Example
Change 1 7-39
TM 55-1510-215-10
Max horsepower available.
Fuel flow and Torque pressure for Max
R/C.
Excess horsepower available for climb.
Known
Pressure altitude
FAT
Weight
Prop RPM
Method
Enter chart at MAX R/C line.
Move right and up to weight.
Move left and read MAX R/C true
airspeed.
7-40
Move right and read MAX R/C indicated
airspeed.
Move down from weight line.
Read horsepower per engine for level
flight at MAX R/C airspeed.
Read Max horsepower available per engine
at MAX R/C airspeed.
Move Down from MAX horsepower
Available.
Read climb fuel flow.
Move left from prop RPM line.
Read torque.
Determine excess horsepower per engine
available for climb - MAX
horsepower available per engine less
horsepower per engine for level flight.
TM 55-1510-215-10
FAT = -10°° C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT FUEL JP4 GEAR UP
(Figure 7-19 Sheet 1 of 2)
(Figure 7-19 Sheet 2 of 2)
Figure 7-19. Cruise FAT -10°C, Sea Level (sheet 1 of 2)
(7-41 blank)/7-42 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -10°C
CRUISE
SINGLE ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-19. Cruise FAT - 10°C, Sea Level (sheet 2 of 2)
Change 1 7-43
TM 55-1510-215-10
FAT = 0°° C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-20 Sheet 1 of 2)
(Figure 7-20 Sheet 2 of 2)
Figure 7-20. Cruise FAT 0°C, Sea Level (sheet 1 of 2)
7-44 Change 1
CRUISE
U-21 G
T74-CP-700
TM 55-1510-215-10
FAT = 0°° C
CRUISE
SINGLE ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT FUEL JP4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-20. Cruise FAT 0°C, Sea Level (sheet 2 of 2)
Change 1 7-45
TM 55-1510-215-10
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
FAT = 10°C
(Figure 7-21 Sheet 1 of 2)
(Figure 7-21 Sheet 2 of 2)
Figure 7-21. Cruise FAT 10°C, Sea Level (sheet 1 of 2)
7-46 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = 10°C
SINGLE ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-21. Cruise FAT 10°C, Sea Level (sheet 2 of 2)
Change 1 7-47
TM 55-1510-215-10
CRUISE
FAT = 20°C
TWIN ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-22 sheet 1 of 2)
(Figure 7-22 sheet 2 of 2)
Figure 7-22. Cruise FAT 20°C, Sea Level (sheet 1 of 2)
7-48 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = 20°C
SINGLE ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-22. Cruise FAT 20°C, Sea Level (sheet 2 of 2)
Change 1 7-49
TM 55-1510-215-10
CRUISE
FAT = 30°C
TWIN ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-23 sheet 1 of 2)
(Figure 7-23 sheet 2 of 2)
Figure 7-23. Cruise FAT 30°C, Sea Level (sheet 1 of 2)
7-50 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = 30°C
SINGLE ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-23. Cruise FAT 30°C, Sea Level (sheet 2 of 2)
Change 1 7-51
TM 55-1510-215-10
CRUISE
FAT = 40°C
TWIN ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-24 sheet 1 of 2)
(Figure 7-24 sheet 2 of 2)
Figure 7-24. Cruise FAT 40°C, Sea Level (sheet 1 of 2)
7-52 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = 40°C
SINGLE ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-24. Cruise FAT 40°C, Sea Level (sheet 2 of 2)
Change 1 7-53
TM 55-1510-215-10
CRUISE
FAT = 50°C
TWIN ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-25 sheet 1 of 2)
(Figure 7-25 sheet 2 of 2)
Figure 7-25. Cruise FAT 50°C, Sea Level (sheet 1 of 2)
7-54 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = 50°C
TWIN ENGINE
PRESSURE ALTITUDE SEA LEVEL
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-25. Cruise FAT 50°C, Sea Level (sheet 2 of 2)
Change 1 7-55
TM 55-1510-215-10
CRUISE
FAT = -20°C
TWIN ENGINE
PRESSURE ALTITUDE 4000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-26 sheet 1 of 2)
(Figure 7-26 sheet 2 of 2)
Figure 7-26. Cruise FAT -20°C, 4000 Ft (sheet 1 of 2)
7-56 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = -20°C
SINGLE ENGINE
PRESSURE ALTITUDE 4000 FEET
FLAPS 0 PERCENT JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-26. Cruise FAT -20°C, 4000 Ft (sheet 2 of 2)
Change 1 7-57
TM 55-1510-215-10
CRUISE
FAT = -10°C
TWIN ENGINE
PRESSURE ALTITUDE 4000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-27 sheet 1 of 2)
(Figure 7-27 sheet 2 of 2)
Figure 7-27. Cruise FAT -10°C, 4000 Ft (sheet 1 of 2)
7-58 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = 10°C
SINGLE ENGINE
PRESSURE ALTITUDE 4000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-27. Cruise FAT -10C, 4000 Ft (sheet 2 of 2)
Change 1 7-59
TM 55-1510-215-10
CRUISE
FAT = 0°C
TWIN ENGINE
PRESSURE ALTITUDE 4000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-28 sheet 1 of 2)
(Figure 7-28 sheet 2 of 2)
Figure 7-28. Cruise FAT 0°C, 4000 Ft (sheet 1 of 2)
7-60 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = 0°C
SINGLE ENGINE
PRESSURE ALTITUDE 4000 FEET
FLAPS 0 PERCENT JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-28. Cruise FAT 0°C, 4000 Ft (sheet 2 of 2)
Change 1 7-61
TM 55-1510-215-10
CRUISE
FAT = 10°C
TWIN ENGINE
PRESSURE ALTITUDE 4000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-29 sheet 1 of 2)
(Figure 7-29 sheet 2 of 2)
Figure 7-29. Cruise FAT 10°C, 4000 Ft (sheet 1 of 2)
7-62 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = 10°C
SINGLE ENGINE
PRESSURE ALTITUDE 4000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7- 29. Cruise FAT 10 °C, 4000 Ft (sheet 2 of 2)
Change 1 7-63
TM 55-1510-215-10
CRUISE
FAT = 20°C
TWIN ENGINE
PRESSURE ALTITUDE 4000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-30 sheet 1 of 2)
(Figure 7-30 sheet 2 of 2)
Figure 7-30. Cruise FAT 20°C, 4000 Ft (sheet 1 of 2)
7-64 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = 20°C
SINGLE ENGINE
PRESSURE ALTITUDE 4000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-30. Cruise FAT 20°C, 4000 Ft (sheet 2 of 2)
Change 1 7-65
TM 55-1510-215-10
CRUISE
FAT = 30°C
TWIN ENGINE
PRESSURE ALTITUDE 4000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-31 sheet 1 of 2)
(Figure 7-31 sheet 2 of 2)
Figure 7-31. Cruise FAT 30°C, 4000 Ft (sheet 1 of 2)
7-66 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = 30°C
SINGLE ENGINE
PRESSURE ALTITUDE 4000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-31. Cruise FAT 30°C, 4000 Ft (sheet 2 of 2)
Change 1 7-67
TM 55-1510-215-10
CRUISE
FAT = 40°C
TWIN ENGINE
PRESSURE ALTITUDE 4000 FT
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7-32. Cruise FAT 40°C, 4000 Ft
7-68 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = -30°C
TWIN ENGINE
PRESSURE ALTITUDE 8000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-33 sheet 1 of 2)
(Figure 7-33 sheet 2 of 2)
Figure 7-33. Cruise FAT -30°C, 8000 Ft (sheet 1 of 2)
(7-69 blank)/7-70 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = -30°C
SINGLE ENGINE
PRESSURE ALTITUDE 8000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-33. Cruise FAT -30°C, 8000 Ft (sheet 2 of 2)
Change 1 7-71
TM 55-1510-215-10
CRUISE
FAT = -20°C
TWIN ENGINE
PRESSURE ALTITUDE 8000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-34 sheet 1 of 2)
(Figure 7-34 sheet 2 of 2)
Figure 7-34. Cruise FAT -20°C, 8000 Ft (sheet 1 of 2)
7-72 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = -20°C
SINGLE ENGINE
PRESSURE ALTITUDE 8000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-34. Cruise FAT -20°C, 8000 Ft (sheet 2 of 2)
Change 1 7-73
TM 55-1510-215-10
CRUISE
FAT = -10°C
TWIN ENGINE
PRESSURE ALTITUDE 8000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-35 sheet 1 of 2)
(Figure 7-35 sheet 2 of 2)
Figure 7-35. Cruise FAT -10°C, 8000 Ft (sheet 1 of 2)
7-74 Change 1
CRUISE
U-21A
T74-CP-700
TM 55-1510-215-10
CRUISE
FAT = -10°C
SINGLE ENGINE
PRESSURE ALTITUDE 8000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-35. Cruise FAT -10°C, 8000 Ft (sheet 2 of 2)
Change 1 7-75
TM 55-1510-215-10
CRUISE
FAT = 0°C
TWIN ENGINE
PRESSURE ALTITUDE 8000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-36 sheet 1 of 2)
(Figure 7-36 sheet 2 of 2)
Figure 7-36. Cruise FAT 0°C, 8000 Ft (sheet 1 of 2)
7-76 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = 0°C
CRUISE
SINGLE ENGINE
PRESSURE ALTITUDE 8000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-36. Cruise FAT 0°C, 8000 Ft (sheet 2 of 2)
Change 1 7-77
TM 55-1510-215-10
FAT = 10°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 8000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-37 Sheet 1 of 2)
(Figure 7-37 Sheet 2 of 2)
Figure 7-37. Cruise FAT 10°C, 8000 Ft (sheet 1 of 2)
7-78 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = 10°C
CRUISE
SINGLE ENGINE
PRESSURE ALTITUDE 8000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-37. Cruise FAT 10°C, 8000 FT (sheet 2 of 2)
Change 1 7-79
TM 55-1510-215-10
FAT = 20°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 8000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7-38. Cruise FAT 20°C, 8000 Ft
7-80 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = 30°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 8000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-39. Cruise FAT 30°C, 8000 Ft
Change 1 7-81
TM 55-1510-215-10
FAT = -30°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 12,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
(Figure 7-40 Sheet 1 of 2)
(Figure 7-40 Sheet 2 of 2)
Figure 7-40. Cruise FAT -30°C, 12,000 (sheet 1 of 2)
7-82 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -30°C
CRUISE
SINGLE ENGINE
PRESSURE ALTITUDE 12,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-40. Cruise FAT -30°C, 12,000 Ft (sheet 2 of 2)
Change 1 7-83
TM 55-1510-215-10
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 12,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
FAT = -20°C
(Figure 7-41 Sheet 1 of 2)
(Figure 7-41 Sheet 2 of 2)
Figure 7-41. Cruise FAT -20°C, 12,000 Ft (sheet 1 of 2)
7-84 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -20°C
CRUISE
SINGLE ENGINE
PRESSURE ALTITUDE 12,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7- 41. Cruise FAT - 20 °C, 12,000 Ft (sheet 2 of 2 )
Change 1 7-85
TM 55-1510-215-10
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 12,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
FAT = -10°C
(Figure 7-42 Sheet 1 of 2)
(Figure 7-42 Sheet 2 of 2)
Figure 7-42. Cruise FAT -10°C, 12,000 Ft (sheet 1 of 2)
7-86 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -10°C
CRUISE
SINGLE ENGINE
PRESSURE ALTITUDE 12,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-42. Cruise FAT -10°C, 12,000 Ft (sheet 2 of 2)
Change 1 7-87
TM 55-1510-215-10
FAT = 0°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 12,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7-43. Cruise FAT 10°C, 12,000 Ft
7-88 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -10°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 12,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-44. Cruise FAT 10°C, 12,000 Ft
Change 1 7-89
TM 55-1510-215-10
FAT = 20°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 12,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7-45. Cruise FAT 20°C, 12,000 Ft
7-90 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -40°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 16,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-46. Cruise FAT -40°C, 16,000 Ft
Change 1 7-91
TM 55-1510-215-10
FAT = -30°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 16,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7-47. Cruise FAT -30°C, 16,000 Ft
7-92 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -20°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 16,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-48. Cruise FAT -20°C 16,000 Ft
Change 1 7-93
TM 55-1510-215-10
FAT = -10°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 16,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7-49. Cruise FAT -10°C, 16,000 Ft
7-94 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = 0°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 16,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-50. Cruise FAT 0°C, 16,000 Ft
Change 1 7-95
TM 55-1510-215-10
FAT = 10°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 16,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
°
7-96 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = 20°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 16,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-52. Cruise FAT 20°C, 16,000 Ft
Change 1 7-97
TM 55-1510-215-10
FAT = -50°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 20,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7-53. Cruise FAT -50°C, 20,000 Ft
7-98 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -40°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 20,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-54. Cruise FAT -40°C, 20,000 Ft
Change 1 7-99
TM 55-1510-215-10
FAT = -30°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 20,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7-55. Cruise FAT -30°C, 20,000 Ft
7-100 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -20°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 20,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7- 56. Cruise FAT - 20 °C, 20,000 Ft
Change 3 7-101
TM 55-1510-215-10
FAT = -10°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 20,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7- 57. Cruise FAT - 10 °C, 20,000 Ft
7-102 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = 0°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 20,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-58. Cruise FAT 0°C, 20,000 Ft
Change 1 7-103
TM 55-1510-215-10
FAT = 10°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 20,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7- 59. Cruise FAT 10 °C, 20,000 Ft
7-104 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -60°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 24,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-60. Cruise FAT -60°C 24,000 Ft
Change 1 7-105
TM 55-1510-215-10
FAT = -50°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 24,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7- 61. Cruise FAT - 50 °C, 24,000 Ft
7-106 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -40°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 24,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-62. Cruise FAT -40°C, 24,000 Ft
Change 1 7-107
TM 55-1510-215-10
FAT = -30°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 24,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7-63. Cruise FAT -30°C, 24,000 Ft
7-108 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -20°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 24,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-64. Cruise FAT -20°C, 24,000 Ft
Change 1 7-109
TM 55-1510-215-10
FAT = -10°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 24,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7-65. Cruise FAT -10°C, 24,000 Ft
7-110 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -60°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 25,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7- 66. Cruise FAT - 60 °C, 25,000 Ft
Change 1 7-111
TM 55-1510-215-10
FAT = -50°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 25,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7-67. Cruise FAT -50°C, 25,000 Ft
7-112 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -40°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 25,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7-68. Cruise FAT -40°C, 25,000 Ft
Change 1 7-113
TM 55-1510-215-10
FAT = -30°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 25,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
Figure 7-69. Cruise FAT -30°C, 25,000 Ft
7-114 Change 1
CRUISE
U-21G
T74-CP-700
TM 55-1510-215-10
FAT = -20°C
CRUISE
TWIN ENGINE
PRESSURE ALTITUDE 25,000 FEET
FLAPS 0 PERCENT FUEL JP-4 GEAR UP
CRUISE
U-21G
T74-CP-700
Figure 7- 70. Cruise FAT - 20° C, 25,000 Ft
Change 1 7-115
TM 55-1510-215-10
Section XV. CLIMB/DESCENT
7-54. Description.
The Climb/Descent Chart (fig. 7-71), shows the rate
of climb or descent at a given weight for a change in
horsepower per engine. Two engine and single-engine
operation is shown.
available is obtained from the appropriate Cruise Chart
(fig. 7-18 thru 7-70).
b. Descent. The horsepower change per engine is
that decrease in power needed to obtain a desired rateof-descent. Determine torque and propeller speed from
the Cruise Charts.
7-55. Use of the Chart.
7-56. Conditions.
The horsepower per engine required for level flight is
obtained from the appropriate Cruise Chart.
a. Climb. The horsepower change per engine is
that additional power needed to obtain the desired rateof-climb at a known weight. The excess horsepower
7-116
a. Power - Maximum climb power.
b. Rate-of-climb/descent will change as altitude
and temperature change. Re-evaluate as altitude and
temperature change.
TM 55-1510-215-10
CLIMB/DESCENT
GEAR UP FLAPS 0 PERCENT
CLIMB DESCENT
U-21G
74-CP-700
Figure 7-71. Climb/Descent
7-117
TM 55-1510-215-10
Section XVI. APPROACH SPEED
7-57. Description.
7-59. Condition.
The Approach Speed Chart (fig. 7-72), shows the
variation of approach speed with weight for flaps up and
flaps down operation. The landing approach speed (Vref )
is the indicated airspeed the aircraft should be at when
50 feet above the runway in landing configuration.
a. Speeds. These speeds have been selected to
provide adequate margins above the stall speeds for the
appropriate flap settings.
7-58. Use of Chart.
The approach speed line shows the indicated
approach airspeed to use for a given aircraft weight.
7-118
b. Landing Chart. Performance scheduled on the
Landing Chart (fig. 7-73), is based on use of these
speeds.
c. Touchdown.
The touchdown shall be
accomplished at speeds less than those scheduled.
Consideration shall be given to the appropriate stall
speed from the Stall Speed Chart (fig. 8-4) and to the
minimum touchdown speed needed to remain within the
recommended area on the Crosswind-Takeoff or
Landing Chart (fig. 7-7).
TM 55-1510-215-10
APPROACH SPEED
GEAR DOWN
APPROACH SPEED
U-21G
T74-CP-700
Figure 7- 72. Approach Speed
7-119
TM 55-1510-215-10
Section XVII. LANDING
7-60. Description.
7-62. Conditions.
The Landing Chart (fig. 7-73), shows the total
ground roll distance required for varying air
temperatures, altitudes, weights and flap positions.
b. Flaps - As selected.
NOTE
c. Landing Gear - DN.
Landing data is provided up to
weights of 9650 pounds in the event
emergency
landing
is
required
immediately after takeoff. Any time
the maximum landing weight of 9168
pounds is exceeded, an appropriate
entry shall be made on DA Form
2408-13.
7-61. Use of Chart.
Performance is based on the approach speed
obtained from the Approach Speed Chart (fig. 7-72).
7-120
a. Power - As required (on final approach).
d. Technique - Establish 800 FT/MIN descent on
final approach at the appropriate airspeed from the
Approach Speed Chart (fig. 7-72). When landing is
assured, reduce power and speed.
e. Runway - Runway conditions for this chart are
based on a dry, level, hard-surface. Conditions other
than these may vary aircraft landing distance. Distance
decreases approximately 4% per 1% uphill gradient.
Distance increases approximately 4% per 1% downhill
gradient.
f. Wind - All data presented are based on calm
wind conditions. Distance decreases approximately 1%
per knot headwind. Distance increases approximately
4% per knot tailwind.
TM 55-1510-215-10
LANDING
CALM WINDS LEVEL, DRY, HARD SURFACE
LANDING
U-21G
T74-CP-700
Figure 7- 73. Landing
7-121/(7-122 blank)
TM 55-1510-215-10
CHAPTER 8
NORMAL PROCEDURES
Section I. MISSION PLANNING
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/passenger briefings. The pilot in command shall
insure compliance with the contents of this manual that
are applicable to the mission.
8-2. Operating Limits and Restrictions.
