TM-55-1510-221-10

TM-55-1510-221-10
TM 55-1510-221-10
TECHNICAL MANUAL
OPERATOR’S MANUAL
FOR
ARMY RC-12H 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
“Approved for public release; distribution is unlimited.”
EMERGENCY PROCEDURES
REFERENCES
HEADQUARTERS, DEPARTMENT
OF THE ARMY
30 DECEMBER 1988
This copy is a reprint which includes current
pages from Changes 1 through 4.
ABBREVIATIONS AND TERMS
ALPHABETICAL INDEX
URGENT
TM 55-1510-221-10
C5
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 28 May 1998
CHANGE
NO. 5
OPERATOR’S MANUAL
FOR
ARMY RC-12H AIRCRAFT
DlSTRlBUTlON STATEMENT A: Approved for public release; distribution is unlimited
TM 55-1510-221-10, 30 December 1988, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by vertical
bar in the margin. An illustration change is indicated by a miniature pointing hand.
2.
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i and ii
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INDEX-3 and INDEX4
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By Order of the Secretary of the Army:
DENNIS J. REIMER
General, United States Army
Chief of Staff
Administrative Assistant to the
Secretary of the Army
04640
DISTRIBUTION:
To be distributed in accordance with Initial Distribution Number (IDN) 310993, requirements for TM 55
1510-221-10.
TM 55-1510-221-10
C4
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 30 April 1993
CHANGE
NO. 4
Operator's Manual
For
Army RC-12H Aircraft
DISTRIBUTION STATEMENT A:
Approved for public release; distribution is unlimited.
TM 55-1510-221-10, 30 December 1988, is changed as follows:
1.
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TM 55-1510-221-10
C4
By Order of the Secretary of the Army:
GORDON R. SULLIVAN
General, United States Army
Chief of Staff
Official:
MILTON H. HAMILTON
Administrative Assistant to the
Secretary of the Army
C4224
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31-E, block no 0993, requirements for TM 55-1510-221-10.
URGENT
TM 55-1510-221-10
C3
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 17 August 1992
CHANGE
NO. 3
Operator's Manual
For
Army RC-12H Aircraft
TM 55-1510-221-10, 30 December 1988, is changed as follows:
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is indicated by a vertical bar in the margin. An illustration change is indicated
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8-29 and 8-30
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2.
By Order of the Secretary of the Army:
GORDON R. SULLIVAN
General, United States Army
Chief of Staff
Official
MILTON H. HAMILTON
Administrative Assistant to the
Secretary of the Army
02244
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31-E, block no. 0993, requirements
for TM 55-1510-221-10.
DISTRIBUTION STATEMENT A:
Approved for public release; distribution is unlimited.
URGENT
TM 55-1510-221-10
C2
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 30 June 1992
CHANGE
NO. 2
Operator's Manual
For
Army RC-12H Aircraft
TM 55-1510-221-10, 30 December 1988, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material
is indicated by a vertical bar in the margin. An illustration change is indicated
by a miniature pointing hand.
2.
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- - - - -
3-76A through 3-76F
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By Order of the Secretary of the Army:
GORDON R. SULLIVAN
General, United States Army
Chief of Staff
Official:
MILTON H. HAMILTON
Administrative Assistant to the
Secretary of the Army
01862
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31-E, block no. 0993, -10 & CL
maintenance requirements for TM 55-1510-221-10.
TM 55-1510-221-10
C1
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 9 November 1990
CHANGE
No. 1
Operator's Manual
For
Army RC-12H Aircraft
TM 55-1510-221-10, 30 December 1988, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material
is indicated by a vertical bar In the margin. An illustration change is indicated
by a miniature pointing hand.
2.
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9-11 and 9-12
9-12.1/9-12.2
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By Order of the Secretary of the Army:
CARL E. VUONO
Official:
General, United States Army
Chief of Staff
THOMAS F. SIKORA
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, -10 & CL Maintenance requirements for RC-12D Airplane, Reconnaissance.
TM 55-1510-221-10
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.
NOISE LEVELS
Sound pressure levels in this aircraft during some operating conditions exceed the Surgeon General’s hearing conservation criteria, as defined in TM MED 501. Hearing protection devices, such as the aviator helmet
or ear plugs shall be worn by all personnel in and around the aircraft during its operation.
STARTING ENGINES
Operating procedures or practices defined in this Technical Manual must be followed correctly. Failure to
do so may result in personal injury or loss of life.
Exposure to exhaust gases shall be avoided since exhaust gases are an irritant to eyes, skin and respiratory
system.
HIGH VOLTAGE
High voltage is a possible hazard around AC inverters, ignition exciter units, and strobe beacons.
USE OF FIRE EXTINGUISHERS IN CONFINED AREAS
Monobromotrifluoromethane (CF3Br) is very volatile, but is not easily detected by its odor. Although non
toxic, it must be considered to be about the same as other freons and carbon dioxide, causing danger to personnel primarily by reduction of oxygen available for proper breathing. During operation of the fire extinguisher, ventilate personnel areas with fresh air. The liquid shall not be allowed to come into contact with the
skin, as it may cause frostbite or low temperature burns because of its very low boiling point.
VERTIGO
The strobe/beacon lights should be turned off during flight through clouds to prevent sensations of vertigo,
as a result of reflections of the light on the clouds.
CARBON MONOXIDE
When smoke, suspected carbon monoxide fumes, or symptoms of lack of oxygen (hypoxia) exist, all personnel shall immediately don oxygen masks, and activate the oxygen system.
FUEL AND OIL HANDLING
Turbine fuels and lubricating oils contain additives which are poisonous and readily absorbed through the
skin. Do not allow them to remain on skin.
SERVICING AIRCRAFT
When conditions permit, the aircraft shall be positioned so that the wind will carry fuel vapors away from
all possible sources of ignition. The fueling unit shall maintain a distance of 20 feet between unit and filler
point. A minimum of 10 feet shall be maintained between fueling unit and aircraft.
a
TM 55-1510-221-10
Prior to refueling, the hose nozzle static ground wire shall be attached to the grounding lugs that are
located adjacent to filler openings.
SERVICING BATTERY
Improper service of the nickel-cadmium battery is dangerous and may result in both bodily injury and
equipment damage. The battery shall be serviced in accordance with applicable manuals by qualified personnel only.
Corrosive Battery Electrolyte (Potassium Hydroxide). 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.
JET BLAST
Occasionally, during starting, excess fuel accumulation in the combustion chamber causes flames to be
blown from the exhausts. This area shall be clear of personnel and flammable materials.
RADIOACTIVE MATERIAL
Instruments contained in this aircraft may contain radioactive material (TB 55-1500-314-25). These items
present no radiation hazard to personnel unless seal has been broken due to aging or has accidentally been broken. If seal is suspected to have been broken, notify Radioactive Protective Officer.
RF BURNS
Do not stand near the antennas when they are transmitting.
OPERATION OF AIRCRAFT ON GROUND
At all times during a towing operation, be sure there is a man in the cockpit to operate the brakes.
Personnel should take every precaution against slipping or falling. Make sure guard rails are installed when
using maintenance stands.
Engines shall be started and operated only by authorized personnel. Reference AR 95-1.
Insure that landing gear control handle is in the DN position.
AUTOPILOT COMPATIBILITY
The RC-12H aircraft is certified with wingtip pods installed. Should the pods be removed, the autopilot
system must be replaced with a standard C-12D autopilot. Effected wiring must also be changed.
b
TM 55-1516-221-10
TECHNICAL MANUAL
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C. 30 DECEMBER 1988
Operator’s Manual
ARMY MODEL RC-12H
REPORTING OR ERRORS AND RECOMMENDING IMPROVEMENTS
You can help Improve this manual. If you find any mistakes or if you know of any way to improve the procedures,
please let use know. Mail your letter. DA Form 2028 (Recommended Changes to Publications and Blank Forms), or
DA Form 2028-2 located in the back of this manual directly to: Commander, U.S. Army Aviation and Missile
Command. ATTN: AMSAM-MMC-LS-LP, Redstone Arsenal. AL 35898-5230. A reply will be furnished directly
to you. You may also send in your comments electronically to our E-mail address at <[email protected]>, or
by fax at (205) 842-6546 or DSN 788-6546. Instructions for sending an Electronic DA Form 2028 may be found at
the back of this manual immediately preceding the hard copy DA Forms 2028.
TABLE OF CONTENTS
PAGE
CHAPTER 1
INTRODUCTION
1-1
CHAPTER 2
AIRCRAFT AND SYSTEMS DESCRIPTION AND
OPERATION
2-1
Section
CHAPTER 3.
Section
I.
II.
III.
IV.
V.
VI.
VII.
VIII.
2-1
2-19
2-20
2-27
2-34
2-42
2-51
2-52
IX.
X.
XI.
XII.
Aircraft
Emergency equipment
Engine and related systems
Fuel systems
Flight controls
Propellers
Utility systems
Heating, ventilation, cooling, and environmental control
system
Electrical power supply and distribution system
Lighting
Flight instruments
Servicing, parking, and mooring
I.
II.
Ill.
IV.
AVIONICS
General
Communications
Navigation
Transponder and radar
3-1
3-1
3-2
3-20
3-77
MISSION EQUIPMENT
4-1
Mission avionics
Aircraft survivability equipment
4-1
4-1
OPERATING LIMITS AND RESTRICTIONS
5-1
CHAPTER 4.
Section
1.
II.
CHAPTER 5.
Section
CHAPTER 6.
I.
II.
Ill.
IV.
V.
VI.
VII.
VIII.
General
System limits
Power limits
Loading limits
Airspeed limits, maximum and minimum
Maneuvering limits
Environmental restrictions
Other limitations
WEIGHT/BALANCE AND LOADING
2-58
2-67
2-70
2-77
5-1
5-1
5-6
5-8
5-6
5-10
5-10
5-11
6-1
Change
5
i
TM 55-1510-221-10
TABLE OF CONTENTS (CONT’D)
PAGE
Section
1.
II.
III.
IV.
V.
6-1
6-1
6-14
6-14
6-14
CHAPTER 7.
PERFORMANCE DATA
7-1
CHAPTER 8.
NORMAL PROCEDURES
8-1
Section
I.
II.
Ill.
IV.
V.
VI.
Mission planning
Operating procedures and maneuvers
Instrument flight
Flight characteristics
Adverse environmental conditions
Crew duties
8-1
8-1
8-23
8-24
8-26
8-29
EMERGENCY PROCEDURES
9-1
Aircraft systems
9-1
APPENDIX A.
REFERENCES
A-1
APPENDIX B.
ABBREVIATIONS AND TERMS
B-1
CHAPTER 9.
Section
INDEX
ii
General
Weight and balance
Fuel/oil
Cargo loading
Center of gravity
Change 5
I.
INDEX-1
TM 55-1510-221-10
CHAPTER 1
INTRODUCTION
1-1. GENERAL.
These instructions are for use by the operator(s).
They apply to the RC-12H aircraft.
1-5. APPENDIX B,
TERMS.
ABBREVIATIONS
AND
Appendix B is a listing of abbreviations and
terms used throughout the manual.
1-2. WARNINGS, CAUTIONS, AND NOTES.
Warnings, cautions, and notes are used to
emphasize important and critical instructions and
are used for the following conditions:
An operating procedure, practice, etc.,
which if not correctly followed, could
result in personal injury or loss of life.
1-6. INDEX.
The index lists, in alphabetical order, every
titled paragraph, figure, and table contained in this
manual. Chapter 7, Performance Data, has an additional index within the chapter.
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.
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.
1-3. DESCRIPTION.
This manual contains the best operating instructions and procedures for the RC-12H aircraft under
most circumstances. The observance of limitations,
performance, and weight/balance data provided is
mandatory. The observance of procedures is mandatory except when modification is required because
of multiple emergencies, adverse weather, terrain,
etc. Your flying experience is recognized, and therefore, basic flight principles are not included. THIS
MANUAL SHALL BE CARRIED IN THE AIRCRAFT AT ALL TIMES.
1-4. APPENDIX A, REFERENCES.
Appendix A is a listing of official publications
cited within the manual applicable to and available
for flight crews.
For information concerning destruction of Army
materiel to prevent enemy use, refer to TM 750-2441-5.
1-9. FORMS AND RECORDS.
Army aviators flight record and aircraft maintenance records which are to be used by crew members are prescribed in DA PAM 738-751 and TM
55-1600-342-23.
1-10. EXPLANATION OF CHANGE SYMBOLS.
Changes, except as noted below, to the text and
tables, including new material on added pages, are
indicated by a vertical line in the outer margin
extending close to the entire area of the material
affected; exception: pages with emergency markings,
which consist of black diagonal lines around three
edges, may have the vertical line or change symbol
placed along the inner margins. Symbols show current changes only. A miniature pointing hand symbol is used to denote a change to an illustration.
However, a vertical line in the outer margin, rather
than miniature pointing hands, is utilized when
there have been extensive changes made to an illustration. Change symbols are not utilized to indicate
changes in the following:
1-1
TM 55-1510-221-10
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 correction changes the meaning of instructive
information and procedures.
1-11. AIRCRAFT DESIGNATION SYSTEM.
The designation system prescribed by AR 70-50
is used in aircraft designations as follows:
EXAMPLE RC-12H
R - Modified mission symbol (Reconnaissance)
1-2
C - Basic mission and type symbol (cargo)
12 - Design number
H - Series symbol
1-12. 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.
1-13. PLACARD ITEMS.
All placard items (switches, controls, etc.) are
shown throughout this manual in capital letters.
TM 55-1510-221-10
CHAPTER 2
AIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATION
Section I. AIRCRAFT
2-1. INTRODUCTION.
2-6. EXHAUST DANGER AREA.
The purpose of this chapter is to describe the
aircraft and its systems and controls which contribute to the physical act of operating the aircraft. It
does not contain descriptions of avionics or mission
equipment, covered elsewhere in this manual. This
chapter contains descriptive information and does
not describe procedures for operation of the aircraft.
These procedures are contained within appropriate
chapters in the manual. This chapter also contains
the emergency equipment installed. This chapter is
not designed to provide instructions on the complete
mechanical and electrical workings of the various
systems; therefore, each is described only in enough
detail to make comprehension of that system sufficiently complete to allow for its safe and efficient
operation.
Danger areas to be avoided by personnel while
aircraft engines are being operated on the ground are
depicted in figure 2-5. Distance to be maintained
with engines operating at idle are also shown. Temperature and velocity of exhaust gases at varying
locations aft of the exhaust stacks are shown for
maximum power. The danger area extends to 40 feet
aft of the exhaust stack outlets. Propeller danger
areas are also shown.
2-2. GENERAL.
The RC-12H is a pressurized, low wing, all
metal aircraft, powered by two PT6A-41 turboprop
engines (fig. 2-1 and 2-11), and has all weather capability. Distinguishable features of the aircraft are the
slender, streamlined engine nacelles, an aft rotating
boom antenna, mission antennas, wing tip pods, a
T-tail and a ventral tin below the empennage. The
basic mission of the aircraft is radio reconnaissance.
Cabin entrance is made through a stair-type door
(fig. 2-2) on the left side of the fuselage.
2-3. DIMENSIONS.
Overall aircraft dimensions are shown in figure
2-3.
2-4. GROUND TURNING RADIUS.
Minimum ground turning radius of the aircraft
is shown in figure 2-4.
2-5. MAXIMUM WEIGHTS.
Maximum takeoff gross weight is 15,000
pounds. Maximum landing weight is 15,000 pounds.
Maximum ramp weight is 15,090 pounds. Maximum zero fuel weight is 11,500 pounds.
2-7. LANDING GEAR SYSTEM.
The landing gear is a retractable, tricycle type,
electrically operated by a single DC motor. This
motor drives the main landing gear actuators
through a gear box and torque tube arrangement,
and also drives a chain mechanism which controls
the position of the nose gear. Positive down-locks
are installed to hold the drag brace in the extended
and locked position. The down-locks are actuated by
overtravel of the linear jackscrews and are held in
position by a spring-loaded overcenter mechanism.
The jackscrew in each actuator holds all three gears
in the UP position, when the gear is retracted. A
friction clutch between the gearbox and the torque
shafts protects the motor from electrical overload in
the event of a mechanical malfunction. A 150ampere current limiter, located on the DC distribution bus under the center floorboard, protects
against electrical overload. Gear doors are opened
and closed through a mechanical linkage connected
to the landing gear. The nose wheel steering mechanism is automatically centered and the rudder pedals relieved of the steering load when the landing
gear is retracted. Air-oil type shock struts, filled with
compressed air and hydraulic fluid, are incorporated
with the landing gear. Gear retraction or extension
time is approximately six seconds.
(1.) Landing Gear Control Switch. Landing gear system operation is controlled by a manually actuated, wheel-shaped switch placarded LDG
GEAR CONTR - UP - DN, located on the left subpanel (fig. 2-6). The control switch and associated
relay circuits are protected by a S-ampere circuit
breaker, placarded LANDING GEAR RELAY on
the overhead circuit breaker panel (fig. 2-26).
2-1
TM 55-1510-221-10
1. Weather radar
2. Air conditioner condenser air outlet
3. Nose avionics compartment access door
4. Windshield wipers
5. AN/APR39 spiral antenna
6. VHF/AM/FM antenna
7. Global positioning system antenna
8. Low band vert bent blade (upper) antenna
9. Transponder antenna
10. VHF comm antenna
11. “P” band antenna
12. Static air source
13. Relief tube drain
14. ELT control switch
15. UHF L-band antenna
16. Emergency light
17. Cargo door
18. Cabin door
19. Strobe light
20. Low band vert bent blade (lower) antenna
21. Navigation light
22. Ice light
23. Outside air temperature gage probe
24. Wide band data link fwd antenna
25. Glideslope antenna
Figure 2-1. General Exterior Arrangement (Sheet 1 of 5)
2-2
AP 011758
TM 55-1510-221-10
26. Navigation light
27. Strobe beacon
28. Dorsal fin (ADF sense) antenna
29. AN/APR-44 antenna
30. Flare dispenser
31. Mid-band dipole antenna
32. Emergency entrance/exit hatch
33. TACAN antenna
34. Nose avionics compartment access door
35. Air conditioner condenser air inlet
36. Radome
37. High band vert & horiz antenna
38. Ice light
39. AN/APR-39 blade antenna
40. High band monopole antenna
41. Static air source
42. Oxygen system servicing door
43. “P” band antenna
44. Wide band data link aft antenna
45. Mid-band dipole antenna
46. Static wick
47. Aft rotating boom antenna
AP 011758.1)
Figure 2-1. General Exterior Arrangement (Sheet 2 of 5)
2-3
TM 55-1510-221-10
1. AN/APR-39 spiral antenna
2. UHF comm & intercept antenna
3. Mid-band dipole antenna
4. Tailet
5. Flare dispenser
6. Low band horiz towel bar antenna
7. AN/APR-44 antenna
8. VOR NAV/LOC antenna
9. Strobe antenna
10. Navigation light
11. ELINT & DF antenna pod
12. Recognition light
13. Stall warning vane
14. High band monopole antenna
15. Bleed air heat exchanger air inlet
16. Pitot tube
17. Wide band data link fwd antenna
18. Taxi light
19. Landing lights
Figure 2-7. General Exterior Arrangement (Sheet 3 of 5)
2-4
AP 011758.2
TM 55-1510-221-10
1. UHF comm & intercept antenna
2. AN/APR-39 spiral antenna
3. Navigation light
4. ELINT & DF antenna pod
5. Fuel filler cap
6. Mid-band dipole antenna
7. AN/APR-39 antenna
8. Marker beacon antenna
9. Wide band data link antenna
10. High band vert & horiz antenna
11. TACAN antenna
12. GPS antenna
13. Low band vert bent blade antenna (upper)
14. High band monopole antenna
15. Strobe beacon
16. Low band horiz towel bar antenna
17. UHF L-band antenna
18. VHF comm antenna
19. Transponder antenna
AP 011758.3
Figure 2-1. General Exterior Arrangement (Sheet 4 of 5)
2-5
TM 55-1510-221-10
1. Radio altimeter antenna
2. VHF/AM/FM antenna
3. Glideslope antenna
4. ADF loop antenna
5. Fuel filler cap
6. Mid-band dipole antenna
7. AN/APR-39 spiral antenna
8. ELINT & DF antenna pod
9. Navigation light
10. UHF comm & Intercept antenna
11. High band monopole antenna
12. INS TACAN antenna
13. AN/APR-44 antenna
14. ELT antenna
15. Static wick
16. Strobe beacon
17. Dorsal fin (ADF sense antenna)
18. Strobe dams
Figure 2-1. General Exterior Arrangement (Sheet 5 of 5)
2-6
AP 011758.4
TM 55-1510-221-10
Figure 2-2. General Interior Arrangement
2-7
TM 55-1510-221-10
Figure 2-3. Principal Dimensions
2-8
Figure 2-4. Ground Turning Radius
2-9
TM 55-1510-221-10
Figure 2-5. Exhaust and Propeller Danger Areas
2-10
TM 55-1510-221-10
CAUTION/ADVISORYANNUNCIATOR PANEL
Figure 2-6. Subpanels
2-11
TM 55-1510-221-10
a. Landing Gear Down Position-Indicator
Lights. Landing gear down position is indicated by
three green lights on the left subpanel, placarded
GEAR DOWN. These lights may be checked by
operating the ANNUNCIATOR TEST switch. The
circuit is protected by a 5-ampere circuit breaker,
placarded LANDING GEAR IND, on the overhead
circuit breaker panel (fig. 2-26).
b. Landing Gear Position Warning Lights.
Two red bulbs, wired in parallel and activated by
microswitches independent of the GEAR DOWN
position indicator lights, are positioned inside the
clear plastic grip on the landing gear control switch.
These lights illuminate whenever the landing gear
switch is in either the UP or DN position and the
gear is in transit. Both bulbs will also illuminate
should either or both power levers be retarded below
approximately 79 to 81% N1 when the landing gear
is not down and locked. To turn the switch lights
OFF during single-engine operation, the power lever
for the inoperative engine must be advanced to a
position which is higher than the setting of the warning horn microswitch. Extending the landing gear
will also turn the lights off. Both red lights indicate
the same warning conditions, but two are provided
for a fail-safe indication in the event one bulb burns
out. The circuit is protected by a 5-ampere circuit
breaker, placarded LANDING GEAR IND, on the
overhead circuit breaker panel (fig. 2-26).
c. Landing Gear Warning Light Test Button.
A test button, placarded HDL LT TEST, is located
on the left subpanel. Failure of the landing gear
switch to illuminate red, when this test button is
pressed, indicates two defective bulbs or a circuit
fault. The circuit is protected by a 5-ampere circuit
breaker, placarded LANDING GEAR RELAY
CONTROL, on the overhead circuit breaker panel
(fig. 2-26).
d. Landing Gear Warning Horn. When either
power lever is retarded below approximately 79 to
81 N, when the landing gear is not down and locked
or if the flaps are extended beyond 40% and the
landing gear is not down and locked, a warning horn
located in the overhead control panel will sound
intermittently. To prevent the warning horn from
sounding during long descents or an ILS approach,
a pressure differential “Q” switch is connected into
the copilot’s static line. The switch prevents the
warning horn from sounding until airspeed drops
below 140 KIAS. An altitude sensing switch is
installed in series with the 140 KIAS “Q“ switch
2-12
which prevents the warning horn from sounding
after climbing through 12,500 feet MSL. The horn
will be engaged when the aircraft descends through
10,500 MSL. The warning horn circuit is protected
by a 5-ampere circuit breaker, placarded LANDING
GEAR WARN, on the overhead circuit breaker
panel (fig. 2-26).
e. Landing Gear Warning Horn Test Switch.
The landing gear warning horn may be tested by the
test switch on the right subpanel (fig. 2-6). The
switch, placarded STALL WARN TEST - OFF LDG GEAR WARN TEST, will sound the landing
gear warning horn and illuminate the landing gear
position warning lights when moved to the momentary LDG GEAR WARN TEST position. The circuit
is protected by a 5-ampere circuit breaker, placarded
LANDING GEAR WARN, on the overhead circuit
breaker panel (fig. 2-26).
f. Landing Gear Safety Switches. A safety
switch on each main landing gear shock strut controls the operation of various aircraft systems that
function only during flight or only during ground
operation. These switches are mechanically actuated
whenever the main landing gear shock struts are
extended (normally after takeoff), or compressed
(normally after landing). The safety switch on the
right main landing gear strut activates the landing
gear control circuits, cabin pressurization circuits
and the flight hour meter flight time function on the
copilot’s clock when the strut is extended. This
switch also activates a down-lock hook, preventing
the landing gear from being raised while the aircraft
is on the ground. The hook, which unlocks automatically after takeoff, can be manually overridden by
pressing down on the red button, placarded DN
LOCK REL located adjacent to the landing gear
switch. If the override is used and the landing gear
control switch is raised, power will be supplied to
the warning horn circuit and the horn will sound.
The safety switch on the left main landing gear strut
activates the left and right engine ambient air shutoff valves when the strut is extended.
g. Landing Gear Alternate Engage Handle.
During 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. 2-7)
located adjacent to the landing gear alternate extension handle (fig. 2-7). To disengage the clutch, pull
the alternate engage handle up and turn clockwise.
To engage the clutch, turn the alternate engage handle counterclockwise and release.
TM 55-1510-221-10
1. Power levers
2. Propeller levers
3. Condition levers
4. FLAP cover
5. Rudder trim
6. Flare dispenser control panel
7. Flare dispenser switch
8. TACAN Control panel
9. NAV 1/NAV 2 control panel
10. Elev trim & rudder boost switches
11. Marker beacon audio control panel
12. Mode selector unit
13. ADF control panel
14. VHF AM/FM control panel
15. HF command set control panel (KHF-950)
16. No. 2 UHF command set
17. Transponder control panel
18. VHF AM control panel
19. No. 1 UHF command set
20. Flight director mode select panel (MS 500)
21. Autopilot pitch and turn control
22. No. 1 HSI course and heading knobs
23. Emergency Ianding gear alternate
engage handle
24. Emergency landing gear extension
ratchet handle
25. Aileron tab control and position indicator
26. Elevator tab control and position Indicator
27. Go-around button
AP 011769
Figure 2-7. Control Pedestal
2-13
TM 55-1510-221-10
Continued pumping of handle after
GEAR DOWN position indicator lights
(3) are illuminated could damage the
drive mechanism, and prevent subsequent
gear retraction.
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.
h. Landing Gear Alternate Extension Handle.
Manual landing gear extension is provided through
a manually powered system as a backup to the electrically operated system. Before manually extending
the gear, make certain that the landing gear switch
is in the down position with the LANDING GEAR
RELAY circuit breaker pulled. Pulling up on the
alternate engage handle, located on the floor, and
turning it clockwise will lock it in that position.
When the alternate engage handle is pulled, the
motor is electrically disconnected from the system
and the alternate drive system is locked to the gearbox and motor. When the alternate drive is locked
in, the chain is driven by a continuous action ratchet
which is activated by pumping the landing gear
alternate extension handle adjacent to the alternate
engage handle.
k. Wheel Brake System. The main wheels are
equipped with multiple-disc hydraulic brakes actuated by master cylinders attached to the rudder pedals at the pilot’s and copilot’s position. Braking is
permitted from either set of rudder pedals. Brake
fluid is supplied to the system from the reservoir in
the nose compartment. The toe brake sections of the
rudder pedals are connected to the master cylinders
which actuate the system for the corresponding
wheels. No emergency brake system is provided.
Repeated and excessive application of brakes, without allowing sufficient time for cooling to accumulate between applications, will cause a loss of braking 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.
(1.) After a manual landing gear extension
has been made, do not stow the handle, move any
landing gear controls, or reset any switches or circuit
breakers. The gear cannot be retracted manually.
2-8. PARKING BRAKE.
(2.) After a practice manual extension, the
alternate handle may be stowed and the landing gear
retracted electrically. Rotate the alternate engage
handle counterclockwise and push it down. Stow the
handle, push in the LANDING GEAR RELAY circuit breaker on the overhead circuit breaker panel
and retract the gear in the normal manner with the
landing gear switch. Refer to Chapter 9 for emergency gear extension procedures.
i. Tires. The aircraft is equipped with dual 22
x 6.75 - 10 tubeless, 8-ply rated, rim-inflation tires
on each main gear. The nose gear is equipped with
a single 22 x 6.75 - 10, 8-ply rated, tubeless, riminflated tire.
j. Steerable Nose Wheel. The aircraft can be
maneuvered on the ground by the steerable nose
wheel system. Direct linkage from the rudder pedals
(fig. 2-8) to the nose wheel steering linkage allows
the nose wheel to be turned 12° to the left of center
2-14
Dual parking brake valves are installed below
the cockpit floor. Both valves can be closed simultaneously by pressing both brake pedals to build up
pressure, then pulling out the handle placarded
PARKING BRAKE, on the left subpanel. Pulling
the handle full out sets the check valves in the system and any pressure being applied by the toe
brakes is maintained. Parking brakes are released
when the brake handle is pushed in. The parking
brake may be set from either cockpit position. Parking brakes shall not be set during flight.
2-9. ENTRANCE AND EXIT PROVISIONS.
NOTE
Two keys are provided in the loose tools
and equipment bag. Both keys will fit the
locks on the cabin door, emergency hatch,
tailcone access door and the right and left
nose avionics compartment doors.
TM 55-1510-221-10
1. Free air temperature gage
2. Oxygen system pressure gages
3. Storm window lock
4. Oxygen regulator control panel
5. Control wheel
6. Sun visor
7. Overhead circuit breaker and control panel
8. Windshield wiper
9. Magnetic compass
10. Rudder pedals
11. Mission control panel
12. Pedestal extension
13. Assist step
AP 012087
Figure 2-8. Cockpit
2-15
TM 55-1510-221-10
a.
Cabin Door.
Structural damage may be caused if more
than one person is on the cabin door at a
time. The door is weight limited to 300
pounds or less.
A swing-down door (fig. 2-9), hinged at the bottom, provides a stairway for normal and emergency
entrance and exit. Two of the steps are movable and
fold flat against the door in the closed position. A
step folds down over the door sill when the door
opens to provide a platform (step) for door seal protection. A plastic encased cable provides support for
the door in the open position, a handhold, and a
convenience for closing the door from inside. A
hydraulic damper permits the door to lower gradually during opening. A rubber seal around the door
seals the pressure vessel while the aircraft is in
flight. The door locking mechanism is operated by
either of the two mechanically interconnected handles, one inside and the other outside the door.
When either handle is rotated, three rotating camtype latches on either side of the door capture posts
mounted on the cargo door. In the closed position,
the door becomes an integral part of the cargo door.
A button adjacent to the door handle must be
depressed before the handle can be rotated to open
the door. A bellows behind the button is inflated
when the aircraft is pressurized to prevent accidental unlatching and/or opening of the door. A small
round window just above the second step permits
observation of the pressurization safety bellows. A
placard adjacent to the window instructs the operator that the safety lock arm is in position around the
bellows shaft which indicates a properly locked
door. Pushing the red button adjacent to the window
will illuminate the inside door mechanism. A
CABIN DOOR annunciator light on the caution/
advisory panel will illuminate if the door is not
closed and all latches fully locked. The cabin door
opening is 21.5 inches wide by 50.0 inches high.
b. Cargo Door. A swing-up door (fig. 2-9)
hinged at the top, provides cabin access for loading
cargo or bulky items. After initial opening force is
applied, gas springs will completely open the cargo
door automatically. The door is counterbalanced
and will remain in the open position. A door support rod is used to hold the door in the open position, and to aid in overcoming the pressure of the
gas spring assemblies when closing the door. Once
closed, the gas springs apply a closing force to assist
in latching the door. A rubber seal around the door
seals the pressure vessel while in flight. The door
2-16
locking mechanism is operated only from inside the
aircraft, and is operated by two handles, one in the
bottom forward portion of the door and the other in
the upper aft portion of the door. When the upper
aft handle is operated per placard instructions, two
rotating cam-type latches on the forward side of the
door and two on the aft side rotate, capturing posts
mounted on the fuselage side of the door opening.
The bottom handle, when operated per placard
instructions, actuates four pin lug latches across the
bottom of the door. A button on the upper aft handle must be pressed before the handle can be
released to open or latch the door. A latching lever
on the bottom handle must be lifted to release the
handle before the lower latches can be opened.