The minimum, maximum, normal and cautionary
operational ranges represent careful aerodynamic and
structural calculations, substantiated by flight test data.
These limitations must be adhered to during all phases
of the mission. Refer to chapter 5, OPERATING LIMITS
AND RESTRICTIONS, for detailed information.
must be within the limits prescribed in chapter 5,
OPERATING LIMITS AND RESTRICTIONS.
8-4. Performance.
Refer to chapter 7, PERFORMANCE DATA, to
determine the capability of the aircraft for the entire
mission. Consideration must be given to changes in
performance resulting from variations in loads,
temperatures, and pressure altitudes. Record the data
on the Performance Planning Card for use in completing
the flight plan and for reference throughout the mission.
8-5. Flight Plan.
A flight plan shall be completed and filed in
accordance with AR 95-1, DOD FLIP, and local
regulations.
8-6. Crew/Passenger Briefings.
8-3. Weight/Balance and Loading.
The aircraft must be loaded, cargo and passengers
secured, and weight and balance verified in accordance
with chapter 6, WEIGHT/BALANCE AND LOADING.
This aircraft is in weight and balance class I and requires
a weight and balance clearance only when loaded in
other than a normal manner in accordance with AR 9516. The aircraft weight and center of gravity conditions
A crew/passenger briefing shall be conducted to
insure a thorough understanding of individual and team
responsibilities. The briefing should include the copilot,
crew chief, passengers and ground crew responsibilities
and the coordination necessary to complete the mission
in the most efficient manner. 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/passenger briefing.
8-1
TM 55-1510-215-10
Section II. OPERATING PROCEDURES AND MANEUVERS
8-7. Operating Procedures and Maneuvers.
a. This section deals with normal procedures and
includes all steps necessary to insure 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.
Your flying experience is recognized;
therefore, basic flight principles are avoided. Only the
duties of the minimum crew necessary for the actual
operation of the aircraft are included.
b. Procedures specifically related to instrument
flight that are different from normal procedures are
covered in this section following normal procedures.
Descriptions of functions, operations, and effects of
controls are covered in Section IV, FLIGHT
CHARACTERISTICS, and are repeated in this section
only when required for emphasis. Checks that must be
performed under adverse environmental conditions such
as desert and cold weather operations supplement
normal procedures checks in this section and are
covered in Section V, ADVERSE ENVIRONMENTAL
CONDITIONS.
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 55-1510-215-CL. To provide for easier
cross referencing the procedural steps are numbered to
coincide with the corresponding numbered steps in this
manual.
8-10. Checklist Callout.
Pilot and crewmembers shall not rely on memory for
verifying prescribed operational checks, except those
immediate action emergency procedures that must be
memorized for safe aircraft emergency operation. Oral
callout and confirmation of checklist items shall be
accomplished by pilot and the crewmembers.
8-11. Preflight Check.
The preflight check is performed as follows:
NOTE
If ferry fuel is used, refer to chapter 2
for
complete
operating
and
emergency procedures.
a. Before Exterior Check.
8-8. Symbols Definition.
Items which apply only to night or only to instrument
flying shall have an “N” or and “I”, respectively,
immediately preceding the check to which it is pertinent.
The symbol “O” shall be used to indicate “if installed”.
Those duties which are the responsibility of the copilot,
will be indicated by a circle around the step number i.e.,
5 Starter and IGN SYS circuit breakers - IN. The
symbol star “H” indicates an operational check is
required.
Operational checks are contained in the
performance section of the condensed checklist. The
symbol asterisk “*” indicates that performance of step is
mandatory for all thru-flights, when there has been no
change in crew. The asterisk applies only to checks
performed prior to takeoff. Placarded items appear in
upper case.
*
1. Publications - Check DA Form 2408-12,
-13, -14, and availability of DOD AVFUEL Identaplate
DD Form 1896, Operator's Manual (-10), Checklist (-CL)
and locally required forms and publications.
2. Oxygen cylinder pressure valves - As
required.
*
3. Oxygen system pressure - Check.
*
4. Keylock switch - OFF.
5. Fuel firewall valves - OPEN and safetied.
CAUTION
8-9. Checklists.
Normal procedures are given primarily in checklist
form and amplified as necessary in accompanying
paragraph form when a detailed description of a
8-2
If high or gusty winds are present,
and the flight controls are unlocked,
control surfaces may be damaged by
buffeting.
TM 55-1510-215-10
*
6. Flight controls - Unlocked.
17. Fire extinguishers (2) - Check as required.
18. Fire axe - Secured.
19. First aid kits (5) - Check.
CAUTION
b. Exterior Check (fig. 8-1).
The elevator trim tab control must
not be forced past the limits which
are marked on the elevator trim
indicator scale.
*
NOTE
Fuel and oil quantity check may be
performed prior to EXTERIOR CHECK
to preclude carrying ladder around
during the inspection.
7. Parking brake - Set.
8. Trim tabs - Zero.
(1) Fuel sample sample for contamination.
9. Avionics MASTER switch - OFF.
*
Check
collective
CAUTION
WARNING
Do not cycle landing gear handle on
the ground.
If fuel is detected at transfer pump or
auxiliary pump seal drain tube, a fire
hazard exists.
10. Gear handle - DN.
11. Battery
minimum).
-
ON
fuel
(2) Left wing.
(Stabilized,
22
volts
12. Strobe beacons - Check illumination.
13. Lighting systems - Check as required.
14. Pitot, stall warning, fuel vent and battery
* H
vent heat systems - Check.
(1) Pitot heat switch - ON (check cover
removed).
1. Skin condition - Check for skin damage
such as buckling, cracking, splitting, distortion, dents, or
fuel leaks.
2. Controls, flaps and trim tab - Check hinge
attachment and trim tab rig.
3. Static wicks - Check for security of
discharge wicks.
4. Wing tip and navigation light - Check.
(2) Stall warning heat switch - ON.
(3) Fuel vent heat switches (2) - ON.
(4) Pitot tube - Check by feel for heat
and free of obstructions.
(5) Stall warning vane - Check by feel
for heat, condition and operation.
(6) Fuel vents (2) - Check by feel for
heat, and free of obstructions.
(7) Battery vent - Check by feel for
heat, and free of obstructions.
(8) Pitot heat switch - OFF.
(9) Stall warning heat switch - OFF.
(10) Fuel vent heat switches (2) - OFF.
15. Battery - OFF.
16. Safety belts, shoulder harnesses, inertia
reels - Check condition and operation.
5. Landing light - Check.
6. Tiedown - Released.
7. Fuel vent - Check flush vent free of
obstructions.
*
8. Wing tank fuel and cap - Check fuel level
visually. Check seal is installed, cap is tight and properly
installed (latch tab aft).
9. Deicer boot - Check bonding secure, boot
free of cuts and cracks, stall stripes installed and
secured.
10. Wing ice lights - Check.
11. Fuel vents (2) - Check heated and
recessed vents free of obstructions.
12. Inverter air intake screen and exhaust port
- Check free of obstructions.
Change 7 8-3
TM 55-1510-215-10
Figure 8-1. Exterior Check
(3) Left main landing gear.
*
*
1. Tire - Check for cuts, bruises, wear,
and proper inflation.
2. Brake assembly - Check brake lines
for damage or signs of leakage, and brake linings for
wear.
*
3. Shock strut - Check for signs of
leakage and 2.75 INCH EXTENSION (MINIMUM) left
and right struts approximately equal.
7. Doors and linkage - Check.
8. Air bypass and oil cooler (rear) Check free of obstructions and oil leaks.
*(O)
9. Firewall fuel filter drain (at inertial
separator duct) - Turn/release. Check for fuel drainage.
(4) Left engine and propeller.
1. Accessory section exhaust vent Check free of obstructions.
4. Torque knee - Check.
2. Starter-generator air intake - Check
free of obstructions.
5. Safety switch - Check.
6. Wheel
Check.
8-4
well
general
condition
-
3. Left cowl locks - Locked.
TM 55-1510-215-10
4. Left exhaust stub - Check for cracks
(6) Left nose avionics compartment.
and security.
*
5. Propeller blades and spinner Check blades for nicks, security of spinner and free
rotation.
*
obstructions.
6. Nacelle air intake - Check free of
7. Nacelle
condition and security.
lip
ice
boot
-
Check
(O)
keyed.
1. Voice security computer - Installed/
(O)
keyed.
2. Transponder computer
(O)
3. Transponder - Set M-2 code.
4. Left
access door - Secured.
*
8. Oil cooler air intake - Check free of
obstructions and leakage.
avionics
compartment
(7) Nose section.
1. Wheel
9. Right cowl locks - Locked.
10. Right exhaust stub - Check for
cracks and security.
nose
- Installed/
well
general
condition
-
Check.
2. Doors and linkage - Check.
3. Nose gear turning stop - Check
CAUTION
condition.
*
4. Tire - Check for cuts, bruises, wear,
and proper inflation.
A cold oil check is unreliable. If the oil
level is LESS THAN 3 QUARTS LOW the
oil quantity is sufficient to operate the
engine for total fuel duration unless
engine oil consumption is excessive or
a leak exists. Do not add oil. However,
if cold oil check indicates MORE THAN
3 QUARTS LOW, motor engine 15 to 20
seconds then recheck and add as
required. Do not overfill.
5. Torque knee - Check.
*
6. Shock strut - Check for signs of
leakage and a 2.75-INCH EXTENSION (MINIMUM).
7. Shimmy
damper
and
attaching
linkage - Check.
8. Taxi light - Check.
9. Radome - Check.
10. Windshield and wipers - Check.
*
11. Engine compartment - Check for
fuel and oil leaks, oil level and positively secure oil cap
(latch tab aft).
*
12. Nacelle tank fuel and cap - Check
fuel level visually. Check seal is installed, cap is tight
and properly installed (latch tab aft).
*(O)
13. Fuel filter drain ring - Pull/release.
Check for fuel drainage.
*
14. Engine compartment access door Locked. Visually check locking hooks.
(5) Fuselage underside.
1. General condition - Check.
11. Ram air intake - Check free of
obstructions.
12. Ram air intake lip ice boot - Check
security and condition.
13. Right nose avionics compartment
access door - Secured.
14. Battery compartment access panel Secured (top and bottom).
(8) Right engine and propeller.
1. Accessory section exhaust vent Check free of obstructions.
2. Antennas - Check for security.
3. Strobe beacon - Check condition.
Change 7 8-5
TM 55-1510-215-10
2. Starter-generator air intake - Check
free of obstructions.
*
8. Air bypass and oil cooler (rear) Check free of obstructions and oil leaks.
3. Left cowl locks - Locked.
4. Left exhaust stub - Check for cracks
and security.
* (O)
9. Firewall fuel filter drain (at inertial
separator duct) - Turn/release. Check for fuel drainage.
*
5. Propeller blades and spinner Check blades for nicks, security of spinner and free
rotation.
*
obstructions.
6. Nacelle air intake - Check free of
7. Nacelle
condition and security.
lip
ice
boot
7. Doors and linkage - Check.
-
Check
*
8. Oil cooler air intake - Check free of
obstructions and leakage.
(10)
Right wing.
1. Inverter air intake
exhaust port - Check free of obstructions.
11. Engine compartment - Check for
fuel and oil leaks, oil level and positively secure oil cap
(latch tab aft).
*
12. Nacelle tank fuel and cap - Check
fuel level visually. Check seal is installed, cap is tight
and properly installed (latch tab aft).
3. Heated battery vent - Check for
general condition and security.
4. Wing ice light - Check.
5. Deicer boot
secure, boot free of cuts and
(9)
Right main landing gear.
*
1. Tire - Check for cuts, bruises, wear,
and proper inflation.
2. Brake assembly - Check brake lines
for damage or signs of leakage, and brake linings for
wear.
*
3. Shock strut - Check for signs of
leakage and 2.75-INCH EXTENSION (MINI-MUM).
4. Torque knee - Check.
5. Safety switch - Check.
6. Wheel well general condition Check.
-
Check
bonding
*
6. Wing tank fuel and cap - Check fuel
level visually. Check seal is installed, cap is tight and
properly installed (latch tab aft).
*
7. Tiedown - Released.
8. Fuel vent - Check flush vent free of
obstruction.
9. Landing light - Check.
*
13. Fuel filter drain ring - Pull/release.
Check for fuel drainage.
*
14. Engine compartment access door Locked. Visually check locking hooks.
and
2. Fuel vents (2) - Check heated,
recessed and flush vents free of obstructions.
9. Right cowl locks - Locked.
10. Right exhaust stub - Check for
cracks and security.
screen
10. Wing tip and navigation light Check.
11. Static wicks - Check for security of
discharge wicks.
12. Controls, flaps, and trim tabs Check hinge attachment and trim tab rig.
13. Skin condition - Check for skin
damage, such as buckling, cracking, splitting, distortion,
dents or fuel leaks.
(11)
Fuselage right side.
1. Skin condition - Check for skin
damage, such as buckling, cracking, splitting, distortion
or dents.
2. Cabin air exhaust vents - Check free
of obstructions.
*
3. Antennas - Check for security.
4. Static port - Check free of dirt or
obstructions.
8-6 Change 7
TM 55-1510-215-10
*
5. Tiedown - Released.
5. Chocks - Removed.
*(O)
6. Tail stand - Removed.
c. Interior Check.
(12) Empennage.
1. Ladder - Stowed.
1. Right horizontal stabilizer deicer
boot - Check bonding secure, boot free of cuts and
cracks.
2. Right horizontal stabilizer - Check
security, cracks and skin condition.
3. Static wicks - Check for security of
discharge wicks.
4. Right elevator and trim tab - Check
security and condition. Verify neutral position of the trim
tab.
5. Navigation
and
beacon
lights
-
Check condition.
6. Rudder
security and condition.
7. Vertical
and
trim
tab
-
Check
*
2. Cargo/loose equipment - Secured.
*
3. Cargo door - LOCK.
*
4. Main entrance door - LOCK.
5. Cabin emergency exit hatch - Secured
(safety seal intact).
*H
6. Crew/passenger briefing
(Section VI).
- As required
8-12. Before Staring Engines.
*
1. Seats, pedals, belts, harnesses - Adjust.
2. Cockpit emergency
Secured (safety seal intact).
entrance/exit
hatch
-
3. Overhead control panel switches - Set.
stabilizer
-
Check
skin
condition.
4. Magnetic compass - Check for fluid, heading,
and correction card.
8. Left elevator and trim tab - Check
security and condition. Verify neutral position of the trim
tab.
5. Free air temperature gage - Note current
reading.
6. Fire detection test switch - OFF.
9. Static wicks - Check for security of
discharge wicks.
10. Left horizontal stabilizer - Check
security, cracks and skin condition.
CAUTION
11. Left horizontal stabilizer deicer boot
- Check bonding secure, boot free of cuts and cracks.
12. Vertical stabilizer deicer boot
Check bonding secure, boot free of cuts and cracks.
Movement of the power levers to the
full forward position with the
condition levers in the FUEL CUTOFF
position may result in bending and
possible damage to common linkage.
-
(13) Fuselage left side.
1. Static port - Check for freedom from
dirt or obstructions.
*
7. Power levers - IDLE.
CAUTION
2. Cabin air exhaust vent - Check free
of obstructions.
3. Skin condition - Check for skin
damage, such as buckling, cracking, splitting, distortion
or dents.
Do not position the power levers into
the REVERSE range while the
engines are shut down. Damage to
the reversing linkage will result.
4. Main entrance and cargo doors Check.
*
8. Propeller levers - HIGH RPM.
Change 7 8-7
TM 55-1510-215-10
*
9. Condition levers - FUEL CUTOFF.
10. Flaps - UP
11. Landing gear emergency clutch disengage lever
- Stowed.
12. Landing gear emergency extension handle Stowed.
13. Fuel system circuit breakers - Check in.
14. Auxiliary fuel pumps - OFF.
15. Transfer pumps - OFF.
16. Crossfeed - CLOSED.
17. Deleted.
*
36. DC GPU - Connect as required.
*
37. Battery - ON.
*
38. Voltage - Check (28 VDC maximum).
*
39. Annunciator panel - Test.
*(N)40. Navigation lights - ON.
*
41. Landing gear handle lights (2) - Test.
42. Landing
Illuminated.
*
gear
down
indicator
lights
(3)
-
43. Keylock switch - ON.
* 44. Fire detection system - Test. Verify warning
horn sounds; simultaneously the red MASTER
WARNING light flashes and FIRE L ENG and FIRE R
ENG lights illuminate in each (3) numbered fire detection
system test switch positions.
18. Engine instruments - Check placards, and
slippage marks.
NOTE
A ground test of the engine fire
detection system insures electrical
continuity within the system.
19. Deleted.
20. Emergency static air source - NORMAL.
21. Copilot's circuit breaker panel - Check circuit
breakers in.
22. Right subpanel circuit breakers - Check in.
23. Heater - OFF.
24. Gear handle - DN.
25. Windshield anti-ice switches - OFF.
26. Inlet air separator - OFF.
27. Left subpanel light switches (4) - OFF.
28. Deice cycle switch - Centered (off).
29. Autofeather switch - OFF.
30. Heat switches (9) - OFF.
31. Landing lights - OFF.
32. Engine ice vanes - As required.
33. Ignition/start switches (2) - OFF.
34. Engine autoignition - OFF.
35. Inverters - OFF.
8-8 Change 7
45. Master warning button - Press, FAULT WARN
light should extinguish.
*
46. Generators - OFF.
*H 47. Auxiliary fuel pumps and crossfeed - Check as
follows:
NOTE
The
auxiliary
fuel
pump
and
crossfeed check must correspond to
the engine being started first.
(1) Fuel fail lights - Illuminated.
NOTE
FUEL
FAIL
lights
may
be
extinguished, if a temperature rise
has expanded trapped fuel creating a
pressure rise within the lines.
Trapped
pressure
should
be
dissipated by briefly pulling the ring
of the FUEL STRAINER DRAIN in
each nacelle. Release the ring, after
enough fuel has drained to relieve
trapped pressure. Both FUEL FAIL,
lights should illuminate.
TM 55-1510-215-10
(2) Crossfeed - CLOSED.
(3) Auxiliary fuel pump - ON.
FAIL light extinguishes.
Check FUEL
NOTE
If both FUEL FAIL lights extinguish,
the crossfeed valve is malfunctioning.
(4) Crossefed - OPEN. Check that FUEL
CROSSFEED light illuminates and the other FUEL FAIL
light extinguishes.
(5) Auxiliary fuel pump - OFF.
CAUTION
Monitor ITT to avoid a hot start. If
there is a rapid rise in ITT, be
prepared to abort the start before
limits are exceeded. During engine
start, the maximum allowable ITT is
1090°C for two seconds. If this limit
is exceeded, use ABORT START
procedure and record the peak
temperature and duration on DA
Form 2408-13.
8-13.* Starting Engines (Battery/GPU).
A GPU will be used for engine starting whenever
possible. When making a battery start, start the right
engine first (starting circuit is shorter).
CAUTION
For a GPU start, it is recommended to
start the left engine first, a the GPU
receptacle
is
located
on
the
underside of the right wing outboard
of the engine nacelle.