These act as additional aids in preventing accidental
opening or unlatching of the door. The cabin and
cargo doors are equipped with dual sensing circuits
to provide the crew remote indication of cabin/cargo
door security. An annunciator light placarded
CABIN DOOR will illuminate if the cabin or cargo
door is open and the BATT switch in ON. If the battery switch is OFF, the annunciator will illuminate
only if the cargo door is not securely closed and
latched. The cargo door sensing circuit receives
power from the hot battery bus. The cargo door
opening is 52.0 inches wide by 52.0 inches high.
Insure that the cabin door is closed and
locked. Operating the cargo door while
the cabin door is open may damage the
door hinge and adjacent structure.
(1.) Opening cargo door.
Avoid side loading of the gas springs to
prevent damage to the mechanism.
1.
Handle access door (lower forward corner of door) -Unfasten
and open.
2.
Handle - Lift hook and move to
OPEN position.
3.
Handle access door - Secure.
4.
Handle access door (upper aft corner of door) - Unfasten and open.
5.
Handle - Press button and lift to
OPEN position then latch in
place.
TM 55-1510-221-10
Figure 2-9. Cabin and Cargo Doors
2-17
TM 55-1510-221-10
6.
Handle access door - Secure.
7.
Door support rod - Attach one
end to cargo door ball stud (on
forward side of door).
8.
Support rod detent pin - Check in
place.
9.
Cabin door sill step - Push out on
and allow cargo door to swing
open. Gas springs will automatically open the door.
10.
Door support rod - Attach free
end to ball stud on forward fuselage door frame.
(2.) Closing cargo door.
Avoid side loading of the gas springs to
prevent damage to the mechanism.
1.
Door support rod - Detach from
fuselage door frame ball stud,
then firmly grasp free end of rod
while exerting downward force to
overcome the pressure of gas
spring assemblies. Then remove
support rod from door as gas
spring assemblies pass over-center
position.
d. Cabin Door Caution Light. As a safety precaution, two illuminated MASTER CAUTION
lights, on the glare shield and a steadily illuminated
CABIN DOOR yellow caution annunciator light on
the caution/advisory panel indicate the cabin door is
not closed and locked. This circuit is protected by
5-ampere circuit breakers placarded ANN PWR and
ANN IND, located on the overhead circuit breaker
panel (fig. 2-26).
2-10. WINDOWS.
a. Cockpit Window. The pilot and copilot
have side windows, a windshield and storm windows, which provide visibility from the cockpit. The
storm windows may be opened on the ground or
during unpressurized flight.
2.
Cargo door - Pull closed, using
finger hold cavity in fixed cabin
door step.
3.
Handle access door (upper aft corner of door) - Unfasten and open.
4.
Handle - Press button and pull
handle down until it latches in
closed position.
5.
Handle access door - Secure.
b. Cabin Windows. The outer cabin windows
have two-ply construction, are of the pressure type
and are integral parts of the pressure vessel. All
cabin windows are painted over except for the window farthest aft on the right side and the window
farthest aft on the left side. All unpainted windows
have flaps which may be raised to permit visibility
or lowered to black out the windows.
6.
Handle access door (lower forward comer of door) -Unfasten
and open.
2-11. SEATS.
7.
Handle - Move to full forward
position.
8.
Safety hook - Check locked in
position by pulling aft on handle.
9.
Handle access door - Secure.
c. Cabin Emergency Hatch. The cabin emergency hatch, placarded EXIT - PULL, is located on
the right cabin sidewall just aft of the copilot’s seat.
2-16
The hatch may be released, from the inside with a
pull-down handle. A flush mounted pull out handle
allows the hatch to be released from the outside. The
hatch is of the non-hinged plug type, which removes
completely from the frame when the latches are
released. The hatch can be key locked from the
inside, to prevent opening from the outside. The
inside handle will unlatch the hatch whether or not
it is locked, by overriding the locking mechanism.
The keylock should be unlocked prior to flight to
allow removal of the hatch from the outside in the
event of an emergency. The key remains in the lock
when the hatch is locked and can be removed only
when the hatch is unlocked. The key slot is in the
vertical position when the hatch is unlocked.
Removal of the key from the lock before flight
assures the pilot that the hatch can be removed from
the outside if necessary.
a. Pilot and Copilot Seats. The pilot and copilot seats (fig. 2-10) are separated from the cabin by
movable curtains. The controls for vertical height
adjustment and fore and aft travel are located under
each seat. The forward and aft adjustment handle is
located beneath the bottom front inboard corner of
each seat. Pulling up on the handle allows the seat
to move fore or aft. The height adjustment handle is
located beneath the bottom front outboard corner of
each seat. Pulling up on the handle, allows the seat
TM 55-1510-221-10
1. Adjustable headrest
2. Seatbelt/shoulder harness buckle
3. Moveable armrest
4. Seat height adjustment (pilot), fore
and aft adjustment (copilot)
5. Seat fore and aft adjustment (pilot),
height adjustment (copilot)
6. Expandable map pocket
AP 004766
Figure 2-10. Pilot and Copilot Seats
to move up and down. Both seats have adjustable
headrests and armrests which will raise and lower
for access to the cockpit. Handholds on either side
of the overhead panels and a fold-away protective
pedestal step are provided for pilot and copilot entry
into the cockpit. For the storage of maps and the
operator’s manual, pilot and copilot seats have an
inboard-slanted, expandable pocket affixed to the
lower portion of the seat back. Pocket openings are
held closed by shock cord tension.
b. Pilot and Copilot Seat Belts and Shoulder
Harnesses. Each pilot and copilot seat is equipped
Section II.
with a lap-type seat belt and shoulder harness connected to an inertia reel. The shoulder harness belt
is of the “Y” configuration with the single strap
being contained in an inertia reel attached to the
base of the seatback. The two straps are worn with
one strap over each shoulder and fastened by metal
loops into the seat belt buckle. The spring loading at
the inertia reel keeps the harness snug but will allow
normal movement required during flight operations.
The inertia reel is designed with a locking device
that will secure the harness in the event of sudden
forward movement or an impact action.
EMERGENCY EQUIPMENT
2-12. DESCRIPTION.
The equipment covered in this section includes
all emergency equipment, except that which forms
part of a complete system. For example, landing
gear system, etc. Chapter 9 describes the operation
of emergency exits and location of all emergency
equipment.
2-13. FIRST AID KITS.
Four first aid kits are included in the survival
kit.
2-19
TM 55-1510-221-10
2-14. HAND-OPERATED FIRE EXTINGUISHER.
Repeated or prolonged exposure to high
concentrations of monobromotrifluoromethane (CF 3 Br) or decomposition
products should be avoided. The liquid
shall not be allowed to come into contact
with the skin, as it may cause frost bite or
low temperature burns because of its very
low boiling point.
One hand-operated fire extinguisher is mounted
below the pilot’s seat and a second extinguisher is
located on the left cabin sidewall, aft of the cabin
Section III.
NOTE
Engine tire extinguisher systems are
described in Section III.
2-15. SURVIVAL KITS.
Tie-down provisions for a survival raft and kit
are provided just forward of the toilet on the right
hand side of the cabin (fig. 2-2).
ENGINES AND RELATED SYSTEMS
2-16. ENGINES.
The aircraft is powered by two PT6A-41 turboprop engines (fig. 2-11). The engine has a three stage
axial, single stage centrifugal compressor, driven by
a single stage reaction turbine. The power turbine, a
two stage reaction turbine, counter-rotating with the
compressor turbine, drives the output shaft. Both
the compressor turbine and the power turbine are
located in the approximate center of the engine with
their shafts extending in opposite directions. Being
a reverse flow engine, the ram air supply enters the
lower portion of the nacelle and is drawn in through
the aft protective screens. The air is then routed into
the compressor. After it is compressed, it is forced
into the annular combustion chamber, and mixed
with fuel that is sprayed in through 14 nozzles
mounted around the gas generator case. A capacitance discharge ignition unit and two spark igniter
plugs are used to start combustion. After combustion, the exhaust passes through the compressor turbine and two stages of power turbine then is routed
through two exhaust ports near the front of the
engine. A pneumatic fuel control system schedules
fuel flow to maintain the power set by the gas generator power lever. The accessory drive at the aft end
of the engine provides power to drive the fuel
pumps, fuel control, the oil pumps, the refrigerant
compressor (right engine), the starter-generator, and
2-20
door. They are of the monobromotrifluromethane
(CF 3Br) 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.
the turbine tachometer transmitter. The reduction
gearbox forward of the power turbine provides gearing for the propeller and drives the propeller
tachometer transmitter, the propeller overspeed governor, and the propeller governor.
2-17. ENGINE COMPARTMENT COOLING.
The forward engine compartment including the
accessory section is cooled by air entering around
the exhaust stack cutouts, the gap between the propeller spinner and forward cowling, and exhausting
through ducts in the upper and lower aft cowling.
2-18. AIR INDUCTION SYSTEMS - GENERAL.
Each engine and oil cooler receives ram air ducted from an air scoop located within the lower section of the forward nacelle. Special components of
the engine induction system protect the power plant
from icing and foreign object damage.
2-19. FOREIGN OBJECT DAMAGE CONTROL.
The engine has an integral air inlet screen
designed to obstruct objects large enough to damage
the compressor.
TM 55-1510-221-10
1. Primary prop governor
2. Torque pressure transmitter
3. Torque pressure switch
4. Torque pressure manifold
5. Exhaust duct
6. TGT temperature probe
7. Fuel flow divider manifold
8. Fire detector
9. Engine mount bolt
10. Engine mount truss assembly
11. Engine air intake screen
12. Ignition exciter
13. Starter-generator
14. Fuel boost pump
15. Air conditioner compressor drive
belt (#2 engine only)
16. Fire detector
17. Air conditioner compressor
(#2 engine only)
18. Bleed air adapter
19. Bleed air line
20. Engine mount
21. Ignition exciter plug
22. Oil scavenge tubes
23. Overspeed governor
24. Prop deice brush block bracket
25. Prop reverse linkage lever
AP 005485.1
Figure 2-11. PT6A-41 Engine (Sheet 1 of 2)
2-21
TM 55-1510-221-10
26. Fuel control unit
27. Fuel control unit control rod
28. Starter generator leads
29. Engine driven fuel pump
30. Power control lever
31. Prop interconnect linkage (aft)
32. Oil pressure transducer
33. Engine mount
34. Fireshield
35. Trim resistor thermocouple
36. Prop interconnect linkage (fore)
37. Prop shaft
38. Tach generator
39. Chip detector
40. Oil pressure tube
41. Fire extinguisher line
42. Ignition exciter plug
43. Engine mount bolt
44. Linear actuator
45. Engine baffle and seal assy
46. Fuel/oil heater
47. Tach-generator (aft)
48. Drain manifold
49. Overhead breather tube
60. Engine truss mounting bolt
AP 005485.2
Figure 2-11. PT6A-41 Engine (Sheet 2 of 2)
2-22
TM 55-1510-221-10
2-20. ENGINE ICE PROTECTION SYSTEMS.
a. Inertial Separator.
After the ice vanes have been manually
extended, they may be mechanically actuated only. No electrical extension or
retraction shall be attempted as damage
to the actuator may result. Linkage in the
nacelle area must be reset prior to operation of the electric system.
An inertial separation system is built into each
engine air inlet to prevent moisture particles from
entering the engine inlet plenum under icing conditions. A movable vane and a bypass door are lowered into the airstream when operating in visible
moisture at 5°C or colder, by energizing electrical
actuators with the switches, placarded ICE VANE RETRACT - EXTEND, located on the overhead
control panel. A mechanical backup system is provided, and is actuated by pulling the T-handles just
below the pilot’s subpanel placarded ICE VANE No. 1 ENG - No. 2 ENG. Decrease airspeed to 160
knots or less to reduce forces for manual extension.
Normal airspeed may then be resumed.
(1.) The vane deflects the ram airstream
slightly downward to introduce a sudden turn in the
airstream to the engine, causing the moisture particles to continue on undeflected, because of their
greater momentum, and to be discharged overboard.
(2.) While in the icing flight mode, the
extended position of the vane and bypass door is
indicated by green annunciator lights, No. 1 VANE
EXT and No. 2 VANE EXT.
(3.) In the non-ice protection mode, the
vane and bypass door are retracted out of the airstream by placing the ice vane switches in the
RETRACT position. The green annunciator lights
will extinguish. To assure adequate oil cooling,
retraction should be accomplished at 15°C and
above. The vanes should be either extended or
retracted; there are no intermediate positions.
(4.) If for any reason the vane does not
attain the selected position within 15 seconds, a yellow No. 1 VANE FAIL or No. 2 VANE FAIL light
illuminates on the caution/advisory panel. In this
event, the manual backup system should be used.
When the vane is successfully positioned with the
manual system, the yellow annunciator lights will
extinguish. During manual system use, the electric
motor switch position must match the manual handle position for a correct annunciator readout.
b. Engine Air Inlet Deice System.
(1.) Description. Hot engine exhaust gas is
utilized for heating the air inlet lips to prevent the
formation of ice. Hot exhaust gas is picked up inside
each engine exhaust stack and carried by plumbing
to the inlet lip. The gas flows through the inside of
the lip to the bottom where it is allowed to escape.
(2.) Fuel heater. An oil-to-fuel heat
exchanger, located on the engine accessory case,
operates continuously and automatically to heat the
fuel sufficiently to prevent ice from collecting in the
fuel control unit. Each fuel control unit is protected
against ice. Fuel control heat is automatically turned
on for all engine operations.
2-21. ENGINE FUEL CONTROL SYSTEM.
a. Description. The basic engine fuel system
consists of an engine driven fuel pump, a fuel control unit, a fuel flow divider, a dual fuel manifold
and fourteen fuel nozzles. The fuel flow divider acts
as a drain valve to clear residual fuel after engine
shutdown.
b. Fuel Control Unit. One fuel control unit is
mounted on the accessory case of each engine. This
unit is a hydro-pneumatic 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. Engine shutdown is accomplished by moving
the appropriate condition lever to the full aft FUEL
CUTOFF position, which shuts off the fuel supply.
2-22. POWER LEVERS.
Moving the power levers into reverse
range without the engines running may
result in damage to the reverse linkage
mechanism.
Two power levers are located on the control
pedestal (fig. 2-7). 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
N 1 speed governor in the fuel control unit. Power is
increased when N1 RPM is increased. The power
2-23
TM 55-1510-221-10
levers also control propeller reverse pitch. Distinct
movement (pulling up and then aft on the power
lever) by the pilot is required for reverse thrust.
Placarding beside the lever travel slots reads
POWER. Upper lever travel range is designated
INCR (increase), supplemented by an arrow pointing forward. Lower travel range is marked IDLE,
LIFT and REVERSE. A placard below the lever
slots reads: CAUTION REVERSE ONLY WITH
ENGINES RUNNING.
2-23. CONDITION LEVERS.
Two condition levers are located on the control
pedestal (fig. 2-7). Each lever starts and stops the
fuel supply, and controls the idle speed for its
engine. The levers have three placarded positions:
FUEL CUTOFF, LO IDLE, and HIGH IDLE. In
the FUEL CUTOFF position, the condition lever
controls the cutoff function of its engine-mounted
fuel control unit. From LO IDLE to HIGH IDLE,
they control the governors of the fuel control units
to establish minimum fuel flow levels. LO IDLE
position sets the fuel flow rate to attain 52 to 55%
(at sea level) minimum N1 and HIGH IDLE position sets the rate to attain 70% minimum N1. The
power lever for the corresponding engine can select
N 1 from the respective idle setting to maximum
power. An increase in low idle N1 will be experienced at high field elevation.
2-24. FRICTION LOCK KNOBS.
Four friction lock knobs (fig. 2-7) are located on
the control pedestal to adjust friction drag. 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.
2-25. ENGINE FIRE DETECTION SYSTEM.
a. Description. A flame surveillance system is
installed on each engine to detect external engine
fire and provide alarm to the pilot. Both nacelles are
monitored, each having a control amplifier and
three detectors. Electrical wiring connects all sensors
and control amplifiers to DC power and to the cockpit visual alarm units. In each nacelle, one detector
monitors the forward nacelle, a second monitors the
upper accessory area, and a third the lower accessory
area. Fire emits an infrared radiation that will be
2-24
sensed by the detector which monitors the area of
origin. Radiation exposure activates the relay circuit
of a control amplifier which causes signal power to
be sent to cockpit warning systems. An activated
surveillance system will return to the standby state
after the fire is out. The system includes a functional
test switch and has circuit protection through the
FIRE DETR circuit breaker. Warning of internal
nacelle fire is provided as follows: the red MASTER
WARNING lights on the glareshield illuminate
accompanied by the illumination of a red warning
light in the appropriate fire control T-handle placarded No. 1 FIRE PULL or No. 2 FIRE PULL (fig.
2-29). Fire detector circuits are protected by a single
5-ampere circuit breaker, placarded FIRE DETR,
located on the overhead circuit breaker panel (fig.
2-26).
b. Fire Detection System Test Switch. O n e
rotary switch placarded FIRE PROTECTION TEST
on the copilot’s subpanel is provided to test the
engine fire detection system. Before checkout, battery power must be on and the FIRE DETR circuit
breaker must be in. Switch position DETR 1, checks
the area forward of the air intake of each nacelle,
including circuits to the cockpit alarm and indication devices. Switch position DETR 2, checks the
circuits for the upper accessory compartment of
each nacelle. Switch position DETR 3, checks the
circuits for the lower accessory compartment of each
nacelle. Each numbered switch position will initiate
the cockpit indications previously described.
c. Erroneous Fire Detection System Indications. During ground test of the engine fire detection
system, an erroneous indication of system fault 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, change
the aircraft heading to enable a valid system check.
2-26. ENGINE FIRE EXTINGUISHER SYSTEM.
a. Description. The fire extinguisher system
utilizes an explosive squib, connected to a valve
which, when opened, allows the distribution of the
pressurized extinguishing agent through a plumbing
network of spray nozzles strategically located in the
fire zones of the engines.
b. Fire Pull Handles. The tire control
T-handles, which are used to arm the extinguisher
system are centrally located on the pilot’s instrument panel (fig. 2-29), immediately below the
glareshield. These controls receive power from the
hot battery bus. The fire detection system will indicate an engine fire by illuminating the master fault
warning light on the pilot’s and copilot’s glareshield
and the respective No. 1 FIRE PULL or No. 2 FIRE
TM 55-1510-221-10
PULL lights in the fire control T-handles. Pulling
the fire control T-handle will electrically arm the
extinguisher system and close the fuel firewall shutoff valve for that particular engine. This will cause
the red light in the PUSH TO EXTINGUISH switch
and the respective red No. 1 FUEL PRESS or No. 2
FUEL PRESS light on the warning annunciator
panel to illuminate. Pressing the lens of the PUSH
TO EXTINGUISH switch (after lifting one side of
its spring-loaded clear plastic guard) will fire the
squib, expelling all the agent in the cylinder at one
time. The respective yellow caution light, No. 1
EXTGH DISCH or No. 2 EXTGH DISCH on the
caution/advisory annunciator panel and the master
fault caution lights on the glareshield will illuminate
and remain illuminated, regardless of the master
switch position, until the squib is replaced.
c. Fire Extinguisher System Test Switch. A
rotary test switch, placarded FIRE PROTECTION
TEST, is located on the copilot’s subpanel. The test
functions, placarded EXTGH - No. 1 - No. 2, are
arranged on the left side of the switch and provide
a test of the pyrotechnic cartridge circuitry. During
preflight, the pilot should rotate the test switch
through the two positions and verify the illumination of the green SQUIB OK light on the PUSH TO
EXTINGUISH switch and the corresponding yellow
No. 1 or No. 2 EXTGH DISCH light on the caution/
advisory annunciator panel.
d. Fire Extinguishing System Supply Cylinder
Gages. A gage, calibrated in PSI, is mounted on each
supply cylinder for determining the level of charge
and should be checked during preflight (Table 2-1).
2-27. OIL SUPPLY SYSTEM.
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 2.3 U.S. gallons. The total system
capacity for each engine, which includes the oil tank,
oil cooler, lines, etc., is 3.5 U.S. gallons. The oil
level is indicated by a dipstick attached to the oil
filler cap. Oil grade, specification and servicing
points, are described in Section IX, Servicing the
warning annun.
b. The oil system of each engine is coupled to
a heat exchanger unit (radiator) of tin-and-tube
design. These exchanger units are the only airframe
mounted part of the oil system and are attached to
the nacelles below the engine air intake. Each heat
exchanger incorporates a thermal bypass which
assists in maintaining oil at the proper temperature
range for engine operation.
2-28. ENGINE CHIP DETECTION SYSTEM.
A magnetic chip detector is installed in the bottom of each engine nose gearbox to warn the pilot of
oil contamination and possible engine failure. The
sensor is an electrically insulated gap immersed in
the oil, functioning as a normally-open switch. If a
large metal chip or a mass of small particles bridge
the detector gap, a circuit is completed, sending a
signal to illuminate a red annunciator panel indicator light placarded No. 1 CHIP DETR or No. 2 CHIP
DETR and the MASTER WARNING lights. Chip
detector circuits are protected by two 5-ampere circuit breakers, placarded No. 1 CHIP DETR and
No. 2 CHIP DETR on the overhead circuit breaker
panel (fig. 2-26).
2-29. ENGINE IGNITION SYSTEM.
Maximum allowable oil consumption is
one quart per 10 hours of engine operation.
a. Description. The basic ignition system consists of a solid state ignition exciter unit, two igniter
plugs, two shielded ignition cables, pilot controlled
IGNITION AND ENGINE START switches and
the ENG AUTO IGN switch. Placing an IGNITION
AND ENGINE START switch to ON (forward) will
cause the respective igniter plugs to spark, igniting
the fuel/air mixture sprayed into the combustion
a. The engine oil tank is integral with the airinlet casting located forward of the accessory gearbox. Oil for propeller operation, lubrication of the
reduction gearbox and engine bearings is supplied
Table 2-1. Engine Fire Extinguisher Gage Pressure
TEMP °C
-40
-29
-18
-06
04
16
27
38
48
PSI
190
to
240
220
to
275
250
to
315
290
to
365
340
to
420
390
to
480
455
to
550
525
to
635
605
to
730
2-25
TM 55-1510-221-10
chamber by the fuel nozzles. The ignition system is
activated for ground and air starts, but is switched
off after combustion light up.
b. Ignition and Engine Start Switches. T w o
three-position toggle switches, placarded IGNITION
AND ENGINE START, are located on the overhead
control panel (fig. 2-12). These switches will initiate
starter motoring and ignition in the ON position, or
will motor the engine in the STARTER ONLY position. The ON switch position completes the starter
circuit for engine rotation, energizes the igniter plugs
for fuel combustion, and activates the No. 1 IGN
ON or No. 2 IGN ON light on the annunciator
panel. In the center position the switch is OFF. Two
5-ampere circuit breakers on the overhead circuit
breaker panel, placarded IGNITOR CONTR No. 1
and No. 2, protect ignition circuits. Two 5-ampere
circuit breakers on the overhead circuit breaker
panel, placarded START CONTR No. 1 and No. 2,
protect starter control circuits (fig. 2-26).
2-30. 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 and a green indicator
light placarded No. 1 IGN ON or No. 2 IGN ON, on
the annunciator panel. Autoignition is accomplished
by energizing the two igniter plugs in each engine.
NOTE
The system should be turned OFF during
extended ground operation to prolong the
life of the igniter plugs.
a. Autoignition Switches. T w o s w i t c h e s ,
located on the overhead control panel (fig. 2-12)
placarded ENG AUTO IGN-ARM control the
autoignition systems. The ARM position initiates a
readiness mode for the autoignition system of the
corresponding engine. The OFF position disarms the
system. Each switch is protected by a corresponding
START CONTR No. 1 or No. 2 5-ampere circuit
breaker on the overhead circuit breaker panel (fig.
2-26).
b. Autoignition Lights. If an armed autoignition system changes from a ready condition to an
operating condition (energizing the igniter plugs in
the engine) a corresponding green annunciator panel
light will illuminate. The annunciator panel light is
placarded No. 1 IGN ON or No. 2 IGN ON and indi-
Figure 2-12. Overhead Control Panel
2-26
TM 55-1510-221-10
cates that the igniters are energized. The autoignition system is triggered from a ready condition to an
operating condition when engine torque drops below
approximately 20%. 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 approximately 20% torque. The
autoignition lights are protected by 5-ampere IGNITOR CONTR No. 1 or No. 2 circuit breakers, located
on the overhead circuit breaker panel (fig. 2-26).
b. Engine Torquemeters. Two torquemeters
on the instrument panel indicate torque applied to
the propeller shafts of the respective engines (fig.
2-29). Each gage shows torque in percent of maximum using 2 percent graduations and is actuated by
an electrical signal from a pressure sensing system
located in the respective propeller reduction gear
case. Torquemeters are protected by individual 0.5ampere circuit breakers placarded TORQUE
METER No. 1 or No. 2 on the overhead circuit
breaker panel (fig. 2-26).
2-31. ENGINE STARTER-GENERATORS.
c. Turbine Tachometers. Two tachometers on
the instrument panel register compressor turbine
RPM (N1) for the respective engine (fig. 2-29).
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.
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 the starter
function, 28 volts DC is required to power rotation.
In the generator function, each unit is capable of
400 amperes DC output. When the starting function
is selected, the starter control circuit receives power
through the respective 5-ampere START CONTR
circuit breaker on the overhead circuit breaker panel
from either the aircraft battery or an external power
source. When the generating function is selected, the
starter-generator provides electrical power. For additional description of the starter-generator system,
refer to Section IX.
2-32. ENGINE INSTRUMENTS.
The engine instruments are vertically mounted
near the center of the instrument panel (fig. 2-29).
a. Turbine Gas Temperature Indicators. Two
TGT gages on the instrument panel are calibrated in
degrees Celsius (fig. 2-29). Each gage is connected to
thermocouple probes located in the hot gases
between the turbine wheels. The gages register the
temperature present between the compressor turbine
and power turbine for the corresponding engine.
Section IV.
2-33. FUEL SUPPLY SYSTEM.
The engine fuel supply system (fig. 2-13) consists
of two identical systems sharing a common fuel
management panel (fig. 2-14) and fuel crossfeed
plumbing (fig. 2-15). Each fuel system consists of
five interconnected wing tanks, a nacelle tank, and
an auxiliary inboard fuel tank. A fuel transfer pump
is located within each auxiliary tank. Additionally,
d. Oil Pressure/Oil Temperature Indicators.
Two gages on the instrument panel register oil pressure in PSI and oil temperature in °C (fig. 2-29). Oil
pressure is taken from the delivery side of the main
oil pressure pump. Oil temperature is transmitted by
a thermal sensor unit which senses the temperature
of the oil as it leaves the delivery side of the oil pressure pump. Each gage is connected to pressure transmitters installed on the respective engine. Both
instruments are protected by 5-ampere circuit breakers, placarded OIL PRESS and OIL TEMP No. 1 or
No. 2, on the overhead circuit breaker panel (fig.
2-26).
e. Fuel Flow Indicators. Two gages on the
instrument panel (fig. 2-29) register the rate of flow
for consumed fuel as measured by sensing units coupled into the fuel supply lines of the respective
engines. The fuel flow indicators are calibrated in
increments of hundreds of pounds per hour. Both
circuits are protected by 0.5-ampere circuit breakers
placarded FUEL FLOW No. 1 or No. 2, on the overhead circuit breaker panel (fig. 2-26).
FUEL SYSTEM
the system has an engine-driven boost pump, a
standby fuel pump located within each nacelle 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. Refer to Section IX for fuel grades and specifications. Fuel tank
capacity is shown in table 2-2. Gravity feed fuel flow
is shown in figure 2-16.
2-27
TM 55-1510-221-10
Figure 2-13. Fuel System Schematic
2-28
TM 55-1510-221-10
1. STANDBY PUMP switch (#1 engine)
2. FUEL quantity indicator (#1 engine)
3. FUEL QUANTITY gaging system control switch
4. FUEL quantity indicator (#2 engine)
5. STANDBY PUMP switch (#2 engine)
6. AUX TRANSFER OVERRIDE switch (#2 engine)
7. CROSSFEED valve switch
8. AUX TRANSFER OVERRIDE switch (#1 engine)
AP 006447
Figure 2-14. Fuel Management Panel
Engine operation using only the enginedriven primary (high pressure) fuel pump
without standby pump or engine-driven
boost pump fuel pressure is limited to 10
cumulative hours. This condition is indicated by illumination of either the No. 1
or No. 2 FUEL PRESS warning annunciator lights and the simultaneous illumination of both MASTER WARNING lights.
Refer to Chapter 9. All time in this category shall be entered on DA Form
2408-13 for the attention of maintenance
personnel.
b. Standby Fuel Pumps. A submerged, electrically-operated standby fuel pump, located within
each nacelle tank, serves as a backup unit for the
engine-driven boost pump. The standby pumps are
switched off during normal system operations. A
standby fuel pump will be operated during crossfeed
operation to pump fuel from one system to the
opposite engine. The correct pump is automatically
selected when the CROSSFEED switch is activated.
Each standby fuel pump has an inertia switch
included in the power supply circuit. When subjected to a 5 to 6 G shock loading, as in a crash situation, the inertia switch will remove electrical power
from the standby fuel pumps. The standby fuel
pumps are protected by two 10-ampere circuit
breakers placarded STANDBY PUMP No. 1 or
No. 2, located the overhead circuit breaker panel (fig.
2-26) and four 5-ampere circuit breakers (2 each in
parallel) on the hot battery bus.
A gear-driven boost pump, mounted on each
engine supplies fuel under pressure to the inlet of
the engine-driven primary high-pressure pump for
engine starting and all normal operations. Either the
engine-driven boost pump or standby pump is capable of supplying sufficient pressure to the enginedriven primary high-pressure pump and thus maintain normal engine operation.
c. Fuel Transfer Pumps. The auxiliary tank
fuel transfer system automatically transfers the fuel
from the auxiliary tank to the nacelle tank without
pilot action. Motive flow to a jet pump mounted in
the auxiliary tank sump is obtained from the engine
fuel plumbing system downstream from the engine
driven boost pump and routed through the transfer
control motive flow valve. The motive flow valve is
energized to the open position by the control system
a.
Engine Driven Boost Pumps.
2-29
TM 55-1510-221-10
Figure 2-15. Crossfeed Fuel Flow
2-30
TM 55-1510-221-10
Figure 2-16. Gravity Feed Fuel Flow
2-31
TM 55-1510-221-10
Table 2-2. Fuel Quantity Data
TANKS
NUMBER
GALLONS
**POUNDS
LEFT ENGINE
Wing Tanks
Nacelle Tank
Auxiliary Tank
5
1
1
135
57
79
877.5
370.5
513.5
RIGHT ENGINE
Wing Tanks
Nacelle Tank
Auxiliary Tank
5
1
1
135
57
79
884.0
370.5
513.5
*TOTALS
14
542
3523.0
* Unusable fuel quantity and weight (4 gallons, 26 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 542
gallons (usable).
to transfer auxiliary fuel to the nacelle tank to be
consumed by the engine during the initial portion of the
flight. When an engine is started, pressure at the engine
driven boost pump closes a pressure switch which, after
a 30 to 50 second tie delay to avoid depletion of fuel
pressure during starting, energizes the motive flow
valve. When auxiliary fuel is depleted, a low level float
switch de-energizes the motive flow valve after a 30 to 60
second time delay provided to prevent cycling of the
motive flow valve due to sloshing fuel. In the event of a
failure of the motive flow valve or the associated control
circuitry, the loss of motive flow pressure when there is
still fuel remaining in the auxiliary fuel tank is sensed by
a pressure switch and float switch, respectively, which
illuminates a caution annunciator light placarded No. 1
NO FUEL XFR or No. 2 NO FUEL XFR. During engine
start, the pilot should note that the NO FUEL XFR lights
extinguish 30 to 50 seconds after engine start. The NO
FUEL XFR lights will not illuminate if auxiliary tanks
are empty. A manual override is incorporated as a
backup for the automatic transfer system. This is
initiated by placing the AUX TRANSFER switch, located
on the FUEL management panel to the OVERRIDE
position. This will energize the transfer control motive
flow valve. The transfer systems are protected by
5-ampere circuit breakers placarded AUXILIARY
TRANSFER No. 1 or No. 2, located on the overhead
circuit breaker panel (fig. 2-26), 2.0 inches high.