CAUTION
If ignition does not occur within 10
seconds after moving the condition
lever to LO IDLE, initiate ENGINE
CLEARING procedure.
If for any
reason
a
starting
attempt
is
discontinued, the entire sequence
must be repeated after allowing the
engine to come to a complete stop.
CAUTION
If the N1 tachometer immediately
indicates above 20%, discontinue the
start.
CAUTION
Do not exceed starter limitation of 40
seconds on, and 60 seconds off for
two starter operations; then 40
seconds on, 30 minutes off.
Change 7 8-9
TM 55-1510-215-10
a.
St art procedure.
21. Fuel control eat switches - ON (LEFT and
RIGHT).
1. Strobe beacon switches (2) - As required.
b.
(N)
Abort Start.
2, Navigation lights - ON.
1. Condition lever - FUEL CUTOFF.
3. Proper - Clear (fig 8-2, Exhaust Danger
Area).
2. Ignition/start switch - STARTER ONLY.
4. Ignition/start switch - On (check IGN ON light
illuminated).
3. ITT - Monitor for drop in temperature.
4. Ignition/start switch - OFF.
5. Condition lever - LO IDLE (after
stabilizes at or above 13 % for 5 seconds).
N1
c.
6. ITT - Monitor (1090°°C for two seconds
maximum for engine being stared. 750°C maximum
for operating engine).
Engine Clearing.
1. Condition lever - FUEL CUTOFF
2. Ignition/start switch - OFF (allow 30 seconds
delay).
7. Ignition/start switch - OFF after ITT has
stabilized. IGN ON light extinguished.
3. Ignition/start switch - STARTER ONLY (for 30
to 40 seconds).
8. Condition lever - HIGH IDLE.
4. Ignition/start switch – OFF.
9. Generator (for battery start) - Reset, then ON.
GEN OUT light extinguished.
10. Aircraft inverter - 2.
extinguished.
Check INV 2 light
12. Oil pressure - Check (40 PSI minimum).
13. Aircraft inverter - OFF.
14. GPU - Disconnect.
15. Generator - Reset, then ON.
OUT light extinguished.
Check GEN
16. Loadmeter - Monitor (0.5 maximum).
17. Second engine - Start 4: through 8 above.
(I) H 3. Windshield anti-ice operation - Check. The
electrothermal windshield will prevent initial ice adhesion
to the glass, but will not remove ice already formed on the
windshield. There, anti-icing switches must be placed ON
before entering icing conditions. When glass temperature
exceeds 43 degrees centigrade, operational checks of the
electrothermal windshields will be unreliable.
(1) Pilot's windshield anti-ice switch - ON
(watch volt-loadmeter for a slight increase).
(2) Copilot's anti-ice switch - ON (confirm
additional meter loads).
Check GEN OUT light
19. Aircraft inverters - 1. Check INV 1 light
extinguished (repeat steps 12 and 13).
20. Condition levers - LO IDLE.
8-10 Change 10
1. Avionics master switch - ON.
2. Radios - ON. Communication and navigation
radios ON/set required; radios and transponder STBY.
11. Deleted
18. Generator - ON.
extinguished.
8-14. *Before Taxiing.
(3) Reposition switches as required for flight.
NOTE
The magnetic compass is erratic when
the windshield anti-ice system is on.
TM 55-1510-215-10
H
4. Autopilot/electric trim system operation - Check
as follows:
(1) Verify
disengaged.
that
all
autopilot
modes
are
(2) Preflight TEST button (autopilot mode
control panel) - Press and hold.
(3) Autopilot mode annunciator panel displays
- Check all illuminated. TRIM display should flash at
least four times, but no more than six times.
(4) Aural trim alert - Listen for sound.
(5) Yaw damp indicator light (autopilot mode
control panel) - Check illuminated.
(6) ARM and ALERT display (altitude selector
panel) - Check displayed.
(7) Computer flag
indicator) - Check in view.
(pilot's
flight
director
(8) Preflight TEST button (autopilot mode
control panel) - Release.
(9) Manual electric trim system - Check as
follows:
1. PITCH TRIM switch (pilot's control
wheel) - Move the left side of the trim switch to the
forward and aft position, while moving the pitch-trim
control wheel. The pitch-trim solenoid should engage
making it more difficult to move the pitch-trim control
wheel, but the electric trim motor should not run. Move
the right side of the trim switch to the forward and aft
positions. The pitch-trim solenoid should not engage
and the electric trim motor should not run.
2. Overpower capability - Check by
moving the pilot's PITCH TRIM switch to the forward and
aft positions while holding the manual pitch-trim control
wheel.
3. TRIM
TEST
switch
(pedestal
extension) - Depress and hold while operating the
electric trim up and down using the pilot's PITCH TRIM
switch (control wheel).
4. TRIM annunciator display (autopilot
annunciator panel) - Check displayed.
5. Aural trim alert - Listen for sound.
6. TRIM
extension) - Release.
TEST
switch
(pedestal
7. Autopilot AP DISC/TR IM INTER
switch (control wheel) - Depress and hold. Attempt to
run the electric trim up and down using the pilot's PITCH
TRIM switch (control wheel). The trim system should not
run either UP or DOWN.
8. Repeat steps 1 through 6 - using the
copilot's PITCH TRIM switch.
9. Pilot's and copilot's PITCH TRIM
switches - Simultaneously move the pilot's switch
forward and the copilot's switch aft. The electric trim
should run up.
(10) Flight director switch (FD) (autopilot mode
control panel) - Press on.
(11) Autopilot switch (AP) (autopilot mode
control panel) - ON.
(12) Control wheel steering
(control wheel) - Depress and hold.
switch
(CWS)
(13) Control wheel - Manually move to neutral
position.
(14) Control wheel steering switch (CWS) Release.
(15) Control wheel - Apply force to all axes.
Determine that autopilot can be overpowered.
(16) Autopilot AP DISC/TRIM INTER (control
wheel) - Depress to disconnect autopilot.
(17) Manual trim - Set for takeoff.
(18) Autopilot switch (AP) (autopilot mode
control panel) - ON.
(19) ROLL TEST switch (control pedestal) Hold to LT position for approximately two seconds.
Autopilot should disconnect and the aural alert should
sound.
(20) Autopilot switch (AP) (autopilot mode
control panel) - ON.
(21) ROLL TEST switch (control pedestal) Hold to RT position for approximately two seconds.
Autopilot should disconnect and the aural alert should
sound.
(22) Autopilot switch (AP) (autopilot mode
control panel) - ON.
Change 5 8-10A
TM 55-1510-215-10
(23) PITCH TEST switch (control pedestal) Hold to UP position for approximately two seconds.
Autopilot should disconnect and the aural alert should
sound.
(24) Autopilot switch (AP) (autopilot mode
control panel) - ON.
(25) PITCH TEST switch (control pedestal) Hold to DN position for approximately two seconds.
Autopilot should disconnect and the aural alert should
sound.
(26) Autopilot switch (AP) (autopilot mode
control panel) - ON.
(27) Vertical trim control (autopilot mode
control panel) - Move to insert a pitch-up command.
(28) Control wheel - Hold to keep from moving
and observe that the trim wheel moves in the nose up
direction after a three second delay.
(29) Control wheel steering
(control wheel) - Depress momentarily.
switch
(36) Autopilot switch (AP) (autopilot
control panel) - OFF.
mode
(37) Flight director switch (FD) (autopilot mode
control panel) - Disengage.
(38) PITCH TRIM switch (control wheel) Move aft until a full nose-up trim position has been
attained, then move switch forward and simultaneously
begin timing. When the full nose-down trim position has
been attained release switch and note time. The time
required for trim system to run from full nose up to full
nose down should be 45 ±9 seconds.
(39) If the autopilot fails the preflight test, the
AUTOPILOT circuit breaker must be pulled. However,
manual electric trim may still be used. If the electric trim
system fails the preflight test, the ELEC TRIM circuit
breaker must be pulled and neither the electric trim nor
the autopilot may be used.
4A. Ground proximity altitude
(GPAAS) - Check as follows:
advisory
system
(CWS)
(1) GPAAS voice advisory VOL control - Full
clockwise.
(30) Vertical trim control (autopilot mode
control panel) - Move to insert a pitch-down command.
(31) Control wheel - Hold to keep from moving
and observe that the trim wheel moves in the nose down
direction after a three second delay.
(2) VOICE OFF switch-indicator - Extinguished.
(3) Audio control panel - Set listening audio
level.
(4) VA FAIL annunciator light - Extinguished.
(32) Control wheel steering switch (CWS)
(control wheel) - Depress and center the control wheel
about the roll axis, then release.
(33) Heading switch (HDG) (autopilot mode
control panel) - Press on.
(34) Heading select bug (horizontal situation
indicator) - Set bug to command a right turn. The control
wheel should rotate clockwise.
(35) Heading select bug (horizontal situation
indicator) - Set bug to command a left turn. The control
wheel should rotate counterclockwise.
Change 9 8-10B
(5) Radio altimeter DH SET control - Set to
200 feet.
(6) Radio altimeter TEST switch - Press and
hold "Minimum, minimum" will be announced once
followed by the illumination of the VA FAIL light.
(7) Radio altimeter TEST switch - Release.
TM 55-1510-215-10
NOTE
THE EXHAUST DANGER AREA DOES NOT INCLUDE PROPELLER WAKE WHICH
INCREASES VELOCITY, AND SIGNIFICANTLY REDUCES TEMPERATURE.
EXHAUST GAS TEMPERATURE AND VELOCITY AT GROUND IDLE IS VERY LOW.
HOWEVER, THE IMMEDIATE AREA OF EXHAUST DISCHARGE SHOULD BE
AVOIDED.
Figure 8-2. Exhaust Danger Area
8-11
TM 55-1510-215-10
H
5. Oxygen system - Check as required.
8-16. Engine Runup.
(1) Cockpit oxygen supply valve (left cockpit
sidewall) - As required.
Turn the aircraft into the wind if possible, and
perform the following checks:
(2) Cabin oxygen supply valve (left cockpit
sidewall) - As required.
*
1. Nose wheel - Center.
NOTE
(3) Oxygen supply pressure gage (left cockpit
sidewall) - Check.
The
nose
wheel
cannot
be
straightened with rudder pedals
when the aircraft is stopped.
(4) Oxygen supply pressure gage (regulator
control panel) - 300 to 400 PSI.
*
2. Parking brake - Set (keep feet on rudder
pedals).
(5) Supply control lever (green) - ON.
(6) Diluter
OXYGEN.
control
lever
(white)
-
100%
NOTE
The parking brake can be set only
from pilot's seat.
(7) Emergency pressure control lever (red) NORMAL.
(8) Oxygen mask hose - Connect to mask
hose connection.
(9) Emergency pressure control lever (red) Set to TEST MASK position while holding mask directly
away from your face, then return lever to NORMAL.
(10) Oxygen mask - Put on and adjust to face.
(11) Emergency pressure control lever (red) Set to TEST MASK position and check mask for leaks,
then return lever to NORMAL.
(12) Flow indicator - Check (during inhalation
blinker appears, during exhalation blinker disappears).
Repeat a minimum of 3 times.
6. Radios - Check intercom, communication and
navigation systems and radar.
7. Taxi clearance - Check.
8. Clock - Set.
9. Altimeters - Set.
*
3. Power levers - IDLE.
*
4. Condition levers - LO IDLE.
*H
5. Fuel transfer pumps - Check as follows:
(1) Transfer test switch - Hold to "R".
(2) Right transfer pump
watching annunciator panel) - ON.
switch
(while
(3) Monitor R FUEL XFR light - Check for
momentary flash.
NOTE
Steady illumination or lack of
momentary flash indicates a transfer
pump system fault. For the system
fault recheck, switch transfer pump
OFF for 5 to 10 seconds and repeat
fuel transfer pump check.
The
momentary flash may not be detected
if the cockpit lights are on.
(4) Repeat check procedure for left transfer
pump system.
10. Parking brake - Release.
6. Flaps - Check operation of 4 panels and flap
position indicator.
8-15. *Taxiing.
1. Brakes - Check.
2. Flight
operation.
8-12 Change 5
instruments
-
Check
settings
and
TM 55-1510-215-10
(5) Power lever  Return to 500 FT-LB
CAUTION
TORQUE.
(6) Repeat steps 4 and 5 using the
other power lever.
Delete.
H
7. Propeller manual feather  Check by
pulling propeller levers aft through detent to FEATHER.
Check that propeller will feather, then advance lever to
HIGH RPM.
*H
8. Engine autoignition system  Check.
(1) Power levers  Advance to ABOVE
450 FT-LB TORQUE.
(2) Autoignition  ARM (check green
IGNITION ARM lights illuminated).
(3) Power levers  Retard to LESS
THAN 350 FT-LB TORQUE (annunciator L and R IGN
ON lights illuminated, green IGNITION ARM light
extinguished).
(4) Autoignition  OFF.
(5) Power levers  IDLE.
H
9. Propeller autofeather system  Check.
(1) Power levers  IDLE.
(2) Autofeather test switch  TEST.
Check AUTOFEATHER lights do not illuminate. If switch
is held in TEST position, propellers will gradually feather.
(3) Power levers  Advance to 500 FTLB TORQUE.
(7) Propeller autofeather test switch 
ARM.
(8) Both power levers  Advance to
88% to 92% N1 (observe ITT and torque limits). Check
both AUTOFEATHER lights illuminated. Retard each
power lever individually below 88% to 92% N1. Check
both AUTOFEATHER lights extinguished.
NOTE
If the aircraft is to be flown with a
defective autofeather system, the
prop feather circuit breaker should
be pulled and the AUTOFEATHER
switch positioned to OFF.
H
10. Overspeed governor  Check a setting
RPM to 2100. Hold PROP GOV TEST switches UP.
RPM should DECREASE to 1980 to 2060. Release test
switches. RPM should return to 2100.
WARNING
Do not exceed it or torque limits.
11. Engine ice vanes (left and right)  Pull to
EXT. Check operation by observing drop in torque
reading, verify that handle remains extended when
released, then push to RET.
NOTE
(4) Autofeather test switch  TEST.
Hold to test position and check both AUTOFEATHER
lights illuminated; retard one power lever. AT 350 to 450
FT-LB TORQUE, check opposite AUTOFEATHER light
extinguished. AT 160 to 290 FT-LB TORQUE, check
both AUTOFEATHER lights extinguished; check
propeller star to feather.
NOTE
Loosen friction knobs prior to
primary governor and, secondary idle
stop check to provide better feel.
H
12. Primary governor  Check. Set 1900
RPM with power levers. Retard propeller levers to
detent position. CHECK FOR 1725 to 1775 RPM then
advance propeller levers to HIGH RPM.
AT 160 to 290 FT-LB TORQUE
setting, AUTOFEATHER light may
flicker.
Change 7 8-13
TM 55-1510-215-10
CAUTION
(1)
18. Surface deice systems  Check by
activating the deice switch to SGL and visually check
inflation and deflation of boots and DE-ICING pressure.
Repeat procedure for MNL switch position.
To prevent damage to reversing
linkage do not force power levers to
full REVERSE with test switch ON.
H
13. Secondary idle stop  Check. Check
with condition levers HIGH IDLE and power levers at
IDLE, then while holding the secondary idle stop test
switches down, move power levers slowly toward
REVERSE in one continuous movement, while
observing that the PRIMARY LOW pitch lights illuminate
and an ARPM RISE of 170 to 250 is obtained. Release
the test switch and RPM should increase. Return power
levers to normal idle position and cancel lights in
annunciator panel by actuating secondary flight idle
switch if they remain illuminated.
NOTE
The deicing system should be
actuated on a daily basis to insure
system function and to exercise the
system distributor valve.
H
19. Inlet air separator system  Check as
follows at 70% N1:
(1) Inlet air separator switch  AUTO.
Observe the following:
(a) Torque should decrease on both
engines.
14. Instrument suction  Check (4.5 to 5.2 in
Hg).
(b) ITT should increase on both
engines.
15. Pneumatic pressure  Check (12 to 20
PSI).
16. Volt-loadmeters  Check bus voltage
BETWEEN 26.5 AND 29.5 VOLTS, and load paralleling
within one increment and a maximum of 0.5 on the
loadmeter scale. Note any discrepancy on DA Form
2408-13.
(I)
17. Propeller deice system  Check. Move
the switch to PROP position and check propeller
ammeter normal operation (14-18 amperes) for 2.5
minutes. Flickering of the propeller deicing ammeter
needle may or may not be detectable during cycle
switching.
(c) MASTER CAUTION lights will
flash, and the PARTICLE SEPARATOR light will
illuminate.
(2) Inlet air separator switch  OFF.
Monitor for deactivation of both left and right systems.
The torque and ITT should return to initial values, the
PARTICLE SEPARATOR light should extinguish, and
the MASTER CAUTION lights should stop flashing after
reset.
20. Condition levers  LO IDLE.
H
21. Deleted.
Page 8-15 Deleted.
H U.S. GOVERNMENT PRINTING OFFICE: 1994-555-121/80069
8-14 Change 10
TM 55-1510-215-10
9. Flight director  As required.
10. Trim  Set.
11. Engine ice vanes  Retracted.
12. Fuel control heat  Check ON.
Deleted.
13. Autofeather switch  Check ARM.
(I)
14. Navigation radios  Set.
15. Flight controls  Check for full travel and
proper response.
16. Windows and doors  Secure.
*8-17. Before Takeoff.
1. Fuel panel  Check switches and circuit
breakers set.
2. Auxiliary fuel pumps  ON.
3. Annunciator panel  Check.
4. Engine and flight instruments  Check.
5. Propeller levers  Check HIGH RPM.
6. Friction locks  Set.
7. Flaps  UP.
WARNING
17. Mirror  Retracted.
(I)
18. Anti-icing/deicing/pitot
required.
heat

*8-18. Line Up.
1. Transponder  As required.
NOTE
To prevent zeroizing the MODE 4
function
of
the
AN/APX-72
transponder, when the landing gear
is down (struts compressed) and
either transponder or aircraft power
is to be turned off, place the CODE
switch
momentarily
in
HOLD
position.
2. Gyro heading  Check.
Insure that the autopilot has been
disengaged and check that the
aircraft manual trim indicator is set to
the takeoff position before takeoff.
Operating the autopilot on the ground
may cause the autotrim to run
because of back force generated by
the elevator downsprings or pilot
induced forces.
8. Autopilot/yaw damp  Disengage.
8-16 Change 7
NOTE
Prior to flight, the gyro magnetic
compass indicator should be crosschecked with the magnetic compass
to verify approximate heading as the
indicator annunciator window will
also be clear (blank) if the heading is
180° out of phase.
3. Power  Stabilized (70%-80% N1).
As
TM 55-1510-215-10
NOTE
AUTOFEATHER system arm lights
should illuminate at 88% to 92% N1.
This is to assure autofeather arming
at takeoff torque.
4. Autoignition  As required.
5. Landing/taxi lights  As required.
8-19. Takeoff.