NOTE
In turbulence or during maneuvers, the NO
FUEL XFR lights may momentarily
illuminate after the auxiliary fuel has
completed transfer.
d. Fuel Gaging System. The total fuel quantity in
the left or right main system or left or right auxiliary
tank is measured by a capacitance type fuel gaging
2-32
Charge 4
system. Two fuel gages, one for the left and one for the
right fuel system, read fuel quantity in pounds. Refer to
Section IX for fuel capacities and weights. A maximum
of 3% error may be encountered in each system.
However, the system is compensated for fuel density
changes due to temperature excursions. In addition to
the fuel gages, yellow No. 1 NAC LOW or No. 2 NAC
LOW lights on the caution/advisory annunciator
panuminate when there is approximately 20 minutes of
fuel per engine remaining (on standard day, at sea level,
normal cruise power consumption rate). The fuel gaging
system is protected by individual 5-ampere circuit
breakers placarded QTY IND and QTY WARN No. 1 or
No. 2, located on the overhead circuit breaker panel (fig.
2-26). A mechanical spiral float gage is installed in each
auxiliary fuel tank to provide an indication of fuel level
when servicing the tank. The gage is installed on the
auxiliary fuel tank cover, adjacent to the filler neck (fig.
2-13). A small sight window in the upper wing skin
permits observation of the gage.
e. Fuel Management Panel. The fuel management
panel is located on the cockpit overhead between the
pilot and copilot. It contains the fuel gages, standby fuel
pump switches, the crossfeed valve switch and a fuel
gaging system control switch and transfer control
switches are also installed.
(1.) Fuel gaging system control switch. A switch on
the fuel management panel (fig. 2-14) placarded FUEL
QUANTITY, MAIN - AUXILIARY, controls the fuel
gaging system. When in the MAIN position the fuel
gages read the total fuel quantity in the left and right
wing fuel system. When in the AUXILIARY position the
fuel gages read the fuel quantity in the left and right
auxiliary tanks only.
TM 55-1510-221-10
(2.) Standby fuel pump switches. T w o
switches, placarded STANDBY PUMP - ON located
on the fuel management panel (fig. 2-14) control a
submerged fuel pump located in the corresponding
nacelle tank. During normal aircraft operation both
switches are off so long as the engine-driven boost
pumps function and during crossfeed operation. The
loss of fuel pressure, due to failure of an engine
driven boost pump will illuminate the MASTER
WARNING lights on the glareshield and will illuminate the No. 1 FUEL PRESS or No. 2 FUEL PRESS
on the warning annunciator panel. Turning ON the
STANDBY PUMP will extinguish the FUEL PRESS
lights. The MASTER WARNING lights must be
manually cleared.
NOTE
Both standby pump switches shall be off
during crossfeed operation.
(3.) Fuel transfer control switches. T w o
switches on the fuel management panel (fig. 2-14),
placarded AUX TRANSFER OVERRIDE - AUTO
control operation of the fuel transfer pumps During
normal operation both switches are in AUTO which
allows the system to be automatically actuated by
fuel flow to the engine. If either transfer system fails
to operate, the fault condition is indicated by two
illuminated MASTER CAUTION lights on the
glareshield and a steadily illuminated yellow No. 1
NO FUEL XFR or No. 2 NO FUEL XFR light on
the caution annunciator panel.
(4.) Fuel crossfeed switch. The fuel crossfeed valve is controlled by a 3-position switch (fig.
2-14), located on the fuel management panel, placarded CROSSFEED - OFF. Under normal flight
conditions the switch is left in the OFF position.
During emergency single engine operation, it may
become necessary to supply fuel to the operative
engine from the fuel system on the opposite side.
The crossfeed system is placarded for fuel selection
with a simplified diagram on the overhead fuel control panel. Place the standby fuel pump switches in
the off position when crossfeeding. A lever lock
switch, placarded CROSSFEED, is moved from the
center OFF position to the left or to the right,
depending on direction of fuel flow. This opens the
crossfeed valve and energizes the standby pump on
the side from which crossfeed is desired. During
crossfeed operation with firewall fuel valve closed,
auxiliary tank fuel will not crossfeed. When the
crossfeed mode is energized, a green FUEL CROSSFEED light on the caution/advisory panel will illuminate. Crossfeed system operation is described in
Chapter 9. The crossfeed valve is protected by a
5-ampere circuit breaker placarded CROSSFEED
located on the overhead circuit breaker panel (fig.
2-26).
f. Firewall Shutoff Valves.
Do not use the fuel firewall shutoff valve
to shut down an engine, except in an
emergency. The engine-driven highpressure fuel pump obtains essential
lubrication from fuel flow. When an
engine is operating, this pump may be
severely damaged (while cavitating) if the
firewall valve is closed before the condition lever is moved to the FUEL CUTOFF position.
The fuel system incorporates a fuel line shutoff
valve mounted on each engine firewall. The firewall
shutoff valves close automatically when the fire
extinguisher T-handles on the instrument panel are
pulled out. The firewall shutoff valves receive electrical power from the main buses and also from the
hot battery bus which is connected directly to the
battery. The valves are protected by circuit breakers
placarded FIREWALL VALVE No. 1 or No. 2 on the
overhead circuit breaker panel (fig. 2-26) and
FIREWALL SHUTOFF No. 1 or No. 2 on the hot
battery bus circuit breaker board.
g. Fuel Tank Sump Drains. A sump drain
wrench is provided in the aircraft loose tools to simplify draining a small amount of fuel from the sump
drain.
(1.) There are five sump drains and one
filter drain in each wing (Table 2-3).
(2.) An additional drain for the extended
range fuel system line extends through the bottom of
the wing center section adjacent to the fuselage.
Anytime the extended range system is in use, a part
of the preflight inspection would consist of draining
a small amount of fuel from this drain to check for
fuel contamination. Whenever the extended range
system is removed from the aircraft and the fuel line
is capped off in the fuselage, the remaining fuel in
the line shall be drained.
h. Fuel Drain Collector System. Each engine is
provided with a fuel drain collector system to return
fuel dumped from the engine during clearing and
shutdown operations back into its respective nacelle
tank. The system draws power from the No. 4 feeder
bus. Fuel transfer is completely automatic. Fuel
from the engine flow divider drains into a collector
tank mounted below the aft engine accessory sec-
2-33
TM 55-1510-221-10
Table 2-3. Fuel Sump Drain Locations
NUMBER
1
1
1
1
1
1
DRAINS
LOCATION
Leading Edge Tank
Integral Tank
Firewall Fuel Filter
Sump Strainer
Gravity Feed Line
Auxiliary Tank
Outboard of nacelle, underside of wing
Underside of wing, forward of aileron
Underside of cowling forward of firewall
Bottom center of nacelle forward of wheel well
Aft of wheel well
At wing root, just forward of the flap
tion. An internal float switch actuates an electric
scavenger pump which delivers the fuel to the fuel
purge line just aft of the fuel purge shutoff valve. A
check valve in the line prevents the backflow of fuel
during engine purging. The circuit breaker for both
pumps is located in the fuel section of the overhead
circuit breaker panel; placarded SCAVENGER
PUMP. A vent line, plumbed from the top of the
collector tank, is routed through an inline flame
arrestor and then downward to a drain manifold on
the underside of the nacelle.
fer of auxiliary fuel, which is automatically controlled, the nacelle tanks are maintained full. A
check valve in the gravity feed line from the outboard wing prevents reverse fuel flow. Normal gravity transfer of the main wing fuel into the nacelle
tanks will begin when auxiliary fuel is exhausted.
The system will gravity feed fuel only to its respective nacelle tank, i.e. left or right (fig. 2-16). Fuel
will not gravity feed through the crossfeed system.
j. Engine Oil-to-Fuel Heat Exchanger. A n
engine oil-to-fuel heat exchanger, located on each
engine accessory case, operates continuously and
automatically to heat the fuel delivered to the engine
sufficiently to prevent the freezing of any water
which it might contain. The temperature of the
delivered fuel is thermostatically regulated to remain
between 21°C and 32°C.
b. Operation With Failed Engine-Driven Boost
Pump or Standby Pump. Two pumps in each fuel
system provide inlet head pressure to the enginedriven primary high-pressure fuel pump. If crossfeed
is used, a third pump, the standby fuel pump from
the opposite system, will supply the required pressure. Operation under this condition will result in an
unbalanced fuel load as fuel from one system will be
supplied to both engines while all fuel from the system with the failed engine driven and standby boost
pumps will remain unused. A triple failure, which is
highly unlikely, would result in the engine driven
primary pump operating without inlet head pressure. Should this situation occur, the affected engine
can continue to operate from its own fuel supply on
its engine-driven primary high-pressure fuel pump.
2-34. FUEL SYSTEM MANAGEMENT.
2-35. FERRY FUEL SYSTEM.
a. Fuel Transfer System. When the auxiliary
tanks are filled, they will be used first. During trans-
Provisions are installed for connection to long
range fuel cells.
i. Fuel Vent System. Each fuel system is
vented through two ram vents located on the underside of the wing adjacent to the nacelle. To prevent
icing of the vent system, one vent is recessed into
the wing and the backup vent protrudes out from
the wing and contains a heating element. The vent
line at the nacelle contains an inline flame arrestor.
Section V. FLIGHT CONTROLS
2-36. DESCRIPTION.
The aircraft’s primary flight control systems
consist of conventional rudder, elevator and aileron
control surfaces. These surfaces are manually operated from the cockpit through mechanical linkage
using a control wheel for the ailerons and elevators,
and adjustable rudder/brake pedals for the rudder.
Both the pilot and copilot have flight controls. Trim
2-34
control for the rudder, elevator and ailerons is
accomplished through a manually actuated cabledrum system for each set of control surfaces. The
autopilot has provisions for controlling the position
of the ailerons, elevators, and rudder. Chapter 3
describes the operation of the autopilot system.
TM 55-1510-221-10
2-37. CONTROL WHEELS
Elevator and aileron control surfaces are operated by manually actuating either the pilot’s or copilot’s control wheel. Switches are installed in the outboard grip of each wheel to operate the elevator trim
tabs. A microphone switch, a chaff dispense switch,
and an autopilot/yaw damp/electric trim disconnect
switch are also installed in the outboard grip of each
wheel. A transponder ident switch is installed on top
of the inboard grip of each control wheel. These control wheels (fig. 2-17) are installed - on each side of
the instrument subpanel. A manually wound 8-day
clock is installed in the center of the pilot’s control
wheel, and a digital electric clock is installed in the
center of the copilot’s control wheel. A map light
switch, and a pitch synchronization and control
wheel steering switch are mounted adjacent to the
clock in each control wheel.
2-38. RUDDER SYSTEM.
a. Rudder Pedals. Aircraft rudder control and
nose wheel steering is accomplished by actuation of
the rudder pedals from either pilot’s or copilot’s station (fig. 2-8). The rudder pedals may be individually adjusted in either a forward or aft position to
provide adequate leg room for the pilot and copilot.
Adjustment is accomplished by depressing the lever
alongside the rudder pedal arm and moving the
pedal forward or aft until the locking pin engages in
the selected position.
b. Yaw Damp System. A yaw damp system is
provided to aid the pilot in maintaining directional
stability and increase ride comfort. The system may
be used at any altitude and is required for flight
above 17,000 feet. It must be deactivated for takeoff
and landing. The yaw damp system is a part of the
autopilot. Operating instructions for this system are
contained in Chapter 3. The system is controlled by
a YAW DAMP switch adjacent to the ELEV TRIM
switch on the pedestal extension.
c. Rudder Boost System. A rudder boost system is provided to aid the pilot in maintaining
directional control resulting from an engine failure
or a large variation of power between the engines.
Incorporated in the rudder cable system are two
pneumatic rudder boosting servos which actuate the
cables to provide rudder pressure to help compensate for asymmetrical thrust.
(1.) During operation, a differential pressure valve accepts bleed air pressure from each
engine. When the pressure varies between the bleed
air systems, the shuttle in the differential pressure
valve moves toward the low pressure side. As the
pressure difference reaches a preset tolerance, a
switch closes on the low pressure side which activates the rudder boost system. This system is
designed only to help compensate for asymmetrical
thrust. Appropriate trimming is to be accomplished
by the pilot. Moving either or both of the bleed air
valve switches on the overhead control panel to
PNEU & ENVIRO - OFF position will disengage the
rudder boost system.
NOTE
Condition levers must be in LOW IDLE
position to perform rudder boost check.
(2.) The system is controlled by a switch
located on the extended pedestal placarded RUDDER BOOST (on) - OFF (fig. 2-7), and is to be
turned on before flight. A preflight check of the system can be performed during the run-up by retarding the power on one engine to idle and advancing
power on the opposite engine until the power difference between the engines is great enough to activate
the switch to turn on the rudder boost system.
Movement of the appropriate rudder pedal (left
engine idling, right rudder pedal moves forward) will
be noted when the switch closes, indicating the system is functioning properly for low engine power on
that side. Repeat the check with opposite power settings to check for movement of the opposite rudder
pedal. The system is protected by a 5-ampere circuit
breaker placarded RUDDER BOOST, located on
the overhead circuit breaker panel (fig. 2-26).
NOTE
With brake deice on, rudder boost may be
inoperative.
2-39. FLIGHT CONTROLS LOCK.
Remove control locks before towing the
aircraft or starting engines. Serious damage could result in the steering linkage if
towed by a tug with the rudder lock
installed.
Positive locking of the rudder, elevator and aileron control surfaces, and engine controls (power
levers, propeller levers, and condition levers) is provided by a removable lock assembly (fig. 2-18) consisting of two pins, and an elongated U-shaped strap
interconnected by a chain. Installation of the control
locks is accomplished by inserting the U-shaped
strap around the aligned control levers from the
2-35
TM 55-1510-221-10
(Pilot)
1. Microphone, intercom, transmit switch
2. Trim, autopilot, yaw damp disconnect switch
3. Pitch-trim switches
4. Chaff dispense switch
5. Transponder ident switch
6. Map light
7. Eight day clock
8. Pitch synchronization and control wheel steering switch
(Copilot)
AP 010329
1. Microphone, intercom, transmit switch
2. Chaff dispense switch
3. Pitch-trim switches
4. Trim, autopilot, yaw damp diconnect switch
5. Transponder ident switch
6. Pitch synchronization and control wheel steering switch
7. Digital clock
8. Map light
Figure 2- 17. Control Wheels
2-36
TM 55-1510-221-10
APOO5445
Figure 2- 18. Control Locks
copilot’s side; then the aileron/elevator locking pin
is inserted through a guide hole in the top of the
pilot’s control column assembly, thus locking the
control wheel. The rudder is held in a neutral position by an L-shaped pin which is installed through
a guide hole in the floor aft of the pilots rudder pedals. The rudder pedals must be centered to align the
hole in the rudder bellcrank with the guide hole in
the floor. Remove the locks in reverse order, i.e.,
rudder pin, control column pin, and power control
clamp.
2-40. TRIM TABS.
Trim tabs are provided for all flight control surfaces. These tabs are manually activated, and are
mechanically controlled by a cable-drum and jackscrew actuator system, except the right aileron tab
which is of the fixed bendable type. Elevator and
aileron trim tabs incorporate neutral, non-servo
action, i.e., as the elevators or ailerons are displaced
from the neutral position, the trim tab maintains an
“as adjusted” position. The rudder trim tab incorporates anti-servo action, i.e., as the rudder is displaced from the neutral position the trim tab moves
in the same direction as the control surface. This
action increases control pressure as rudder is
deflected from the neutral position.
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.
2-7). The amount of elevator tab deflection, in
degrees from a neutral setting, is indicated by a position arrow.
b. Electric Elevator Trim. The electric elevator trim system is controlled by an ELEV TRIM ON - OFF/RESET switch located on the pedestal,
dual element thumb switches on the control wheels,
a trim disconnect switch on each control wheel and
a circuit breaker on the overhead circuit breaker
panel. The ON - OFF/RESET switch must be in the
ON position to operate the system. The dual element thumb switch is moved forward for trimming
nose down, aft for nose up, and when released
returns to the center (off) position. Any activation of
the trim system through the copilot’s trim switch
can be over ridden by activation of the pilot’s
switch. Operating the pilot’s and copilot’s switches
in opposing directions simultaneously results in the
pilot having priority.
A preflight check of the switches should be
accomplished before flight by moving the switches
individually on both control wheels. No one switch
alone should operate the system; operation of eleva-
2-37
TM 55-1510-221-10
tor trim should occur only by movement of pairs of
switches. The trim system disconnect is a bi-level,
pushbutton, momentary type switch, located on the
outboard grip of each control wheel. Depressing the
switch to the first of two levels disconnects the
autopilot and yaw damp system, and the second
level disconnects the electric trim system. The system can be reset by moving the ELEV TRIM switch
toggle on the pedestal (fig. 2-7) to OFF RESET position, then back to ELEV TRIM (on) again.
c. Aileron Trim Tab Control. The aileron trim
tab control, placarded AILERON TAB - LEFT,
RIGHT, is on the control pedestal and will adjust
the left aileron trim tab only (fig. 2-7). The amount
of aileron tab deflection, from a neutral setting, as
indicted by a position arrow, is relative only and is
not in degrees. Full travel of the tab control moves
the trim tab 7-1/2 degrees up and down.
d. Rudder Trim Tab Control. The rudder trim
tab control knob, placarded RUDDER TAB -LEFT,
RIGHT, is on the control pedestal, and controls
adjustment of the rudder trim tab (fig. 2-7). The
amount of rudder tab deflection, in degrees from a
neutral setting, is indicated by a position arrow.
2-41. WING FLAPS.
The all-metal slot-type wing flaps are electrically
operated and consist of two sections for each wing.
These sections extend from the inboard end of each
aileron to the junction of the wing and fuselage.
During extension, or retraction, the flaps are operated as a single unit, each section being actuated by
a separate jackscrew actuator. The actuators are
driven through flexible shafts by a single, reversible
electric motor. Wing flap movement is indicated in
percent of travel by a flap position indicator on the
forward control pedestal. Full flap extension and
retraction time is approximately 11 seconds. The
flap control switch is located on the control pedestal.
Section VI.
2-42. DESCRIPTION.
A three-blade aluminum propeller is installed on
each engine. The propeller is of the full feathering,
constant speed, counterweighted, reversible type,
controlled by engine oil pressure through single
action, engine driven propeller governors. The propeller is flange-mounted to the engine shaft. Centrifugal counterweights, assisted by a feathering spring,
move the blades toward the low RPM (high pitch)
position and into the feathered position. Governor
boosted engine oil pressure moves the propeller to
2-38
No emergency wing flap actuation system is provided. With flaps extended beyond the APPROACH
position, the landing gear warning horn will sound,
unless the landing gear is down and locked. The circuit is protected by a 20-ampere circuit breaker,
placarded FLAP MOTOR, located on the overhead
circuit breaker panel (fig. 2-26).
a. Wing Flap Control Switch. Flap operation
is controlled by a three-position switch with a flapshaped handle on the control pedestal (fig. 2-7). The
handle of this switch is placarded FLAP and switch
positions are placarded: FLAP - UP, APPROACH,
and DOWN. The amount of downward extension of
the flaps is established by position of the flap switch,
and is as follows: UP - 0%, APPROACH - 40%, and
DOWN -100%. Limit switches, mounted on the
right inboard flap, control flap travel. The flap control switch, limit switch, and relay circuits are protected by a 5-ampere circuit breaker, placarded
FLAP CONTR located on the overhead circuit
breaker panel (fig. 2-26). Flap positions between UP
and APPROACH cannot be selected. For intermediate flap positions between 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. In the
event that any two adjacent flap sections extend 3 to
5 degrees out of phase with the other, a safety mechanism is provided to discontinue power to the flap
motor.
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 located on the control pedestal (fig.
2-7). The approach and full down or extended flap
position is 14 and 34 degrees, respectively. The flap
position indicator is protected by a 5-ampere circuit
breaker, placarded FLAP CONTR, located on the
overhead circuit breaker panel (fig. 2-26).
PROPELLERS
the high RPM (low pitch) hydraulic stop and reverse
position. The propellers have no low RPM (high
pitch) stops; this allows the blades to feather after
engine shutdown. Low pitch propeller position is
determined by the low pitch stop which is a mechanically actuated, hydraulic stop. Beta and reverse
blade angles are controlled by the power levers in
the beta and reverse range.
TM 55-1510-221-10
2-43. FEATHERING PROVISIONS.
Both manual and automatic propeller feathering
systems are provided. Manual feathering is accomplished by pulling the corresponding propeller lever
aft past a friction detent. To unfeather, the propeller
lever is pushed forward into the governing range. An
automatic feathering system, will sense loss of
torque and will feather an unpowered propeller.
Feathering springs will feather the propeller when it
is not turning.
a. Automatic Feathering. The automatic
feathering system provides a means of immediately
dumping oil from the propeller servo to enable the
feathering spring and counterweights to start feathering action of the blades in the event of an engine
failure. Although the system is armed by a switch on
the overhead control panel, placarded AUTOFEATHER - ARM - OFF - TEST, the completion of the
arming phase occurs when both power levers are
advanced above 90% N1 - at which time both indicator lights on the caution/advisory annunciator panel
indicate a fully armed system. The annunciator
panel lights are green and are placarded No.1
No.2
engine)
and
AUTOFEATHER (left
AUTOFEATHER (right engine). The system will
remain inoperative as long as either power lever is
retarded below 90% N1 position, unless TEST position of the AUTOFEATHER SWITCH is selected to
disable the power lever limit switches. The system is
designed for use only during takeoff and landing and
should be turned off when establishing cruise climb.
During takeoff or landing, should the torque for
either engine drop to an indication between 16 21%, the autofeather system for the opposite engine
will be disarmed. Disarming is confirmed when the
No.1 AUTOFEATHER or No.2 AUTOFEATHER
annunciator light of the opposite engine becomes
extinguished. If torque drops further, to a reading
between 9 and 14%, oil is dumped from the servo of
the affected propeller allowing a feathering spring to
move the blades into the feathered position. Feathering also causes the No.1 AUTOFEATHER or
No.2 AUTOFEATHER annunciator light of the
feathered propeller to extinguish. At this time, both
the No.1 AUTOFEATHER and No.2 AUTOFEATHER lights are extinguished, the propeller of the
defective engine has feathered, and the propeller of
the operative engine has been disarmed from the
autofeathering capability. Only manual feathering
control remains for the second propeller.
b. Propeller Autofeather Switch. Autofeathering is controlled by an AUTOFEATHER switch on
the overhead control panel (fig. 2-12). The threeposition switch is placarded ARM, OFF and TEST,
and is spring loaded from TEST to OFF. The ARM
position is used only during takeoff and landing.
The TEST position of the switch, enables the pilot
to check readiness of the autofeather systems, below
88% to 92% Nl , and is for ground checkout purposes
only.
c. Autofeather Lights. Two green lights on the
caution/advisory annunciator panel are placarded
AUTOFEATHER No. 1 and AUTOFEATHER
No.2. When illuminated, the lights indicate that the
autofeather system is armed. Both lights will be
extinguished if either propeller has been autofeathered or if the system is disarmed by retarding a
power lever. Autofeather circuits are protected by
one 5-ampere circuit breaker placarded AUTO
FEATHER, located on the overhead circuit breaker
panel (fig. 2-26).
2-44. PROPELLER GOVERNORS.
Two governors, a constant speed governor, and
an overspeed governor, control propeller RPM. The
constant speed governor, mounted on top of the
reduction housing, controls the propeller through its
entire range. The propeller control lever operates the
propeller by means of this governor. If the constant
speed governor should malfunction and request
more than 2000 RPM, the overspeed governor cuts
in at 2080 RPM and dumps oil from the propeller
to keep the RPM from exceeding approximately
2080 RPM. A solenoid, actuated by the PROP GOV
TEST switch located on the overhead control panel
(fig. 2-12), is provided for resetting the overspeed
governor to approximately 1830 to 1910 RPM for
test purposes. If the propeller sticks or moves too
slowly during a transient condition causing the propeller governor to act too slowly to prevent an overspeed condition, the power turbine governor, contained within the constant speed governor housing,
acts as a fuel topping governor. When the propeller
reaches 106% of N2 RPM, the fuel topping governor
limits the fuel flow to the gas generator, reducing N 1
RPM, which in turn prevents the propeller RPM
from exceeding approximately 2120 RPM. During
operation in the reverse range, the power turbine
governor is reset to approximately 95% of propeller
RPM before the propeller reaches a negative pitch
angle. This insures that the engine power is limited
to maintain a propeller RPM of somewhat less than
that of the constant speed governor setting. The
constant speed governor therefore, will always sense
an underspeed condition and direct oil pressure to
the propeller servo piston to permit propeller operation in beta and reverse ranges.
2-45. PROPELLER TEST SWITCHES.
Two two-position switches on the overhead control panel (fig. 2-12), are provided for operational
testing of the propeller systems. Placarding above
2-39
TM 55-1510-221-10
the switches reads PROP GOV TEST. Each switch
controls test circuits for the corresponding propeller.
In the test position, the switches are used to test the
function of the corresponding overspeed governor.
Refer to Chapter 8, for test procedure. Propeller test
circuits are protected by one 5-ampere circuit
breaker placarded PROP GOV located on the overhead circuit breaker panel (fig. 2-26).
2-46. PROPELLER SYNCHROPHASER.
a. Operation. The propeller synchrophaser
automatically matches the RPM of the right propeller (slave propeller) to that of the left propeller (master propeller) and maintains the blades of one propeller at a predetermined relative position with the
blades of the other propeller. To prevent the right
propeller from losing excessive RPM if the left propeller is feathered while the synchrophaser is on, the
synchrophaser has a limited range of control from
the manual governor setting. Normal governor operation is unchanged but the synchrophaser will continuously monitor propeller RPM and reset the governor as required. A magnetic pickup mounted in
each propeller overspeed governor and adjacent to
each propeller deice brush block transmits electric
pulses to a transistorized control box installed forward of the pedestal. The right propeller RPM and
phase will automatically be adjusted to correspond
to the left. To change RPM, adjust both propeller
controls at the same time. This will keep the right
governor setting within the limiting range of the left
propeller. If the synchrophaser is on but is unable to
adjust to the right propeller to match the left, the
actuator has reached the end of its travel. To
recenter, turn the switch off, synchronize the propellers manually, and turn the switch back on.
b. Control Box. The control box converts any
pulse rate differences into correction commands,
which are transmitted to a stepping type actuator
motor mounted on the right engine cowl forward
support ring. The motor then trims the right propeller governor through a flexible shaft and trimmer
assembly to exactly match the left propeller. The
trimmer, installed between the governor control arm
and the control cable, screws in or out to adjust the
governor while leaving the control lever setting constant. A toggle switch installed adjacent to the synchrophaser turns the system on. With the switch off,
the actuator automatically runs to the center of its
range of travel before stopping to assure normal
function when used again. To operate the system,
synchronize the propeller in the normal manner and
turn the synchrophaser on. The system is designed
for in-flight operations and is placarded to be off for
take-off and landing. Therefore, with the system on
2-40
and the landing gear extended, the master caution
lights will illuminate and a yellow light on the caution/advisory annunciator panel, PROP SYNC ON,
will illuminate.
c. Synchroscope. The propeller synchroscope,
provides an indication of synchronization of the
propellers. If the right propeller is operating at a
higher RPM than the left, the face of the synchroscope, a black and white cross pattern, spins in a
clockwise rotation. Left, or counterclockwise, rotation indicates a higher RPM of the left propeller.
This instrument aids the pilot in obtaining complete
synchronization of propellers. The system is protected by a 5-ampere circuit breaker placarded
PROP SYNC, located on the overhead circuit
breaker panel (fig. 2-26).
2-47. PROPELLER LEVERS.
Two propeller levers on the control pedestal (fig.
2-7), placarded PROP, are used to regulate propeller
speeds. Each lever controls a primary governor,
which acts to regulate propeller speeds within the
normal operation range. The full forward position of
the levers is placarded TAKEOFF, LANDING AND
REVERSE - and also HIGH RPM. The full aft position of the levers is placarded FEATHER. When a
lever is placed at HIGH RPM, the propeller may
attain a static RPM of 2,000 depending upon power
lever position. As a lever is moved aft, passing
through the propeller governing range, but stopping
at the feathering detent, propeller RPM will correspondingly decrease to the lowest limit. Moving a
propeller lever aft past the detent into FEATHER
will feather the propeller.
2-48. PROPELLER REVERSING.
Do not move the power levers into
reverse range without the engine running.
Damage to the reverse linkage mechanisms will occur.
Propeller reversing on unimproved surfaces should be accomplished carefully to
prevent propeller erosion from reversed
airflow and, in dusty conditions, to prevent obscuring the operator’s vision.
TM 55-1510-221-10
To prevent an asymmetrical thrust condition, propeller levers must be in HIGH
RPM position prior to propeller reversing.
The propeller blade angle may be reversed to
shorten landing roll. To reverse, propeller levers
must be positioned at HIGH RPM (full forward),
and the power levers are lifted up to pass over the
IDLE detent, then pulled aft into REVERSE setting.
One yellow caution light, placarded REV NOT
READY, on the caution/advisory annunciator panel,
Section VII.
2-50. 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 pilot’s subpanel,
manually controls airflow to the windshield. When
pulled out, defrosting air is ducted to the windshield. As the control is pushed in, there is a corresponding decrease in airflow.
b. Automatic Operation.
1.
Vent blower switches - As required.
2. Cabin temperature mode selector
switch - AUTO.
3. Cabin temperature control rheostat As required.
4. Cabin air, copilot air, pilot air, and
defrost air controls - As required.
c.
Manual Operation.
1.
Pilot air, copilot air - IN.
2.
Cabin air and defrost air controls - Out
3. Cabin temperature mode selector
switch - MAN HEAT.
4.
Cold air outlets - As required.
5. Manual temperature switch - As
required.
alerts the pilot not to reverse the propellers. This
light illuminates only when the landing gear handle
is down, and if propeller levers are not at HIGH
RPM (full forward). This circuit is protected by a
5-ampere circuit breaker, placarded LANDING
GEAR RELAY, located on the overhead circuit
breaker panel (fig. 2-26).
2-49. PROPELLER TACHOMETERS.
Two tachometers on the instrument panel register propeller speed in hundreds of RPM (fig. 2-29).
Each indicator is slaved to a tachometer generator
unit attached to the corresponding engine.
UTILITY SYSTEMS
d. Manual Operation. If the automatic temperature control should fail to operate, the temperature (of defrost air and cabin air) may be controlled
manually by setting the CABIN TEMP MODE control switch to MANUAL COOL position, then using
the MANUAL TEMP DECREASE-INCREASE
switch to set the desired temperature. This control is
located on the overhead control panel (fig. 2-l 2).
2-51.
SURFACE DEICING SYSTEM.
a. Description. Ice accumulation is removed
from each inboard and outboard wing leading edge,
both horizontal stabilizers, the taillets, and certain
mission antennas by the flexing of deicer boots
which are pneumatically actuated. Bleed air from
the engine compressor is used to inflate the deicer
boots and to supply vacuum, through the ejector system, for boot hold down during flight. A pressure
regulator protects the system from over inflation.
When the system is not in operation, a distributor
valve applies vacuum to the boots for hold down. A
selector switch allows automatic single cycle operation or manual operation. To assure operation of the
system in the event of failure of one engine, a check
valve is incorporated in the bleed air line from each
engine to prevent loss of pressure through the compressor of the inoperative engine.
Wing ice lights allow the crew to detect ice formations. Ice protection of the engine is provided by
inertial separation. Automatically cycled electrothermal anti-icing boots are installed on the propeller blades. The engine air inlet leading edge lip is
anti-iced by engine exhaust bleed. The fresh air
inlets are located in sheltered areas and require no
deice protection.
2-41
TM 55-1510-221-10
Operation of the surface deice system in
ambient temperatures below -40°C can
cause permanent damage to the deice
boots.
b. Operation.
(1.) Deice boots are intended to remove
ice after it has formed rather than prevent its formation. For the most effective deicing operation, allow
at least l/2 inch of ice to form on the boots before
attempting ice removal. Very thin ice may crack and
cling to the boots instead of shedding.
NOTE
Never cycle the system rapidly, this may
cause the ice to accumulate outside the
contour of the inflated boots and prevent
ice removal.
(2.) A three position switch on the overhead control panel placarded SURF DEICE MANUAL - OFF - SINGLE CYCLE AUTO, controls the
deicing operation. The switch is spring loaded to
return to the OFF position from SINGLE CYCLE
AUTO or MANUAL. When the SINGLE CYCLE
AUTO position is selected, the distributor valve
opens to inflate the wing boots. After an inflation
period of approximately 6 seconds, an electronic
timer switches the distributor to deflate the wing
boots and a 4 second inflation begins in the horizontal stabilizer boots. When these boots have inflated
and deflated, the cycle is complete.
(3.) If the switch is held in the MANUAL
position, the boots will inflate simultaneously and
remain inflated until the switch is released. The
switch will return to the OFF position when
released. After the cycle, the boots will remain in the
vacuum hold down condition until again actuated
by the switch.
(4.) Either engine is capable of providing
sufficient bleed air for all requirements of the surface deicer system. Check valves in the bleed air and
vacuum lines prevent backflow through the system
during single-engine operation. Regulated pressure is
indicated on a gage, placarded PNEUMATIC PRESSURE, located on the copilots subpanel.
2-52. ANTENNA DEICING SYSTEM.
a. Description. The antenna deice system
removes ice accumulation from the dipole mission
2-42
antennas. The system consists of two ejector distributor valves, a regulator, manifold, and flexible tubing. Control is accomplished through a timing circuit and an antenna deice switch located on the
overhead control panel (fig. 2-12). Erosion resistant
tape is applied to the surface of mission blade antennas not having deice boots.
b. Antenna Deice System Switch. The antenna
deice system is controlled by a switch placarded
ANT DEICE, SINGLE - OFF - MANUAL located
on the overhead control panel (fig. 2-12). The switch
is spring loaded to return to the OFF position from
the SINGLE or MANUAL position. When the
switch is set to the single position, the system will
run through one timed inflation-deflation cycle.
When the switch is held in the MANUAL position
the boots will inflate and remain inflated until the
switch is released.
c. Forward Wide Band Data Link Antenna
Radome Anti-Ice. The forward wide band data link
antenna radome anti-ice system utilizes engine bleed
air to prevent the formation of ice on the radome.
The system is controlled by a switch placarded
RADOME located on the overhead control panel.
The circuit is protected by a circuit breaker placarded RADOME, located on the overhead circuit
breaker panel (fig. 2-26).
d. 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 l/8 to l/4 inch of ice to form on the boots
before attempting ice removal. Very thin ice may
crack and cling to the boots instead of shedding.
NOTE
Never cycle the system rapidly, this may
cause the ice to accumulate outside the
contour of the inflated boots and prevent
ice removal.
2-53. PROPELLER ELECTROTHERMAL ANTI-ICE
SYSTEM.
a. Description. Electrothermal anti-ice boots
are cemented to each propeller blade to prevent ice
formation or to remove the ice from the propellers.
Each thermal boot consists of one outboard and one
inboard heating element, and receives electrical
power from the deice timer. This timer sends current to all propeller deice boots and prevents the
boots from overheating by limiting the time each
element is energized. Four intervals of approximately 30 seconds each complete one cycle. Current
TM 55-1510-221-10
consumption is monitored by a propeller ammeter
on the copilot’s subpanel. Two 20-ampere circuit
breakers placarded PROP ANTI-ICE LEFT and
RIGHT and 5-ampere propeller control circuit
breaker placarded CONTR on the overhead circuit
breaker panel (fig. 2-26), protect the propeller electrothermal deice system during manual operation. A
25 ampere circuit breaker placarded PROP AUTO,
protects the system in automatic operation.
b. Automatic Operation. A control switch on
the overhead control panel placarded PROP - OFF
- AUTO is provided to activate the automatic system. A deice ammeter above the pedestal registers
the amount of current (14 to 18 amperes) passing
through the system being used. During AUTO operation, power to the timer will be cut off if the current rises above 25 amperes. Current flows from the
timer to the brush assembly and then to the slip
rings installed on the spinner backing plate. The slip
rings carry the current to the deice boots on the propeller blades. Heat from the boots reduces the grip
of the ice which is then thrown off by centrifugal
force, aided by the air blast over the propeller surfaces. Power to the two heating elements on each
blade, the inner and outer element, is cycled by the
timer in the following sequence: right propeller outer
element, right propeller inner element, left propeller
outer element, left propeller inner element. Loss of
one heating element circuit on one side does not
mean that the entire system must be turned off.
Proper operation can be checked by noting the correct level of current usage on the ammeter. An intermittent flicker of the needle approximately each 30
seconds indicates switching to the next group of
heating elements by the timer.
c. Manual Operation. The manual propeller
deice system is provided as a backup to the automatic system. A control switch located on the overhead control panel, placarded PROP - INNER OUTER, controls the manual override relays. When
the switch is in the OUTER position, the automatic
timer is overridden and power is supplied to the
outer heating elements of both propellers simultaneously. The switch is of the momentary type and
must be held in position until the ice has been dislodged from the propeller surface. After deicing with
the outer elements, the switch is to be held in the
INNER position to perform the same function for
the inner elements of both propellers. The loadmeters will indicate approximately a 5% increase of
load per meter when manual propeller deice is operating. The propeller deice ammeter will not indicate
any load in the manual mode of operation.
2-54. PITOT AND STALL WARNING HEAT SYSTEM.
Pitot heat should not be used for more
than 15 minutes while the aircraft is on
the ground. Overheating may damage the
heating elements.
a. Pitot Heat. Heating elements are installed
in both pitot masts, located on the nose. Each heating element is controlled by an individual switch
placarded PITOT - ON - LEFT or RIGHT, located
on the overhead control panel (fig. 2-12). It is not
advisable to operate the pitot heat system on the
ground except for testing or for short intervals of
time to remove ice or snow from the mast. Circuit
protection is provided by two 7 l/2 ampere circuit
breakers, placarded PITOT HEAT, on the overhead
circuit breaker panel (fig. 2-26). The “true airspeed
temp probe“ heat control circuit is also protected by
this circuit breaker. If either left or right pitot heat
is on, the“ true airspeed temp probe“ heat will be
on.
NOTE
The “TRUE AIRSPEED TEMP PROBE“
is connected to the autopilot air data
computer.
The heating elements protect the stall
warning lift transducer vane and face
plate from ice, however, a buildup of ice
on the wing may change or disrupt the
airflow and prevent the system from accurately indicating an imminent stall.
b. Stall Warning Heat. The lift transducer is
equipped with anti-icing capability on both the
mounting plate and the vane. The heat is controlled
by a switch located on the overhead control panel
placarded STALL WARN. The level of heat is minimal for ground operation but is automatically
increased for flight operation through the landing
gear safety switch. Circuit protection is provided by
a 15-ampere circuit breaker, placarded STALL
WARN, on the overhead circuit breaker panel (fig.
2-26).
2-43
TM 55-1510-221-10
2-55. STALL WARNING SYSTEM.
The stall warning system consists of a transducer, a lift computer, a warning horn, and a test
switch. Angle of attack is sensed by aerodynamic
pressure on the lift transducer vane located on the
left wing leading edge. When a stall is imminent, the
output of the transducer activates a stall warning
horn. The system has preflight test capability
through the use of a switch placarded STALL
WARN TEST - OFF - LDG GEAR WARN TEST
on the right subpanel. Holding this switch in the
STALL WARN TEST position actuates the warning
horn by moving the transducer vane. The circuit is
protected by a 5-ampere circuit breaker, placarded
STALL WARN, on the overhead circuit breaker
panel.
2-56. BRAKE DEICE SYSTEM.
a. Description. A heated-air brake deice system may be used in flight with gear retracted or
extended, or on the ground. When activated, hot air
is diffused by means of a manifold assembly over
the brake discs on each wheel. Manual and automatic controls are provided. There are two primary
occasions which require brake deicing. The first is
when an aircraft has been parked in a freezing atmosphere allowing the brake systems to become contaminated by freezing rain, snow, or ice, and the aircraft must be moved or taxied. The second occasion
is during flight through icing conditions with wet
brake assemblies presumed to be frozen, which must
be thawed prior to landing to avoid possible tire
damage and loss of directional control. Hot air for
the brake deice system comes from the compressor
stage of both engines obtained by means of a solenoid valve attached to the bleed air system which
serves both the surface deice system and the pneumatic systems operation.
b. Operation. A switch on the overhead control panel, placarded BRAKE DEICE, controls the
solenoid valve by routing power through a control
module box under the aisleway floorboards. When
the switch is on, power from a 5-ampere circuit
breaker on the overhead circuit breaker panel is
applied to the control module. A lo-minute timer
limits operation and avoids excessive wheel well
temperatures when the landing gear is retracted. The
control module also contains a circuit to the green
BRAKE DEICE ON annunciator light, and has a
resetting circuit interlocked with the gear uplock
switch. When the system is activated, the BRAKE
DElCE ON light should be monitored and the control switch selected OFF after the light extinguishes
- otherwise, on the next gear extension the system
will restart without pilot action. The control switch
should also be selected OFF, if deice operation fails
2-44
to self-terminate after 10 minutes. If the automatic
timer has terminated brake deicer operation after
the last retraction of the landing gear, the landing
gear must be extended in order to obtain further
operation of the system.
(1.) The L BL AIR FAIL or R BL AIR
FAIL annunciator lights may momentarily illuminate during simultaneous operation of the surface
deice and brake deice systems at low N 1 speeds. If
the lights immediately extinguish, they may be disregarded.
(2.) During certain ambient conditions,
use of the brake deice system may reduce available
engine power, and during flight will result in a TGT
rise of approximately 20°C. Appropriate performance charts should be consulted before brake deice
system use. If specified power cannot be obtained
without exceeding limits, the brake deice system
must be selected off until after takeoff is completed.
TGT limitations must also be observed when setting
climb and cruise power. The brake deice system is
not to be operated above 15°C ambient temperature.
The system is not to be operated for longer than 10
minutes (one deicer cycle) with the landing gear
retracted. If operation does not automatically terminate after approximately 10 minutes following gear
retraction, the system must be manually selected off.
During periods of simultaneous brake deice and surface deice operation, maintain 85% N1 or higher. If
inadequate pneumatic pressure is developed for
proper surface deicer boot inflation, select the brake
deice system off. Both sources of pneumatic bleed
air must be in operation during brake deice system
use. Select the brake deice system off during singleengine operation. Circuit protection is provided by
a 5-ampere circuit breaker, placarded BRAKE
DEICE, on the overhead circuit breaker panel (fig.
2-26).
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. Two external fuel vents
are provided on each wing. One is recessed to prevent ice formation; the other is electrically heated
and is controlled by two toggle switches on the overhead control panel placarded FUEL VENT ON,
LEFT and RIGHT (fig. 2-12). They are protected by
two 5-ampere circuit breakers, placarded FUEL
VENT HEAT, RIGHT or LEFT, located on the
overhead circuit breaker panel (fig. 2-26). Each fuel
control unit is protected against ice. The pneumatic
governor for each fuel control unit is electrically
heated, and protected by two 7 1/2-ampere circuit
TM 55-1510-221-10
breakers located on the overhead circuit breaker
panel placarded FUEL CONTR HEAT, LEFT or
RIGHT (fig. 2-26).
To prevent overheat damage to electrically heated anti-ice jackets, the FUEL
VENT heat switches should not be turned
ON unless cooling air will soon pass over
the jackets.
b. Normal Operation. For normal operation,
switches for the FUEL VENTS anti-ice circuits are
turned ON as required during the BEFORE TAKEOFF procedures (Chapter 8).
2-58. WINDSHIELD ELECTROTHERMAL ANTIICE SYSTEM.
a. Description. Both pilot and copilot windshields are provided with an electrothermal anti-ice
system. Each windshield is part of an independent
electrothermal 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, two relays, a control switch, and two circuit
breakers. Two switches, placarded WSHLD ANTIICE NORMAL - OFF - HI - PILOT, COPILOT,
located on the overhead control panel (fig. 2-l 2)
control system operation. Each switch controls one
electrothermal windshield system. The circuits of
each system are protected by a 5-ampere circuit
breaker and a 50-ampere circuit breaker which are
not accessible to the flight crew. The 50-ampere circuit breakers are located in the power distribution
panel under the floor ahead of the main spar. The
5-ampere circuit breakers are located on panels forward of the instrument panel.
b. Normal Operation. Two levels of heat are
provided through the three position switches placarded NORMAL in the aft position, OFF in the center position, and HI after lifting the switch over a
detent and moving it to the forward position. In the
NORMAL position, heat is provided for the major
portion of each windshield. In the HI position, heat
is provided at a higher watt density to a smaller portion of the windshield. The lever lock feature prevents inadvertent switching to the HI position during system shutdown.
surization to the cabin at a rate of approximately 10
to 17 pounds per minute. The flow control unit of
each engine controls the bleed air from the engine to
make it usable for pressurization by mixing ambient
air with the bleed air depending upon aircraft altitude and ambient temperature. On takeoff, excessive
pressure bumps are prevented by landing gear safety
switch actuated solenoids incorporated in the flow
control units. These solenoids, through a time delay,
stage the input of ambient air flow by allowing
ambient air flow introduction through the left flow
control unit first, ten seconds later, air flow through
the right flow control unit. The bleed air switches,
located on the overhead control panel (fig. 2-12)
operate an integral electric solenoid which controls
the bleed air to the firewall shutoff valves.
b. Pressure Differential. The pressure vessel is
designed for a normal working pressure differential
of 6.0 PSI, which will provide a cabin pressure altitude of 3870 feet at an aircraft altitude of 20,000
feet, and a nominal cabin altitude of 9840 feet at an
aircraft altitude of 31,000 feet.
c. Cabin Altitude and Rate-of-Climb Controller. A control panel is installed on the copilot’s side
of the subpanel (fig, 2-6) for operation of the system.
A knob, placarded INC RATE controls the rate of
change of pressurization. A control, placarded
CABIN CONTROLLER is used to set the desired
cabin altitude. For proper cabin pressurization, the
CABIN CONTROLLER should be set 500 feet
above cruise altitude. For landing select 500 feet
above field pressure altitude. The selected altitude is
displayed on a mechanically coupled dial above the
control, placarded CABIN ALT-FT. Mechanically
coupled to the cabin altitude dial, placarded
ACFTX 1000. This dial indicates the maximum altitude the aircraft may be flown at to maintain the
desired cabin altitude without exceeding the design
pressure differential. A switch, placarded CABIN
PRESS DUMP-PRESS-TEST, is provided to control
pressurization. The switch is spring loaded to the
PRESS position. In the DUMP position, the safety
valve will be opened and the cabin will be depressurized to the aircraft altitude. In the PRESS position,
cabin altitude is controlled by the CABIN CONTROLLER control. In the TEST position, the landing gear safety switch is bypassed to enable testing of
the pressurization system on the ground. Operating
instructions are contained in Chapter 8.
2-59. PRESSURIZATION SYSTEM.
d. Cabin Rate-of-Climb Indicator. An indicator, placarded CABIN CLIMB, is installed on the
copilot’s side of the instrument panel (fig. 2-29). The
cabin rate-of-climb controller is calibrated in thousands-of-feet per-minute change in cabin altitude.
a. Description. A mixture of bleed air from
the engines, and ambient air, is available for pres-
e. Cabin Altitude Indicator. An indicator,
placarded CABIN ALT, is installed in the instru-
2-45
TM 55-1510-221-10
ment panel (fig. 2-29) above the cabin rate-of-climb
indicator. The longer needle indicates aircraft altitude in thousands-of-feet on the outside dial. The
shorter needle indicates pressure differential in PSI
on the inner dial. Maximum differential is 6.1 PSI.
integral electric solenoid which controls the bleed air
to the firewall shutoff valves. A normally open solenoid operated by the landing gear safety switch controls the introduction of ambient air flow to the
cabin on takeoff.
f. Outflow Valve. A pneumatically operated
outflow valve, located on the aft pressure bulkhead,
maintains the selected cabin altitude and rate-ofclimb commanded by the cabin rate-of-climb and
altitude controller on the copilot’s instrument panel.
As the aircraft climbs, the controller modulates the
outflow valve to maintain a selected cabin rate of
climb and increases the cabin differential pressure
until the maximum cabin pressure differential is
reached. At a cabin altitude of 12,500 feet a pressure
switch mounted on the back of the overhead control
panel completes a circuit to illuminate a red warning
annunciator light, ALT WARN, to warn of operation requiring oxygen. This light is protected by a
5-ampere breaker, placarded PRESS CONTR.
(1.) The unit receives bleed air from the
engine into an ejector which draws ambient air into
the nozzle of the venturi. The mixed air is then
forced into the bleed air line routed to the cabin.
g. Pressurization Safety Valve. Before takeoff,
the safety valve is open with equal pressure between
the cabin and the outside air. The safety valve closes
on liftoff if the CABIN PRESS CONTR switch on
the instrument panel is in the PRESS mode. The
safety valve adjacent to the outflow valve provides
pressure relief in the event of failure of the outflow
valve. This valve is also used as a dump valve and
is opened by vacuum which is controlled by a solenoid valve operated by the cabin pressure dump
switch adjacent to the controller. It is also wired
through a landing gear safety switch. If either of
these switches is open, or the vacuum source or electrical power is lost, the safety valve will close to
atmosphere except at maximum differential pressure
of 6.1 PSI. A negative pressure relief diaphragm is
also incorporated into the outflow and safety valves
to prevent outside atmospheric pressure from
exceeding cabin pressure during rapid descent.
h. Drain. A drain in the outflow valve static
control line is provided for removal of accumulated
moisture. The drain is located behind the lower sidewall upholstery access panel in the baggage section
of the aft compartment.
i. Flow Control Unit. A flow control unit forward of the firewall in each nacelle controls bleed air
flow and the mixing of ambient air to make up the
total air flow to the cabin for pressurization, heating,
and ventilation. The bleed air switches located on
the overhead control panel (fig. 2-12) operates an
2-46
(2.) Bleed air flow is controlled automatically. When the aircraft is on the ground, circuitry
from the landing gear safety switch prevents ambient air from entering the flow control unit to provide maximum heating.
(3.) The bleed air firewall shutoff valve in
the control unit is a spring loaded, bellows operated
valve that is held in the open position by bleed air
pressure. When the electric solenoid is shut off, or
when bleed air diminishes on engine shutdown (in
both cases the pressure to the firewall shutoff valve
is cut off), the firewall valve closes.
2-60. OXYGEN SYSTEM.
a. Description. The oxygen system (fig. 2-19)
is provided primarily as an emergency system, however, the system may be used to provide supplemental (first aid) oxygen. Two 70 cubic foot capacity
oxygen supply cylinders charged with aviator’s
breathing oxygen are installed in the unpressurized
portion of the aircraft behind the aft pressure bulkhead. The pilot and copilot positions are equipped
with diluter demand type regulators, which mix the
proper amount of oxygen for a given amount of air
at altitude. Also a first aid oxygen mask is provided
in the cabin. Oxygen system pressure is shown by
two gages placarded OXYGEN SUPPLY PRESSURE, located aft of the pilot’s oxygen regulator
control panel. Two pressure reducers, located in the
unpressurized portion of the aircraft behind the aft
bulkhead, lower the pressure in the system to 400
PSI, and route oxygen to the regulator control panels. Both cylinders are interconnected, so refilling
can be accomplished through a single tiller valve
located on the aft right side of the fuselage exterior.
A pressure gage is mounted in conjunction with the
filler valve, and each cylinder has a pressure gage.
Table 2-4 shows oxygen flow planning rates vs altitude. Table 2-5 shows oxygen duration capacities of
the system.
TM 55-1510-221-10
--------------LOW PRESSURE SYSTEM
HIGH PRESSURE SYSTEM
APOO6599
Figure 2- 19. Oxygen System Schematic
2-47
TM 55-1510-221-10
Table 2-4. Oxygen Flow Planning Rates vs Altitude
(All Flows In LPM Per Mask At NTPD)
CABIN PRESSURE
ALTITUDE IN FEET
CREW MASK
NORMAL
(DILUTER
DEMAND)
(1)
CREW MASK
100%
(1)
PASSENGER
MASK
3 1,000
30,000
29,000
28,000
27,000
26,000
25,000
24,000
23,000
22,000
2 1 ,000
20,000
19,000
18,000
17,000
16,000
15,000
14,000
13,000
12,000
11,000
10,000
-0-( 2)
-0-( 2)
-0-( 2)
-0-( 2)
-0-( 2)
-0-( 2)
-0-( 2)
-0-( 2)
-0-( 2)
-0-( 2)
-0-( 2)
3.6
3.9
4.2
4.5
4.8
5.1
5.4
5.8
6.1
6.5
6.9
4.2
4.4
4.7
5.0
5.3
5.6
5.9
6.2
6.6
6.9
7.2
7.6
7.9
8.3
8.7
9.1
9.5
10.0
10.4
10.9
11.3
11.9
3.7 (3)
3.7 (3)
3.7 (3)
3.7 (3)
3.7 (3)
3.7 (3)
3.7
3.7
3.7
3.7
3.7
3.7
3.7
3.7
3.7
3.7
3.7
3.7
3.7
3.7
3.7
3.7
NOTES:
(1) Based on minute volume of 20 LPM-BTPS (Body Temperature and Pressure Saturated).
(2) Use 100% oxygen at or above 20,000 feet.
(3) Not recommended for other than emergency descent use above 25,000 feet.
If average climb or descent flows are desired, add the values between altitudes and divide by the number
of values used.
For example, to determine the average rate for a uniform descent between 25,000 feet and 15,000 feet
perform the following.
(5.9 + 6.2 + 6.6 + 6.9 + 7.2 + 7.6 + 3.9 + 4.2 + 4.5 + 4.8 + 5.1) ÷ 11 = 5.7 LPM
This method is preferred over averaging the extremes an some flow characteristics vary in such a way as
to yield as incorrect answer.
2-48
TM 55-1510-221-10
Table 2-5. Oxygen Duration In Minutes 140 Cubic Foot System
TWO MAN
CREW
TWO MAN
CREW PLUS
ONE PASS
CABIN
PRESSURE
ALTITUDE
CREW
MASK
CONDITION
TOTAL
FLOW
LPM-NTPD
DURATION
IN
MINUTES (1)
3 1,000
25,000
20,000
20,000
15,000
15,000
10,000
100%
100%
100%
NORMAL
100%
NORMAL
100%
8.4
11.8
15.2
7.4
19.0
10.2
23.8
384.0
273.3
212.2
448.0
169.7
316.2
135.5
10,000
3 1,000
25,000
20,000
20,000
15,000
15,000
10,000
10,000
NORMAL
100%
100%
100%
NORMAL
100%
NORMAL
100%
NORMAL
13.8
12.1
15.5
18.9
10.9
22.7
13.9
27.5
17.5
233.7
266.6
208.1
170.0
295.9
142.1
232.1
117.3
184.3
(1) For 100% capacity of useable oxygen, 3,226 L.
2-61. OXYGEN DURATION EXAMPLE PROBLEM
WANTED
Duration in minutes of oxygen at 100% capacity.
WANTED
Duration of oxygen for complement at other
cabin pressure altitude, at less than 100% capacity.
KNOWN
KNOWN
Two man crew plus one passenger, cabin pressure altitude = 15,000 feet, crew masks, normal,
100% capacity.
METHOD
Find “two man crew plus one pass” line, move
right then down to 15,000 - “normal” read “232.1”
minutes.
WANTED
Duration of oxygen for previous example data at
84% of capacity.
KNOWN
232.1 minutes duration at l00%, 84% capacity,
total aircraft flow = 13.9 LPM.
METHOD
Multiply 232.1 X 0.84 = 194.9 minutes. or Multiply 3,226 X 0.84 = 2709.8, divide by 13.9 LPM =
194.9 minutes.
Cylinder at 84% capacity, 100% capacity =
3,226 L, cabin pressure altitude = 21,000 feet.
1 crew mask = 7.2 LPM ( 100%) 1 passenger
mask = 3.7 LPM
METHOD
Multiply 3,226 L X 0.84 = 2.709.8 L, multiply
2 crew X 7.2 LPM = 14.4 LPM, multiply 1 passenger X 3.7 LPM, add 14.4 LPM crew plus 3.7 LPM
passenger = 18.1 LPM. Divide 3,226 L by 18.1
LPM = 178.2 minutes.
2-62. OXYGEN CYLINDER CAPACITY EXAMPLE
PROBLEM
WANTED
a. Percent of capacity at known pressure and
temperature.
b.
Pressure when temperature decreases.
2-49
TM 55-1510-221-10
STABILIZED CYLINDER TEMPERATURE
500
1,500
1,000
CYLINDER PRESSURE (PSIG)
2,000
AP010340
Figure 2-20. Cylinder Capacity vs Pressure and Temperature
KNOWN
Pressure = 1,600 PSIG stabilized cylinder temperature is estimated at 20º C decreased stabilized
cylinder temperature is estimated at -30º C.
GEN, and a green supply control lever placarded
ON - OFF. The diluter control lever selects either
normal or 100% oxygen, but acts to select only when
the emergency pressure control lever is in the NORMAL position.
METHOD
a. Enter 1600 PSIG move up to 20º C line,
move right to 84%.
b. Move left on 84% line to -30º C line, move
down to 1250 PSIG.
WANTED
100% capacity pressure at known temperature.
KNOWN
Temperature = -30º C.
METHOD
Move left along 100% line to -30º C line and
move down to 1420 PSIG.
(1.) Regulator control panels. Each regulator control panel contains a blinker-type flow indicator, a 500 PSI pressure gage, a red emergency pressure control lever placarded EMERGENCY NORMAL - TEST MASK, a white diluter control
lever placarded 100% OXYGEN - NORMAL OXY-
2-50
When not in use, the diluter control lever
should be left in the 100% OXYGEN
position to prevent regulator contamination.
(2.) 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 the NORMAL position,
the lever allows delivery of normal or 100% oxygen,
depending upon the selection of the diluter control
lever. In TEST MASK position, 100% oxygen at
positive pressure is delivered to check hose and
mask integrity.
(3.) The 500 PSI oxygen pressure gage
provided on the oxygen control panels should never
TM 55-1510-221-10
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. When oxygen
is in use, a check of the supply pressure
should be made at intervals during flight
to note the quantity available and to
approximate the supply duration. The
outside temperature is reduced as an aircraft ascends to higher altitudes. Oxygen
cylinders thus cooled by temperature
change will show a pressure drop. This
type of drop in pressure will raise again
upon return to a lower or warmer altitude. A valid cause for alarm would be
the rapid loss of oxygen pressure when the
aircraft is in level flight or descending;
should this condition arise, descend as
rapidly as possible to altitude which does
not require the use of oxygen.
WARNING
Pure oxygen will support combustion. Do
not smoke while oxygen is in use.
WARNING
If any symptoms occur suggestive of the
onset of hypoxia, immediately set the
emergency pressure control lever to the
EMERGENCY position and descend
below 10,000 feet. Whenever carbon
monoxide or other noxious gas is present
or suspected, set the dilutor control lever
to 100% OXYGEN and continue breathing undiluted oxygen until the danger is
past.
e. 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 pro-
vided with a cord for connecting with the microphone jack. To test mask and hose integrity, the
individual places the supply control lever on the regulator control panel to the ON position, puts on and
adjusts his mask, selects TEST MASK position, and
checks for leaks.
f. Normal Operation. Oxygen pressure is
maintained at all times to the regulator control panels if the cylinder shut-off valves are on and if there
is pressure in the cylinders. Each individual places
the supply lever (green) on his regulator control
panel to the ON position, and the diluter lever
(white) to the NORMAL OXYGEN position.
g. Emergency Operation. For emergency operation, the affected crew member selects the EMERGENCY position of the emergency pressure control
lever (red) on his regulator control panel. This selection provides 100% oxygen at a positive pressure,
regardless of the position of the diluter control lever
on his panel.
h. First Aid Operation. A first aid oxygen
mask is installed in the aft cabin area as a supplemental or emergency source of oxygen. The mask is
stowed behind an overhead cover placarded FIRST
AID OXYGEN -PULL. Removing the cover allows
the mask to drop out of the container, exposing a
manual control valve, which releases oxygen to the
mask when placed in the ON position. After using
the mask, the manual valve in the container must be
turned OFF before stowing the mask and replacing
the cover.
2-63. WINDSHIELD WIPERS.
a. Description. Two electrically operated
windshield wipers, are provided for use at takeoff,
cruise and landing speed. A rotary switch (fig. 2-12)
placarded WINDSHIELD WIPER, located on the
overhead control panel, selects mode of windshield
wiper operation. An information placard above the
switch states: DO NOT OPERATE ON DRY
GLASS. Function positions on the switch, as read
clockwise, are placarded: PARK - OFF - SLOW FAST. When the switch is held in the 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 lo-ampere circuit
breaker, placarded WSHLD WIPER, located on the
overhead circuit breaker panel (fig. 2-26).
2-51
TM 55-1510-221-10
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 position will stop the windshield wipers, without returning them to the normal
inactive position.
2-64. FERRY CHAIR.
container located below the seat in the cabinet
assembly. This non-flushing system uses a dry chemical preparation to deodorize the stored waste. A toilet tissue dispenser is contained in a slide-out compartment on the forward side of the toilet cabinet. A
box of disposable waste container liners and a box
of chemicial deodorant packets are also stored in the
cabinet.
b. Operation. During use, a removable, throwaway plastic liner is attached to the waste container.
After use, dry chemical deodorant obtained from the
storage cabinet is deposited on the waste and the
hinged lid sections are closed over the cavity. After
each flight, the waste container must be removed,
emptied, relined, replaced in the cabinet and other
toilet items are resupplied as needed,
2-67. SUN VISORS.
For ferry purposes, a forward facing chair with
a lap belt is attached to floor tracks at fuselage station 211.87 (fig. 2-2).
2-65. CIGARETTE LIGHTERS AND ASH TRAYS.
The pilot and copilot have individual cigarette
lighters and ash trays mounted in escutcheons outboard of their seats. The cigarette lighters are protected by a 5-ampere circuit breaker, placarded
CIGAR LIGHTER, on the overhead circuit breaker
panel (fig. 2-26).
2-66. CHEMICAL TOILET.
a. Description. A side-facing chemical toilet
(figure 2-2) is installed in the aft cabin area. Two
hinged lid half-sections must be raised to gain access
to the toilet. Waste is stored within a removable
Section VIII.
2-69.
A sun visor is provided for the pilot and copilot
respectively (fig. 2-8). 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.
2-68. RELIEF TUBE.
One relief tube is provided, located immediately
aft of the cabin door on the left side of the fuselage.
HEATING, VENTILATION, COOLING, AND ENVIRONMENTAL Control System
HEATING SYSTEM.
a. Description. Warm air for heating the cockpit and mission avionics compartments and warm
windshield defrosting air is provided by bleed air
from both engines. Engine bleed air is combined
with ambient air in the heating and pressurization
flow control unit in each nacelle. If the mixed bleed
air is too warm for cockpit comfort, it is cooled by
being routed through an air-to-air heat exchanger
located in the forward portion of each inboard wing.
If the mixed bleed air is not too warm, the air-to-air
heat exchangers are bypassed. The mixed bleed air
is then ducted to a mixing plenum, where it is mixed
with cabin recirculated air. The warm air is then
2-52
When adjusting the sun visors, grasp only
by the top metal attachment to avoid
damage to the plastic shield.
ducted to the cockpit outlets, windshield defroster
outlets, and to the floor outlets in the mission avionics compartment. The environmental system is
shown in figure 2-21. placarded WSHLD W1
(1.) Bleed airflow control unit. A bleed air
flow control unit, located forward of the firewall in
each engine nacelle controls the flow of bleed air
and the mixing of ambient air to make up the total
airflow to the cabin for heating, windshield defrosting, pressurization and ventilation. The unit is fully
pneumatic except for an integral electric solenoid
firewall shutoff valve, controlled by the bleed air
switches located on the overhead control panel (fig.