WARNING
Minimum run takeoff is a contingency
maneuver and may be performed
below V mc and below power off stall
speed. Control of the aircraft may be
lost if engine failure occurs at or
immediately following liftoff or until
the best angle of climb speed can be
obtained.
c. Minimum Run Takeoff. Minimum run takeoffs
are performed with APPROACH flaps. After the lineup
check is completed, set takeoff power and release
brakes. As soon as elevator control becomes effective
(approximately 50 KIAS) raise the nose wheel clear of
the runway. Continue to apply elevator pressure in a
smooth continuous motion until clear of the runway, then
lower nose and accelerate to desired climb speed.
Monitor ITT and torque. Retract the landing gear when
flight is assured.
d. Obstacle Clearance Climb. Follow procedures
as outlined for a minimum run takeoff, to the point of
actual liftoff. When flight is assured, retract the landing
gear and establish a wings-level climb attitude,
maintaining the computed best angle-of-climb speed
(V x ). Climb at this speed until clear of the obstacle.
After the obstacle is cleared, accelerate to an airspeed
equal to or exceeding best rate-of-climb airspeed (V y ).
Retract flaps after attaining single-engine best rate-ofclimb airspeed (V yse).
8-20. After Takeoff.
CAUTION
CAUTION
To prevent damage to the landing
gear retraction mechanism, brakes
should not be applied to slow down
the rotation of the tires while
retracting the landing gear or after
gear is up. A rubberized drag brake
shoe is provided in the wheel well to
stop the wheels from rotating after
retraction.
Retract landing gear below 130 KCAS
(127 KIAS).
Additional airspeed
imposes excessive air loads on the
gear retraction mechanism.
1. Gear  UP.
2. Flaps  UP.
3. Climb power  Set.
4. Auxiliary fuel pumps  OFF.
To aid in planning the takeoff and to obtain
maximum aircraft performance refer to chapter 7.
a. Normal Takeoff. Apply takeoff power. Maintain
directional control with nosewheel steering and rudder
while maintaining wings level with ailerons. Use of
differential power at the beginning of takeoff roll will
assist in directional control. Rotate at (V r ) and allow the
aircraft to accelerate to the desired climb airspeed.
Rotation should be at a rate that will allow liftoff at the
liftoff speed (V lof ).
b. Crosswind Takeoff. Position the aileron control
into the wind at the start of the takeoff roll. Leading with
upwind power at the beginning of the takeoff roll will
assist in maintaining directional control.
5. Autofeather system  OFF.
6. Flight director/yaw damp  As required.
7. Wings and nacelles  Check for fuel and
oil leaks.
8. Landing/taxi lights  As required.
8-21. Climb  Normal.
Refer to chapter 7 for normal climb speeds.
Change 8 8-17
TM 55-1510-215-10
8-22. Climb  Max Rate.
3. Flaps  APPROACH below 174 KCAS
(173 KIAS).
Refer to chapter 7 for maximum rate-of-climb
speeds.
4. Gear  DN below 156 KCAS (154 KIAS).
8-23. Cruise Checks.
NOTE
Refer to chapter 7, for the necessary flight planning
information. Refer to chapter 2 for necessary fuel
system management. The following procedures are
applicable to all cruise requirements.
Deleted.
5. Flaps  DOWN below 130 KCAS (127
KIAS).
6. Airspeed  130 KCAS (127 KIAS).
CAUTION
Maximum turbulent air penetration
speed is 169 KCAS (168 KIAS).
1. Power  Set. Do not exceed maximum
cruise-torque settings in chapter 7.
2. Wings and nacelles  Check for fuel and
oil leaks.
8-24. Fuel System Crossfeed.
Refer to chapter 9 for fuel system crossfeed
emergency procedures.
8-25. Descent.
Descent Max Rat e (Clean).
1. Power  IDLE.
2. Propellers  HIGH RPM.
3. Gear  UP.
4. Flaps  UP.
5. Airspeed  208 KCAS (206 KIAS)
(maximum).
b.
Perform the following checks prior to traffic pattern
entry:
1. Seat belts and shoulder harnesses 
Secure (passengers checked).
2. Fuel panel  Check.
3. Parking brake handle  In.
3. Deleted.
a.
8-26. Descent-Arrival Check.
Descent  Max Angle (Landing Configuration).
1. Power  IDLE.
2. Propellers  HIGH RPM.
8-18 Change 8
4. Inlet air separator  As required.
5. Engine ice vanes  As required.
8-27. Before Landing.
1. Auxiliary fuel pumps  ON.
2. Autofeather  ARM.
3. Flaps  APPROACH below 174 KCAS
(173 KIAS).
4. Gear  DN below 156 KCAS (154 KIAS).
Check lights.
5. Autopilot/yaw damp  Disengaged.
6. Landing lights  As required.
NOTE
If the landing gear is down and
locked, the gear position indicator
lights on the control pedestal will
illuminate green and the internal red
light in the handle will be
extinguished. The warning horn will
be silent when power is reduced.
TM 55-1510-215-10
8-28. Obstacle Clearance Approach.
Fly a traffic pattern completing the before landing checks
as previously described. Lower flaps on base leg as
required and complete final landing check after turning
final. Adjust pitch, power and flaps to maintain the
desired approach angle and airspeed so as to arrive at
the intended landing point at minimum touchdown speed
consistent with existing conditions.
WARNING
Obstacle clearance approach speeds
may be below Vmc.
8-29. Landing.
NOTE
100% flaps must be used to obtain
the minimum landing distances in
chapter 7.
1. Gear  Recheck DOWN (check lights).
2. Propellers  As required (HIGH RPM, if
beta or reverse is required).
NOTE
Propeller reverse below 40 knots will
increase propeller blade erosion.
Exercise caution when reversing on
surfaces where loose sand, gravel, or
dust are present.
a. Normal Landing. Refer to chapter 7 for landing
data. Plan the final approach to arrive at runway
threshold at recommended approach speed. As the
aircraft touches down, gently lower the nosewheel to the
runway and use reversing, brakes, or beta range, as
required.
b. Crosswind Landing. Crab or slip into the wind to
correct for drift during final approach. The crab is
changed to a slip for roundout and touchdown.
Differential power may be used to aid in aircraft control
during approach and landing. Refer to chapter 7 for
landing data.
c. Power Approach/Precision Landing. Set power,
flaps, and trim as required to maintain the desired
airspeed and descent angle.
d. Touch And Go Landing. If a touch and go
landing is to be performed, allow the nose wheel to
touch the runway and perform the following:
1. Flaps  As required.
2. Trim  Set.
3. Power  Max allowable.
8-30. Go-Around.
When a go-around is started before the LANDING
check, use power as required to climb to, or maintain,
desired altitude and airspeed. If the go-around 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 ensuing that the aircraft will not touch the
ground. Retract the flaps to APPROACH, adjusting pitch
attitude simultaneously to avoid altitude loss. Accelerate
to best rate-of-climb airspeed (Vy), retracting flaps fully
after attaining the Vref speed used for the approach and
perform the following:
1. Power  As required.
2. Gear  UP.
3. Flaps  UP.
4. Landing lights  OFF.
5. Climb Power  Set.
6. Yaw damp  As required.
8-31. After Landing (clear of the runway).
CAUTION
Prolonged use of landing lights on
the ground may cause heat damage
to the plastic landing light shield due
to the lack of cooling airflow.
1. Landing/taxi lights  As required.
2. Propellers  HIGH RPM.
3. Flaps  UP.
Change 7 8-19
TM 55-1510-215-10
4. Auxiliary fuel pumps  OFF.
5. Autoignition  OFF.
14. Master switch  Down.
15. Oxygen regulator
NORMAL, 100%, OFF.
6. Anti-icing/deicing  OFF.
control
levers

16. Keylock switch  OFF.
7. Inlet air separator  OFF.
8-33. Before Leaving Aircraft.
8. Engine ice vanes  As required.
1. Wheels  Chocked.
9. Radar/transponder  Standby.
NOTE
NOTE
To prevent zeroizing the MODE 4
function of the transponder, when the
landing gear is down and the struts
compressed, place the CODE control
momentarily in HOLD position before
either transponder or aircraft power
is turned off. CODE HOLD condition
will only be removed when the struts
are extended and the MASTER
control is not in OFF position.
(O)
10. Voice security  Zeroize.
8-32. Engine Shutdown.
1. Parking brake  Set.
2. Landing/taxi lights  OFF.
3. Heater  OFF.
4. Vent blower  OFF.
5. Avionics master switch  OFF.
6. Autofeather switches  OFF.
7. Heat switches (9)  OFF.
8. Inverters  OFF.
9. Propellers  FEATHER.
10. Conditions Levers  FUEL CUTOFF (ITT
610°C or less).
11. Transfer pumps  OFF.
12. Crossfeed  CLOSED.
13. Beacon/lighting systems  OFF.
8-20 Change 5
Parking brakes are primarily intended
for short term usage. Temperature
changes or system leakage may
cause inadvertent brake release.
2. Parking brake  As required.
3. Flight controls  Locked.
(O)
4. Voice security computer  Removed.
(O)
5. Transponder computer  Removed.
(O)
6. Transponder  Check zeroized.
7. Windows and doors  Closed.
8. Walk around inspection  Completed.
NOTE
If strong winds are anticipated while
the aircraft is unattended, the
propellers should be secured to
prevent their windmilling with zero
engine oil pressure.
9. DA Form 2408-12 and -13  Completed.
NOTE
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.
10. Aircraft  Secure.
TM 55-1510-215-10
Section III. INSTRUMENT FLIGHT
8-34. General.
NOTE
This aircraft is certified for operation under
instrument flight conditions. Flight handling, stability
characteristics, and range are the same during
instrument flight conditions as when under visual flight
conditions. Adequate navigation and communication
equipment is installed to provide instrument flight.
NOTE
For proper instrument operation do
not taxi until all gyro warning flags
are masked from view.
8-35. Instrument Flight Procedures.
Refer to FM 1-5, FM 1-30, FLIP, AR 95-1, FAR Part
91, and procedures described in this manual.
a. Instrument Takeoff.
Complete the checks
prescribed in this chapter. Align the aircraft with the
center of the runway. Allow it to roll straight ahead for a
few feet before stopping so the nose wheel will be
straight. Hold the brakes and note the heading to be
maintained during takeoff roll. Follow takeoff procedures
dictated by local conditions.
NOTE
A slight amount of pitch error in the
indication of the attitude indicator will
result
from
accelerations
or
decelerations. It will appear as a
slight climb indication after an
acceleration and a slight dive
indication after deceleration.
The
error will be most noticeable at the
time the aircraft breaks ground
during takeoff.
b. Instrument Climb.
Instrument climbs should
normally be performed at 143 KCAS (140 KIAS). Refer
to chapter 7 for information regarding fuel consumption
and rate-of-climb.
When safe altitude and climb
airspeed are attained, complete the after takeoff checks
as prescribed in this chapter.
Climb or cruise at 10 to 15 KIAS
above normal speeds when icing
conditions exist. Reducing the angle
of attack minimizes the accumulation
of ice on all surfaces.
c. Instrument Cruise. There are no unusual flight
characteristics during cruise in instrument meteorological
conditions.
d. Speed Range. Stability and flight characteristics
are normal through the full speed range during
instrument flight operations.
Power settings during
instrument flight should be conducted in accordance with
the power charts in chapter 7.
e. Communication and Navigation Equipment. No
special technique due to aircraft configuration is
required.
f. Instrument Descent. When a descent at slower
than cruise speed is desired, slow the aircraft to the
desired speed before initiating the descent. Normal
descent or radar controlled descent to traffic pattern
altitude can be made using cruise air speed. Normally,
descent will be made with the aircraft in a clean
configuration, maintaining cruise speed by reducing
power as required.
The aircraft is completely
controllable in a high rate of descent with the landing
gear and wing flaps extended. A high rate of descent
may be obtained using the drag on the propeller in HIGH
RPM.
g. Holding. Recommended holding airspeed is 143
KCAS (140 KIAS). Maintain cruise RPM and reduce
power as required. For extended holding patterns,
consult the appropriate fuel and power chart in chapter
7. For descents in the holding pattern, decrease power
and maintain the holding pattern airspeed.
h. Instrument Approaches.
Various procedures
can be used for instrument approaches.
Whether
letdowns are made in clean or in landing
configuration are at the discretion of the pilot.
Change 8 8-21
TM 55-1510-215-10
Turbulence or action to be accomplished when the
aircraft reaches the lower altitude will govern the type of
letdown. When a low ceiling exists, do not extend flaps
to full down (100%) until runway is in sight and landing is
assured. If it is necessary to make an emergency
instrument descent, the rate during the last 1000 feet
prior to landing should be adjusted to less than 500 feet
per minute.
8-36. Night Flying.
NOTE
Autoignition shall be used during
night operations above 14,000 feet.
Night flying is very closely related to instrument
flying and may often be conducted almost entirely under
instrument conditions. Before takeoff, it is imperative to
ensure that all lights, instruments, and avionic equipment
are functioning properly.
Generally, interior lighting
should be kept to the minimum amount which will still
allow complete visibility of all instruments and gages.
Excessive cockpit lighting decreases outside visibility.
Avoid using landing lights when in thick haze, smoke, or
fog, as reflected light will reduce visibility and may affect
depth perception. During ground operations, the aircraft
should be taxied slowly. Once rolling it is difficult to
judge actual ground speed and excessive speeds may
be developed without realizing it.
Section IV. FLIGHT CHARACTERISTICS
8-37. Normal Flight Characteristics.
The flight characteristics of the aircraft are normal
throughout its speed range.
to prevent yaw will also prevent a tendency to roll. A
pitching tendency will develop if the aircraft is held in the
stall, resulting in the nose dropping sharply, then pitching
to above the horizon; this cycle is repeated until recovery
is made.
8-38. Stalls.
Ample stall warning in the form of buffeting or
mechanical stall warning, is usually given when a stall
approaches.
This buffet warning will occur at
approximately 5 to 10 knots above stall speed with the
aircraft in clean configuration. With full wing flaps, the
buffet will occur almost simultaneously with the stall. If a
forward loading condition exists, minimum control speed
may be encountered rather than full stall. If correct stall
recovery technique is used, very little altitude will be lost
during the stall recovery. The terms "power on" and
"power-off" will be employed during discussion of
specific stall situations. For the purpose of this section,
the term "power-on" shall mean that both engines and
propellers of the aircraft are operating normally, and are
responsive to pilot control. The term "power-off" shall
mean that both engines are operating at idle power.
a. Power-On Stalls.
(1) The power-on stall attitude is very steep.
A light buffet precedes stalls, and the first indication of
approaching stall is a decrease in control effectiveness,
accompanied by a continuous tone from the stall
warning horn (refer to Stall Warning System, chapter 2).
The stall itself is characterized by a rolling tendency if
the aircraft is allowed to yaw. The proper use of rudder
8-22
(2) 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. In
normal power-on stalls, a steep climb attitude is
required. Unless the attitude is maintained, the aircraft
will generally "buffet" instead of stall. 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.
The roll can be
prevented by proper rudder pressure. Power-on stall
characteristics are not greatly affected by landing gear
and wing flap position, except that stalling speed is
reduced in proportion to the degree of wing flap
extension.
b. Power-Off Stalls.
The roll tendency is
considerably less pronounced in power-off stalls (in
either takeoff, landing, or clean configuration), and is
more easily prevented or corrected by proper use of
rudder and aileron. The nose will generally drop straight
through, with some tendency to pitch up again if
recovery is not made immediately. With wing flaps
down, there is little or no roll tendency and stalling speed
is 7 to 8 knots slower than with wing flaps up. Figure 8-4
shows the indicated power-off stall speeds with landing
gear and wing flaps UP or DOWN.
TM 55-1510-215-10
STALL SPEED
POWER OFF
STALL SPEED
U-21G
T74-CP-700
Figure 8-4. Stall Speed
8-23
TM 55-1510-215-10
c. Accelerated Stalls.
Accelerated stalls are
caused by increasing the aircraft's weight due to
centrifugal force in a turn or an abrupt pull-out from a
dive. The stall speed is increased by the square root of
the G factor multiplied by the normal stalling speed. 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.
8-39. Spins.
CAUTION
Do not pull out of the resulting dive
too abruptly as this could cause
excessive wing loads and a possible
secondary stall.
Intentional spins are not permitted. If a spin is
inadvertently entered, use the following recovery
procedures accomplishing steps 1 through 3 as near
simultaneously as possible.
1. Power levers  IDLE.
2. Apply full rudder, opposite the direction of
spin rotation.
NOTE
Spin tests have not been conducted
on this aircraft.
The recovery
technique is based on the best
available information.
8-40. Diving.
Maximum diving airspeed (red line) is 208 KCAS
(208 KIAS, Vmo) as shown in the Flight Envelope Chart,
chapter 5.
Flight characteristics are 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.
8-41. Maneuvering Flight.
The maximum speed at which abrupt full control
travel can be applied without exceeding the design load
factor of the aircraft is 169 KCAS (168 KIAS). The
Flight Envelope Chart as shown in chapter 5 is a plot of
acceleration versus speed. This chart is prepared for
maximum gross weight and shows the speeds at which
maneuvers are restricted and unrestricted, as related to
load limit factors.
8-42. Flight Controls.
3. Push control wheel full forward and
neutralize ailerons.
The aircraft is stable under all normal flight
conditions.
4. Maintain control positions in Steps 1
through 3 until rotation stops then neutralize all controls
and execute a smooth pullout.
8-43. Level Flight Characteristics.
8-24 Change 8
All flight characteristics are conventional throughout
the level flight speed range.
TM 55-1510-215-10
Section V. ADVERSE ENVIRONMENTAL CONDITIONS
8-44. Introduction
The purpose of this section 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 section is primarily narrative;
only those check lists that cover specific procedures
characteristic of weather operations are included. The
checklist in Section II provides for adverse
environmental operation.
8-45. Cold Weather Operation.
Operating difficulties may be encountered during
extremely cold weather, unless proper steps are taken
prior to or immediately after flight. All official 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. Refer to chapter 2, Section XII.
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. Check tires for
proper inflation and assure that they are not frozen to
wheel chocks or to the ground. Use ground heaters to
free frozen tires. When heat is applied to release tires,
the temperature should not exceed 160°F (71°C). In
extreme emergencies, tires may be inflated to 1.5 times
the normal pressure to break adhesions to ice, then
restored to normal pressure. Use external power source
for starting engines.
Pre-heating of engine oil is
unnecessary.
b. Engine Starting. Observe engine starter time
limits. If there is no rise in engine oil pressure after 30
seconds or pressure drops below minimum (40 PSI),
shut down the engine.
c. Warm-Up and Ground Test.
Warm-up
procedures and ground test will follow those outlined in
Section II.
d. Taxiing Instructions. Avoid taxiing through slush
and water if possible. Water and slush splashed on the
wings, antennas and empennage may freeze, increasing
weight and drag and possibly limit control surface
movement.
Deleted.
e. Before Takeoff.