TM 55-1510-221-10
Figure 2-2 1. Environmental System
2-53
TM 55-1510-221-10
2-l 2) and a normally open solenoid valve operated
by the right landing gear safety switch.
(2.) Pneumatic bleed air shutoff valve. A
pneumatic shutoff valve is provided in each nacelle
to control the flow of bleed air to the surface,
antenna and brake deice systems. These valves are
controlled by the bleed air valve switches located on
the overhead control panel (fig. 2-12).
(3.) Bleed air valve switches. The bleed air
flow control unit shutoff valve and pneumatic bleed
air shutoff valves are controlled by two switches
placarded BLEED AIR VALVE - OPEN - ENVIRO
OFF - PNEU & ENVIRO OFF, located on the overhead control panel (fig. 2-12). When set to the open
position, both the environmental flow control unit
shutoff valve and the pneumatic shutoff valve are
open; when set to the ENVIRO OFF position, the
environmental flow control unit shutoff valve is
closed, and the pneumatic bleed air valve is open; in
the PNEU & ENVIRO OFF position, both are
closed. For maximum cooling on the ground, turn
the bleed air valve switches to the ENVIRO OFF
position.
(4.) Cabin temperature mode selector
switch. A switch placarded CABIN TEMP MODE MAN COOL - MAN HEAT -OFF - AUTO - A/C
COLD OPN - -25°C to 10°C located on the overhead control panel, controls cockpit and mission
avionics compartment heating and air conditioning.
When the cabin temperature mode selector switch is
set to the AUTO position, the heating and air conditioning systems are automatically controlled. Control signals from the temperature control box are
transmitted to the bleed air heat exchanger bypass
valves. Here the temperature of the air flowing to
the cabin is regulated by the bypass valves controlling the amount of air bypassing the heat exchangers. When the temperature of the cabin has reached
the temperature setting of the cabin temperature
control rheostat, the automatic temperature control
allows hot air to bypass the air-to-air exchangers
admitting hot air into the cabin. When the bypass
valves are in the fully closed position, allowing no
air to bypass the heat exchangers, the air conditioner
begins to operate, providing additional cooling.
When the cabin temperature mode selector switch is
set to the A/C COLD OPN position, the air conditioning system is in continuous operation. The cabin
temperature control rheostat, in conjunction with
the cabin temperature control sensor, provides regulation of cockpit and mission equipment compartment temperature. Bleed air heat is added as
required to maintain the temperature selected by the
cabin temperature control rheostat.
(5.) Cabin temperature control rheostat. A
control knob placarded CABIN TEMP - INCR,
2-54
located on the overhead control panel (fig. 2-12),
provides regulation of cabin temperature when the
cabin temperature mode selector switch is set to the
AUTO or the A/C COLD OPN position. A temperature sensing unit in the cabin, in conjunction with
the setting of the cabin temperature control rheostat,
initiates a heat or cool command to the temperature
controller for the desired cockpit or mission avionics
compartment environment.
(6.) Manual temperature control switch. A
switch placarded MANUAL TEMP - INCR DECR, located on the overhead control panel (fig.
2-12), controls cockpit and mission avionics compartment temperature with the cabin temperature
mode selector switch set to the MAN HEAT positions. The manual temperature control switch controls cockpit and mission avionics temperature by
providing a means of manually changing the amount
that the bleed air bypass valves are opened or
closed. To increase cabin temperature the switch is
held to the INCR position. To decrease cabin temperature, the switch is held to the DECR position.
Approximately 30 seconds per valve is required to
drive the bypass valves to the fully open or fully
closed position. Only one valve moves at a time.
(7.) Forward vent blower switch. The forward vent blower is controlled by a switch placarded
VENT BLOWER - AUTO - LO - HI, located on the
overhead control panel (fig. 2-12). In the auto position the fan will run at low speed except when the
cabin temperature mode selector switch is set to the
OFF position, in this case the blower will not operate.
(8.) Aft vent blower switch. The aft vent
blower is controlled by a switch placarded AFT
VENT BLOWER - OFF - AUTO - ON, located on
the overhead control panel (fig. 2-l 2). The single
speed blower operates automatically through the
cabin temperature mode selector switch when the aft
vent blower switch is placed in the AUTO position.
The blower runs continuously when the switch is
placed in the ON position, In the OFF position, the
blower will not operate.
b. Automatic Heating Mode.
1.
Bleed air valve switches - OPEN, LEFT and
RIGHT.
2. Cabin temperature mode selector switch AUTO.
3. Cabin temperature control rheostat - As
required.
4. Cabin, cockpit and defrost air knobs - As
required
TM 55-1510-221-10
c. Manual heating mode.
1. Bleed air valve switches - OPEN, LEFT and
RIGHT.
When compressor operation has been terminated by
limit switch activation, the system should be thoroughly checked before returning it to service.
2.
Cabin temperature mode selector switch MAN HEAT.
3.
Vent blower switches - As required.
4.
Manual temperature switch - As required.
5.
Cabin, cockpit and defrost air knobs - As
required.
(4.) Thermal sense switch. A thermal sense
switch is installed on the forward evaporator. This
sense switch actuates a hot gas bypass valve which
bypasses a portion of the refrigerant from the forward evaporator, thereby preventing icing of the
evaporator.
2-70.
AIR CONDITIONING SYSTEM.
a. Description. Cabin air conditioning is provided by a refrigerant gas vapor cycle refrigeration
system consisting of a belt driven, engine mounted
compressor, installed on the No.2 engine accessory
pad, refrigerant plumbing, N1 speed switch, high and
low pressure protection switches, condenser coil,
condenser-under-pressure switch, condenser blower,
forward and aft evaporator, receiver-dryer, expansion valve and a bypass valve. The plumbing from
the compressor is routed through the right inboard
wing leading edge to the fuselage and then forward
to the condenser coil, receiver-dryer, expansion
valve, bypass valve, and forward evaporator, which
are located in the nose of the aircraft. A 7 1/2ampere circuit breaker placarded AIR COND
CONTR, located on the overhead control panel (fig.
2-12), protects the compressor clutch circuit.
(1.) Forward evaporator. The forward
evaporator and blower supplies the cockpit, forward
ceiling outlets, and forward floor outlets. The forward evaporator blower has a high speed which can
be selected by setting the VENT BLOWER switch,
located on the overhead control panel (fig. 2-12), to
the HI position. The forward vent blower is protected by a circuit breaker on the DC power distribution panel, located below the aisleway floor forward of the main spar.
(2.) Aft evaporator. The aft evaporator and
blower are located in the fuselage center aisle equipment bay aft of the rear spar. Environmental air is
circulated through the evaporator in either manual
or automatic control mode. The rear evaporator
supplies the aft ceiling outlets, rear floor outlets, and
toilet compartment. Rear evaporator blower is protected by a circuit breaker on the DC power distribution panel located below the aisleway floor forward of the main spar.
(3.) High and low pressure limit switches.
High and low pressure limit switches are provided to
prevent compressor operation beyond operational
limits. When the low or high pressure switches are
activated, compressor operation will be terminated.
(5.) Condenser blower. A vane-axial
blower draws air through the condenser on the
ground as well as in flight. The current limiter for
the blower is located on the DC distribution panel
below the aisleway floor forward of the main spar.
When the cabin temperature mode selector switch is
set to the A/C COLD OPN position, the condenser
blower will be off, and will remain off until the condenser blower control high pressure switch senses a
compressor discharge pressure equal to the pressure
it is set to. The condenser blower will then remain
in operation until the low pressure switch senses that
the system pressure has dropped to the pressure it is
set to.
(6.) Air conditioning cold operation bypass
valve. Selecting the A/C COLD OPN mode on the
cabin temperature mode selector switch permits the
operation of the air conditioning system by overriding the refrigerant low pressure switch. This allows
the air conditioning system to operate in the manual
mode. Starting the compressor in this optional mode
at low ambient temperature will decrease the operational life of the compressor by five hours each time
the air conditioning system is started using this
mode (A/C COLD OPN). If the air conditioning system has been operating in the normal mode during
flight, and due to decreasing ambient temperatures
make it necessary to switch to the A/C cold operation mode, there will be no degradation in the mean
time between failure for the compressor.
(7.) Air conditioner cold operation advisory
annunciator light. A green advisory annunciator
light placarded A/C COLD OPN, located on the
caution/advisory annunciator panel (fig. 2.6), illuminates when the air conditioning system is operating
in cold mode, or when ambient temperatures require
switching to cold mode if air conditioning system
operation is to be continued.
b. Normal Operation.
(1.) Automatic cooling mode.
1.
Bleed air valve switches - OPEN,
LEFT and RIGHT.
2. Cabin temperature mode selector
switch - AUTO.
2-55
TM 55-1510-221-10
3. Cabin temperature control rheostat - As required.
4. Cabin, cockpit and defrost air
knobs - As required.
(2.) Manual cooling mode.
1.
Bleed air valve switches - OPEN,
LEFT and RIGHT.
NOTE
For maximum cooling on the ground, set
the bleed air valve switches to the
ENVIRO OFF position.
2.
Cabin temperature mode selector
switch - MAN COOL.
(3.) Air conditioning cold operation mode.
(Used if ambient temperature is between 10°C and
-25º C).
NOTE
Setting the cabin temperature mode selector switch to the A/C COLD OPN position at ambient temperatures below -25°C
may cause the air conditioning system to
exceed the compressor low pressure limit
switch setting, terminating compressor
operation, and thereby rendering the system inoperative for the remainder of the
flight.
1.
tributed through the main ducting system to all outlets. Ventilation air, ducted to each individual eyeball cold air outlet, can be directionally controlled
by moving the ball in the socket. Volume is regulated by twisting the outlet to open or close the
valve.
2-72. ENVIRONMENTAL CONTROLS.
An environmental control section on the overhead control panel (fig. 2-12) provides for automatic
or manual control of the system. This section contains all the major controls of the environmental
function including bleed air valve switches, a vent
blower control switch, an aft vent blower switch, a
manual temperature switch for control of the heat
exchanger valves, a cabin temperature level control,
and the cabin temp mode selector switch for selecting automatic heating or cooling or manual heating
or cooling. Four additional manual controls on the
main instrument subpanels may be utilized for partial regulation of cockpit comfort when the cockpit
partition door is closed and the cabin comfort level
is satisfactory.
a. Heating Mode.
(1.)
If the cockpit is too cold:
1.
Pilot and copilot air knobs - As
required.
2.
Defrost air knob - As required.
3.
Cabin air knob - Pull out in small
increments. Allow 3 -5 minutes
after each adjustment for system
to stabilize.
Bleed air valve switches - OPEN,
LEFT and RIGHT.
2. Cabin temperature mode selector
switch - A/C COLD OPN.
(2.)
If the cockpit is too hot:
1.
Cabin air knob - As required.
3. Cabin temperature control rheostat - As required.
2.
Pilot and copilot air knobs - In as
required.
4. Cabin, cockpit and defrost air
knobs - As required.
3.
Defrost air knob - In as required.
b. Cooling Mode:
2-71. UNPRESSURIZED VENTILATION.
Ventilation is provided by two sources. One
source is through the bleed air heating system in
both the pressurized and unpressurized mode. The
second source of ventilation is obtained from ram
air through the condenser section in the nose
through a check valve in the vent blower plenum.
Ventilation from this source is in the unpressurized
mode only with the CABIN PRESS DUMP switch
in the DUMP position. The check valve closes during pressurized operation. Ram air ventilation is dis-
2-56
(1.)
(2.)
If the cockpit is too cold:
1.
Pilot and copilot air knob - In as
required.
2.
Defrost air knob - In as required.
3.
Overhead cockpit outlets - As
required.
If the cockpit is too hot:
1.
Pilot and copilot air knobs - Out
as required.
TM 55-1510-221-10
2.
Cabin air knob. Close in small
increments. Allow 3 - 5 minutes
after each adjustment for system
to stabilize. If CABIN AIR knob
is completely
closed before
obtaining satisfactory cockpit
comfort, it may be necessary to
place the aft vent blower switch in
the ON position to activate the
aft evaporator to recirculate cabin
air.
c. Automatic Mode Control. When the AUTO
mode is selected on the cabin temperature mode
selector switch, the heating and air conditioning systems are automatically controlled. When the temperature of the cabin has reached the selected setting,
the automatic temperature control allows heated air
to bypass the air-to-air exchangers in the wing center
section. By-pass air is hot. Heat exchanger air is
cooled to approximately 30°F above ambient (outside) air. The warm bleed air is mixed with the
cooled air. The rear evaporator picks up recirculated
cabin air only.
NOTE
The automatic mode control works as
described, when the cabin is being cooled
by bleed air. However, when the cabin is
heated with bleed air and the selected
temperature is reached, hot bleed air
routes through the heat exchanger for
cooling in order to maintain the desired
temperature.
(1.) When the automatic control drives
the environmental system from a heat mode to a
cooling mode, the bypass valves close. When the left
bypass valve reaches a fully closed position, the
refrigeration system will begin cooling, provided the
right engine N1 speed is above 65%. When the
bypass valve is opened to a position approximately
30º from full open, the refrigeration system will turn
Off.
(2.) The CABIN TEMP - INCR control
provides regulation of the temperature level in the
automatic mode. A temperature sensing unit in the
cabin, in conjunction with the control setting, initiates a heat or cool command to the temperature
controller for desired cockpit and cabin environment.
d. Manual Mode Control. With the cabin temperature mode selector in the MAN HEAT or MAN
COOL position, regulation of the cabin temperature
is accomplished manually with the MANUAL
TEMP switch.
(1.) In the MAN HEAT mode, the automatic system is overridden and the system is controlled by opening and closing the bypass valves
(two) with the MANUAL TEMP - INCR - DECR
switch. To increase cabin temperature, hold the
switch at the INCR position, to decrease cabin temperature, hold the switch in the DECR position.
Allow approximately 30 seconds per valve to drive
the bypass valves to the fully open or fully closed
position. Only one valve moves at a time.
(2.) With the cabin temperature selector
switch in the MAN COOL position, the automatic
temperature control system is bypassed. When the
left bypass valve reaches a fully closed position, the
refrigeration system will begin cooling, provided the
right engine N1 speed is above 65%. When the
bypass valve is opened to a position approximately
30º from full open, the refrigeration system will turn
off. Hold the MANUAL TEMP switch in the DECR
position approximately one minute to fully close airto-air heat exchanger bypass valves.
(3.) Bleed air entering the cabin is controlled by bleed air valve switches placarded BLEED
AIR VALVE - OPEN - ENVIRO OFF -PNEU &
ENVIRO OFF. When the switch is in the OPEN
position, the environmental flow control unit and
the pneumatic valve are open. When the switch is in
the ENVIRO OFF position, the environmental flow
control unit is closed and the pneumatic bleed air
valve is open; in the PNEU & ENVIRO OFF position, both are closed. For maximum cooling on the
ground, turn the bleed air valve switches to the
ENVIRO OFF position.
(4.) The forward vent blower is controlled
by a switch placarded VENT BLOWER - AUTO LOW - HI. The HI and LOW positions regulate the
blower to two speeds of operation. IN the AUTO
position, the fan will run at low speed except when
the CABIN TEMP mode selector switch is placed in
the OFF position. In the OFF position, the blower
will not operate.
(5.) The aft vent blower is controlled by a
switch placarded AFT VENT BLOWER - OFF AUTO - ON. The single speed blower operates automatically through the CABIN TEMP mode selector
when the AFT VENT BLOWER switch is placed in
the AUTO position. The blower runs continuously
when the switch is placed in the ON position. In the
OFF position, the blower will not operate.
2-57
TM 55-1510-221-10
Section IX. ELECTRICAL POWER SUPPLY AND DISTRIBUTION SYSTEM
2-73. DESCRIPTION.
The aircraft employs both direct current (DC)
and alternating current (AC) electrical power. The
DC electrical power supply (fig. 2-22) is the basic
power system energizing most aircraft circuits. Electrical power is used to start the engines, to power the
landing gear and flap motors, and to operate the
standby fuel pumps, ventilation blower, lights and
electronic equipment. AC power is obtained from
DC power through inverters. The single phase AC
power system is shown in figure 2-23, and the three
phase AC power system is shown in figure 2-24. The
three sources of DC power consist of one 20 cell 34ampere/hour battery and two 400-ampere startergenerators. DC power may 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 (refer to Section XII for GPU
requirements). The starter-generators are controlled
by generator control units. The output of each generator passes through a cable to the respective generator bus (fig. 2-22). Other buses distribute power to
aircraft DC loads, and derive power from the generator buses. The generators are paralleled to balance
the DC loads between the two units. When one of
the generating systems is not on- line, and no fault
exists, all aircraft DC requirements may be supplied
either by the other on-line generating system, or by
an external power source, but not by both. Most DC
distribution buses are connected to both generator
buses but have isolation diodes to prevent power
crossfeed between the generating systems, when connection between the generator buses is lost. Thus,
when either generator is lost because of a ground
fault, the operating generator will supply power for
all aircraft DC loads except those receiving power
from the inoperative generator’s bus which cannot
be crossfed. When a generator is not operating,
reverse current and over-voltage protection is automatically provided. Two inverters operating from
DC power produce the required single-phase AC
power. Three phase AC electrical power for inertial
navigation system and mission avionics is supplied
by two DC powered mission inverters. AC power
may be applied through an external power receptacle
located on the left nacelle. The mission power system is shown in figure 2-25.
2-74. DC POWER SUPPLY.
One nickel-cadmium battery furnishes DC
power when the engines are not operating. This 24volt, 34-ampere/hour battery, located in the right
wing center section, is accessible through a panel on
the top of the wing. DC power is produced by two
2-58
starterengine-driven 28 volt, 400-ampere
generators. Controls and indicators associated with
the DC supply system are located on the overhead
control panel (fig. 2-12) and consist of a single battery switch (BATT), two generator switches (No.1
GEN and No.2 GEN), and two volt-loadmeters.
a. Battery Switch. A switch, placarded BATT
(fig. 2-12) is located on the overhead control panel
under the MASTER SWITCH. The BATT switch
controls DC battery 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 BATT
switch is forced OFF.
NOTE
With battery or external power removed
from the aircraft electrical system, due to
fault, power cannot be restored to the system until the BATT switch is moved to
OFF/RESET, then ON.
b. Generator Switches. Two switches (fig.
2-12), placarded No. 1 GEN and No.2 GEN are
located on the overhead control panel under the
MASTER SWITCH. The toggle switches control
electrical power from the designated generator to
paralleling circuits and the bus distribution system.
Switch positions are placarded RESET, ON and
OFF. RESET is forward (spring-loaded back to ON),
ON is center, and OFF is aft. When a generator is
removed from the aircraft electrical system, due
either to fault or from placing the GEN switch in the
OFF position, the affected unit cannot have its output restored to aircraft use until the GEN switch is
moved to RESET, then ON.
c. Master Switch. All electrical current may be
shut off using the MASTER SWITCH gangbar (fig.
2-12) which extends above the battery and generator
switches. The MASTER SWITCH gangbar is moved
forward when a battery or generator switch is turned
on. When moved aft, the bar forces each switch to
the OFF position.
d. Volt-Loadmeters. Two meters (fig. 2-12),
on the overhead control panel display voltage readings and show the rate of current usage from left and
right generating systems. Each meter is equipped
with a spring-loaded pushbutton switch which when
manually pressed will cause the meter to indicate
main bus voltage. Each meter normally shows the
output amperage reading from the respective generator, unless the pushbutton switch is pressed to
obtain the bus voltage reading. Current consumption
TM 55-1510-221-10
Figure 2-22. DC Electrical System (Sheet 1 of 3)
2-59
TM 55-1510-221-10
#1 AVIONICS BUS
PILOT AUDIO
TRANSPONDER
UHF
TACAN
INS CONTROL
HF RCVR
#1 VHF
#1 VOR
#1 RMI
AFCS DIRECT
AP PWR
VHF/AM/FM
#2 AVIONICS BUS
SERVO DC
RADIO RELAY
COPILOT ALT
BU VOW
COPILOT AUDIO
RADAR
RADAR-NAV
#2 VOR
ADF
#2 RMI
#1 DUAL FED BUS
ANN IND
#1 CHIP DETR
#1 QTY IND
#1 QTY WARN
#1 OIL TEMP
STALL WARN
LANDING GEAR IND
#1 STANDBY PUMP
#1 OIL PRESS
LEFT BLEED AIR WARN
#1 AUXILIARY TRANSFER
RADOME ANTI-ICE
#1 ENG AIR SCOOP HEAT
#2 DUAL FED BUS
ANN PWR
#2 CHIP DETR
#2 QTY IND
#2 QTY WARN
#2 OIL TEMP
BATT CHARGE
FIRE DETR
LANDING GEAR WARN
#2 STANDBY PUMP
X2 OIL PRESS
RIGHT BLEED AIR WARN
#2 AUXILIARY TRANSFER
#2 ENG AIR SCOOP HEAT
ENG AIR SCOOP HEAT MONITOR
#3 DUAL FED BUS
A
LEFT PROP ANTI-ICE
LEFT FUEL VENT HEAT
#1 FIREWALL VALVE
#1 ICE VANE CONTR
WSHLD WIPER
SURF DEICE
LEFT PITOT HEAT
CROSSFEED
#1 START CONTR
PROP SYNC
STALL WARN HEAT
BRAKE DEICE
RIGHT PITOT HEAT
#2 START CONTR
AUTOFEATHER
HF POWER
X4 DUAL FED BUS
A
RIGHT PROP ANTI-ICE
RIGHT FUEL VENT HEAT
#2 FIREWALL VALVE
#2 ICE VANE CONTR
PROP GOV
SCAVENGER PUMP
PROP ANTI-ICE AUTO
LEFT FUEL CONTR HEAT
#1 PRESS WARN
#1 IGNITOR CONTR
PROP ANTI-ICE CONTR
RIGHT FUEL CONTR HEAT
#2 PRESS WARN
#2 IGNITOR CONTR
APO06453.2
Figure 2-22. DC Electrical System (Sheet 2 of 3)
2-60
TM 55-1510-221-10
#5 DUAL FED BUS
ELEC TRIM
LANDING GEAR RELAY
ICE LIGHTS
INST INDIRECT LIGHTS
TEMP CONTR
FLAP MOTOR
BCN LIGHTS
LANDING LIGHTS
LEFT BLEED AIR CONTR
PROVISIONS
PILOT TURN & SLIP
SUBPANEL & CONSOLE LIGHTS
RECOGNITION LIGHTS
AIR COND CONTR
#6 DUAL FED BUS
RUDDER BOOST
EMERG LIGHTS
OVHD LIGHTS
PRESS CONTR
CIGAR LIGHTER
AVIONICS MASTER CONTR
FLAP CONTR
FLT INST LIGHTS
RIGHT BLEED AIR CONTR
TAXI LIGHT
COPILOT TURN & SLIP
NAV LIGHTS
CABIN LIGHTS
CARGO DOOR HEAT
HOT BATTERY BUS
#1 FIREWALL SHUTOFF VALVE
#1 ENGINE FIRE EXTINGUISHER
#1 STANDBY FUEL PUMP
TRANSPONDER
CABIN LIGHT
BATTERY RELAY
#2 FIREWALL SHUTOFF VALVE
#2 ENGINE FIRE EXTINGUISHER
#2 STANDBY FUEL PUMP
CRYTO HOLD
Figure 2-22. DC Electrical System (Sheet 3 of 3)
2-61
TM 55-1510-221-10
APOO6454
Figure 2-23. Single Phase AC Electrical System
2-62
Figure 2-24. Three Phase AC Electrical System
TM 55-1510-221-10
2-63
TM 55-1510-221-10
APOO6456
Figure 2-25. Mission Equipment DC Power System
2-64
TM 55-1510-221-10
is indicated as a percentage of total output amperage
capacity for the generating system monitored.
e. Battery Volt-Amp Meter. The mission control panel (fig. 4-l), located on the right inside fuselage sidewall adjacent to the copilot’s seat, has a battery-amperage meter that displays battery voltage on
the left side of the meter and battery current on the
right side of the meter. Minimum battery voltage for
engine start is 22 VDC.
f. Battery Monitor. Nickel-cadmium
battery
overheating will cause the battery charge current to
increase if thermal runaway is imminent. The aircraft has a charge-current sensor which will detect a
charge current. The charge current system senses
battery current through a shunt in the negative lead
of the battery. Any time the battery charging current
exceeds approximately 7 amperes for 6 seconds or
longer, the yellow BATTERY CHARGE annunciator light and the master fault caution light will illuminate. Following a battery engine start, the caution
light will illuminate approximately six seconds after
the generator switch is placed in the ON position.
The light will normally extinguish within two to five
minutes, indicating that the battery is approaching a
full charge. The time interval will increase if the battery has a low state of charge, the battery temperature is very low, or if the battery has previously been
discharged at a very low rate (i.e., battery operation
of radios or lights for prolonged periods). The caution light may also illuminate for short intervals
after landing gear and/or flap operation. If the caution light should illuminate during normal steadystate cruise, it indicates that conditions exist that
may cause a battery thermal runaway. If this occurs,
the battery switch shall be turned OFF and may be
turned back ON only for gear and flap extension and
approach to landing. Battery may be used after a 15
to 20 minute cool down period.
g. Generator Out Warning Lights. Two caution/advisory annunciator panel lights inform the
pilot when either generator is not delivering current
to the aircraft DC bus system. These lights are placarded No.1 DC GEN and No.2 DC GEN. Illumination of the two MASTER CAUTION lights and
either fault light indicates that either the identified
generator has failed or voltage is not sufficient to
keep it connected to the power distribution system.
The GPU shall be adjusted to regulate at
28 volts maximum to prevent damage to
the aircraft.
h. DC External Power Source. External DC
power can be applied to the aircraft through an
external power receptacle on the underside of the
right wing leading edge just outboard of the engine
nacelle. The receptacle is installed inside of the wing
structure and is accessible through a hinged access
panel. DC power is supplied through the DC external power plug and applied directly to the battery
bus after passing through the external power relay.
Turn off all external power while connecting the
power cable to, or removing it from, the external
power supply receptacle. The holding coil circuit of
the relay is energized by the external power source
when the BATT switch is in the ON position. The
GPU shall be adjusted to regulate at 28 volts maximum to prevent damage to the aircraft battery.
i. Security Keylock Switch, The aircraft has a
security keylock switch (fig. 2-12) installed on the
overhead control panel, placarded OFF - ON. The
switch is connected to the battery relay circuit and
must be ON when energizing the battery master
power switch. The key cannot be removed from the
lock when in the ON position.
j. Circuit Breakers. The overhead circuit
breaker panel (fig. 2-26) contains circuit breakers for
most aircraft systems. The circuit breakers on the
panel are grouped into areas which are placarded as
to the general function they protect. A DC power
distribution panel is mounted beneath the aisleway
floor forward of the main spar. This panel contains
higher current rated circuit breakers and is not
accessible to the flight crew under normal conditions.
2-75. AC POWER SUPPLY.
a. Single Phase AC Power Supply. AC power
for the aircraft is supplied by inverter units, numbered No. 1 and No.2 (fig. 2-23) which obtain operational current from the DC power system. Both
inverters are rated at 750 volt-amperes and provide
single-phase output only. Each inverter provides 115
volt and 26 volt, 400 Hz AC output. The inverters
are protected by circuit breakers mounted on the
DC power distribution panel beneath the aisleway
floor. Controls and indicators of the AC power system are located on the overhead control panel and
on the caution/advisory annunciator panel.
( 1 . ) A C P o w e r WARNING/CAUTION
Lights. Illumination of the two MASTER CAUTION lights and the illumination of an annunciator
caution light No.1 INVERTER or No.2 INVERTER
indicates an inverter failure.
(2.) Instrument AC Light. A red warning
light and two MASTER WARNING lights located
2-65
TM 55-1510-221-10
Figure 2-26. Overhead Circuit Breaker Panel
on the warning annunciator panel, placarded INST
AC, will illuminate if all instrument AC busses
should fail.
(3.) Inverter Control Switches.
Two
switches, placarded INVERTER No.1 and No.2 on
the overhead control panel (fig. 2-12) give the pilot
control of the single-phase AC inverters.
(4.) Volt-Frequency Meters. Two voltfrequency meters (fig. 2-12) are mounted in the
overhead control panel to provide monitoring capability for both 115 VAC buses. Normal display on
the meter is shown in frequency (Hz). To read voltage, press the button located in the lower left corner
of the meter. Normal output of the inverters will be
indicated by 115 VAC and 400 Hz on the meters.
b. Three Phase AC Power Supply. Three phase
AC electrical power for operation of the inertial navigation system and mission avionics is supplied by
two DC powered 3000 volt-ampere solid state three
phase inverters.
(1.) Three phase inverter control switches.
Two three position switches placarded #1 INV-OFFON-RESET and NO. 2 INV-OFF-ON-RESET,
located on the mission control panel (fig. 4-1) controls three phase inverter operation.
2-66
(2.) Three phase volt/frequency meters.
Two three phase volt/frequency meters, mounted on
the mission control panel (fig. 4-l), monitor and display the voltage and frequency outputs of the three
phase inverters.
(3.) Three phase loadmeters. Two three
phase loadmeters, mounted on the mission control
panel (fig. 4-l), monitors inverter output level.
(4.) Three phase AC off annunciator light.
An indicator light placarded 3!zsl! AC OFF, located
on the misson annunciator panel (fig. 4-l), indicates
a problem with one of the three phase AC power
busses.
(5.) Three phase AC external power. External three phase AC power for operation of the inertial navigation system or mission equipment, can be
applied to the aircraft through an external power
receptacle located on the underside of the left wing
leading edge just outboard of the engine nacelle (fig.
2-l). The receptacle is installed inside the wing
structure and is accessible through a hinged access
panel. The AC electrical system is automatically isolated from the external power source if the external
power is over or under voltage, over or under frequency, or has an improper phase sequence.
TM 55-1510-221-10
(a.) External AC power annunciator
Light. An annunciator light placarded EXT AC
PWR ON, located on the mission annunciator panel
(fig. 4-1) indicates that external AC power is connected to the 3 phase buses. The EXTERNAL
POWER annunciator in the advisory annunciator
panel indicates that an AC GPU plug is mated to
the AC external power receptacle.
(b.) External AC power control
switch. A switch placarded EXT POWER-OFF-ONRESET, located on the mission control panel (fig.
4-l), controls application of three phase AC power
to the aircraft.
Section X. LIGHTING
2-76. EXTERIOR LIGHTING.
a. Description. Exterior lighting (fig. 2-27)
consists of a navigation light on the aft end of the
aft section of the vertical stabilizer, one navigation
light on the outside of each wing tip pod, two strobe
beacons, one on top of the vertical stabilizer and one
on the underside of the fuselage center section, dual
landing lights and a taxi light mounted on the nose
gear assembly, a recognition light located in each
wing tip, and two ice lights, one light flush mounted
in each nacelle, positioned to illuminate along the
leading edge of each outboard wing.
b. Navigation Lights. The navigation lights
are protected by a 5-ampere circuit breaker placarded NAV on the overhead circuit breaker panel
(fig. 2-26). Control of the lights is provided by a
switch placarded NAV-ON on the overhead control
panel (fig. 2-l 2).
c. Strobe Beacons. The strobe beacons are
dual intensity units. They are protected by a 15ampere circuit breaker placarded BCN on the overhead circuit breaker panel (fig. 2-26). Control of the
lights is provided by a switch placarded BEACON DAY - NIGHT (fig. 2-12). Placing the switch in the
DAY position will activate the high intensity white
section of the strobe lights for greater visibility during daytime operation. Placing the switch in the
NIGHT position activates the lower intensity red
section of the strobe lights.
d. Landing/Taxi Lights. Dual landing lights
and a single taxi light are mounted on the nose gear
assembly. The lights are controlled by switches, placarded LANDING and TAXI, located in the
LIGHTS section of the pilot’s subpanel. The landing
light circuit is protected by a 5-ampere circuit
breaker placarded LANDING, located on the overhead circuit breaker panel (fig. 2-26). The taxi light
circuit is protected by a 5-ampere circuit breaker
placarded TAXI, located on the overhead circuit
breaker panel (fig. 2-26). Landing/Taxi lights are
turned off when the landing gear is retracted. The
landing lights and taxi light power circuits are protected by 35-ampere and 15-ampere circuit breakers,
respectively, on the DC power distribution panel
located below the aisleway floor forward of the main
spar.
e. Ice Lights. The ice lights circuit is protected by a 5-ampere circuit breaker placarded ICE
on the overhead circuit breaker panel (fig. 2-26).
Control of the lights is provided by a switch placarded ICE - ON on the overhead control panel (fig.
2-12). Prolonged use during ground operation may
generate enough heat to damage the lens.
f. Recognition Lights. A switch placarded
RECOG - ON, located in the pilot’s subpanel
LIGHTS section (fig. 2-6), controls the white recognition light in each wing tip. When requested, this
steady, bright light is used for identification in the
traffic pattern. The recognition lights circuit is protected by a 7 l/2 ampere RECOG circuit breaker
located on the overhead circuit breaker panel (fig.
2-26).
2-77. INTERIOR LIGHTING.
Lighting systems are installed for use by the
pilot and copilot. The lighting systems in the cockpit
are provided with intensity controls on the overhead
control panel. A switch placarded MASTER PANEL
LIGHTS - ON, on the overhead control panel (fig.
2-12), provides overall on-off control for all engine
instrument lights, pilot and copilot instrument
lights, overhead panel lights, console and subpanel
lights and the outside air temperature light.
a.
Cockpit Lighting.
(1.) Flight instrument lights. Each individual flight instrument contains internal lamps for illumination. The circuit is protected by a 7 l/2-ampere
circuit breaker placarded FLT INST on the overhead circuit breaker panel (fig. 2-26). Control is provided by two rheostat switches placarded PILOT
INST LIGHTS - OFF - BRT and COPILOT INST
LIGHTS - OFF - BRT on the overhead control
panel (fig. 2- 12). Turning the control clockwise from
OFF turns the lights on and increases their brilliance.
2-67
TM 55-1510-221-10
1. Wing navigation light
2. Emergency light
3. Strobe beacon
4. Tail navigation light
5. Recognition lights
6. Ice light
7. Taxi light
6. Landing lights
Figure 2-27. Exterior Lighting
2-68
AP 011764
TM 55-1510-221-10
(2.) Instrument indirect fights. Three lights
are mounted in the glareshield overhang along the
top edge of the instrument panel and provide overall
instrument panel illumination. The circuit is protected
by a 5-ampere circuit breaker placarded INST
INDIRECT on the overhead circuit breaker panel (fig.
2-26). Control is provided by a rheostat switch
placarded INST INDIRECT LIGHTS - OFF - BRT on
the overhead control panel (fig. 2-12). Turning the
control clockwise from OFF turns the lights on and
increases their brilliance.
(3.) Engine
lights.
instrument
Each
individual engine instrument contains internal lamps for
illumination. The circuit is protected by a 7 l/2- ampere
circuit breaker placarded FLT INST on the overhead
circuit breaker panel (fig. 2-26). Control is provided by
a rheostat switch placarded ENGINE INST LIGHTS
OFF BRT on the overhead control panel (fig. 2-12).
Turning the control clockwise from OFF turns the lights
on and increases their brilliance.
NOTE
The floodlight is connected to the hot battery
bus and will not be turned off by the battery
switch; therefore, it must be turned OFF when
the aircraft is shutdown to prevent discharging
the battery.
(4.) Flood light. A single overhead flood
light is installed. It provides overall illumination of the
entire cockpit area. The circuit is protected by a
5-ampere circuit breaker mounted beneath the battery
and connected to the emergency battery bus. Control is
provided by a rheostat switch placarded - OVERHEAD
FLOODLIGHT-OFF-BRT on the overhead control
panel (fig. 2-12). Turning the control clockwise from
OFF turns the light on and increases its brilliance.
(5.) Overhead panel lights. Lamps on the
overhead circuit breaker panel, control panel, and fuel
management panel are protected by a 7 1/2-ampere
circuit breaker placarded OVHD on the overhead circuit
breaker panel (fig. 2-26). Control is provided by a
rheostat switch placarded OVERHEAD PANEL
LIGHTS - OFF - BRT on the overhead control panel
(fig. 2-12). Turning the control clockwise from OFF
turns the lights on and increases their brilliance.
(6.) Subpanel and console lights. Lights on
the pilot’s and copilot’s subpanels, console edge lighted
panels, mission control panel, and pedestal extension
panels are protected by a 7 l/2-ampere circuit breaker
placarded SUBPNL & CONSOLE on the overhead
circuit breaker panel (fig. 2-26). Control is provided by
two rheostat switches placarded SUBPANEL or
CONSOLE LIGHTS - OFF - BRT on the overhead
control panel (fig. 2- 12). Turning the control clockwise
from OFF turns the lights on and increases their
brilliance.
(7.) Outside air temperature light. Two post
lights are mounted adjacent to the outside air
temperature gage on the left cockpit sidewall trim panel.
The circuit is protected by a 71/2-ampere circuit breaker
placarded FLT INST on the overhead circuit breaker
panel (fig. 2-26). Control is provided by a pushbutton
switch adjacent to the gage. No intensity control is
provided.
b. Cabin Lighting.
(1.) Threshold and spar cover lights. A
threshold light is installed just above floor level on the
left side of the cabin just inside the cabin door. A spar
cover light is installed on the left side of the sunken
aisle immediately aft of the main spar cover. Both
circuits are protected by a 5-ampere circuit breaker
mounted beneath the battery and connected to the
emergency battery bus. Both lights are controlled by the
switch mounted adjacent to the threshold light. If the
lights are illuminated, closing the cabin door will
automatically extinguish them.
(2.) Cabin aisle lights. Three cabin aisle
lights are installed in the cabin aisle. Control is provided
by the CABIN LIGHTS switch on the right subpanel.
Control is provided by the CABIN LIGHT switch on
the right subpanel.
(3.) Cabin spot lights. A spot light is mounted
to each cabin aisle light. Each spot light is individually
controlled by a rheostat placarded OFF-ON-BRT on the
back of the light. There is a momentary ON switch in
the center of the rheostat. Each light is capable of
producing a red or white spotlight by turning the
selector on the front of the light. To remove the light
from the stationary position, pull down on the light. The
light is connected to the light housing by an 11 inch
coiled cord that extends to approximately 50 inches.
(4.) Cabin door latching mechanism light. A
light is provided to check the cabin door latching
mechanism. It is controlled by a red pushbutton switch
located adjacent to the round observation window,
which is just above the second step.
2-78. EMERGENCY LIGHTING.
a. Description. An independent battery operated
lighting system is installed. The system is actuated
automatically by shock, such as a forced landing. It
provides adequate lighting inside and outside the
fuselage to permit the crew to read instruction
2-69
TM 55-151 O-221 -10
placards and locate exits. An inertia switch, when
subjected to a 2 G shock, will illuminate interior lights in
the cockpit, forward and aft cabin areas, and exterior
lights aft of the emergency exit and aft of the cabin door.
The battery power source is automatically recharged by
the aircraft electrical system.
ORIDE OFF - RESET - AUTO - TEST. Should the
system accidentally actuate, the emergency lights will
illuminate. Placing the switch in the momentary OFF
RESET position will extinguish the lights. To test the
system, place the switch in the TEST position. The
lights should illuminate. Moving the switch to the OFF
- RESET position will turn the system off and reset it.
b. Operation. An emergency lights override switch,
located on the overhead control panel (fig. 2-12), is
provided to turn the system off if it is accidentally
actuated. The switch is placarded EMERG LIGHTS
Section Xl. FLIGHT INSTRUMENTS
2-79.
PITOT AND STATlC SYSTEM.
a. Description, The pitot and static system (fig.
2-28) supplies static pressure to two airspeed indicators,
the copilot altimeter, the air data computer (ADC), two
vertical velocity indicators, and also ram air to the
airspeed indicators and the ADC. This system consists
of two pitot masts (one located on each side of the lower
portion of the nose), static air pressure ports in the
aircraft’s exterior skin on each side of the aft fuselage,
and associated system plumbing. The pitot mast is
protected from ice formation by internal electric heating
elements.
b. Alternate Static Air Source. An alternate static
air line, which terminates just aft of the rear pressure
bulkhead, provides a source of static air for the pilot’s
instruments in the event of source failure from the pilot’s
static air line. A control on the pilot’s subpanel placarded
PILOTS STATIC AIR SOURCE, may be actuated to
select either the NORMAL or ALTERNATE air source by
a two position selector valve. The valve is secured in the
NORMAL position by a spring clip. Refer to Chapter 7
for airspeed indicator and altimeter calibration
information when using the alternate air source.
operation, The indicator dials are calibrated in knots
from 40 to 300. A striped pointer automatically displays
the maximum allowable airspeed at the aircraft’s
present altitude.
2-82. COPILOT’S ENCODING ALTIMETER.
The copilot’s altimeter on the upper right side of the
instrument panel (fig. 2-29) is a self-contained unit
consisting of a precision pneumatic altimeter combined
with an altitude encoder. The display face indicates
while, simultaneously, the encoder transmits pressure
altitude information to the INS and GPS. Altitude is
displayed by a 10,000 foot counter, a 1000 foot counter, a
100 foot counter, and a single needle pointer which
indicates hundreds of feet on a circular scale in 20 foot
intervals. The needle pointer is also coupled to the 100
foot drum counter so that both move at the same time.
Below an altitude of 10,000 feet, a diagonal striped
symbol will appear on the 10,000 foot counter. A
barometric pressure setting knob is provided to insert
the desired altimeter setting in inches Hg or millibars. If
AC power to the altitude encoder is lost, an OFF flag will
appear in the upper center portion of the instrument face
to indicate that the encoder is inoperative and the
system is not reporting altitude to ground stations.
2-80. TURN-AND-SLIP INDICATORS.
2-83. PILOT’S ALTlMETER INDICATOR.
Turn-and-slip indicators are installed separately on
the pilot and copilot sides of the instrument panel (fig.
2-29). These indicators are gyroscopically operated.
They use DC power and are protected by 5-ampere
circuit breakers placarded TURN & SLIP PILOT or
COPILOT on the overhead circuit breaker panel (fig.
2-26).
2-81. AIRSPEED INDICATORS.
Airspeed indicators are installed separately on the
pilot and copilot sides of the instrument panel (fig. 2-29).
These indicators require no electrical power for
2-70
Change 4
The pilot’s altimeter, on the upper left side of the
instrument panel (fig. 2-29), is a servoed unit under
control of the Air Data Computer and is part of the Flight
Director/Autopilot System. It lacks encoding capability,
but displays altitude as described for the copilot’s
instrument. Operating instructions are provided in
chapter 3. When the BAR0 knob is adjusted to ground
supplied instructions, the updated altitude pressure is
routed to the Air Data Computer. The ADC recomputes
all data on hand, sends corrected altitude pressure
information to the Flight Director and autopilot, sends
servo commands to correct the display on the pilot’s
TM 55-1510-221-10
Figure 2-28. Pitot and Static System
2-71
TM 55-1510-221-10
1. Airspeed indicator
2. Attitude director indicator
3. Flight director annunciator
panel
4. Gyro fast erect switch
5. Master caution/warning annunciator
6. Marker beacon dimmer control
7. Marker beacon indicator lights
8. Pilot’s altimeter
9. Warning annunciator panel
10. Push-to-exinguish/squib OK
annunciators
11. Fire pull handles
12. Radar warning control panel (AN/APR-39)
13. Torque indicators
14. Prop Tachometers
15. Turbine gas temperature indicators
16. Radar signal detecting set indicator
(AN/APR-39)
17. Accelerometer
18. RMI
19. Copilot’s gyro horizon indicator
20. Altitude select controller
21. Copilot’s altimeter
22. Vertical speed indicator
AP 011765.1
Figure 2-29. lnstrument Panel (Sheet 1 of 2)
2-72
TM 55-1510-221-10
23. PILOT SELECT annunciator
24. Course indicator selector switch
25. RMI select switch
26. Compass #1 and #2 switch
27. Microphone select switch
28. Gyro INCREASE-DECREASE switch
29. Gyro SLAVE-FREE switch
30. Turn & slip indicator
31. Compass sync annunciator
32. Copilot’s horizontal situation
indicator
33. Cabin altitude indicator
34. Cabin rate-of-climb indicator
35. HSI readout dim control
36. INS control display Indicator
37. TACAN range indicator
38. Fuel flow gages
39. Oil pressure and temp gages
40. Turbine tachometers
41. Radar warning control panel (AN/APR-44)
42. Weather radar indicator
43. Propellers synchroscope
44. Propeller synchronizer switch
45. Radio altimeter indicator
AP 011765.2
46. Pilot’s horizontal situation
indicator
Figure 2-29. lnstrument Panel (Sheet 2 of 2)
2-73
TM 55-1510-221-10
altimeter, and supplies altitude information to the
transponder.
2-87. STANDBY MAGNETIC COMPASS.
WARNING
2-84. VERTICAL VELOCITY INDICATORS.
Vertical velocity indicators are installed separately on the pilot and copilot sides of the instrument panel (fig. 2-29). They indicate the speed at
which the aircraft ascends or descends based on
changes in atmospheric pressure. The indicator is a
direct reading pressure instrument requiring no electrical power for operation.
2-85. ACCELEROMETER.
The accelerometer, located on the instrument
panel registers and records positive and negative G
loads imposed on the aircraft. One hand moves in
the direction of the G load being applied while the
other two, one for positive G loads and one for negative G loads, follow the indicating pointer to its
maximum travel. The recording pointers remain at
the respective maximum travel positions of the G’s
being applied, providing a record of maximum G
loads encountered. Depressing the push-to-reset
knob at the lower left comer of the instrument
allows the recording pointers to return to the normal
position.
2-86. OUTSIDE AIR TEMPERATURE (OAT)
GAGE.
The outside air temperature gage, mounted outboard of the pilot’s seat, (fig. 2-8), indicates the outside air temperature in degrees Celsius.
Inaccurate indications on the standby
magnetic compass will occur while windshield heat and/or air conditioning is
being used.
The standby magnetic compass is located below
the overhead fuel management panel and to the
right of the windshield divider. It may be used in
the event of failure of the compass system, or for
instrument cross check. Readings should be taken
only during level flight since errors may be introduced by turning or acceleration. A compass correction chart indicating deviation is located on the
magnetic compass.
2-88. MISCELLANEOUS INSTRUMENTS.
a. Annunciator Panels. Three annunciator
panels are installed. One is a warning panel with red
fault identification lights, and the others are caution/
advisory panels with yellow and green identification
lights. The warning panel is mounted near the center
of the instrument panel below the glareshield (fig.
2-29) and one caution/advisory panel is located on
the center subpanel (fig. 2-6). The mission annunciator panel is located on the copilot’s sidewall. Some
normal flight operations involve indications from
the mission control panel (fig. 4-l). Illumination of
a red warning light signifies the existence of a hazardous condition requiring immediate corrective
action.
Table 2-6. Warning Annunciator Panel Legend
2-74
NOMENCLATURE
COLOR
NO.1 FUEL PRESS
NO.2 FUEL PRESS
L BL AIR FAIL
RED
RED
RED
R BL AIR FAIL
RED
ALT WARN
INST AC
AP TRIM FAIL
NO.1 CHIP DETR
NO.2 CHIP DETR
AP DISC
RED
RED
RED
RED
RED
RED
WARNING ANNUNCIATOR
CAUSE FOR ILLUMINATION
Fuel pressure failure on left side
Fuel pressure failure on right side
Left bleed air warning line has melted or failed, indicating
possible loss of No.1 engine bleed air
Right bleed air warning line has melted or failed, indicating
possible loss of No. 2 engine bleed air
Cabin altitude exceeds 12,500 feet
No AC power to engine instruments
Trim inoperative or running opposite direction commanded
Contamination of No.1 engine oil detected
Contamination of No.2 engine oil detected
Autopilot has disengaged.
TM 55-1510-221-10
Table 2- 7. Caution/Advisory Annunciator Panel Legend
(Sheet 1 of 2)
NOMENCLATURE
CAUTION/ADVISORY ANNUNCIATOR
CAUSE FOR ILLUMINATION
COLOR
No.1 DC GEN
No. 1 INVERTER
REV NOT READY
Yellow
Yellow
Yellow
No.2
No.2
No.1
No.1
Yellow
Yellow
Yellow
Yellow
INVERTER
DC GEN
EXTGH DISCH
NAC LOW
CABIN DOOR
No.2 NAC LOW
Yellow
Yellow
No.2 EXTGH DISCH
No. 1 VANE FAIL
Yellow
Yellow
BATTERY CHARGE
PROP SYNC ON
No.2 VANE FAIL
Yellow
Yellow
Yellow
DUCT OVERTEMP
IFF
No.1 NO FUEL XFR
Yellow
Yellow
Yellow
No.2 NO FUEL XFR
Yellow
No.1 LIP HEAT
Yellow
No.2 LIP HEAT
Yellow
INS
Yellow
No.1 LIP HEAT ON
No.2 LIP HEAT ON
A/C COLD OPN
Green
Green
Green
No.1 VANE EXT
FUEL CROSSFEED
AIR COND N, LOW
No.2 VANE EXT
No.1 IGN ON
Green
Green
Green
Green
Green
No.1 engine generator off the line
No. 1 inverter inoperative
Propeller levers are not in the high RPM, low pitch position,
with the landing gear extended
No.2 inverter inoperative
No.2 engine generator off line
No. 1 engine fire extinguisher discharged
No.1 engine has 20 minutes fuel remaining at sea level, normal cruise power consumption rate
Cabin/door open or not secure
No.2 engine has 20 minutes fuel remaining at sea level, normal cruise power consumption rate
No.2 engine fire extinguisher discharged
No.1 engine ice vane malfunction. Ice vane has not attained
proper position
Excessive charge rate on battery
Synchrophaser turned on with landing gear extended
No.2 engine ice vane malfunction. Ice vane has not attained
proper position
Excessive bleed air temperature in environmental heat ducts
Transponder fails to reply to a valid mode 4 interogation
Auxiliary fuel tank on side of No. 1 engine not transferring fuel
into nacelle tank
Auxiliary fuel tank on side of No.2 engine not transferring fuel
into nacelle tank
Failure of lip heat valve to conform to selected position or in
transit
Failure of lip heat valve to conrom to selected position or in
transit
Inertial navigation system’s cooling fan is off or an INS malfunction that illuminates the WARN annunciator on the CDU
No.1 engine air scoop heat switch is on
No.2 engine air scoop heat switch is on
Air conditioner is operating in cold mode, or ambient temperatures require switching to cold mode if air conditioner operation is to be continued
No. 1 ice vane extended
Crossfeed valve open
No.2 engine RPM too low for air conditioning load
No.2 ice vane extended
No.1 engine ignition/start switch on No.1 engine autoignition
switch armed and engine torque below 20 percent
2-75
TM 55-1510-221-10
Table 2-7. Caution/Advisory Annunciator Panel Legend
(Sheet 2 of 2)
NOMENCLATURE
CAUTION/ADVISORY ANNUNCIATOR
COLOR
CAUSE FOR ILLUMINATION
L BL AIR IFF
EXTERNAL POWER
R BL AIR OFF
No.2 IGN ON
Green
Green
Green
Green
No. 1 AUTOFEATHER
Green
No.2 AUTOFEATHER Green
BRAKE DEICE ON
Green
Left environmental bleed air valve closed
External power connector plugged in
Right environmental bleed air valve closed
No.2 engine ignition/start switch on, No.2 engine autoignition
switch armed and engine torque below 20 percent
No.1 engine autofeather armed with power levers advanced
above 90% N1
No.2 engine autofeather armed with power levers advanced
above 90% N1
Brake deice system activated
a. Annunciator Panels. Three annunciator
panels are installed. One is a warning panel with red
fault identification lights, and the others are caution/
advisory panels with yellow and green identification
lights. The warning panel is mounted near the center
of the instrument panel below the glareshield (fig.
2-29) and one caution/advisory panel is located on
the center subpanel (fig. 2-6). The mission annunciator panel is located on the copilot’s sidewall. Some
normal flight operations involve indications from
the mission control panel (fig. 4-l). Illumination of
a red warning light signifies the existence of a hazardous condition requiring immediate corrective
action. A yellow caution light signifies a condition
other than hazardous requiring pilot attention. A
green advisory light indicates a functional situation.
Table 2-6, 2-7, and 2-8 provides a list of causes for
illumination of the individual annunciator lights. In
frontal view both panels present rows of small,
opaque rectangular indicator lights. Word printing
on each indicator identifies the monitored function,
situation, or fault condition, but cannot be read
until the light is illuminated. The bulbs of all annunciator panel lights are tested by activating the
ANNUNCIATOR TEST switch, located on the right
subpanel near the caution/advisory panel. The system is protected by two 5-ampere circuit breakers
placarded ANN PWR and ANN IND on the overhead circuit breaker panel (fig. 2-26). The annunciator system lights are dimmed when the MASTER
PANEL LIGHTS switch is ON and the pilot’s flight
instrument lights are illuminated. The lights are
automatically reset to maximum brightness if:
(1.) The main aircraft power (both DC
generators) are OFF.
2-76
(2.) T h e I N S T I N D I R E C T L I G H T S
switch is rotated clockwise.
(3.) T h e M A S T E R P A N E L L I G H T S
switch is off.
(4.) T h e M A S T E R P A N E L L I G H T S
switch is ON and the PILOT INST LIGHTS switch
is OFF.
(5.) Master warning light (red). A MASTER WARNING light is provided for both the pilot
and the copilot and is located on each side of the
glareshield (fig. 2-29). Any time a warning light illuminates, the MASTER WARNING light will illuminate, and will stay illuminated until the MASTER
WARNING light is pressed to reset the circuit. If a
new condition occurs, the light will be reactivated,
and the applicable annunciator panel light will illuminate.
(6.) Master caution light (yellow). A MASTER CAUTION light is provided for both the pilot
and copilot located adjacent to the MASTER
WARNING LIGHT. Whenever a caution light illuminates, the MASTER CAUTION will illuminate,
and will stay illuminated until the condition is corrected and/or the MASTER CAUTION light is
pressed to reset the circuit. If a new condition
occurs, the light will be reactivated and the appropriate annunciator panel lights will illuminate.
b. Clocks. One manually-wound 8-day clock is
mounted in the center of the pilot’s control wheel
and an electric digital clock is mounted in the center
of the copilot’s control wheel.
TM 55-1510-221-10
Table 2-8. Mission Control Panel Annunciator Legend
NOMENCLATURE
MSN OVERTEMP
CRYPT0 ALERT
PWR SPLY FAULT
CALL
3 Ø AC OFF
BAT FEED FAULT
MISSION POWER
LINK MODE
RADOME HOT
LINK SYNC
SPCL EQPT OVRD
DIPLEXER PRESS
TWTA STANDBY
ANT MALF
NO INS UPDATE
TDOA OVERTEMP
LB PS OVERTEMP
TDOA FAULT
LB PS FAULT
ELINT FAULT
ANT STOWED
ANT OPERATE
RADOME HEAT
MISION AC ON
INS UPDATE
TDOA PWR ON
MISSION DC ON
WAVE GUIDE
EXT AC PWR ON
EXT DC PWR ON
MISSION ANNUNCIATOR
CAUSE FOR ILLUMINATION
COLOR
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Yellow
Green
Green
Green
Green
Green
Green
Green
Green
Green
Green
Section XII.
Mission equipment is overheating.
Coded messages being received.
Mission power out of tolerance.
Reciving transmission on VOW.
Three phase AC power fault.
Ground fault detected in battery or external power line.
Mission power is off.
WBDL fault in link or contact.
Radome heat is too high.
WBDL has synchronization fault.
Mission power switch is in override.
Diplexer has lost pressurization.
WBDL is in standby mode.
Boom antenna is out of position.
INS update is not in process.
TDOA equipment is overheating.
LB PS equipment is overheating.
TDOA system has fault.
LB PS has fault.
ELINT system has fault.
Boom antenna is in horizontal position.
Boom antenna is in vertical position.
Radome heat is on.
Mission AC power is on.
INS update in process.
TDOA power is on.
Mission DC power is on.
Wave guide is pressurized.
External AC power is on.
External DC power is on.
SERVICING, PARKING, AND MOORING
2-89. GENERAL.
The following paragraphs include the procedures
necessary to service the aircraft except lubrication.
The lubrication requirements of the aircraft are covered in the aircraft maintenance manual. Table 2-9,
2-10, 2-11 and 2-12 are used for identification of
fuel, oil, etc. used to service the aircraft. The servicing instructions provide procedures and precautions
necessary to service the aircraft.
2-77
TM 55-1510-221-10
DETAIL C
DETAIL A
DETAIL B
1. Air conditioning compressor
2. External power receptacle
3. Hand fire extinguisher
4. Battery 24 VDC
5. Oxygen system filler
6. Oxygen cyilnders 2 (64 cu ft bottles)
7. Electric toilet
8. Fuel filler cap (typical left and right)
9. Landing gear tires
10. Engine fire extinguisher
11. Engine oil filler cap (typical left and right)
12. Wheel brake fluid reservoir
AP011766
Figure 2-30. Servicing Locations
2-78
TM 55-1510-221-10
Table 2-9. Approved Military Fuels, Oil, Fluids, and Unit Capacities.
SPECIFICATION
SYSTEM
Fuel
Engine Oil
Hydraulic Brake System
Oxygen System
Toilet Chemical
MIL-T-5624 (JP4 and JP-5)
MIL-L-23699
MIL-H-5606
MIL-O-27210
Monogram DG-19
2-90. FUEL HANDLING PRECAUTIONS.
Table 2-2, Fuel Quantity Data, lists the quantity and
capacity of fuel tanks in the aircraft. Service the fuel
tanks after each flight to keep moisture out of the tanks
and to keep the bladder type cells from drying out.
Observe the following precautions:
WARNING
During warm weather open fuel caps slowly
to prevent being sprayed with fuel.
WARNING
When aviation gasoline is used in a turbine
engine, extreme caution should be used
when around the combustion chamber and
exhaust area to avoid cuts or abrasions. The
exhaust deposits contain lead oxide which
will cause lead poisoning.
CAPACITY
546 U.S. Gals.
14 U.S. Quarts per engine
1 U.S. Pint
128 Cubic Feet
3 Ounces
such as drills or buffers, in or near the aircraft during
fueling.
b. Keep fuel servicing nozzles free of snow, water,
and mud at all times.
c. Carefully remove snow, water, and ice from the
aircraft fuel filler cap area before removing the fuel filler
cap (fig. 2-30). Remove only one aircraft filler cap at any
one time, and replace each one immediately after the
servicing operation is completed.
d. Wipe all frost from fuel filler necks before
servicing.
e. Drain water from fuel tanks, filter cases, and
pumps prior to first flight of the day. Preheat, when
required, to insure free fuel drainage.
f. Avoid dragging the fueling hose where it can
damage the soft, flexible surface of the deicer boots.
CAUTION
g. Observe NO SMOKING precautions.
Proper procedures for handling JP-4 and
JP-5 fuel cannot be over stressed. Clean,
fresh fuel shah be used and the entrance of
water into the fuel storage or aircraft fuel
system must be kept to a minimum.
h. Prior to transferring the fuel, insure that the
hose is grounded to the aircraft.
i. Wash off spilled fuel immediately
CAUTION
When conditions permit, the aircraft shall be
positioned so that the wind will carry the
fuel vapors away from all possible sources of
ignition. The fuel vehicle shall be positioned
to maintain a minimum distance of 10 feet
from any part of the aircraft, while
maintaining a minimum distance of 20 feet
between the fueling vehicle and the fuel filler
point.
j. Handle the fuel hose and nozzle cautiously to
avoid damaging the wing skin.
k. Do not conduct fueling operations within 100 feet
of energized airborne radar equipment or within 300 feet
of energized ground radar equipment installations.
l. Wear only nonsparking shoes near aircraft or
fueling equipment, as shoes with nailed soles or metal
heel plates can be a source of sparks.
a. Shut off unnecessary electrical equipment on the
aircraft, including radar and radar equipment. The
master switch may be left on, to monitor fuel quantity
gages, but shall not be moved during the fueling
operation. Do not allow operation of any electrical tools,
Change 4
2-79
TM 55-1510-221-10
Table 2- 10. Approved Fuels
SOURCE
PRIMARY OR
STANDARD FUEL
US MILITARY FUEL
NAT0 Code No.
COMMERCIAL FUEL
ASTM-D-l 655)
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
Phillips Petroleum
Shell Oil
Sinclair
Standard Oil Co.
JP4 (MIL-T-5624)
F-40 (Wide Cut Type)
JET B
American JP-4
American JP-4
Arcojet B
Chevron
Texaco
Union Oil
Foreign Fuel
Belgium
Canada
Denmark
France
Germany (West)
Greece
Italy
Netherlands
Norway
Portugal
Turkey
United Kingdom (Britain)
Chevron B
Texaco Avjet B
Union JP-4
NATO F-40
BA-PF-2B
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
J P-4 M IL-T-5624
JP-4 MIL-T-5624
JP-4 MIL-T-5624
D.Eng RD 2454
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
ALTERNATE
FUEL
JP-5 (MIL-T-5624)
F-44 (High Flash Type)
JET A
JET A-l
American Type A
NATO F-34
American Type A
Arcojet A-1
Arcojet A
Richfield A
Richfield A-1
B.P.A.T.K.
Caltex Jet A-1
CITGO A
Conoco Jet-50
Conoco Jet-60
Gulf Jet A
Gulf Jet A-1
EXXON A
EXXON A-1
Mobil Jet A
Mobil Jet A-1
Philjet A-50
Aeroshell 640
Aeroshell 650
Superjet A
Superjet A-1
Jet A Kerosene
Jet A-1
Kerosene
Chevron A-50
Chevron A- 1
Avjet B
Avjet A-1
76 Turbine Fuel
NATO F-44
3-6P-24e
UTL-9130-007/UTL9130-010
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-L-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-L-27686 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.
2-80
TM 55-1510-221-10
Table 2-11. Standard Alternate and Emergency Fuels
ENGINE
ARMY STANDARD FUEL
ALTERNATE TYPE
PT6A
MIL-T-5624
Grade JP-4
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).
2-91.
FILLING FUEL TANKS.
a. Army Standard Fuels. Army standard fuel
is JP-4.
WARNING
b. Alternate Fuels, Army Alternate fuels are
JP-5 and JP-8.
Prior to removing the fuel tank filler cap,
the hose nozzle static ground wire shall be
attached to the grounding lugs that are
located adjacent to the filler opening.
C. Emergency Fuel. Avgas is emergency fuel
and subject to 150 hour time limit.
Fill tanks as follows:
a.
Attach bonding cables to aircraft.
b. Attach bonding cable from hose nozzle to
ground socket adjacent to fuel tank being filled.
Do not insert fuel nozzle completely into
fuel cell due to possible damage to bottom
of fuel cell. Nozzle should be supported
and inserted straight down to prevent
damage to the anti-siphon valve.
c. Fill main tank before filling respective auxiliary tanks unless less than a full fuel load is
desired.
d. Secure applicable fuel tank filler cap. Make
sure latch tab on cap is pointed aft.
e.
Disconnect bonding cables from aircraft.
2-92. DRAINING MOISTURE FROM FUEL SYSTEM.
To remove moisture and sediment from the fuel
system, 12 fuel drains are installed (plus one for the
ferry system, when installed).
2-93. FUEL TYPES.
Approved fuel types are as follows:
2-94. USE OF FUELS.
Fuel is used as follows:
a. 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. 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.
b. Use of Kerosene Fuels. The use of kerosene
fuels (JP-5 type) in turbine engines dictates the need
for observance of special precautions. Both ground
starts and air restarts at low temperature may be
more difficult due to low vapor pressure. Kerosene
fuels having a freezing point of minus 40 degrees C
(minus 40 degrees F) limit the maximum altitude of
a mission to 28,000 feet under standard day conditions.
c. Mixing of Fuels in Aircraft Tanks. When
changing from one type of authorized fuel to
another, for example JP-4 to JP-5, it is not necessary
to drain the aircraft fuel system before adding the
new fuel.
d. Fuel Specifications. Fuel 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 refueling with
NATO F-40 (JP-4) or Commercial ASTM Type B
fuels. Whenever this condition occurs, the engine
2-81
TM 55-1510-221-10
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-l fuels. Most commercial turbine
engines will operate satisfactorily on either kerosene
or JP-4 type fuel. The difference in specific gravity
may possibly require fuel control adjustments; if so,
the recommendations of the manufacturers of the
engine and airframe are to be followed.