Turn on the propeller
electrothermal deicer system and all anti-icing systems,
just prior to takeoff. Also accomplish the BEFORE
TAKEOFF steps in Section II.
f. Takeoff. Takeoff procedures for cold weather
operations are the same as for normal takeoff. Follow
the procedures in Section II. Deep snow on the runway
may cause enough drag to prevent takeoff. Light snow
or ice will decrease the traction of the tires. Use of
brakes and nose wheel steering may be ineffective.
g. During Flight. It may be advisable to cycle the
landing gear a few times after takeoff to dislodge ice
accumulated from spray of slush or water on runway.
Trim tabs and controls should also be exercised
periodically to prevent freezing. If visible moisture is
inadvertently encountered, antiicing systems should be
activated.
h. Descent. Use normal procedures in Section II
for descent.
NOTE
If icing conditions are encountered
during approach, slightly higherthan-normal airspeed should be
maintained in order to compensate
for the additional increased drag,
additional weight, and higher than
normal stall speeds.
i. Landing.
normal landing.
Use procedures in Section II for
NOTE
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.
j. Engine Shutdown. Follow the normal engine
shutdown procedures in Section II.
Change 7 8-25
TM 55-1510-215-10
k. Before Leaving the Aircraft. Park the aircraft in
a warm area if possible. Should this be impossible, after
wheel chocks are in place release the brakes to prevent
freezing and lock the control surfaces. Condensation
will be minimized if aircraft fuel tanks are filled. Remove
any accumulation of dirt and ice from the landing gear
shock struts. Install protective covers to guard against
possible collection of snow and ice.
g. Before Leaving the Aircraft. Install all protective
covers to keep out blowing sand and dust. If blowing
sand and dust do not present a hazard, leave the cockpit
vent/storm windows open to provide ventilation. Take
extreme care to prevent sand or dust from entering the
fuel and oil systems during servicing.
8-47. Hot Weather Operation.
8-46. Desert 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 increase
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.
a. Preparation for Flight. Check the position of the
aircraft in relation to other aircraft. Propeller blown sand
blast can damage other aircraft. Wipe the landing gear
shock struts free of dust and sand with a clean cloth.
Remove all aircraft protective covers. Check instrument
panel and general interior for dust and sand
accumulation. Open main entrance door and cockpit
vent/storm windows to ventilate the aircraft. Operate all
moveable control surfaces.
b. Engine Starting. Perform the normal procedures
in Section II.
c. Warm-Up Ground Tests. Perform the normal
procedures in Section II. To minimize possibility of
damage to the engine during desert operation use
inertial separators (activate ICE VANE).
d. Taxiing. When practical avoid taxiing over
sandy terrain. Propeller and engine deterioration may
result from the impingement of sand and gravel. Use
minimum wheel braking action since brake cooling is
retarded at high ambient temperatures.
e. Takeoff. No special technique or procedures
are required during takeoff.
f. Landing. No special technique or procedures
are required during landing. (Refer to Section II for
normal engine shutdown procedures.)
8-26
CAUTION
A limitation based on pressure
altitude and ambient temperature
prohibits aircraft takeoff under
certain high ambient temperature
conditions.
(Takeoff Temperature
Limitation Chart in chapter 5).
During hot weather operations, the principle
difficulties encountered are high interstage turbine
temperature (ITT) during engine starting and takeoff,
over-heating of brakes, and longer takeoff and landing
rolls. 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.
a. Preparation for Flight. No special technique or
procedures are required for flight preparation.
b. Engine Starting.
Engine starting under
conditions of high ambient temperatures may produce a
higher than normal ITT during the start. The ITT 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 and Ground Tests. Perform normal
procedure in Section II.
d. Taxiing. Use the wheel brakes as little as
possible since cooling is retarded at high ambient
temperatures.
e. Takeoff. No special technique or procedures
are required during takeoff.
TM 55-1510-215-10
f. During Flight.
No special technique
procedures are required during normal operations.
or
g. Descent. No special technique or procedures
are required during descent.
h. Landing. No special technique or procedures
are required during landing.
constant power settings are vital to proper flight
technique. Establish recommended penetration speed
and proper attitude prior to entering turbulent air.
Maintaining a pre-established attitude will result in a
fairly constant airspeed.
Complete the following
procedures when there is a possibility of encountering
turbulent air due to thunderstorm activity.
1. Establish penetration airspeed.
i. Engine Shutdown.
Refer to Section II for
ENGINE SHUTDOWN procedure.
2. Secure loose equipment.
3. No smoking.
CAUTION
If fuel tanks are completely filled, fuel
expansion may cause overflow,
thereby creating a fire hazard.
j. Before Leaving the Aircraft. Install wheel chocks
and release the wheel brakes to prevent warpage of the
brake discs.
Leave windows and doors open as
necessary for ventilation.
4. Tighten
harnesses.
seat
belts
and
shoulder
5. Turn cockpit and cabin lights on to
minimize the blinding effect of lightning.
6. Turn on pitot heat and stall warning heat
when visible moisture is encountered.
7. Check
flight
instruments
for
proper
indication.
8-48. Turbulent Air and Thunderstorms.
8. Advise crew members of approach to
storm.
WARNING
Due to the comparatively light wing
loading, control in severe turbulence
and moderately heavy thunderstorms
is critical. Since turbulence imposes
heavy loads on the aircraft structure
make all necessary changes in
aircraft attitude with the minimum
amount of control pressures to avoid
excessive loads on the aircraft's
structure.
Plan the flight to avoid areas of severe turbulence
and thunderstorms. During night or instrument flight
conditions it is not always possible to detect individual
storm areas or find the in-between clear areas. If such
turbulence is to be penetrated, it will be necessary to
counter rapid changes in attitude and accept major
indicated altitude variations.
9. Maintain constant power settings and
pitch attitude regardless of airspeed or altitude
indications. Concentrate of maintaining a level attitude.
NOTE
If possible, adjust to the desired
power setting prior to entering a
storm area. Generally, this power
setting will be lower than that used
for cruise operations.
10. Maintain original heading. Make no turns
unless absolutely necessary.
11. Do not chase the airspeed indicator; doing
so will only result in extreme variations of aircraft attitude
and the possibility of exceeding safe airspeed.
8-49. Penetration.
12. Apply as little pressure as possible on the
control surfaces to prevent excessive strains on the
aircraft structure.
The maximum safe penetration speed in severe
turbulence is 169 KCAS (168 KIAS). Pitch attitude and
13. Erratic
expected.
altimeter
indications
may
be
Change 8 8-27
TM 55-1510-215-10
8-50. Ice and Rain.
WARNING
Autoignition shall be used before
entering anticipated or actual icing
conditions (10°C or below in visible
moisture).
WARNING
Extend the engine ice vanes when
temperature is 5°°C or below before
entering visible moisture. If the ice
formation is allowed to progress to a
critical point, the loss of intake air
may make it impossible for the
engine to run at normal power.
a. If flights into known icing conditions are made,
all anti-icing and deicing systems of the aircraft should
be properly utilized in accordance with the procedures
detailed in chapter 2. Rain presents no particular
problems other than restricted visibility.
b. The aircraft is equipped with anti-icing or deicing
systems to prevent the formation of ice on the pitot tube,
stall warning, fuel vents, heater air intake, engine air
intake, and propeller blades. Deicer boots are provided
to remove ice from the wing and tail leading edges, and
an electrothermal windshield anti-ice system is provided
to prevent ice from forming on the windshield.
Windshield defrosters are installed to alleviate conditions
resulting from frost or light ice. Windshield wipers are
installed for rain at lower airspeeds. An engine ice vane
is provided for removal of ice and rain from the engine
intake air. An autoignition system automatically provides
combustion re-ignition if flameout should occur due to
icing conditions, etc. If severe icing conditions are
encountered ascend or descend to altitudes where these
conditions do not prevail. Operation of deice and antiice equipment discussed here is described in chapter 2.
(1) Icing.
Icing occurs because of
supercooled water vapor such as fog, clouds, or rain.
The most severe formation will generally occur at
approximately -5°C.
8-28
(2) Taxiing. Extreme care must be exercised
when taxiing on ice or slippery runways. Excessive use
of either brakes or power may result in an uncontrollable
skid.
(3) 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.
(4) Climb. Climb at 10 to 15 KIAS above
normal climb speeds when icing conditions exist.
Reducing the angle of attack minimizes the
accumulation of ice on all surfaces.
(5) Cruising flight. Prevention of ice formation
is more effective and satisfactory than attempts to
dislodge the ice after it has formed. If icing conditions
are encountered on the anti-icing systems prior to the
first sign of ice formation. Refer to chapter 2 for use of
deicer boots. 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. At any time visible moisture
is present, turn on the pitot and stall warning heat. Refer
to chapter 2 for additional information.
NOTE
Do not operate deicer boots
continuously. Continuous operation
tends to balloon the ice over the
boots. Allow at least 1/2 inch of ice
to accumulate on the boots, then
activate the deicer boots to remove
the ice. Repeat this procedure as
required.
(6) 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.
NOTE
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 55-1510-215-10
Section VI. CREW DUTIES
8-51. Crew/Passenger Briefing.
(3) Seat belts.
The following is a guide that should be used in
accomplishing required crew/passenger briefings, when
a unit crew/passenger briefing is not available. Items
that do not pertain to a specific mission may be omitted.
(4) Movement in aircraft.
(5) Internal communications.
(6) Security of equipment.
a. Crew Introduction.
(7) Smoking.
b. Equipment.
(8) Oxygen.
(1) Personal to include ID tags.
(9) Refueling.
(2) Professional.
(10) Weapons.
(3) Survival.
(11) Protective masks.
c. Flight Data.
(12) Parachutes.
(1) Route.
e. Emergency Procedures.
(2) Altitude.
(1) Emergency exits.
(3) Time en route.
(2) Emergency equipment.
(4) Weather.
(3) Emergency landing/ditching procedures.
d. Normal Procedures.
(4) Bail out.
(1) Entry and exit of aircraft.
(2) Seating.
8-29/(8-30 blank)
TM 55-1510-215-10
CHAPTER 9
EMERGENCY PROCEDURES
Section I. AIRCRAFT SYSTEMS
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 contained in
the
condensed
checklist
TM
55-1510-215-CL.
Emergency operations of avionics equipment are
covered when appropriate in chapter 3, Avionics.
9-2. Immediate Action Emergency Checks.
Those checks that must be performed immediately
in an emergency procedure are underlined.
These
immediate action emergency checks must be committed
to memory.
NOTE
The urgency of certain emergencies
requires immediate and instinctive
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 LAND 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 to 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.
9-4. Emergency Exits and Equipment
The main cabin entrance door is used for normal or
emergency exit. A removable window (cabin emergency
exit hatch) on the right side of the fuselage (fig. 9-1), and
the cockpit emergency entrance/exit hatch are used for
emergency exit only. The main cabin entrance door, the
cabin emergency exit hatch (removable window), and
the cockpit emergency entrance/exit hatch provide
emergency escape routes from the aircraft either on the
ground or when ditching.
9-5. Emergency Entrance.
In the event that entry to the aircraft through the
main entrance door is impossible or impractical, entry
may be made through the alternate emergency exits
which are the cabin window emergency exit hatch (fig.
9-2), or the cockpit emergency entrance/exit hatch. In
order to open the cabin window emergency exit hatch,
either the right rear window must be broken through or
the fuselage cut open along the dotted line stenciled
CUT HERE FOR EMERGENCY RESCUE. The cockpit
emergency entrance/exit hatch (fig. 9-1) may be
removed either externally or from inside the cockpit by
actuating either the external or the inside latch handle.
9-6. Engine.
a. Flight Characteristics Under Partial Power
Conditions. Single-engine operations, due to an engine
malfunction, are sometimes preceded by symptoms
which will enable you to take preventative action.
Complete engine failure most often occurs due to fuel
starvation. This type of emergency situation is seldom
due to mechanical causes, however, failure due to
carelessness or improper operating techniques is not
uncommon. This may be avoided by constant attention
to engine torque, turbine temperatures, and fuel and oil
pressure and temperatures. These operating limitations
are established and discussed in chapter 5. If an engine
malfunction is indicated, land as soon as practical. It is
essential to have a thorough understanding of the
principles affecting single-engine performance and the
limitations resulting from the unbalance of power. Two
principal factors govern flight safety on one engine:
AIRSPEED and POWER; directional control being in
proportion to airspeed. There are no unusual flight
9-1
TM 55-1510-215-10
CABIN EMERGENCY EXIT HATCH JETTISON PROCEDURE
(Figure 9-1 sheet 1 of 2)
(Figure 9-1 sheet 2 of 2)
Figure 9-1. Emergency Exits and Equipment (sheet 1 of 2)
9-2 Change 7
TM 55-1510-215-10
COCKPIT EMERGENCY ENTRANCE/EXIT HATCH-INTERIOR JETTISON PROCEDURE
AV 094952.2
Figure 9-1. Emergency Exits and Equipment (sheet 2 of 2)
9-3
TM 55-1510-215-10
*TO
REMOVE
THE
COCKPIT
EMERGENCY
ENTRANCE/EXIT
HATCH FROM OUTSIDE, TURN
PROTRUDING EXTERNAL LATCH
HANDLE AND LIFT AWAY RELEASED
DOOR.
(1) Engine malfunction during takeoff run
(abort). Below takeoff airspeed (Vlof), use the
following procedure:
NOTE
Single-engine reversing may be applied if
required. Use extreme caution if takeoff
surface is not hard and dry.
1. POWER levers - IDLE.
2. Braking - As required.
NOTE
Braking action is impaired if wheels are
allowed to skid.
NOTE
If insufficient runway remains
stopping perform steps 3 through 5.
NOTE
** EMERGENCY ENTRANCE MAY
ALSO BE MADE BY BREAKING IN
THE WINDOW ON THE CABIN
EMERGENCY EXIT HATCH (FIRST
LARGE SQUARE WINDOW SECTION
FORWARD OF THE CUT HERE FOR
EMERGENCY RESCUE AREA) AND
ACTUATING THE HATCH RELEASE
MECHANISM.
AV 095259
Figure 9-2. Emergency Entrance
characteristics during single-engine operation as
long as airspeed is maintained at or above minimum
control speed (Vmc) and stall speed (Vs). Minimum
control speed is the minimum airspeed for which
lateral and/or directional control can be maintained
with the flaps retracted, dead engine propeller
feathered, live engine set to takeoff power and 5°
maximum bank angle maintained towards the live
engine (landing gear UP or DN). 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 retracting the landing gear (if extended) and
establishing the appropriate single-engine best rateof-climb speed. Refer to chapter 7 for single-engine
climb performance.
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 takeoff speed has been attained. Refer to
chapter 7 for takeoff performance.
9-4 Change 7
3. CONDITION
for
levers - FUEL
CUTOFF.
4. Firewall
Shutoff
valves
-
CLOSED.
5. MASTER SWITCH - DOWN.
(2) Engine Malfunction After Liftoff.
Engine Malfunction After Liftoff. If an engine fails
after becoming airborne, maintain single-engine best
rate-of-climb speed (Vyse) or, if airspeed is below
(Vyse), maintain whatever airspeed is attained
between liftoff (Vlof) and (Vyse) until sufficient
altitude is attained to trade altitude for airspeed and
accelerate to (Vyse).
(3) Engine malfunction after liftoff (abort).
Perform the following and land in a wings-level
attitude:
1. POWER levers - REDUCE.
2. Gear - DOWN.
3. Complete normal landing.
NOTE
If able to land on remaining runway
check gear down, use brakes and reverse
thrust as required.
(4) Engine malfunction after takeoff. If an
engine malfunctions, after becoming airborne, and
best single-engine climb speed has been attained,
proceed as follows:
TM 55-1510-215-10
NOTE
Do not retard the malfunctioning
engine power lever, or turn the
autofeather
system
OFF,
until
propeller rotation is completely
stopped. To do so will deactivate the
autofeather circuit and prevent
automatic feather.
5. Propeller
FEATHER.
3. Flaps - UP.
4. Engine clean up - Perform.
c. Engine Malfunction During Flight. Refer to
chapter 7 for single-engine cruise information. If one
engine malfunctions during flight, maintain directional
control and proceed as follows:
NOTE
If the situation permits, analyze the
malfunction and take appropriate
action to restore power.
1. Autopilot/yaw damp  Disengage.
2. Power  As required.
NOTE
Deleted.
3. Dead engine  Identify.
4. Power lever (dead engine)  IDLE.

7. Gear  UP.
8. Flaps  UP.
9. Power  Set (for single-engine
cruise).
10. Engine clean up  Perform.
d. Engine Clean Up. Clean up procedure to
be used after engine malfunction or an
unsuccessful restart is as follows:
1. Power - Maximum allowable.
2. Gear- UP.
engine)
6. Condition lever (dead engine) 
FUEL CUTOFF.
NOTE
If the autofeather system fails to
operate, identify the dead engine,
retard the power lever of suspected
engine to confirm identification, then
feather the propeller for the dead
engine.
(dead
1. Auxiliary fuel pump (dead engine)
 OFF.
2. Crossfeed  CLOSED (if no restart
is to be attempted).
3. Fuel firewall valve (dead engine) 
CLOSED (if no restart is to be attempted).
4. Generator (dead engine)  OFF.
5. Electrical load  Monitor.
6. Autoignition (dead engine)  OFF.
7. Fuel control heat (dead engine) 
OFF, (if no restart is to be attempted).
e. Engine
Malfunction
During
Final
Approach. If an engine malfunctions during final
approach (after LANDING CHECK), continue
approach using the following procedure:
1. Power  As required.
2. Gear  DN.
NOTE
100% flaps must be used to obtain the
minimum landing distance in chapter 7.
Change 7 9-5
TM 55-1510-215-10
f. Engine Restart During Flight (Using Starter).
Successful restarts may be achieved at all altitudes
normally flown. Do not attempt to restart the engine
unless it can be determined that no damage to the
engine will result and no additional hazard will be
created. If in doubt, continue single-engine flight and
land the aircraft as soon as practicable.
11. Condition lever - LO IDLE.
12. ITT and N1 - Monitor (1090°C maximum).
NOTE
When N1 is below 13%, starting
temperatures tend to be higher than
normal. To preclude overtemperature
(1090°C or above) during engine
acceleration to idle speed, momentarily
move the condition lever into FUEL
CUTOFF position as necessary.
CAUTION
Do not exceed starter limitation of 40
seconds on, and 60 seconds OFF for two
starter operations, then 40 seconds on,
30 minutes OFF.
NOTE
The avionics master switch and radar
should not normally be turned on, until
the generators are on. This will help
protect the solid-state circuitry.
13. Ignition/start switch - OFF (when N1 is above
50%, or start attempt is discontinued).
14. Engine clean up - Perform (if restart is
unsuccessful).
15. Oil pressure - Check (40 PSI minimum).
16. Generator - RESET, then ON.
1. Electrical load - Reduce to minimum.
17. Propeller - Synchronize.
2. Firewall shutoff valve - OPEN.
3. Power lever (dead engine) - IDLE.
4. Propeller (dead engine) - FEATHER.
5.Condition lever (dead engine) - FUEL
CUTOFF.