2-95. SERVICING OIL SYSTEM
An integral oil tank occupies the cavity formed
between the accessory gearbox housing and the compressor inlet case on the engine. The tank has a calibrated oil dipstick and an oil drain plug. Avoid spilling oil. Any oil spilled must be removed
immediately. Use a cloth moistened in solvent to
remove oil. Overfilling may cause a discharge of oil
through the accessory gearbox breather until a satisfactory level is reached. Service oil system as follows:
1.
Open the access door on the upper cowling
to gain access to the oil filler cap and dipstick.
6.
If oil level is over 2 quarts low, motor or run
engine as required, and service as necessary.
7.
Install and secure oil tiller cap.
8.
Check for any oil leaks.
2-96. SERVICING HYDRAULIC BRAKE SYSTEM
RESERVOIR.
Gain access to brake hydraulic system reservoir.
1.
2. Remove brake reservoir cap and till reservoir to washer on dipstick with hydraulic
fluid.
3.
Install brake reservoir cap.
2-97. INFLATING TIRES.
Inflate tires as follows:
1.
Inflate nose wheel tires to a pressure
between 55 and 60 PSI.
2. Inflate main wheel tires to a pressure
between 73 and 77 PSI.
2-98. SERVICING THE CHEMICAL TOILET.
A cold oil check is unreliable. If possible,
check oil within 10 minutes after engine
shutdown. If over 10 minutes have
elapsed, motor the engine (starter only)
for 15-20 seconds, then recheck. If over
10 hours have elapsed, start the engine
and run for 2 minutes, then recheck. Add
oil as required. Do not overfill.
2.
Remove oil filler cap.
3.
Insert a clean funnel, with a screen incorporated, into the filler neck.
4.
Replenish with oil to within 1 quart below
MAX mark or the MAX COLD on dipstick
(cold engine). Fill to MAX or MAX HOT
(hot engine).
5.
Check oil filler cap for damaged preformed
packing, general condition and locking.
The toilet should be serviced during routine
ground maintenance of the aircraft following any
usage. The waste storage container should be
removed, emptied, its disposable plastic liner
replaced, and the container replaced in the toilet
cabinet. Toilet paper, waste container plastic liners,
and dry chemical deodorant packets should also be
resupplied within the toilet cabinet as needed.
2-99.
TEM.
SERVICING THE AIR CONDITIONING SYS-
Servicing the air conditioning system consists of
checking and maintaining the correct refrigerant
level, compressor oil level, belt tension and condition, system leak detection, and replacement of the
evaporator air filters. It is imperative that the maintenance of the air conditioning system, except for filter replacement, be accomplished only by qualified
refrigerant system technicians. Flexible fiberglass filters cover the evaporator coils and should be
replaced after 300 hours of operation. Install filters
as follows:
a.
Insure that oil filler cap is correctly
installed and securely locked to prevent
loss of oil and possible engine failure.
2-82
Forward Evaporator Filter Replacement:
1.
Remove the access door in the nose
wheel well keel under the refrigerant
plumbing.
TM 55-1510-221-10
2. Pull the filter down and out of the
retaining springs on the evaporator
coil.
3.
NOTE
Do not apply anti-icing, deicing and
defrosting fluid to exposed aircraft surfaces if snow is expected. Melting snow
will dilute the defrosting fluid and form a
slush mixture which will freeze in place
and become difficult to remove.
Fold the new filter to insert it through
the access doors. The filter; must be
carefully inserted between the coil
assembly and the refrigerant plumbing
under the retaining springs.
4. Install the access doors.
b. Aft Evaporator Filter Replacement.
1.
Remove the carpet and floor panel
behind the rear spar, and remove the
cover of the evaporator plenum.
2.
Remove the old filter from behind the
retaining springs on the evaporator
coil.
3. Insert the new filter between the
retainer springs and the evaporator
coil.
4. Install the plenum cover, floor panel,
and carpet.
c. 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 permit protective covers to be
removed with a minimum of sticking. To prevent
freezing rain and snow from blowing under protective covers and diluting the fluid, insure that protective covers are fitted tightly. As a deicing measure,
keep exposed aircraft surface wet with fluid for protection against frost.
2-100. 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 antiicing, deicing, and defrosting fluid mixed in
accordance with table 2-12.
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
Table 2- 12. Recommended Fluid Dilution Chart
AMBIENT
TEMPERATURE
(ºF)
PERCENT
DEFROSTING
FLUID BY VOLUME
PERCENT WATER
BY
VOLUME
80
20
30º and above
70
30
20º
60
40
10º
55
45
0º
50
50
-10º
45
55
-20º
40
60
-30º
NOTES:
1. Use anti-icing and deicing fluid (MIL-A-8243) or commercial fluids.
2. Heat Mixture to a temperature of 82º to 93°C (180º to 200°F).
FREEZING POINT
OF MIXTURE (“F)
(APPROXIMATE)
10º
0º
-15º
-25º
-35º
-45º
-55º
2-83
TM 55-1510-221-10
intervals to assure complete coverage.
Removal of ice accumulations using undiluted defrosting fluid is expensive and slow.
If tires are frozen to ground, use undiluted
defrosting fluid to melt ice around tire.
Move aircraft as soon as tires are free.
4.
2-101.
APPLICATION OF EXTERNAL POWER.
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. 3inimum GPU requirement is 400 amperes
continuous and 1800 amperes for one
tenth of a second.
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 underside of the
right wing leading edge just outboard of the engine
nacelle. An external AC Power receptacle is installed
on the underside of the left wing leading edge just
outboard of the engine nacelle.
2-102. SERVICING OXYGEN SYSTEM.
The oxygen system furnishes breathing oxygen
to the pilot, copilot and first aid position Oxygen
cylinder location is shown in figure 2-19.
a.
(1.) Keep oxygen regulators, cylinders,
gages, valves, fittings, masks, and all other components of the oxygen system free of oil, grease, gasoline, and all other readily combustible substances.
The utmost care shall be exercised in servicing, handling, and inspecting the oxygen system.
(2.) Do not allow foreign matter to enter
oxygen lines.
(3.) Never allow electrical equipment to
come in contact with the oxygen cylinder.
(4.) Never use oxygen from a cylinder
without first reducing its pressure through a regulator.
b. Replenishing Oxygen System.
1.
Remove oxygen access door on outside
of aircraft (fig. 2- 19).
2.
Remove protective cap on oxygen system filler valve.
3. Attach oxygen hose from oxygen servicing unit to filler valve.
If the oxygen system pressure is below
200 PSI, do not attempt to service system. Make an entry on DA Form
2408-13.
4.
Insure that supply cylinder shutoff
valves on the aircraft are open.
5.
Slowly adjust the valve position so that
pressure increases at a rate not to
exceed 200 PSIG per minute.
6.
Close pressure regulating valve on oxygen servicing unit when pressure gage
on oxygen system indicates the pressure obtained using the Oxygen System
Servicing Pressure Chart (fig. 2-31).
Oxygen System Safety Precautions.
Keep fire and heat away from oxygen
equipment. Do not smoke while working
2-84
with or near oxygen equipment, and take
care not to generate sparks with carelessly
handled tools when working on the oxygen system.
TM 55-1510-221-10
OUTSIDE AIR TEMPERATURE, °C
Figure 2-31. Oxygen System Servicing Pressure
2-85
TM 55-1510-221-10
NOTE
To compensate for loss of aircraft cylinder pressure as the oxygen cools to ambient temperature after recharging, the cylinder should be charged initially to
approximately 10% over prescribed pressure. Experience will determine what initial pressure should be used to compensate for the subsequent pressure loss upon
cooling. A small top-off will create little
heat. A complete recharge will create substantial heating.
The final stabilized cylinder pressure should be
adjusted for ambient temperature per figure 2-31.
7. Disconnect oxygen hose from oxygen
servicing unit and tiller valve.
8. Install protective cap on oxygen filler
valve.
9. Install oxygen access door.
2-103. 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. Parking, covers, ground handling and towing
equipment are shown in figure 2-32.
a. General Ground Handling Procedure. Accidents resulting in injury to personnel and damage to
equipment can be avoided or minimized by close
observance of existing safety standard and recognized ground handling procedures. Carelessness or
insufficient knowledge of the aircraft or equipment
being handled can be fatal. The applicable technical
manuals and pertinent directives should be studied
for familiarization with the aircraft, its components,
and the ground handling procedures applicable to it,
before attempting to accomplish ground handling.
b. Ground Handling Safety Practice. Aircraft
equipped with turboprop engines require additional
maintenance safety practices. The following list of
safety practices should be observed at all times to
prevent possible injury to personnel and/or damaged
or destroyed aircraft:
(1.) Keep intake air ducts free of loose
articles such as rags, tools, etc.
(2.)
Stay clear of exhaust outlet areas.
(3.) During ground runup, make sure the
brakes are firmly set.
2-86
(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.) When moving the aircraft, do not
push on propeller deicing boots. Damage to the
heating elements may result.
c. Moving Aircraft on Ground. Aircraft on the
ground shall be moved in accordance with the following:
(1.) Taxiing. Taxiing shall be in accordance with chapter 8.
When the aircraft is being towed, a qualified person must be in the pilot’s seat to
maintain control by use of the brakes.
When towing, do not exceed nose gear
turn limits. Avoid short radius turns, and
always keep the inside or pivot wheel
turning during the operation. Do not tow
aircraft with rudder locks installed, as
severe damage to the nose steering linkage
can result. When moving the aircraft
backwards, do not apply the brakes
abruptly. Tow the aircraft slowly, avoiding sudden stops, especially over snowy,
icy, rough, soggy, or muddy terrain. In
arctic climates, the aircraft must be towed
by the main gears, as an immense breakaway load, resulting from ice, frozen tires,
and stiffened grease in the wheel bearings
may damage the nose gear.
TM 55-1510-221-10
Do not tow or taxi aircraft with deflated
shock struts.
(2.) Towing. Towing lugs are provided on
the upper torque knee fitting of the nose strut. When
it is necessary to tow the aircraft with a vehicle, use
the vehicle tow bar. In the event towing lines are
necessary, use towing lugs on the main landing gear.
Use towing lines long enough to clear nose and/or
tail by at least 15 feet. This length is required to prevent the aircraft from overrunning the towing vehicle or fouling the nose gear.
d.. Ground Handling Under Extreme Weather
Conditions. Extreme weather conditions necessitate
particular care in ground handling of the aircraft. In
hot, dry, sandy, desert conditions, special attention
must be devoted to finding a firmly packed parking
and towing area. If such areas are not available, steel
mats or an equivalent solid base must be provided
for these purposes. In wet, swampy areas, care must
be taken to avoid bogging down the aircraft. Under
cold, icy, arctic conditions, additional mooring is
required, and added precautions must be taken to
avoid skidding during towing operations. The particular problems to be encountered under adverse
weather conditions and the special methods
designed to avoid damage to the 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.)
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.
2. Pull parking brake handle out. This
will cause the parking brake valves to
lock the hydraulic fluid under pressure
in the parking brake system, thereby
retaining braking action.
3. Release brake pedals.
Do not set parking brakes when the
brakes are hot during freezing ambient
temperatures. Allow brakes to cool before
setting parking brakes.
4. To release the parking brakes push in
on the parking brake handle.
b. The control lock (fig. 2-18) holds the
engine and propeller control levers in a secure position. It also holds the elevators and rudder at neutral
position and the ailerons in a staggered attitude, one
slightly "up" and the other slightly "down". Install
the control locks as follows:
1.
With engine and propeller control
levers in secure position, slide lock
onto control pedestal to prevent operation of levers.
2.
Install elevator and aileron lockpin vertically through pilot’s control column
to lock control wheel.
3.
Install rudder lock pin through flapper
door forward of pilot’s seat, making
sure rudder is in neutral position.
4.
Reverse steps 1 through 3 above to
remove control lock. Store control lock.
2-104. PARKING.
Parking is defined as the normal condition
under which the aircraft will be secured while on the
ground. This condition may vary from the temporary expedient of setting the parking brake and
chocking the wheels to the more elaborate mooring
procedures described under Mooring. The proper
steps for securing the aircraft must be based on the
time the aircraft will be left unattended, the aircraft
weights, the expected wind direction and velocity,
and the anticipated availability of ground and air
crews for mooring and/or evacuation. When practical head the aircraft into the wind, especially if
strong winds are forecast or if it will be necessary to
leave the aircraft overnight. Set the parking brake
and chock the wheels securely. Following engine
shutdown, position and engage the control locks.
Depress both brakes.
2-105. INSTALLATION OF PROTECTIVE COVERS.
The crew will insure that the aircraft protective
covers are installed.
NOTE
Cowlings and loose equipment will be
suitably secured at all times when left in
an unattended condition.
2-106. MOORING.
The aircraft is moored to insure its immovability, protection, and security under various weather
2-87
TM 55-1510-221-10
Figure 2-32. Parking, Covers, Ground Handling, and Towing Equipment
2-88
TM 55-1510-221-10
2-33. 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. Locate anchor rods
at point shown in figure 2-33. When
anchor kits are not available, use metal
stakes or deadman type anchors, providing they can successfully sustain a
minimum pull of 3000 pounds.
conditions. The following paragraphs give, in detail,
the instructions for proper mooring of the aircraft.
a. Mooring Provisions. Mooring points (fig.
2-33) are provided beneath the wings and tail. Additional mooring cables may be attached to each landing gear. General mooring equipment and procedures necessary to moor the aircraft, in addition to
the following, are given in TM 55-1500-204-25/1.
(1.) Use mooring cables of 1/4 inch diameter aircraft cable and clamp (clip-wire rope), chain
or rope 3/8 inch diameter or larger. Length of the
cable or rope will be dependent upon existing circumstances. Allow sufficient slack in ropes, chains,
or cable to compensate for tightening action due to
moisture absorption of rope or thermal contraction
of cable or chain. Do not use slip knots. Use bowline
knots to secure aircraft to mooring stakes.
(2.) Chock the wheels.
b. Mooring Procedures for High Winds. Structural damage can occur from high velocity winds;
therefore, if at all possible, the aircraft should be
moved to a safe weather area when winds above 75
knots are expected. Moored aircraft condition is
shown in figure 2-33. If aircraft must be secured use
the following steps:
1.
After aircraft is properly located, place
nose wheel in centered position. Head
aircraft into the wind, or as nearly so
as is possible within limits determined
by locations of fixed mooring rings.
When necessary, a 45 degree variation
of direction is considered to be satisfactory. Locate each aircraft at slightly
more than wing span distance from all
other aircraft. Position nose mooring
point approximately 3 to 5 feet downwind from ground mooring anchors.
2. Deflate nose wheel shock strut to
within 3/4 inch of its fully deflated
position.
3.
Fill all fuel tanks to capacity, if time
permits.
4. Place wheel chocks fore and aft of
main gear wheels and nose wheel. Tie
each pair of chocks together with rope
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 shown in figure
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
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. 2-32).
9. Secure propellers to prevent windmilling (fig. 2-32).
10. Disconnect battery.
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. The
storm appears to pass two times, each
time with a different wind direction.
This will necessitate turning the aircraft after the first passing. sli.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.
2-89
TM 55-1510-221-10
NOTE
IF STRONG WINDS ARE ANTICIPATED OR AIRCRAFT IS
TO BE LEFT UNATTENDED, PROPELLER RESTRAINT,
PITOT MAST, AND INTAKE COVERS MUST BE INSTALLED, AND THE FLIGHT CONTROLS LOCK
ENGAGED.
BEFORE TOWING, THE PROPELLER RESTRAINT MUST
BE INSTALLED WITH ONE PROPELLER BLADE IN THE
DOWN POSITION AS SHOWN.
THE USE OF DOUBLE OR SINGLE MOORING POINTS
FOR NOSE AND/OR WING TIEDOWNS IS DETERMINED
BY LOCAL OPTION DEPENDING ON TYPE AND AVAILABILITY OF AIRCRAFT SECURING EQUIPMENT.
USE ROPE ONLY (NYLON TYPE IF AVAILABLE) FOR
NOSE TIEDOWN (DETAIL A). ATTACH ROPE(S) TO AIRCRAFT AND GROUND MOORING POINTS IN A MANNER
THAT WILL PREVENT ROPE DAMAGE TO AIRCRAFT
COMPONENTS.
Figure 2-33. Mooring the Aircraft
2-90
TM 55-1510-221-10
CHAPTER 3
AVIONICS
Section I.
3-1. INTRODUCTION.
Except for mission avionics, this chapter covers
all avionics equipment installed in the RC-12H aircraft. It provides a brief description of equipment
covered, the technical characteristics and locations.
It covers systems and controls and provides the
proper techniques and procedures to be employed
when operating the equipment. For more detailed
operational information consult the vendor manuals
that accompany the aircraft loose tools.
3-2. AVIONICS EQUIPMENT CONFIGURATION.
The aircraft avionics covered consists of three
groups of electronic equipment. The communication
group consists of the interphone, UHF command,
backup VOW, VHF/AM-FM, VHF command and
HF command systems. The navigation group provides the pilot and copilot with the instrumentation
required to establish and maintain an accurate flight
course and position, and to make an approach on
instruments under Instrument Meteorological Conditions (IMC). The navigation group includes equipment for determining altitude, attitude, position,
destination, range and bearing, heading reference,
groundspeed, and drift angle. The transponder and
radar group includes an identification, position,
emergency tracking system, a radar system to locate
potentially dangerous weather areas, and a radar system to differentiate between friendly and unfriendly
search radar.
NOTE
All avionics equipment require a 3-minute warmup period. The weather radar
has an automatic time delay of 3 to 4
minutes.
3-3. POWER SOURCE.
a. DC Power. DC power for the avionics
equipment is provided by four sources: the aircraft
GENERAL
battery, left and right generators, and external
power. Power is routed through two 50-ampere circuit breakers to the avionics power relay which is
controlled by the AVIONICS MASTER POWER
switch on the overhead control panel (fig. 2-12).
Individual system circuit breakers and the associated avionics busses are shown in fig. 2-22. With the
switch in the ON (forward) position, the avionics
power relay is de-energized and power is applied
through both the AVIONICS MASTER POWER
No. 1 and No.2 circuit breakers to the individual avionics circuit breakers on the overhead circuit
breaker panel (fig. 2-26). In the OFF (aft) position,
the relay is energized and power is removed from
avionics equipment. When external power is applied
to the aircraft, the avionics power relay is normally
energized, removing power from the avionics equipment. To apply external power to the avionics
move the AVIONICS MASTER
equipment,
POWER switch to the EXT PWR position. This will
de-energize the avionics power relay and allow
power to be applied to the avionics equipment.
b. Single-Phase AC Power. AC power for the
avionics equipment is provided by two inverters.
The inverters supply 115-volt and 26-volt singlephase AC power when operated by the INVERTER
No. 1 or No.2 switches (fig. 2-12). Either inverter is
capable of powering all avionics equipment requiring AC power. AC power from the inverters is
routed through fuses in the nose avionics compartment.
c. Three-Phase AC Power. Three phase AC
electrical power for operation of the inertial navigation system and mission avionics is supplied by two
DC powered 3000 volt-ampere solid state three
phase inverters. The three phase inverters are controlled by two three-position switches located on the
mission control panel (fig. 4-1) placarded No. 1 INV
- OFF - ON - RESET and #2 INV - OFF - ON RESET.
3-1
TM 55-1510-221-10
Section II. COMMUNICATIONS
3-4. DESCRIPTION.
The communications equipment group consists
of an interphone system connected to individual
audio control panels for the pilot and copilot which
interface with VHF, UHF, BU VOW, VHF AM-FM
and HF communication units.
3-5. MICROPHONES, SWITCHES AND JACKS.
Boom and oxygen mask microphones can be utilized in the aircraft.
a. Microphone Switches. The pilot and copilot
are provided with individual microphone control
switches, placarded INTPH-XMIT-MIC, attached to
respective control wheels. A foot-actuated microphone switch is also positioned on the floorboards
forward of each pilot’s seat.
b. Controls and Functions.
(1.) Microphone control wheel switches
(fig. 2-17). Keys selected facility.
(a.) INTPH (depressed to first
detent). Keys interphone facility, disregards position
of transmitter selector switch.
(b.) XMIT (depressed full down).
Keys facility selected by transmitter select switch.
(2.) Floorboard microphone switches. Controls connection of selected microphone to audio
system.
depressed.
(a.) Held
selected microphone to audio system.
Connects
(b.) Released. Disconnects selected
microphone from audio system.
c. Microphone jack selector switches. T w o
switches, placarded MIC HEADSET - OXYGEN
MASK, are located on the extreme left and extreme
right of the instrument panel (fig. 2-29). These
switches provide a means of selecting whether the
headset microphone jack or the oxygen mask microphone jack is connected to the audio system.
d. Controls and Functions.
(I.) MIC HEADSET - OXYGEN MASK
switch. Selects microphone jack to connect to audio
system.
(a.) MIC HEADSET. Connects headset microphone to audio system.
3-2
(b.) OXYGEN M A S K . C o n n e c t s
microphone in oxygen mask to audio system.
3-6. AUDIO CONTROL PANELS.
a. Description. Separate but identical audio
control panels (fig. 3-1), serve the pilot and copilot.
The controls and switches of each panel provide the
user with a means of selecting desired reception and
transmission sources, and also a means to control
the volume of audio signals received for interphone,
communication and navigation systems. The user
selects between the UHF, VHF, BU VOW, VHF
AM-FM and HF transceivers. The audio control
panels are protected by respective 2-ampere AUDIO
PILOT and AUDIO COPILOT circuit breakers
located on the overhead circuit breaker panel (fig.
2-26).
b. Controls and Functions.
(I.) Master VOL control. Controls sidetone volume to headset. Also serves as final volume
adjustment for received audio from any source
before admission to headset.
(2.) Radios audio monitor controls. Each
is combination rotary control and on-off push-pull
switch, permitting both receiver selection and volume adjustment.
(a.) No. 1. On connects user’s headset
to audio from VHF-AM transceiver No. 1.
(b.) No.2. On connects user’s headset
to audio from the VHF/AM/FM transceiver.
(c.) No.3. On connects user’s headset
to audio from No. 1 UHF transceiver in use.
(d.) No.4. On connects user’s headset
to audio from HF or VOW transceivers.
(e.) No.5 On connects user’s headset
to audio from No. 2 UHF (BU VOW) transceiver.
(3.) NAV receiver audio monitor controls.
Combination volume control and “ON-OFF“
switches for NAV receivers.
(a.) NAV-A. On connects user’s headset to audio from VOR-1, VOR-2 or marker beacon
set in use.
(b.) NAV-B. On connects user’s headset to audio from TACAN or ADF set in use.
(4.) Microphone impedence select switch.
Two-position, thumb-actuated switch. Enables selec-
TM 55-1510-221-10
Figure 3-1. Audio Control Panel (Typical Pilot, Copilot)
tion of interface circuit with best impedence match
to microphone used.
(a.) The impedance of MIC 1 position is 5 Ohms.
(b.)
tion is 150 Ohms.
The impedance of MIC 2 posi-
(5.) Transmitter-interphone selector switch.
Connects microphone and headset to selected radio
transmitter or interphone line routing received
audio to headset. Bypasses control of respective
receiver audio switch.
(a.) PVT. Position not used.
(b.) ICS. Activates pilot-to-copilot
intercom.
(c.) No.1. Permits audio reception
from VHF-AM No. 1 transceiver. Routes key and
microphone signals to VHF-AM No. 1 transceiver.
(d.) No.2. Permits audio reception
from VHF/AM/FM transceiver. Routes key and
microphone signals to VHF/AM/FM transceiver.
(e.) No.3. Permits audio reception
from No. 1 UHF transceiver. Routes key and mic
signals to No. 1 UHF transceiver.
(f.) No.4. Permits audio reception
from HF or VOW transceivers. Routes key or microphone signals to transceiver.
(g.) No.5 Permits audio reception
from No. 2 UHF (BU VOW). Routes key and and
microphone signals to transceiver.
(6.) ICS select switch. Controls activation
of microphones.
(a.) HOT MIC. Admits speech to
interphone system without need to key selected
microphone.
(b.) NORM. Blocks speech from
interphone system unless selected microphone is
keyed.
(c.) ICS OFF. Deactivates interphone system.
c. Normal Operation.
(1.) Turn-on procedure: Both audio control panels are activated when electrical power is
applied to aircraft.
3-3
TM 55-1510-221-10
NOTE
It is presumed the AVIONICS MASTER
POWER switch is ON, and that normally
used avionics circuit breakers remain set.
The circuit breakers of routinely used avionic systems are normally left set.
(2.) Receiver operating procedure:
1.
6. If ICS OFF is selected intercom
function is switched off.
7. Volume control (selected transceiver) - Set for comfort.
d. Emergency Operation. Not applicable.
e. Shutdown Procedure.
1.
Receiver audio switches (audio
control panel) - As required.
2. Master volume control (audio
control panel) - Do not use. (Adjust volume control of system
being used.)
NOTE
Audio select switches and volume controls are routinely left in positions of normal use.
3. Move each receiver audio switch
ON then OFF, separately, to verify audio presence in headphones
for each system. Adjust volume.
(3.) Transmitter operating procedure:
1.
Transmitter-interphone
selector
switch (audio control panel) - Set
for transceiver desired.
2. Microphone jack selector switch
(instrument panel, fig. 2-29) - As
desired.
2. Leave controls and circuit breakers positioned for normal operation.
3-7. MARKER BEACON AUDIO CONTROL
PANEL (FIG. 3-2).
a. Description. The marker beacon audio control panel, located on the pedestal extension (fig.
2-7) allows the pilot or copilot to control the volume
of the marker beacon (MKR BCN). It also has controls for the selection of ADF voice or range filters
and MKR BCN HI-LO sensitivity.
b. Controls and Functions.
(1.)
ADF filter.
ADF Filter Switches. Controls selected
(a.) FILTER V-OFF switch. Selects
filter to block voice transmissions from ADF ground
station.
3. Control wheel microphone switch
(control wheel) - XMIT.
(b.) FILTER R-OFF switch. Selects
filter to block range transmissions from ADF ground
station.
4. Microphone switch (floorboard) Depress to transmit.
(2.) MKR BCN volume control. Adjusts
volume of marker beacon radio signals received.
(4.) Intercommunication procedure:
1. Transmitter-interphone
selector
switch (audio control panel) - ICS.
2. Microphone jack selector switch
(instrument panel, fig. 2-29) - As
desired.
3. ICS select switch (audio control
panel, fig. 3-1) - As desired
(NORM-HOT MIC-ICS OFF).
4. If HOT MIC is selected - Talk
when ready.
5.
3-4
AVIONICS MASTER POWER
switch (overhead control panel,
fig. 2-12) - OFF.
If NORM microphone is selected
- Depress microphone switch and
transmit.
(3.) MKR BCN HI-LO switch. Selects sensitivity of marker beacon receiver.
3-8. UHF COMMAND SET (AN/ARC-184).
a. Description. The UHF command set is a
line-of-sight radio transceiver which provides transmission and reception of amplitude modulated
(AM) signals in the ultra high frequency range of
225.000 to 399.975 MHz for a distance range of
approximately 50 miles. Channel selection is spaced
at 0.025 MHz. A separate receiver is incorporated to
provide monitoring capability for the UHF guard
frequency (243.0 MHz). UHF audio output is
applied to the audio panel where it is routed to the
headsets.
TM 55-1510-221-10
Figure 3-2. Marker Beacon Audio Control Panel
NOTE
The PRESET channel selector and manual frequency selectors are inoperative
when the mode selector is set to GUARD
position. The receiver-transmitter will be
set to the emergency frequency only.
The transmitter and receiver sections of the
UHF unit operate independently but share the same
power supply and frequency control circuits. Separate cables route transmit and receive signals to their
respective receiver/transmitter.
Complete provisions only are installed for a
TSEC/KY-28 voice security device to be located on
the LH forward avionics rack behind the pilot. The
UHF command set is protected by the 71/2 ampere
UHF circuit breaker on the overhead circuit breaker
panel (fig. 2-26). Figure 3-3 illustrates the UHF
command set. The associated blade type antenna is
shown in figure 2-1.
b. Controls and Functions. UHF control panel
(fig. 3-3):
(1.) Manual frequency selector/indicator
(hundreds). Selects and indicates hundreds digit of
frequency (2 or 3) in MHz.
(2.) Manual frequency selector/indicator
(tens). Selects and indicates tens digit of frequency
(0 through 9) in MHz.
(3.) Manual frequency selector/indicator
(units). Selects and indicates units digit of frequency
(0 through 9) in MHz.
(4.)
set channel.
Preset channel indicator. Displays pre-
(5.) Manual frequency selector/indicator
(tenths). Selects and indicates tenths digit of frequency (0 through 9) in MHz.
(6.) Preset channel selector. Selects one of
20 preset channel frequencies.
(7.) Manual frequency selector (hundredths
and thousandths). Selects hundredths and thousandths digits of frequency (00, 25, 50, or 75) in
MHz.
(a.) Mode selector. Selects operating
mode and method of frequency selection.
1 . MANUAL.
Enables the
manual selection of any one of 7,000 frequencies.
2. PRESET. Enables selection
of any one of 20 preset channels.
3-5
TM 55-1510-221-10
Figure 3-3. UHF Control panel (AN/ARC- 164)
3 . GUARD. Selection automatically tunes the main receiver and transmitter to
the guard frequency and the guard receiver is
enabled.
(d.) ADF. Not used.
c. Normal Operation.
(1.) Turn on procedure:
(b.) SOUELCH switch. Turns main
receiver squelch on or off.
(c.)
NOTE
VOL control. adjusts volume.
(8.) TONE pushbutton. When pressed,
transmits a 1,020 Hz tone on the selected frequency.
It is presumed aircraft power is on and
normally used avionic circuit breakers
remain depressed.
(9.) Function selector. Selects operating
function.
1.
Avionics master power switch
(overhead panel, fig. 2-12) - ON.
2.
Function select switch (UHF control panel, fig. 3-3) -MAIN or
BOTH position, as required.
(a.) OFF. Turns set off.
(b.) MAIN. Selects normal transmission with reception on main receiver.
(c.) BOTH. Selects normal transmission with reception on both the main receiver and
the guard frequency receiver.
3-6
TM 55-1510-221-10
NOTE
2.
UHF control panel (fig. 3-3) - Set
required frequency using either
PRESET CHAN control or MANUAL frequency select controls.
If function selector is at MAIN setting,
only the normal UHF communications
will be received. If selector is at BOTH
position, emergency communications on
the guard channel and normal UHF communications will both be received.
3. Microphone jack selector switch
(instrument panel, fig.2-29) - As
desired.
(2.) Receiver operating procedure:
4. Microphone switch - Depress to
transmit.
1.
(3.)
UHF audio monitor switch (No.3,
audio control panel) -ON, or
selector
transmitter-interphone
switch (audio control panel) No.3 position.
(6.) Shutdown procedure: Function selector switch (UHF control panel, fig. 3-3) -OFF.
d. UHF Command Set Voice Security Operation (KY-28).
2. Volume control (UHF control
panel) - Mid position.
NOTE
To use preset frequency (UHF control
Disregard operating procedures involving
the voice security control-indicator if unit
is not installed.
panel):
1.
Mode selector switch - PRESET
position.
2.
Preset channel selector switch Rotate to desired channel.
( 4 . ) T o use non-preset frequency (UHF
control panel):
1.
Mode selector switch - MANUAL
position.
2.
Manual frequency selectors (5) Rotate each knob to set desired
frequency digits.
NOTE
The PRESET channel selector and manual frequency selectors are inoperative
when the mode selector switch is set to
the GUARD position.
(I.) Turn-on procedure:
1.
2. Function switch (UHF control
panel) - BOTH.
(2.) Receiver operating procedure:
1.
To adjust volume when audio is not being
received, turn squelch switch OFF, adjust
volume for comfortable noise level, then
turn squelch switch ON.
3. Set required frequency using preset channel control or manual frequency selector.
4. Volume control - Adjust.
NOTE
To adjust volume when radio signals are
not being received, turn squelch switch
OFF, adjust volume for comfortable noise
level, then turn squelch disable switch
ON.