6. Auxiliary fuel pumps (2) - ON.
7. Crossfeed - OPEN.
8. Fuel control heat - ON.
9. Ignition/start switch - On. (Check IGN ON light
illuminated, N1 over 13% and stabilized for
approximately 5 seconds).
10. ITT (live engine) - Monitor (750°C maximum).
18. Power - As required (after engine is stabilized
at idle).
19. Electrical equipment - As required.
20. Auxiliary fuel pumps (2) - As required.
21. Crossfeed - As required.
g. Engine Restart During Flight (No Starter Assist,
Engine and Propeller Windmilling). A restart without actuating
the starter may be accomplished. The reason for engine
malfunction should be determined before attempting a restart.
If altitude permits, diving the aircraft will increase engine RPM
and assist in restart.
NOTE
The avionics master switch and radar
should not normally be turned on, until
the generators are on. This will help
protect the solid-state circuitry.
CAUTION
If ignition is not evident in 10 seconds
after moving the condition lever to LO
IDLE, position the condition lever to
FUEL CUTOFF and the IGNITION/START
switch to OFF.
Accomplish ENGINE
CLEARING procedure (chapter 8).
If
another restart attempt is to be made,
repeat the entire ENGINE RESTART
procedure.
9-6 Change 7
1.
2.
3.
4.
5.
Electrical load - Reduce to minimum.
Firewall shutoff valve - OPEN.
Power Lever (dead engine) - IDLE.
Propeller (dead engine) - HIGH RPM.
Condition lever (dead engine) - FUEL
CUTOFF.
6. Auxiliary fuel pumps (2) - ON.
TM 55-1510-215-10
7. Crossfeed - OPEN.
20. Autoignition - As required.
8. Generator (dead engine) - OFF.
21. Electrical equipment - As required.
9. Fuel control heat - ON.
22. Auxiliary fuel pumps (2) - As required.
10. Airspeed
(minimum).
-
143
KCAS
(140
KIAS)
23. Crossfeed - As required.
11. Altitude - Below 20,000 feet.
12. Autoignition - ARM.
CAUTION
If ignition is not evident in 10
seconds after moving the condition
lever to LO IDLE, position the
condition lever to FUEL CUTOFF and
the autoignition switch to OFF. If
another restart attempt is to be made,
repeat the entire ENGINE RESTART
procedure.
13. Condition lever - LO IDLE.
NOTE
When N1 is below 13%, starting
temperatures tend to be higher than
normal.
To
preclude
overtemperature (1090°C or above)
during engine acceleration to idle
speed,
momentarily
move
the
condition lever into FUEL CUTOFF
position as necessary.
14. ITT and N1 - Monitor (1090°C maximum).
15. Engine clean up - Perform if restart is
unsuccessful.
16. Oil pressure - Check (40 PSI minimum).
17. Generator - RESET, then ON (when N1 is
above 50%)
18. Propeller - Synchronize.
h. Deleted.
i. 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 appropriate airspeed per the Maximum
Glide Distance Chart (fig. 9-3), with the gear and flaps
up. Turn off all electrical equipment to conserve battery
power for extending the gear and flaps, but leave the
master switch ON.
Refer to the Maximum Glide
Distance Chart (fig. 9-3).
j. Landing With One or More Engines Inoperative.
NOTE
With one engine inoperative level
flight cannot be assured with full
flaps and the landing gear down.
(1) Single-engine landing. Fly a normal
pattern and perform Single Engine Before Landing check
as appropriate. Extend flaps beyond the approach
position only if required, and then only after there is no
possibility of a go-around. Plan for a slightly higher than
normal approach, allowing for sufficient straight-away on
final, so minor alignment, speed, and altitude corrections
can be accomplished without excessive low altitude
maneuvering. Do not extend full-flaps until gear is down
and locked. Make a normal touchdown, reducing power
during flareout. Avoid excessive or abrupt changes in
power. A feathered propeller will result in less drag and
may cause the aircraft to "float" during landing. After
touchdown, apply brakes and propeller reversing as
required.
NOTE
Single-engine reversing may be
applied if required.
Use extreme
caution if landing surface is not hard
and dry.
19. Power - As required.
Change 8 9-7
TM 55-1510-215-10
MAXIMUM GLIDE DISTANCE
GLIDE DISTANCE
U-21G
T74-CP-700
ASSOCIATED CONDITIONS:
1.
2.
3.
4.
POWER OFF (PROPELLERS FEATHERED)
GEAR AND FLAPS UP
ZERO WIND
AIRSPEED AND WEIGHT COMBINATIONS
SHOWN
PRODUCE
MAXIMUM
GLIDE
DISTANCE
WEIGHT - LBS.
9650
9000
8000
7000
6000
Figure 9-3. Maximum Glide Distance
9-8
BEST GLIDE SPEED
KIAS
112
108
102
94
87
TM 55-1510-215-10
(2) Forced landing - no power.
If
sufficient altitude remains after reaching a suitable
landing area, a circular pattern will provide best
observation of surface conditions, wind speed and
direction. When the condition of the terrain has been
noted and the landing area selected, set up a
rectangular pattern and extend the gear when landing is
assured. Fly the base leg as necessary to control point
of touchdown. With both propellers feathered, the
normal tendency is to overshoot due to less drag. In the
event a positive gear down and locked indication cannot
be determined, prepare for a gear up landing. Refer to
LANDING EMERGENCIES (para. 9-1).
(3) Single-engine descent-arrival check.
Perform the following checks prior to traffic pattern entry:
1. Seat
belts
and
shoulder
harnesses - Secure (crew/passengers checked).
2. Fuel panel - Check.
3. Parking brake handle - In.
4. Inlet air separator - As required.
5. Engine ice vanes - As required.
(4) Single-engine before landing check.
Perform the following checks upon pattern entry:
1. Auxiliary fuel pump (live engine)
- ON.
2. Flaps - APPROACH below 174
KCAS (173 KIAS).
NOTE
To prevent zeroizing the Mode 4
function of the transponder, when the
landing gear is down (struts
compressed) and either transponder
or aircraft power is to be turned off,
place the CODE switch momentarily
in HOLD position.
3. Gear - DN below 156 KCAS
(154 KIAS).
NOTE
100% flaps must be used to obtain
the minimum landing distances in
chapter 7.
1. Gear - Recheck DN (check
lights.
2. Propeller
(live
engine)
-
HIGH RPM.
WARNING
Once the flaps are full DOWN on a
single-engine landing, do not attempt
a go-around.
k. Single Engine Go-Around. The decision to
go around must be made as early as possible.
Elevator forces at the start of a go-around are very
high and a considerable amount of rudder control
will be required at low airspeeds.
Retrim as
required. If rudder application is insufficient or
applied too slowly, directional control cannot be
maintained. If control difficulties are experienced,
reduce power on the operating engine immediately.
Ensure that the airplane will not touch the ground
before retracting the landing gear. Retract the flaps
only as safe airspeed permits (APPROACH until
Vref used for APPROACH, then UP). Perform single
go-around as follows:
NOTE
Once flaps are fully extended, a
single engine go-around may not be
possible when close to ground under
conditions of high gross weights
and/or high density altitude.
1. Power
-
Maximum
allowable.
2. Gear - UP.
3. Flaps - UP.
4. Power - As required.
4. Landing lights - ON.
(5) Single-engine landing check.
5. Landing/Taxi lights - OFF.
6. Deleted.
l. Chip Detector Warning Light On.
The
illumination of either chip detector light on the
annunciator panel will require the following
immediate procedures:
Change 8 9-9
TM 55-1510-215-10
1. Engine instruments - Monitor.
2. Land as soon as practical.
(1) If propeller RPM increases and
engine torque decreases - Secure engine as soon as
practical.
9-7. Propeller.
WARNING
a. Propeller Failure. If an overspeed condition
occurs that cannot be controlled with the propeller
lever, or by reducing power, feather the propeller.
If the propeller feathers by itself complete the
feathering procedure.
The engine with the
disabled propeller may be operated to provide
electrical power for systems needed in flight. The
procedure to follow in the event of propeller failure
is as follows:
1. Power (failed propeller) - IDLE.
2. Propeller (failed propeller) - FEATHER.
While operating with the PROP GOV
IDLE STOP circuit breaker pulled, the
secondary low pitch stop system is
inoperative.
Should the primary
hydraulic low pitch stop fail, the
propeller will reverse when power is
reduced.
(2) If propeller RPM decreases and
engine torque increases - Pull PROP GOV IDLE STOP
circuit breaker immediately (on right subpanel).
3. Condition lever - As required.
NOTE
4. Engine cleanup - As required.
Propeller RPM decrease and torque
increase indicates malfunction of the
secondary low pitch stop system
which will cause the propeller to
feather within 10 to 60 seconds. In
this event, PULL the PROP GOV IDLE
STOP circuit breaker and continue
flight to destination. Malfunction of
the secondary low pitch stop must be
corrected before further flight.
b. Primary Pitch Light On. The procedure to
follow in the event of illumination of a PRI PITCH
light is as follows:
1. Propeller RPM and engine torque Monitor.
2. The action to be taken depends on torque
and propeller speed:
(3) If propeller RPM and torque remain
stable, reset the PROP GOV IDLE STOP circuit breaker.
WARNING
9-8. Fire.
Propeller RPM increase and torque
decrease when a light illuminates
indicates a malfunction in the
primary governor system and/or the
hydraulic low pitch stop.
The
overspeed governor will prevent RPM
from exceeding 2288. Do not pull the
PROP GOV IDLE STOP circuit
breaker, as this could cause propeller
reversal. Secure engine as soon as
practical.
9-10 Change 7
a. Engine Fire. The following procedures indicate
the action to be taken in case of engine fire.
NOTE
No engine fire extinguisher systems
are installed.
(1) Engine/nacelle fire during start or
ground operation. If a nacelle fire should develop during
start or ground operations, proceed as follows:
TM 55-1510-215-10
1. Firewall shutoff valves - CLOSED.
1. Fight the fire.
2. Master switch - Down.
2. Land immediately if fire continues.
3. Propellers - FEATHER.
WARNING
CAUTION
Land the aircraft immediately,
or bail out, depending on
circumstances
and
seriousness of the fire.
Deleted.
(2) Engine fire - during flight. In the event an
engine fire develops during flight, shut down the affected
engine in the following manner:
1. Firewall shutoff valve - CLOSED.
2. Power- IDLE.
3. Propeller - FEATHER.
4. Condition lever - FUEL CUT-OFF.
5. Auxiliary fuel pump - OFF.
d. Electrical Fire.
All circuits are
protected by circuit breakers, which tends to
prevent an electrical fire. Upon noting the
existence of an electrical fire, immediately
turn off all affected electrical circuits, if
known. If electrical fire source is unknown,
proceed as follows:
(O)
As required.
1. Crew oxygen masks -
(O)
2. Passenger
masks - As required.
6. Transfer pump - OFF.
oxygen
3. Master switch - Down.
7. Crossfeed - CLOSED.
4. All electrical switches - OFF
b. Fuselage Fire. If a fuselage fire occurs, perform
the following:
5. Battery - ON.
6. Generators - RESET, then
WARNING
The
extinguisher
agent
(Bromochlorodifluoromethane) in the
fire extinguisher can produce toxic
effects if inhaled.
ON.
7. Essential equipment ON (individually until fire source is isolated).
Change 7 9-11
TM 55-1510-215-10
NOTE
pressure 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.
With battery switch and generator
switches turned OFF, landing gear,
wing flaps, and auxiliary and transfer
pump circuits will be inoperative. If
deemed reasonably safe, turn on
battery or generator switch long
enough to extend the landing gear
and then turn it off again. Otherwise
extend the landing gear manually and
land with the wing flaps up.
b. Boost Pump Failure. If the L or R FUEL FAIL
light illuminates in flight proceed as follows:
1. Auxiliary fuel pump (affected engine) ON.
e. Smoke and Fume Elimination. To eliminate
smoke and fumes from the aircraft, perform the
following:
c. Fuel Leaks/Syphoning. The action to be taken
in the event of fuel leaks or syphoning depends upon the
origin and severity of fuel loss.
(O)1. Crew oxygen masks - On.
(O)2. Passenger masks - On. The copilot should
confirm that all passengers are receiving
supplemental oxygen.
3. Cockpit
required.
vent/storm
windows
2. Fuel fail light - Check extinguished.
-
Open
(1) If fuel filler cap syphoning occurs proceed
as follows:
as
1. Airspeed - 123 KCAS (120 KIAS).
2. Land as soon as practicable.
9-9. Fuel System.
(2) Deleted.
a. Fuel System Emergencies. The engine will
operate with an engine driven fuel boost pump
failure. However, failure of an engine-driven (high
pressure) fuel pump will result in engine fuel
starvation. Perform the SINGLE-ENGINE procedure
and land as soon as practical.
CAUTION
d. Fuel
System
Crossfeed
Single-Engine
Operation.
During single-engine operation the fuel
supply for the dead engine may be used to supply the
live engine by routing the fuel through the crossfeed
system. Use the following procedure for fuel crossfeed
during single-engine operation:
1. Fuel
firewall
valve
(dead
engine)
CLOSED.
2. Auxiliary fuel pump (dead engine) - ON.
Do not use fuel firewall valve to shut
down engine operation, except in an
emergency. The engine-driven high-
9-12 Change 8
3. Crossfeed - OPEN.
-
TM 55-1510-215-10
4. Fuel
crossfeed
light
-
Check
9-10. Electrical System.
illuminated.
NOTE
With a fuel firewall valve closed,
the respective FUEL FAIL light may
be illuminated, but the indication
will lack significance due to the
closed position of the firewall
valve.
WARNING
If battery overheats, do not open
battery compartment or attempt to
disconnect or remove battery.
Battery fluid will cause burns, and
overheated battery could cause
thermal burns and may explode.
5. Transfer pump (dead engine) ON.
6. Auxiliary fuel pump (live engine) Check OFF (side receiving crossfeed).
7. Crossfeed and fuel quantity Monitor.
e. Fuel Transfer - Manual. A no fuel transfer
condition is indicated by illumination of a yellow R
FUEL XFR or L FUEL XFR light on the annunciator
panel, and the simultaneous flashing of both MASTER
CAUTION lights. Manual refill of a nacelle tank is
conducted with the TRANSFER PUMP switch ON,
and is initiated by holding the respective TRANSFER
TEST switch in the selected position (L or R). The
switch may or may not be released after fuel transfer
has started.
If the TRANSFER TEST switch is
released, fuel transfer will be stopped by either 24 or
57-gallon float switch. If the TRANSFER TEST switch
is held ON, the transfer pump will continue to run
(overriding the float switches) until all available fuel is
transferred. The pump will then automatically shut off
and the appropriate FUEL XFR light will illuminate on
the annunciator panel, indicating that no more fuel is
available for transfer. However, if transferable fuel
remains in the wing tanks after the nacelle tank is
filled to capacity, continuing to hold the TRANSFER
TEST switch ON will only circulate fuel between the
nacelle tank and the wing tanks.
NOTE
With fuel remaining in the wing
tanks, illumination of the FUEL
XFR light, after manual activation
of the TRANSFER TEST switch,
indicates transfer pump failure.
CAUTION
Circuit breakers should not be
reset more than once until the
cause of circuit malfunction has
been determined and corrected.
a. Electrical Power System Failure. In the event
both generators and the battery are shut off, the flaps
will be inoperative, and the landing gear must be
extended manually. All systems (except the stall
warning horn) which utilize DC power for operation
(including the AC power system) will be inoperative.
CAUTION
Only one attempt should be made
to restore an inoperative generator
to use.
b. Generator Failure. Either generator is capable
of carrying the full load of normal aircraft electrical
equipment. If one generator becomes inoperative, all
non-essential electrical equipment should be used
judiciously to avoid overloading the remaining
generator. Loads in excess of single generator output
will drain the battery, with the resultant loss of reserve
and emergency power. The choice of equipment to be
used will be determined by existing conditions. The
use of accessories which create a very high drain,
should be avoided. If both generators are shut off due
either to generator failure, ground fault or engine
failure, all equipment should be turned off to preserve
battery power for extending the landing gear and wing
flaps. The general location of failure within the aircraft
electrical system may be indicated by illumination of
the respective GEN OUT annunciator panel red light,
and the simultaneous flashing of the MASTER
WARNING lights.
9-13
TM 55-1510-215-10
Illumination indicates one of three conditions, generator
failure, generator overvoltage, or a faulted generator
feeder cable. If the indicated generator cannot be
returned to the line, reduce load on the remaining
generator to a volt-loadmeter reading of 1.0 or less for
continuous operation.
c. Ground Fault - Generator Feeder Cable.
Ground fault condition in the generator to generator-bus
tie cable is indicated by flashing MASTER WARNING
lights, and an L or R GEN OUT light illuminated.
1. Ground fault circuit breaker - Reset
(one time).
2. Affected generator - RESET, then
ON.
(2) A
battery-generator
bus
overload
condition is indicated by divergent volt-loadmeter
readings and tripped BAT RELAY, BUS OVERLOAD,
and GND FAULT circuit breakers. Reset the BAT
RELAY circuit one time on the copilot's circuit breaker
panel (fig. 2-19).
(1) The following procedure will be used in
the event of a ground fault:
CAUTION
1. Generator - RESET, then ON.
2. Generator (GEN OUT light remains
illuminated) - OFF.
3. Electrical equipment - OFF
required to reduce generator load to 1.0 or less).
Do not reset the BUS OVERLOAD
circuit breaker more than one time. If
it trips after a reset, no attempt
should be made to reset the affected
generator using the GEN switch.
(as
(2) A ground fault condition in the generatorbus tie-cable is indicated by unbalanced volt-loadmeter
reading. Pull and reset GND FAULT circuit breaker on
the copilot's circuit breaker panel (fig. 2-19).
NOTE
In no case, will a single ground fault
on either generator cable, or the tie-in
cable between the generator buses,
cause the loss of more than one
power source.
d. Bus Overload - Generator Buses. A generator
bus overload condition is indicated by flashing MASTER
WARNING lights, L or R GEN OUT light illuminated,
GND FAULT circuit breaker tripped, zero reading on the
volt-loadmeter for the affected generator, and a high
reading on the operating generator.
(1) The following procedure will be used in
the event of a bus overload:
CAUTION
NOTE
In no event will a single bus overload
cause the loss of more than one
power source.
e. Both Gen Out Lights Illuminated.
1. Generators - RESET, then ON.
2. Generators (GEN OUT lights remain
illuminated) - OFF.
3. All nonessential electrical equipment
- OFF.
4. Land as soon as practicable.
f. Inverter Failure. Either inverter No. 1 or No. 2
is capable of supplying the full amount of normally
required single-phase AC power. Inverter failure is
indicated by the illumination of the respective yellow
annunciator light INV 1 or INV 2 along with the
simultaneous flashing of both MASTER CAUTION lights.
g. Inverter Light Illuminated.
1. Inverter - Select other inverter.
Do not reset the GND FAULT circuit
breaker more than one time. If it trips
after a reset, no attempt should be
made to reset the affected generator
using the GEN switch.
9-14 Change 7
NOTE
Deleted.
TM 55-1510-215-10
2. Inverter control circuit breakers Reset.
NOTE
Allow at least a 30 second cooling
period before resetting inverter
control circuit breaker.
3. Inverter lights remain illuminated Return to original inverter.
4. Inverter lights still remain illuminated
9-10.1.
Flight
Controls
Malfunction
(Unscheduled Electric Elevator Trim)
In the event of unscheduled electric elevator
trim, perform the following:
1. AP DISC/TRIM INTER switch
- Depress and hold.
2. ELEC TRIM circuit breaker Pulled.
9-11. Door Open Light Illuminated.
- Inverter off.
1. Do not attempt to close door.
2. Land as soon as practicable.
CAUTION
9-11.1. Split Flap Condition.
Deleted.
5. TACAN - OFF.
6. Land as soon as practicable.
h. Feeder Bus Failure. Equipment receiving power
from a failed feeder bus should not be returned to
service. The electrical load can be reduced by turning
OFF all non-essential electrical equipment.
i. Battery Monitor Light Illuminated. Illumination of
the battery light indicates excessive charge of battery.
This is a normal function after a battery start or after
lowering gear and flaps. However, illumination during
normal steadystate cruise flight indicates that conditions
exist that may cause a battery thermal runaway.
Proceed as follows:
1. Battery switch - OFF.
A split flap condition exists when the flaps are
in an asymmetrical configuration. Initial indication
of a split flap condition is a rolling motion during
flap extension or retraction.
Procedure for
recovery is as follows:
1. Aileron/rudder - As required.
2. Power - Asymmetric power as
required to maintain aircraft control.
3. Flaps - Extend/retract
symmetric configuration. if possible.
to
9-12. Emergency Descent.
Emergency descent is a maximum effort in
which damage to the aircraft must be considered
secondary to getting the aircraft down.
For
emergency descent, proceed as follows:
2. Loadmeter - Check.
CAUTION
NOTE
A DECREASE reading of 0.25 or more
graduation (approx.
one needle
width) verifies a deteriorating battery.
Leave battery switch OFF for
remainder of flight.
3. Battery condition good - Battery
switch ON.
4. Battery condition unsatisfactory Battery ON for flap and landing gear extension only.
5. Battery - OFF.
Do
not
practice
EMERGENCY
DESCENT procedure.
Damage to
landing gear door will result.
1.
2.
3.
4.
5.
KIAS) maximum.
Power-IDLE.
Propellers - HIGH RPM.
Gear-DN.
Flaps - APPROACH.
Airspeed - 156 KCAS (154
Change 8 9-14A/(9-14B blank)
TM 55-1510-215-10
9-13. Landing Emergencies.
a. Landing Gear System Failure. Should one or
more of the landing gear fail to extend, or fail to indicate
a safe condition, the following steps should be taken
before proceeding manually to extend the gear.
WARNING
The heater must be shut down prior
to all emergency landings to
minimize fire hazards.
Change 8 9-15
TM 55-1510-215-10
1. Gear control circuit breaker - Check.
NOTE
2. Gear
indicator
circuit
breaker
-
Check.
The landing gear cannot be retracted
manually.
CAUTION
If the LDG GEAR POWER circuit
breaker is tripped, allow approximately
three minutes cooling time before
resetting.
3. Gear power circuit breaker - Check.
4. Gear indicators - Check.
5. Gear handle - UP, then DN.
c. Gear-Up Landing.
The main landing gear
wheels protrude slightly from the wheel well in the gearup position and will roll when the aircraft is landed with
the gear retracted. Due to decreased drag with the gear
up, the tendency will be to overshoot the approach. It is
recommended that the landing be made with full flaps on
a hard surface runway, preferably paved. In landing on
soft ground or dirt, sod has a tendency to roll up into
chunks, damaging the underside of the aircraft's
structure. When making a gear-up landing, proceed as
follows:
1. Crew/passenger emergency briefing
- Complete.
2. Loose equipment - Stow.
6. Gear position - Check (use air-to-air
or air-to-ground fly-by method for visual landing gear
position verification).
b. Landing Gear Emergency Extension.
The
landing gear may be extended manually if the electrical
portion of the extension mechanism should fail. As
airspeed is reduced, it is correspondingly easier to
manually actuate the emergency extension handle.
When making an emergency gear extension, proceed as
follows:
1. Airspeed - Below 156 KCAS (154
KIAS).
3. Seat belts and harnesses - Secure.
4. Gear emergency clutch disengage
lever - Disengage.
5. Gear emergency extension handle Stow.
6. Gear control circuit breaker - In.
7. Gear handle - UP.
8. Flaps - As required.
2. Gear power circuit breaker - Out
(pulled).
9. Non-essential electrical equipment 3. Gear handle - DN.
4. Gear emergency clutch disengage
lever - Pull up and turn clockwise.
CAUTION
Do not pump handle after GEAR
DOWN position indicator lights (3)
are illuminated. Further movement of
the handle could damage the drive
mechanism.
5. Gear emergency extension handle Pump the handle up and down until the three GEAR
DOWN green lights illuminate. In the event of complete
electrical failure, pump until resistance is felt.
9-16 Change 8
OFF.
10. Condition levers - FUEL CUTOFF
(on ground when able).
11. Master switch - Down.
d. Landing With Main Gear Down, Nose Gear Up
or Unlocked. If a landing gear indicator shows an unsafe
indication every possible means should be used to
determine the position of the landing gear. If a landing is
to be made with the nose gear up or unlocked and the
main gear down, then it is recommended that the landing
be made on a prepared runway. After touchdown, hold
the nose off as long as possible allowing the nose to
smoothly touch down with minimum speed. A full flap
landing is recommended. After touchdown, flaps should
be retracted.
TM 55-1510-215-10
Adjusting to full nose down elevator trim at this point will
aerodynamically
increase
up
elevator
control
effectiveness, thus, allowing a much lower nose
touchdown speed. Use the following procedure as
applicable:
1. Crew/passenger emergency briefing
- Complete.
toward the flat tire side. Directional control can be
maintained with wheel braking and reverse thrust. If
aware that a main gear tire 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,
nose wheel stability will be reduced and application of
brakes should be used only as required to maintain
positive control. Use the following procedures:
2. Loose equipment - Stow.
1. Land on side of runway favoring
good tire.
3. Seat belts and harnesses - Secured.
2. Brake - On good wheel only.
4. Non-essential electrical equipment OFF.
3. Flat nose tire - Use light braking.
5. Condition levers - FUEL CUTOFF
(on ground when able).
6. Master switch - Down.
e. Landing With One Main Gear Up or Unlocked. If
one main landing gear fails to extend and the opposite
gear extends normally, a break in the drive mechanism
to the unextended gear has occurred. If all efforts to
retract the extended gear fail, land the aircraft on a
runway or on firm, hard surface in preference to loose
dirt or grass. Touchdown smoothly, well over to the
same side of the runway as the extended gear to allow
room for loss of directional control. Holding the opposite
wing high and nose straight, lightly apply brake to the
unsafe gear. This jar may result in locking the unsafe
gear. If not, allow the opposite wing to lower slowly.
Evacuate the aircraft as soon as possible. Use the
following procedures as applicable:
1. Crew/passenger emergency briefing
- Complete.
2. Loose equipment - Stow.
3. Seat belts and harnesses - Secured.
4. Non-essential electrical equipment OFF.
5. Condition levers - FUEL CUTOFF
(on ground when able).
6. Master switch - Down.
f. Landing With Nose Gear Don, Main Gear Up or
Unlocked.
Complete the same approach and
procedures as used for GEAR-UP LANDING.
9-14. Ditching.
a. General. The following ditching procedures are
based on the experiences of pilots who have
successfully ditched other multiengine aircraft. The
success of those ditchings were the result of all
crewmembers carrying out the correct ditching
procedures. Ditching commenced from low altitudes
does not always allow time for more than minimum
preparation and planning and may not permit relying on
the checklist.
Therefore it is essential that each
crewmember be thoroughly familiar with ditching
procedures and assigned responsibilities. Further, the
pilot in command must insure that all passengers have
been briefed on ditching procedures and understand
how to use installed survival equipment. If at all possible
ditching should be made while power is still available on
both engines. However, if one engine has failed, the
ditching should be accomplished in as near symmetrical
condition as possible. An engine and/or wing fire is
probably the most serious condition from the standpoint
of structural integrity and lateral control.
A fire
concentrated within the wing or nacelle will be sustained
by fuel or oil and will destroy effective use of flaps and
ailerons in a very short time. With such a fire, except in
extremely high wind conditions the aircraft should be
ditched parallel to the primary swell system. Model test
and actual ditchings of various aircraft indicate that
ditchings into the wall of water created by the major
swell is roughly analogous to flying into a mountain.
Accordingly, a careful evaluation of sea condition is
essential to successful ditching. While descending,
begin analyzing the sea condition as soon as the surface
can be seen clearly (2000 feet or more if possible). The
primary swell can readily be distinguished from high
altitude and will be seen first. At low altitude it may be
hidden beneath another system plus a surface chop, but
from altitude the largest and most dangerous system will
be the first one recognized.
g. Landing With Flat Tire. When landing is made
with a flat tire on one main gear. the aircraft may turn
9-17
TM 55-1510-215-10
The wind driven sea, if any, will be recognized by the
appearance of white caps.
anticipated and available power is
insufficient to maintain desired rate of
descent, consideration should be
given to utilization of approach flaps. It
is difficult to judge height above a calm
sea, and use of landing lights at night
make the estimate even more difficult
b. Ditching Under IMC Conditions. Where IMC
conditions or night operations preclude visual
determination of sea conditions, forecast data should be
utilized, and the ditching must be made on instruments.
With no surface reference, the aircraft must be flown into
the water on heading, and in a fixed attitude which
combines safe control speed and rate of descent.
Whenever possible, ditching should be made as close as
safety permits to coastlines or in the vicinity of surface
vessels to improve the rescue situation. If radios are still
operational, attempt to contact coastal stations or surface
vessels for current wind, sea swell, and altimeter setting.
WARNING
Do not remove raft from its carrying
case inside the aircraft.
NOTE
CAUTION
Ditch parallel to and near the crest of
the swell unless there is a strong
crosswind of 20 knots or more. In
strong winds ditch heading should be
more into the wind and slightly
across the swell, planning to
touchdown on the upslope of the
swell near the top (fig. 9-6). Wave
motion is indicative of wind direction,
but the swell does not necessarily
move with the wind. Water surface
conditions
are
indicative
of
windspeed, as related below:
SURFACE CONDITION
Few white crests
Many white crests
Streaks of foam from crests
Spray blown from tops of
waves
WIND
SPEED KNOTS
10-15
15-25
25-35
35-45
NOTE
Full flaps are recommended for power
on ditching.
However, impaired
directional control caused by engine
loss or aileron damage is another
factor to be considered in determining
ditching configuration. At low gross
weights, it is possible to retain an
optimum ditching airspeed, desired
rate of descent, and directional control
by using full flaps and power as
necessary on the operating engine. If
direction
control
problems
am
9-18
Keep life raft away from any damaged
surfaces which might tear it.
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.
c. Ditching Procedure.
(1) Ditching procedure with power
1. Announce intention to ditch and time
to impact.
2. Distress message - Transmit
3. Transponder - Emergency.
4. Life vest - Put on and adjust (do not
inflate).
5. Seat belts/harnesses
(passengers in braced position).
6. Gear - UP.
7. Flaps –Down.
-
Secure
TM 55-1510-215-10
NOTE
Rate of descent, 100 ft. per min.
during final stages of the approach
(approximately last 300 feet).
8. Airspeed - 103 KCAS (100 KIAS).
(2) Ditching Procedure Without Power:
e. After Leaving Aircraft. When safely clear of the
aircraft, pull inflation ring to inflate the raft. Tie down first
aid kit in the center of the raft to prevent it from being
lost in case the raft capsizes. After all personnel have
been evacuated move raft out from under any part of the
aircraft which might strike them as it sinks. Remain in
the vicinity of the aircraft as long as it remains afloat.
9-15. Bailout.
1. Announce intention to ditch and time
to impact.
2. Distress message - Transmit 3.
Transponder - Emergency.
4. Life vest - Put on and adjust (do not
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 the Emergency Exits and Equipment
illustration (fig. 9-1). Proceed as follows if bailout
becomes necessary:
inflate).
1. Radio - Distress procedure (If time
5. Seat belts/harnesses
(passengers in braced position).
-
Secure
permits).
2. Voice security and transponder -
6. Gear - UP.
ZEROIZE.
7. Flaps - APPROACH.
3. Airspeed - Reduce.
8. Airspeed - 100 KIAS.
4. Flaps - DOWN.
d. Water Entry. It is essential that an attempt be
made to control the attitude of the aircraft throughout
ditching until all motion stops. Evacuate the aircraft
through the emergency exit or main entrance door.
Take the life raft and first aid kit.
5. Trim - As required.
6. Main entrance door - OPEN.
7. Abandon the aircraft.
Change 8 9-19/(9-20 blank)
TM 55-1510-215-10
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 95-16
AR 380-40
AR 385-40
AR 700-26
FAR Part 91
FM 1-5
FM 1-30
TB 55-1500-314-25
TB 55-9150-200-24
TB AVN 23-13
TB MED 501
TM 11-6140-203-14-2
TM 38-750
TM 55-405-9
TM 55-410
TM 55-1510-200-PM
TM 55-1500-204-25/1
TM 55-1510-215-23
TM 750-244-1-5
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
Instrument Flying and Navigation for Army Aviators
Meteorology for Army Aviators
Handling, Storage, and Disposal of Army Aircraft Components Containing Radioactive
Materials
Engine and Transmission Oils, Fuels and Additives for Army Aircraft
Anti-icing, Deicing and Defrosting Procedures for Parked Aircraft
Noise and Conservation of Hearing
Operator's Organizational, Direct Support, General Support and Depot
Maintenance Manual Including Repair Parts and Special Tools List: Aircraft NickelCadmium Batteries
Army Maintenance Management System
Army Aviation Maintenance Manual: Weight and Balance
Aircraft Maintenance, Servicing and Ground Handling Under Extreme Environmental
Conditions
Phased Maintenance Checklist
General Aircraft Maintenance Manual
Aviation Unit and Aviation Intermediate Maintenance Manual, Army Models U-21G, RU21E and RU-21H
Procedures for the Destruction of Aircraft and Associated Equipment to Prevent Enemy
Use
A-1/(A-2 blank)
TM 55-1510-215-10
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.
B-1. 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.
Vlof
Lift off speed (takeoff airspeed).
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.
B-1
TM 55-1510-215-10
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.
Vx
Best angle-of-climb speed.
Vxse
Best single-engine angle of climb speed.
Vy
Best rate-of-climb speed.
Vyse
The best single-engine rate of climb speed.
B-2. Meteorological Terminology.
Altimeter Setting
Barometric pressure corrected to sea level.
°C
Degrees Celsius.
°F
Degrees Fahrenheit.
FAT
Free air temperature is the free air static temperature, obtained either from
ground meteorological sources or from inflight temperature indications adjusted
for compressibility effects.
Indicated Pressure
The number actually read from an altimeter when, the barometric scale
Altitude
(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 Centigrade;
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).
B-2
TM 55-1510-215-10
B-3. Power Terminology.
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. These powers are 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).
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.
B-4. Control and Instrument Terminology.
Condition Lever (Fuel
Shut-off Lever)
The fuel shut-off lever actuates a valve in the fuel control unit which controls the
flow of fuel at the fuel control outlet and regulates the idle range from LO to
HIGH.
Interstage
Turbine Temperature
(ITT)
Thermocouple probes wired in parallel indicate the temperature between the
compressor and power turbines.
N1 Tachometer (Gas
Generator RPM)
The tachometer registers the RPM of the gas generator with 100%
representing a gas generator speed of 37,500 RPM.
Power Lever (Gas
Generator N1 RPM)
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.
Propeller Control
Lever (N 2 RPM)
This lever requests the control to maintain RPM at a selected value and, in
the maximum decrease RPM position, feathers the propeller.
Propeller Governor
This Governor will maintain the selected propeller speed requested by the
propeller control lever.
Change 7 B-3
TM 55-1510-215-10
Torquemeter
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.
B-5. 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
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.
Climb Speed
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 lift off.
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 maximum rate of climb of 100 feet per minute can be
attained for existing aircraft weight.
Takeoff Weight
The weight of the aircraft at lift off from the runway.
B-6. 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.
B-4
TM 55-1510-215-10
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-7. Miscellaneous Abbreviations.
Deg
Degrees
DN
Down
FT
Foot or feet
FT-LB
Foot-pounds
GAL
Gallons
HR
Hours
kHz
Kilohertz
B-5
TM 55-1510-215-10
LB
Pounds
MAX
Maximum
MHz
Megahertz
MIN
Minimum
NAUT
Nautical
NM
Nautical miles
PSI
Pounds per square inch
R/C
Rate of climb
B-6
TM 55-1510-215-10
INDEX
Subject
Paragraph, Figure,
Table Number
A
AC Power Supply ................................................... 2-70
ADF Radio Set (KR 87) ............................... 3-26, F3-16
Adverse Environmental Conditions
Introduction ......................................................... 8-44
Aerial Delivery System ............................................ 6-18
After Landing........................................................... 8-31
After Takeoff .......................................................... 8-20
Air Cargo Features ................................................. 6-17
Aircraft Compartment and Stations .................. 6-3, F6-1
Aircraft Designation System .................................... 1-12
Aircraft - General ...................................................... 2-2
Aircraft - Introduction ................................................. 2-1
Aircraft Weighings .................................................... 6-7
Air Induction Systems - General ............................. 2-20
Airspeed Indicators ................................................. 2-75
Airspeed Limitations ................................................ 5-17
Aircraft Personnel Cargo Features .......................... 6-14
Airspeed Position Error Correction .......................... 7-16
Altimeter, Copilots ................................................... 2-78
Altimeter Position Error Correction .......................... 7-17
Altitude Limitations ................................................. 5-28
Altitude Selector Controller .......................... 3-23, F3-13
Ambient Temperature Takeoff
Limitation ............................................................ 5-13
Anti-Icing, Deicing and Defrosting
Protection ........................................................... 2-90
Anti-Icing, Deicing and Defrosting
Treatment ........................................................... 2-91
Appendix A, References .................................... 1-4, A-1
Appendix B, Abbreviations and Terms ............... 1-5, B-1
Application of External Power ................................. 2-92
Approved Fuels ............................................. 2-38, T2-5
Approved Oils ........................................................ T2-4
Army Aviation Safety Program................................... 1-7
Audio Control Panel ............................................... F3-1
Autoignition System ................................................ 2-32
Autoignition During Night Operation ........................ 5-29
Autopilot Limitations ............................................... 5-9A
Avionics Equipment Configuration ............................ 3-2
Avionics - Introduction .............................................. 3-1
Avionics - Power Source ........................................... 3-3
B
Bailout .................................................................... 9-15
Balance Definitions ................................................... 6-9
Bank and Pitch Limits ............................................. 5-27
Before Landing ....................................................... 8-27
Before Leaving Aircraft ........................................... 8-33
Subject
Paragraph, Figure
Table Number
B
Before Starting Engines .......................................... 8-12
Before Takeoff ....................................................... 8-17
Before Taxiing ........................................................ 8-14
Bus Overload Protection System ............................. 2-72
C
Cabin Door Warning Light ...................................... 2-11
Cargo Center-of-Gravity
Computation ............................................... 6-20, T6-1
Cargo Center-of-Gravity
Planning .............................................................. 6-19
Cargo Moment Chart .............................................. F6-7
Cargo Uploading .................................................... 6-27
Center-of-Gravity
Limitations................................................... 6-28, F6-8
Chart C - Basic Weight and
Balance Record ......................................... 6-10, F6-2
Checklists ................................................................ 8-9
Checklist Callout .................................................... 8-10
Cigarette Lighters and Ash Trays ............................ 2-63
Circuit Breaker and Fuse Panels .......................... F2-18
Climb - Max Rate ................................................... 8-22
Climb - Normal ....................................................... 8-21
Cockpit ................................................................... F2-5
Cockpit Vent/Storm Window Speed ........................ 5-23
Cold Weather Operation ......................................... 8-45
Communications - Description .................................. 3-4
Condition Levers .................................................... 2-26
Control Pedestal and Landing Gear
Emergency Extension Controls ........................... F2-6
Control Wheels ....................................................... 2-40
Control Wheels and Control Locks ........................ F2-16
Copilot's Altimeter .................................................. 2-78
Copilot's Attitude Director
Indicator......................................... 2-80A, 3-20, F3-11
Copilot's Horizontal Situation
Indicator ..................................................... 3-17, F3-9
Course Deviation Indicator (KI 204) ........................ 3-18
Crew/Passenger Briefings ............................... 8-6, 8-51
Crossfeed Fuel Flow ............................................ F2-13
Cruise Checks......................................................... 8-23
D
DC Power Supply ................................................... 2-69
Definition of Landing Terms ...................................... 9-3
Defrosting System .................................................. 2-52
Descent .................................................................. 8-25
Change 5 INDEX-1
TM 55-1510-215-10
INDEX
Paragraph, Figure,
Table Number
Subject
D
Descent - Arrival Check ......................................... 8-26
Desert Operation .................................................. 8-46
Destruction of Army Materiel to
Prevent Enemy Use.............................................. 1-8
Dimensions ................................................... 2-3, F2-3
Ditching ................................................................ 9-14
Diving ................................................................... 8-40
Door Open Light Illuminated .................................. 9-11
E
Electrical Power Supply and Distribution
System - Description ....................................2-68, F2-19
Electrical System (Emergencies) ............................ 9-10
Electrical System Schematic ............................... F2-19
Electronic Equipment Configuration .......................... 3-2
Emergency Descent .............................................. 9-12
Emergency Entrance ...................................... 9-5, F9-2
Emergency Equipment - Description ...................... 2-14
Emergency Exits and Equipment .................... 9-4, F9-1
Emergency Procedures - Aircraft Systems ............... 9-1
Engine and Related Systems Description...............................................2-18, F2-10
Engine (Emergencies) ............................................. 9-6
Engine Chip Detection System ............................... 2-30
Engine Compartment Cooling ................................. 2-19
Engine Fuel Control System ................................... 2-24
Engine Fire Detection System................................. 2-28
Engine Ignition System........................................... 2-31
Engine Instruments ............................................... 2-34
Engine Limitations ................................................. 5-10
Engine Operating Limitations .................................. T5-1
Engine Runup........................................................ 8-16
Engine Shutdown .................................................. 8-32
Engine Starter - Generators ................................... 2-33
Entrance and Exit Provisions ........................ 2-10, F2-8
Exceeding Operational Limits .................................. 5-3
Exhaust Danger Area ............................................ F8-2
Explanation of Change Symbols ............................ 1-11
Exterior Area Check .............................................. F8-1
Exterior Lighting ..........................................2-73, F2-20
Subject
Paragraph, Figure,
Table Number
F
First Aid Kits ......................................................... 2-16
Flight Controls ...................................................... 8-42
Flight Controls - Description ................................... 2-39
Flight Controls Malfunction ................................. 9-10.1
Flight Control System (KFC 250) ............................ 3-27
Flight Controls Lock .............................................. 2-42
Flight Director ...................................... 2-80, 3-17, F3-7
Flight Envelope Chart ............................................ F5-3
Flight Plan ............................................................... 8-5
Flight Under IMC (Instrument
Meteorological Conditions) ................................. 5-31
Floor Loading Limits .............................................. 5-16
FM Liaison Set (AN/ARC-131) ........................ 3-7, F3-2
Foreign Object Damage Control ............................. 2-21
Forms and Records ............................................... 1-10
Free Air Temperature Gage.................................... 2-81
Friction Lock Knobs................................................ 2-27
Fuel and Oil Data .................................................. 6-13
Fuel Load ............................................................. 6-12
Fuel, Lubricants, Specifications,
and Capacities ................................................... T2-3
Fuel Management .......................................2-36, F2-12
Fuel Moment Chart ............................................... F6-4
Fuel Quantity Data ................................................ T2-1
Fuel System (Emergencies) ..................................... 9-9
Fuel System Anti-Icing ........................................... 2-57
Fuel System Crossfeed ......................................... 8-24
Fuel System - Description ............................2-35, F2-11
G
General Exterior Arrangement ............................... F2-1
General Interior Arrangement ................................ F2-2
Go-Around ............................................................ 8-30
Gravity Fuel Flow ................................................ F2-14
Ground Fault Protection System ............................. 2-71
Ground Handling ................................................... 2-94
Ground Proximity Altitude
Advisory System................................................. 3-39
Ground Turning Radius .................................. 2-4, F2-4
Gyro Magnetic Compass System............................ 3-15
F
Feathering Provisions ............................................ 2-46
Ferry Fuel System .......................................2-37, F2-15
Fire ......................................................................... 9-8
Fire Axe ............................................................... 2-17
Fire Extinguisher, Hand Operated .......................... 2-15
INDEX-2 Change 9
H
Hand-Operated Fire Extinguisher ........................... 2-15
Health Indicator Test ............................................. F8-3
Heating and Ventilation Systems ........................... 2-67
HF Communication Set (KHF 950) ......................... 3-11
Hot Weather Operation........................................... 8-47
TM 55-1510-215-10
INDEX
Paragraph, Figure,
Table Number
Subject
I
Ice and Rain ......................................................... 8-50
Immediate Action Emergency Checks ...................... 9-2
Inflating Tires ........................................................ 2-89
Inlet Air Separator System ..................................... 2-22
Installation of Protective Covers ............................. 2-96
Instrument Glass Alignment Marks ........................... 5-7
Instrument Flight - General .................................... 8-34
Instrument Flight Procedures ................................. 8-35
Instrument Marking Color
Codes ................................................................. 5-6
Instrument Markings ....................................... 5-5, F5-1
Instrument Panel ................................................. F2-22
Interior Lighting...................................................... 2-74
Interphone System .................................................. 3-6
Introduction - Description ......................................... 1-3
Introduction - General .............................................. 1-1
L
Landing ................................................................ 8-29
Landing Emergencies ............................................ 9-13
Landing Gear Extension Speed .............................. 5-20
Landing Gear Retraction Speed ............................. 5-21
Landing Gear System ............................................. 2-6
Level Flight Characteristics .................................... 8-43
Line Up ................................................................ 8-18
Load Planning ....................................................... 6-23
Loading Procedure................................................. 6-24
M
Magnetic Compass ............................................... 2-82
Maneuvering Flight ............................................... 8-41
Maneuvers ........................................................... 5-26
Marker Beacon/Glideslope Receiver
(R-884A/ARN-58) ............................................... 3-22
Maximum Allowable Airspeed ................................ 5-18
Maximum Design Maneuvering
Speed................................................................ 5-25
Maximum Glide Distance ....................................... F9-3
Maximum Weights .................................................. 2-5
Microphone Switches .............................................. 3-5
Minimum Crew Requirements .................................. 5-4
Minimum Single-Engine Control
Airspeed ............................................................... 5-24
Miscellaneous Instruments .................................... 2-84
Mission Planning ..................................................... 8-1
Mooring ......................................................2-97, F2-26
Paragraph, Figure,
Table Number
Subject
N
NAV 1 Receiver (KNS 81) ............................3-24, F3-14
NAV 2 Receiver (KN 53) ..............................3-25, F3-15
Navigation - Description ........................................ 3-13
Night Flying .......................................................... 8-36
Normal Flight Characteristics ................................. 8-37
Normal Procedures - Performance ........................... 8-4
O
Obstacle Clearance Approach ............................... 8-28
Oil Supply System ................................................ 2-29
Operating Limits and Restrictions General ............................................................... 5-2
Operating Limits and Restrictions Normal Procedures .............................................. 8-2
Operating Limits and Restrictions Purpose .............................................................. 5-1
Operating Procedures and Maneuvers ..................... 8-7
Overhead Control Panel ...................................... F2-21
Overtemperature and Overspeed Limitations .......... 5-11
Oxygen Duration Table .......................................... T2-2
Oxygen Requirements ........................................... 5-30
Oxygen System ..........................................2-59, F2-17
Oxygen System Servicing Pressure ..................... F2-24
P
Parking.................................................................. 2-95
Parking Brake Handle .............................................. 2-9
Parking, Covers, Ground Handling, and
Towing Equipment ........................................... F2-25
Passenger Seats .................................................. 5-33
Pedestal ............................................................... F2-6
Penetration ........................................................... 8-49
Performance Charts - Data Basis ............................. 7-9
Performance Data - Chart Explanation ..................... 7-5
Performance Data - Color Coding ............................ 7-7
Performance Data - Description ............................... 7-1
Performance Data - General .................................... 7-3
Performance Data - General Conditions ................. 7-11
Performance Data - Index ........................................ 7-6
Performance Data - Limits ....................................... 7-4
Performance Data - Purpose ................................... 7-2
Performance Data - Reading the Charts ................... 7-8
Performance Data - Specific
Conditions ......................................................... 7-10
Performance Discrepancies ................................... 7-12
Performance Information........................................ 7-15
Change 5 INDEX-3
TM 55-1510-215-10
INDEX
Paragraph, Figure,
Table Number
Subject
P
Performance Planning Card .......................... 7-13, F7-1
Performance Planning Sequence............................ 7-14
Personnel Loading and Unloading ................. 6-15, F6-5
Personnel Load Consumption ................................ 6-16
Personnel Moments Chart ..................................... F6-6
Phased Maintenance Checklist ................................ 1-9
Pilot's Encoding Altimeter ................... 2-77, 3-21, F3-21
Pilot's Flight Director
Indicator ......................................... 2-80, 3-19, F3-10
Pilot's Horizontal Situation
Indicator ................................................... 3-16, F3-8
Pitot and Stall Warning Heat
System ............................................................. 2-56
Pitot and Static System .......................................... 2-55
Power Definitions for Engine
Operation........................................................... 5-12
Power Levers ....................................................... 2-25
Power Plant Ice Protection Systems ....................... 2-23
Preflight Check ..................................................... 8-11
Preparation of Cabin for Loading ............................ 6-22
Preparation of General Cargo ................................ 6-21
Propeller (Emergencies) .......................................... 9-7
Propeller Electrothermal Deicer
System ............................................................. 2-54
Propeller Governors .............................................. 2-47
Propeller Governor Test Switches .......................... 2-48
Propeller Levers..................................................... 2-49
Propeller Limitations ................................................ 5-8
Propeller Reversing ............................................... 2-50
Propellers - Description ......................................... 2-45
Propeller Tachometers .......................................... 2-51
R
Radio Altimeter Indicator .............................3-22, F3-12
Radio Magnetic Indicators .................... 2-83, 3-14, F3-6
Radio Telephone (KT 96) ...................................... 3-12
Rear View Mirror ................................................... 2-65
Recommended Fluid Dilutions Chart ...................... T2-7
Relief Tubes ......................................................... 2-64
Rudder Pedals ...................................................... 2-41
Paragraph, Figure,
Table Number
Subject
S
Servicing Hydraulic Brake System
Reservoir............................................................ 2-88
Servicing Locations .............................................. F2-23
Servicing Oil System.............................................. 2-87
Servicing Oxygen System ..................................... 2-93
Servicing, Parking, and Mooring General ............................................................. 2-85
Spins ................................................................... 8-39
Split Flap Condition ............................................ 9-11.1
Stalls ................................................................... 8-38
Stall Speed Chart .................................................. F8-4
Stall Warning System............................................. 2-61
Standard, Alternate, and
Emergency Fuels ............................................... T2-6
Starter Limitations ................................................... 5-9
Steerable Nose Wheel ............................................ 2-7
Starting Engines (Battery/GPU) .............................. 8-13
Subpanels ............................................................ F2-7
Sun Visors ............................................................ 2-66
Surface Deicer System .......................................... 2-53
Symbols Definition .................................................. 8-8
T
Takeoff.................................................................. 8-19
Takeoff Temperature Limitations ............................ F5-2
Taxiing .................................................................. 8-15
Tie-down Devices ................................................. 6-26
Transponder (KT 76A) .................................3-28, F3-18
Transponder Set (AN/APX-72) .....................3-29, F3-19
Trim Tabs ............................................................. 2-43
Turbulence Penetration Speed................................ 5-19
Turbulent Air and Thunderstorms ........................... 8-48
Turn-and-Bank Indicators ..............................2-76, 3-21
U
UHF Command Set (AN/ARC-51BX) ............... 3-9, F3-4
Use of Words Shall, Should and May ..................... 1-13
S
Seating Provisions ................................................ 2-62
Seats .......................................................... 2-13, F2-9
Securing Loads ..................................................... 6-25
Servicing Fuel System .......................................... 2-86
INDEX-4 Change 7
V
Vertical Velocity Indicators...................................... 2-79
VHF Command Sets .................................... 3-10, F3-5
Voice Security System (TSEC/KY-28) ...................... 3-8
TM 55-1510-215-10
INDEX
Subject
Paragraph, Figure,
Table Number
W
Warning Pages .......................................................... a
Warning, Cautions, and Notes ................................. 1-2
Weather Radar Set (AN/APN-215(V) 1) .................. 3-30
F3-20
Weight and Balance - Charts and Forms ................... 6-5
Weight and Balance Clearance Form
F, DD Form 365F (Transport) ..................... 6-11, F6-3
Weight and Balance - Purpose ................................. 6-4
Weight and Balance - Responsibility ........................ 6-6
Weight/Balance and Loading, Class ......................... 6-2
Weight/Balance and Loading, Extent
of Coverage ......................................................... 6-1
Subject
Paragraph, Figure,
Table Number
W
Weight/Balance and Loading - Normal
Procedures .......................................................... 8-3
Weight Definitions ................................................... 6-8
Weight Limitations ................................................. 5-15
Wind Limitations ................................................... 5-32
Wheel Brake System................................................ 2-8
Windows .............................................................. 2-12
Windshield Electrothermal Anti-Ice
System ............................................................. 2-58
Windshield Wipers ................................................ 2-60
Wing Flaps ........................................................... 2-44
Wing Flap Extension Speeds ................................. 5-22
Change 5 INDEX-5/(INDEX-6 blank)
TM 55-1510-215-10
By Order of the Secretary of the Army:
Official:
E. C. MEYER
General, United States Army
Chief of Staff
ROBERT M. JOYCE
Major General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, Operator Maintenance Requirements for U-21 aircraft.
I U.S. GOVERNMENT PRINTING OFFICE : 1994 0 - 300-421 (82212)
TM 55-1510-215-10
TM 55-1510-215-10
TM 55-1510-215-10
TM 55-1510-215-10
The Metric System and Equivalents
Linear Measure
Liquid Measure
1 centiliter = 10 milliliters = .34 fluid ounce
1 deciliter = 10 centiliters = 3.38 fluid ounces
1 liter = 10 deciliters = 33.81 fluid ounces
1 dekaliter = 10 liters = 2.64 gallons
1 hectoliter = 10 dekaliters = 27.42 gallons
1 kiloliter = 10 hectoliters = 264.18 gallons
1 centimeter = 10 millimeters = .39 inch
1 decimeter= 10 centimeters = 3.94 inches
1 meter = 10 decimeters = 39.37 inches
1 dekameter = 10 Meters = 32.8 feet
1 hectometer = 10 dekameters = 328.08 feet
1 kilometer = 10 hectometers = 3,280.8 feet
Square Measure
Weights
1 sq. centimeter = 100 sq millimeters = .155 sq. inch
1 sq. decimeter= 100 sq centimeters = 125.5 sq. inches
1 sq. meter (centare) = 100 sq decimeters = 10.76 sq. feet
1 sq. dekameter (are) = 1,076.4 sq. feet
1 sq. hectometer (hectare) = 100 sq. dekameters = 2.47 acres
1 sq. kilometer = 100 sq. hectometers = .386 sq. mile
1 centigram = 10 milligrams = .15 grain
1 decigram = 10 centigrams = 1.54 grains
1 gram = 10 decigram = 0.35 ounce
1 dekagram = 10 Grams = .35 ounce
1 hectogram = 10 dekagrams = 3.52 ounces
1 kilogram = 10 hectograms = 2.2 pounds
1 quintal = 100 kilograms = 220.46 pounds
1 metric ton = 10 quintals = 1.1 short tons
Cubic Measure
1 cu. centimeter = 1000 cu. millimeters = .06 cu. inch
1 cu. decimeter = 1000 cu. centimeters = 61.02 cu. inches
1 cu. meter = 1000 cu. decimeters = 35.31 cu. feet
Approximate Conversion Factors
To change
To
inches
feet
yards
miles
square inches
square feet
square yards
square miles
acres
cubic feet
cubic yards
fluid ounces
pints
quarts
gallons
ounces
pounds
short tons
pound-feet
pounds-inches
centimeters
meters
meters
kilometers
square centimeters
square meters
square meters
square kilometers
square hectometers
cubic meters
cubic meters
milliliters
liters
liters
liters
grams
kilograms
metric tons
newton-meters
newton-meters
Multiply by
2.540
.305
.914
1.609
6.451
.093
.836
2.590
.405
.028
.765
29.573
.473
.946
3.785
28.349
.454
.907
1.365
.11375
To change
To
Multiply by
ounce-inches
centimeters
meters
meters
kilometers
square centimeters
square meters
square meters
square kilometers
square hectometers
cubic meters
cubic meters
milliliters
liters
liters
liters
grams
kilograms
metric tons
newton-meters
inches
feet
yards
miles
square inches
square feet
square yards
square miles
acres
cubic feet
cubic yards
fluid ounces
pints
quarts
gallons
ounces
pounds
short tons
.007062
.394
3.280
1.094
.621
.155
10.764
1.196
.386
2.471
35.315
1.308
.034
2.113
1.057
.264
.035
2.205
1.102
Celsius Temperature
°C
Temperature (Exact)
°F
Fahrenheit Temperature
5/9 (after subtracting 32)
TM 55-1510-215-10
PIN: 032081-010
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