5.
4.
Squelch - As desired.
(5.) Transmitter operating procedure:
1. Transmitter-interphone selector
(audio panel control panel, fig.
3-1) - No.3 position.
Mode selector switch (UHF control panel, fig. 3-3) - As required.
2. Transmitter-interphone selector
switch (audio control panel, fig.
3-1) - No.3 position, or No.3
radio monitor control - On.
3. Volume - Adjust.
NOTE
Power switch (Voice Security
panel, fig 3-6) - ON.
Squelch switch - As required.
(3.) Transmitter
(PLAIN):
operating
procedure
selector
1. Transmitter-interphone
switch (audio control panel) No.3 position.
3-7
TM 55-1510-221-10
2.
Plain/cipher switch (voice security
control panel) -PLAIN.
3-8A. UHF COMMAND SET (AN/ARC-164).
3.
Microphone switch - Press.
a. Description. The UHF command set is a lineof-sight radio transceiver which provides transmission and reception of amplitude modulated (AM) signals in the ultra high frequency range of 225.000 to
399.975 MHz for a distance range of approximately
50 miles. Channel selection is spaced at 0.025 MHz.
A separate receiver is incorporated to provide monitoring capability for the UHF guard frequency (243.0
MHz). UHF audio output is applied to the audio
panel where it is routed to the headsets.
(4.) Transmitter operating procedure (CIPHER):
1.
Transmitter-interphone
selector
switch (audio control panel) No.3 position.
2.
Plain/cipher switch (voice security
control panel, fig. 3-6) - CIPHER.
(CIPHER indicator will be illuminated as long as switch is in
CIPHER position.)
3. RE-X/REG switch (voice security
control panel) - REG.
4. Microphone
switch Press
momentarily (interrupted tone
from voice security unit should
no longer be heard).
NOTE
No traffic will be passed if the interrupted
tone is still heard after pressing and
releasing the microphone switch.
5.
Microphone switch - Press (do not
talk). Wait until beep is heard
then speak into microphone.
(5.) Shutdown procedure:
1.
Function selector switch (UHF
control panel)- OFF.
2.
Power switch (Voice security control panel) - OFF.
e. UHF Command Set - Emergency Operation:
NOTE
Transmission on emergency frequencies
(guard channels) is restricted to emergencies only. An emergency frequency of 121.
500 MHz is also available on the VHF
command radio set.
NOTE
The PRESET channel selector and manual
frequency selectors are inoperative when
the mode selector is set to GUARD position. The receiver-transmitter will be set
to the emergency frequency only.
Existing capabilities of the HAVE QUICK moditied radio are preserved to the maximum extent possible when it is operated in the normal (non-hopping)
mode. No new procedures are required for normal
radio operation.
To operate in the AJ mode, the radio must first be
initialized. This initialization requires the setting of
two control entries into the radio, Word-of-Day
(WOD) and Time-of-Day (TOD). The WOD defines
the choice of frequency hopping pattern for the day.
The WOD choice is a managerial function and the
same WOD may be used for one or more days. The
TOD must be loaded into the clock contained within
the radio.
The transmitter and receiver sections of the UHF
unit operate independently, but share the same
power supply and frequency control circuits. Separate cables, route transmit and receive signals to their
respective receiver/transmitter.
The UHF command set is protected by the 7 1/2ampere UHF circuit breaker in the overhead circuit
breaker panel (fig. 2-6). Figure 3-2 illustrates the
UHF command set. The associated blade type
antenna is shown in figure 2-1.
b. Controls and Functions. UHF control panel
1. Transmitter-interphone selector switch
(audio control panel) - No.3 position.
(fig. 3-2):
2. Mode selector switch (UHF control
panel) - GUARD.
(1) Manual frequency selector/indicator
(hundreds). Selects and indicates hundreds digit of
frequency (2 or 3) in MHz.
3. Microphone switch - Press.
3-8
Change 4
TM 55-1510-221-10
(2) Manual frequency selector/indicator
(tens). Selects and indicates tens digit of frequency (0
through 9) in MHz.
(3) Manual frequency selector/indicator
(units). Selects and indicates units digit of frequency
(0 through 9) in MHz.
(4) Preset channel indicator. Displays preset
channel.
(d) ADF. Not used.
c. Normal Operation.
(1) Turn on procedure:
NOTE
It is presumed aircraft power is on and
normally used avionic circuit breakers
remain depressed.
1. Avionics master power switch - ON.
(5) Manual frequency selector/indicator
(tenths). Selects and indicates tenths digit of frequency (0 through 9) in MHz.
(6) Preset channel selector. Selects one of 20
preset channel frequencies.
(7) Manual frequency selector (hundredths
and thousandths). Selects hundredths and thousandths digits of frequency (00, 25, 50, or 75) in
MHz.
2. Function select switch - MAIN or
BOTH position, as required.
NOTE
If function selector is at MAIN setting,
only the normal UHF communications
will be received. If selector is at BOTH
position, emergency communications on
the guard channel and normal UHF communications will both be received.
(2) Receiver operating procedure:
(8) Mode selector. Selects operating mode
and method of frequency selection.
1. Transmitter-interphone selector
switch - No. 3 position.
(a) MANUAL. Enables the manual selection of any one of 7,000 frequencies.
2. UHF audio monitor switch - ON, No.
3 position.
(b) PRESET. Enables selection of any one
of 20 preset channels.
(c) GUARD. Selection automatically
tunes the main receiver and transmitter to the guard
frequency and the guard receiver is enabled.
(9) SQUELCH switch. Turns main receiver
squelch on or off.
3. Volume control - Mid position.
(3) To use preset frequency:
1. Mode selector switch - PRESET position.
2. Preset channel selector switch Rotate to desired channel.
(4) To use non-preset frequency:
(10) VOL control. Adjusts volume.
(11) TONE pushbutton. When pressed, transnits a 1,020 Hz tone on the selected frequency.
(12) Function selector. Selects operating
‘unction.
(a) OFF. Turns set off.
(b) MAIN. Selects normal transmission
with reception on main receiver.
(c) BOTH. Selects normal transmission
with reception on both the main receiver and the
guard frequency receiver.
1. Mode selector switch - MANUAL
position.
2. Manual frequency selectors (5) Rotate each knob to set desired frequency digits.
NOTE
The PRESET channel selector and manual
frequency selectors are inoperative when
the mode selector switch is set to the
GUARD position.
3. Volume - Adjust.
Change 4
3-8.1
TM 55-1510-221-10
Figure 3-2. UHF Control Panel (AN/ARC-164)
NOTE
To adjust volume when audio is not being
received, turn squelch switch OFF, adjust
volume for comfortable noise level, then
turn squelch switch ON.
4. Squelch - As desired.
(5) Transmitter operating procedure:
1. Transmitter-interphone selector - No.
3 position.
2. UHF control panel - Set required frequency using either PRESET CHAN
control or MANUAL frequency select
controls.
3. Microphone jack selector switch - As
desired.
4. Microphone switch - Depress to transmit.
(6) Shutdown procedure: Function selector
switch (fig. 3-2) - OFF.
3-8.2
Change 4
3-9. VOICE ORDER WIRE (AN/ARC-194).
A radio set identical in type and performance to
the UHF command set (fig. 3-3) is located in the
pedestal, to serve as voice order wire. This set provides the pilot and copilot with secure 2-way voice
communications. Complete provisions only are provided for a KY-58 voice security device. The voice
order wire set is protected by a 7 1/2 ampere VOW
circuit breaker on the overhead circuit breaker panel
(fig. 2-26). The voice order wire shares an antenna
mounted on the aircraft belly with the transponder
(lower antenna, fig 2-1).
TM 55-1510-221-10
3-10. VHF-AM COMMUNlCATlONS (VHF-206).
a. Description. VHF-AM communications
provide transmission and reception of amplitude
modulated signals in the very high frequency range
of 116.000 to 151.975 MHz for a range of approximately 50 miles, varying with altitude. A dual head
control panel (fig. 3-4) is mounted on the pedestal
extension, accessible to both the pilot and copilot.
The panel. provides two sets of control indicators,
frequency indicators, frequency select knobs, a single volume control, and a single selector switch for
quick frequency changing. Transmission audio is
routed by pilot and copilot No.1 transmitter selector
switches located on the audio control panel (fig.
3-1). Received audio is routed by pilot and copilot
No. 1 receiver audio switches (fig. 3-1), to the respective headsets. The VHF radio is protected by a 10ampere VHF circuit breaker on the overhead circuit
breaker panel (fig. 2-26). The associated antenna is
shown in figure 2-1.
b. Controls/Indicators and Functions.
(1.) Frequency indicator. Indicates set
operating frequency.
(2.) Control frequency indicators. Indicates
frequency selected (left or right active).
(3.) COMM TEST switch. Overrides automatic squelch circuit.
(4.) Frequency selectors. Select desired set
operating frequency.
(5.) TRANS switch. Selects right or left
frequency control selectors.
(6.) VOL-OFF control. Adjusts volume of
received audio, turns set ON or OFF.
c.
VHF-AM Set - Normal Operation.
(1.) Turn-on procedure: Volume control Turn clockwise (ON).
Change 4
3-8.3/(3-8.4 blank)
TM 55-1510-221-10
Figure 3-4. VHF-AM Control Panel (VHF-20B)
(2.) Receiver operating procedure:
1.
Transmitter-interphone
selector
switch (audio control panel, fig.
3-1) - No.1 position, or radio
monitor control No.1 -ON.
2. Frequency selector - Set desired
frequency.
3.
(3.)
2.
Transmitter-inter-phone selector
switch (audio control panel, fig.
3-1) - No.1 position.
Microphone switch - Press.
(4.) Shutdown procedure: Volume control
- Turn counterclockwise (OFF).
d.
selector
Transmitter-interphone
switch (audio control panel, fig.
3-1) - No. 1 position.
2. Frequency selector (VHF control
panel, fig. 3-4) -121.500 MHz
(emergency frequency).
3. Microphone switch - Press.
Volume control - As required.
Transmitter operating procedure:
1.
1.
VHF-AM Set Emergency Operation.
NOTE
Transmission on emergency frequency
(121.500 MHz) is restricted to emergencies only. Emergency frequency 243.000
MHz (guard channel) is also available on
the UHF command radio.
3-11. VHF AM-FM COMMAND SET (AN/ARC199).
a. Description. The VHF AM-FM Command
Set provides for normal and secure 2-way AM voice
communication in the very high frequency range of
116.000 to 151.975 MHz and FM voice communication in the 30.000 to 87.975 MHz band. Twenty
channels may be preset. Audio signals are applied
through the No.2 position of the transmitterinterphone selector switches and through the No.2
receiver audio switches on the pilot’s and copilot’s
audio control panels (fig. 3-1). Complete provisions
only are installed for a TSEC/KY-28 voice security
device. Circuits are protected by a lo-ampere VHF
AM-FM circuit breaker on the overhead circuit
breaker panel (fig. 2-26). Figure 3-5 illustrates the
VHF AM-FM control panel. The associated antenna
is shown in figure 2-1.
3-9
TM 55-1510-221-10
Figure 3-5. VHF AM-FM Control Panel (AN/ARC-186)
3-10
TM 55-1510-221-10
b. Controls/Indicators and Functions.
(1.) 10 MHz selector. Selects receivertransmitter frequency in increments of 10 MHz
from 30 to 150 MHz. Clockwise rotation increases
frequency.
(2.) 10 MHz indicator. Indicates manually
selected receiver-transmitter frequency in 10 MHz
increments from 30 to 150 MHz.
(3.) 1.0 MHz selector. Selects receivertransmitter frequency in 1.0 MHz increments.
Clockwise rotation increases frequency.
(4.) I.0 MHz indicator. Indicates manually selected receiver-transmitter frequency in 1.0
MHz increments.
(5.) 0.1 MHz indicator. Indicates manually selected receiver-transmitter frequency in 0.1
MHz increments.
(6.) 0.1 MHz selector. Selects receivertransmitter frequency in 0.1 MHz increments.
Clockwise rotation increases frequency.
(a.) NB. Limits selectivity to narrowband intermediate frequency.
(b.) WB. Limits selectivity to wideband intermediate frequency of FM band.
Momentary
(c.) MEM LOAD.
switch. If pressed, loads manually selected frequency
in preset channel memory.
(15.) AM SQUELCH control. Screwdriver
adjustable potentiometer. Squelch fully overridden
at full counterclockwise position. Clockwise rotation
increases input signal required to open squelch.
(16.) Mode selector switch. Three-position
rotary switch.
(a.)
OFF.
Shuts
off
receiver-
transmitter.
(b.) TR. Selects transmit/receive
modes.
(c.) DF. Not operational.
(I 7.) SQ-DIS-TONE select switch. Threeposition switch.
(7.) 0.025 MHz indicator. Indicates manually selected receiver-transmitter frequency in 0.025
MHz increments,
(a.) Center. Selects squelch function.
(b.) SQ-DIS. Shuts off squelch func-
(8.) 0.025 MHz selector. Selects receivertransmitter frequency in 0.025 MHz increments.
Clockwise rotation increases frequency.
tion.
(9.) Preset CHAN indicator. Indicates
selected preset channel.
(18.) Frequency control/emergency select
switch. Three-position switch.
(10.) Preset CHAN selector. Selects preset
channel from 1 to 20. Clockwise rotation increases
number selected.
selection.
(11.) Preset channels freq. list. Writing area
to keep track of preset channels,
selection.
(12.) LOCKOUT FM-AM switch. Screwdriver adjustable three-position switch. Warning
tone announces lockout.
(a.)
Center. Selects AM or FM band.
(b.)
AM. Shuts off AM band.
(c.)
FM. Shuts off FM band.
(13.) FM SQUELCH control. Screwdriver
adjustable potentiometer. Squelch fully overdriven
at full counterclockwise position. Clockwise rotation
increases input signal required to open squelch.
(14.) WB-NB-MEM LOAD switch. Threeposition switch.
(c.) TONE. Transmits tone of
approximately 1000 Hz.
(a.) PRE. Enables preset channel
(b.)
MAN. Enables manual frequency
(c.) EMER-AM-FM. Selects a prestored guard channel.
(19.) VOL control. Clockwise rotation
increases volume.
c. Normal Operation.
(1.) Turn-on procedure: Mode selector
switch (VHF AM-FM control panel, fig. 3-5) -TR.
(2.)
Receiver operating procedure:
1.
Frequency control emergency
selector switch (fig. 3-5) -MAN or
PRE, as desired.
2. Transmitter-interphone selector
switch (audio control panel, fig.
3-11
TM 55-1510-221-10
3-1) - No.2 position, or radio
monitor control No.2 -ON.
are provided to install either the KY-28
or the KY-58 voice security system on the
LH fwd avionics rack behind the pilot
(fig. 2-2).
3. Manual frequency/preset channel
selectors - Set desired frequency.
4.
Volume control - As required.
(3.) Transmitter operating procedure:
1.
Transmitter-interphone
selector
switch (audio control panel, fig.
3-1) - No.2 position.
2. Microphone switch - Press.
(4.) Shutdown procedure: Mode selector
switch (fig. 3-5) - OFF.
d.
VHF AM-FM Emergency Operation:
(1.) Emergency AM Mode:
1. Transmitter-interphone
selector
switch (audio control panel, fig.
3-1) - No.2 position.
2.
Mode selector switch - TR.
control/emergency
3. Frequency
selector switch - EMER AM.
NOTE
Selecting EMER AM or FM automatically
disables secure speech function and
enables normal voice communication.
4. Microphone switch - Press.
(2.) Emergency FM Mode:
1.
2.
Transmitter-interphone
selector
switch (audio control panel, fig.
3-1) - No.2 position.
Mode selector switch - TR.
control/emergency
3. Frequency
selector switch - EMER FM.
4. Microphone switch - Press.
(3.) Shutdown Procedure: Shutdown mode
select switch - OFF.
3-12. VOICE SECURITY SYSTEM TSEC/KY-28
(PROVISIONS ONLY)
NOTE
Voice security system TSEC/KY-58 may
be installed in lieu of voice security system TSEC/KY-28. Complete provisions
3-12
a. Description. The KY-28 voice security system provides secure (ciphered) two-way voice communications for the pilot and copilot in conjunction
with the UHF and VHF/AM/FM command sets,
and the backup VOW set. System circuits are protected by the VHF, VHF/AM/FM, RADIO RELAY,
and BU VOW circuit breakers on the overhead circuit breaker panel (fig. 2-26). Figure 3-6 illustrates
the KY-28 voice security (CIPHONY) control indicator.
b. Controls/Indicators and Functions. Voice
security control/indicator (LH fwd avionics rack
behind the pilot) (fig. 2-2).
(1.) POWER ON switch. Turns set on or
Off.
NOTE
The POWER ON switch must be in ON
position for FM liaison or secure mission
operations in either the plain or cipher
mode.
(2.) POWER ON indicator. Illuminates
when POWER ON switch is placed in ON (up) position.
(3.) PLAIN indicator. Illuminates when
PLAIN/CIPHER switch is in PLAIN position.
(4.) PLAIN/CIPHER switch.
(a.) PLAIN. Enables unciphered
communications on FM liaison set.
(b.) CIPHER. Enables ciphered communications on FM liaison set.
(5.) RE-X/REG switch.
(a.) RE-X. Enables re-transmission
of ciphered communications at a distant location.
(b.) REG. Enables normal cipher or
plain communications.
(6.) ZEROIZE switch. Normally OFF.
Place in ON position during emergency situations to
neutralize and make inoperative the associated
cipher equipment.
(7.) CIPHER indicator. Illuminates when
PLAIN/CIPHER switch is in CIPHER position.
TM 55-1510-221-10
Figure 3-6. Voice Security Control Indicator (C-8157/ARC)
c.
ation.
5. PLAIN/CIPHER switch (voice
security panel) - As required.
VHF/AM/FM Set and Voice Security Oper-
(1.) Turn-on procedure: P O W E R O N
switch (Voice security panel, fig. 3-6) - ON.
(3.) Transmit
1.
NOTE
The POWER ON switch must be in ON
position, regardless of the mode of the
operation, whenever the voice security
(CIPHONY) KY-28 is installed in the aircraft.
operating
procedure
(PLAIN):
selector
Transmit/interphone
(audio panel) - No. 2 position.
2. PLAIN/CIPHER switch (Voice
security panel) - PLAIN.
3. Microphone switch - Press.
(4.) Transmit operating procedure (CIPHER):
1.
control
SQUELCH
(VHF/
AM/FM panel) - As required.
selector
Transmit/inter-phone
(audio panel) - No. 2 position.
2.
selector
2. Transmitter-interphone
(audio panel, fig. 3-1) -#2 position. or Audio monitor control #2
- ON.
PLAIN/CIPHER switch (Voice
security panel) - CIPHER. Indicator will be on while switch is in
CIPHER position.)
3.
RE-X/REG switch (Voice security
panel) - As required. (Set RE-X
position only if distant station is
using re-transmitting equipment.)
4.
Press
switch Microphone
momentarily (interrupted tone
(2.) Receive operating procedure:
1.
3. Mode selector (VHF/AM/FM
panel) - TR.
(VHF/
selectors
4. Frequency
AM/FM panel) - As required.
3-13
TM 55-1510-221-10
from voice security unit should
no longer be heard).
NOTE
No traffic will be passed if the interrupted
tone is still heard after pressing and
releasing the microphone switch.
5.
Microphone switch- Press (do not
talk). Wait until beep is heard,
then speak into microphone.
(5.) Shutdown procedure:
1.
Mode selector (VHF/AM/FM
panel) - OFF.
2.
POWER ON switch (Voice security panel) - OFF.
(5.) RE-X-REG. Two-position switch.
(a.) RE-X. Enables re-transmission
of ciphered communications at a distant location.
(b.) REG. Enables normal cipher or
plain communications.
(6.) ZEROIZE switch. Normally OFF.
Place in ON position during emergency situations to
neutralize and make inoperative the associated
cipher equipment.
(7.) CIPHER indicator. Illuminates when
PLAIN-CIPHER switch is in CIPHER position.
VHF AM-FM Set and Voice Security Oper-
c.
ation.
(1.) Turn-on procedure: Power switch
(voice security panel, fig. 3-6) - ON.
3-13. VOICE SECURITY SYSTEM TSEC/KY-58
(PROVISIONS ONLY).
a. Description. The TSEC/KY-58 voice security system provides secure (ciphered) two-way voice
communications for the pilot and copilot in conjunction with the UHF and VHF AM-FM command
sets, and the voice order wire set. The control indicator is located in the forward avionics rack behind
the pilot. System circuits are protected by the VHF,
VHF AM-FM and BU VOW circuit breakers on the
overhead circuit breaker panel (fig. 2-26).
NOTE
The power switch must be in ON position, regardless of the mode of the operation,
whenever
the voice security
(CIPHONY) KY-58 is installed in the aircraft.
(2.) Receive operating procedure:
1.
Squelch control (VHF AM-FM
panel, fig. 3-5) - As required.
2.
Transmitter-interphone
selector
switch (audio control panel, fig.
3-1) - No.2 position, or radio
monitor control No.2 -ON.
NOTE
3.
The power switch must be in ON position
for FM or secure mission operations in
either the plain or cipher mode.
Mode selector switch (VHF
AM-FM control panel, fig. 3-5) TR.
4.
(VHF
Frequency
selectors
AM-FM control panel) - As
required.
5.
Plain-cipher switch (voice security
control panel) - As required.
b. Controls/Indicators and Functions.
(1.) POWER ON switch. Turns set on or
off.
(2.) POWER ON indicator. Illuminates
when POWER ON switch is placed in ON (up) position.
(3.) PLAIN indicator. Illuminates when
PLAIN-CIPHER switch is in PLAIN position.
(4.) PLAIN-CIPHER. Selects unciphered
or ciphered communications on FM set.
(a.) PLAIN. Enables unciphered
communications on FM set.
(b.) CIPHER. Enables ciphered communications on FM set.
3-14
(3.) Transmitter
operating
procedure
(PLAIN):
1.
Transmitter-interphone
selector
switch (audio control panel, fig.
3-1) - No.2 position.
2. Plain-cipher switch (Voice security control panel) -PLAIN.
3. Microphone switch - Press.
TM 55-1510-221-10
(4.) Transmitter operating procedure (CIPHER):
selector
1. Transmitter-interphone
(audio control panel, fig. 3-1) No.2 position.
2.
Plain-cipher switch (voice security
panel) - CIPHER. (Indicator will
be illuminated while switch is in
CIPHER position.)
The HF system has two methods of frequency
selection. The first method is called direct tuning
(frequency agile). The second is a channelized operation in which desired operating frequencies are preset, stored and referenced to a channel number.
b. Controls/Indicators and Functions (HF
Control Panel, jig. 3-7.
(1.)
FREQ display. Displays frequency
selected.
3. RE-X-REG switch (voice security
panel) - As required. (Set RE-X
position only if distant station is
using re-transmitting equipment.)
(2.) MODE display.
LSB, AM, or USB mode.
switch Press
4. Microphone
momentarily (interrupted tone
from voice security unit should
no longer be heard.)
selected.
NOTE
No traffic will be passed if the interrupted
tone is still heard after pressing and
releasing the microphone switch.
5.
Microphone switch - Press (do not
talk). Wait until beep is heard,
then speak into microphone.
(3.)
Displays selected
CHANNEL display. Displays channel
(4.) Light sensor. The light sensor is a
photocell which adjust brigntness of the display.
(5.) MODE switch. The mode switch is a
momentary pushbutton switch that selects LSB, AM
or USB.
(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.
(5.) Shutdown procedure:
1.
Mode selector switch (VHF
AM-FM panel) - OFF.
2. Power switch (voice security
panel) - OFF.
3-14. HF COMMUNICATION SET (KHF-950).
a. Description. The HF command set (fig. 3-7)
provides long-range voice communications within
the frequency range of 2.0 to 29.99 MHz and
employs either standard amplitude modulation
(AM), lower sideband (LSB), or upper sideband
(USB) modulation. The distance range of the set is
approximately 2,500 miles and varies with atmospheric conditions. With the capability to preset and
store 99 frequencies for selection during flight, the
system also allows for selection of other frequencies
manually (direct tuning), or reprogramming of any
preset frequency. The system will automatically
match the antenna by keying the microphone. Power
to the system is routed through a 25 ampere circuit
breaker placarded HF PWR. The receiving portion
of the system is protected by a 5 ampere circuit
breaker placarded HF REC. Both circuit breakers
are located on the overhead circuit breaker panel.
NOTE
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 (semi-duplex), 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. T h i s
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
3-15
TM 55-1510-221-10
Figure 3-7. HF Control Panel (KCU-951)
NOTE
channel control from 1 through 99, displayed at the
right end of the display window.
(9.) STO (Store) recessed switch. Stores
displayed data when programming preset channels.
Applies
(10.) OFF- VOLUME control.
power to the unit and controls the audio output
level.
(11.) SQUELCH control. Provides variable
squelch threshold control.
(12.) CLARIFIER control. Provides 250 Hz
of local oscillator adjustment.
c. Normal Operation.
(1.) Turn on procedure:
NOTE
It is presumed aircraft power is on and
normally used avionic circuit breakers
remain depressed.
3-16
Aircraft can be configured for either HF
or VOW on position 4 of Audio control
panel (fig. 3-1).
1.
AVIONIC MASTER POWER
switch - ON.
2. OFF-VOLUME switch - Turn
clockwise out of OFF position.
Adjust volume as desired.
(2.) Frequency operation (Simplex only):
OFF-VOLUME Switch - Turn clockwise out of OFF
position. Adjust volume as desired.
TM 55-1510-221-10
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.
NOTE
ch 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.
(5.) Receiver operating procedure:
button
2.
Select the channel to be preset.
4. Set the desired frequency. (Refer
to frequency tuning)
5. Push and release STO button
once.
NOTE
(3.) Direct frequency tuning (Simplex
FREQ/CHAN
(FREQ).
Stow the cursor if a frequency
digit is flashing.
3. Set the desired operating mode
(LSB,USB,or AM).
only).
1.
1.
out
2. Select desired mode (USB,LSB,or
AM).
3. Select digit to be changed (outer
knob), digit (cursor) will flash.
4. Select numerical value of digit
(inner knob).
"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.
5. Stow cursor (or repeat procedure
for additional changes).
6. To return to an operating mode,
push the PGM button.
6. Tune antenna coupler (press
microphone button).
(a.) Simplex operation: Setting a
channel up for simplex operation (receive and transmit on the same frequency).
(4.) Channel Programming.
1.
FREQ/CHAN button in (cursor
stowed).
2.
PGM button in (PGM displayed).
3.
Select channel to be preset.
4.
Set mode (LSB,USB or AM).
5.
Set desired frequency. (Refer to
frequency tuning)
6.
Push and release STO button
twice.
NOTE
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
The first press of the STO button stores the frequency in the receive position and the second press
stores the same frequency in the transmit position.
The second push also stores the cursor.
If another channel is to be reset, use the smaller
concentric knob to select the channel and repeat the
3-17
TM 55-1510-221-10
steps for selecting a new frequency. The cursor was
automatically stowed. To return to one of the operating modes, push the PGM button again.
(b.) Semi-duplex operation: Setting a
channel for semi-duplex (transmit on one frequency
and receive on another).
1.
Select channel to be preset.
2. Set desired frequency. (Refer to
frequency selection)
3.
Set mode (LSB,USB, or AM).
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 an operating mode,
push the PGM button.
NOTE
The mode for each channel (LSB, USB or
AM) is stored along with the frequency. If
the mode is changed, the system will
receive and transmit in the mode selected
for transmit.
d.
Shutdown. Off/Volume switch - OFF.
e. HF Command Set - Emergency operation.
Not applicable.
3-18
3-15. EMERGENCY LOCATOR TRANSMITTER
(ELT).
a. Description. An emergency locator transmitter is provided to assist in locating an aircraft
and crew in the event an emergency landing is
necessitated. The output frequency is 121.5 and 243
MHz simultaneously. Range is approximately lineof-sight. The transmitter unit has separate function
control switches located on one end of the case. In
the event the impact switch has been inadvertently
actuated, the beacon can be reset by firmly pressing
the pushbutton RESET switch on the front of the
case. The RESET switch and a 3-position toggle
switch, placarded ARM, OFF and ON, also on the
transmitter case, may be actuated by inserting one
finger through a small, round, spring-loaded door on
the left side of the aft fuselage (fig. 3-8). The transmitter unit is accessible through a service panel
located on the bottom of the aft fuselage.
b. Controls and Functions.
(1.) RESET switch. When pressed, resets
transmitter.
(2.) Function switch. Selects operating
mode of set.
(a.) ARM. Arms set to be actuated
by impact switch (normal mode).
(b.) OFF. Turns set off.
(c.) ON. Manually activates transmitter for test or emergency purposes.
TM 55-1510-221-10
Figure 3-8. Emergency Locator Transmitter (Narco 03716-0300)
3-19
TM 55-1510-221-10
Section III. NAVIGATION
3-16. DESCRIPTION.
The navigation equipment group provides the
pilot and copilot with instrumentation required to
establish and maintain an accurate flight course and
position, and to make an approach on instruments
under Instrument Meteorological Conditions (IMC).
The navigation configuration includes equipment
for determining attitude, position, destination range
and bearing, heading reference and groundspeed.
3-17. RADIO MAGNETIC INDICATORS (RMI).
a. Description. The pilot and copilot are each
provided with identical radio magnetic indicators
(RMI) (fig. 3-9) located on the instrument panel
(fig. 2-29). Each unit serves as a navigational aid for
the respective user and, by means of individual
source select switches, will display aircraft magnetic
or directional gyro heading and VOR, TACAN, INS
or ADF bearing information. The pilot’s RMI is
protected by the 1-ampere No.1 RMI circuit breaker
on the overhead circuit breaker panel (fig. 2-26) and
the 3.0-ampere F13 fuse on the No.1 junction box.
The copilot’s RMI is protected by the 1-ampere
No.2 RMI circuit breaker on the overhead circuit
breaker panel and the 3.0-ampere F9 fuse on the
No.1 junction box.
b. Controls and Functions.
(1.) Pilot’s COMPASS No. 1 - No.2 switch.
Selects desired source of magnetic heading information for display on pilot’s HSI and copilot’s RMI.
(a.) No.1. Selects compass system
No.1 for display control.
(b.) No.2. Selects compass system
No.2 for display control.
COMPASS
(2.) Copilot’s
No. l-No.2
switch. Selects desired source of magnetic heading
information for display on copilot’s HSI and pilot’s
RMI.
(a.) No.1. Selects compass system
No.1 for display.
(b.) No.2. Selects compass system
No.2 for display.
(3.) RMI select switch. Selects which of
two signals will be displayed on respective RMI sin-
Figure 3-9. Radio Magnetic Indicator (RMI) (332C-10)
3-20
TM 55-1510-221-10
gle- needle pointer, if single-needle switch is in the
VOR-TACAN position.
(6.) Single needle pointer. Indicates bearing selected by single needle switch.
(a.) VOR 1. Selects VOR 1 bearing
signals for display.
(7.) Single needle switch. Selects desired
signal to be displayed on single needle pointer.
(b.) TACAN. Selects TACAN bearing
signal for display.
mation.
c.
Indicators and Functions (RMI, fig. 3-9).
(1.) Double needle pointer. Indicates bearing selected by double needle switch.
(2.) Compass card. Indicates aircraft heading at top of dial.
(3.) Heading index. Reference point for
aircraft heading.
(4.)
pass signal.
Warning flag. Indicates loss of com-
(5.) Double needle switch. Selects desired
signal to be displayed by double needle pointer.
(a.)
ADF. Selects ADF bearing infor-
mation.
(b.) VOR. Selects VOR 2 bearing
information.
(a.) INS. Selects INS bearing infor(b.) VOR-TACAN. Selects signal as
determined by RMI select switch on instrument
panel, either VOR 1 or TACAN.
3-18. HORIZONTAL SITUATION INDICATORS.
a. Description. The pilot and copilot have separate HSI instruments on respective instrument
panel sections (fig. 3-10 and 3-11). Each HSI combines displays to provide a map-like presentation of
the aircraft position with respect to magnetic heading. Each indicator displays aircraft heading, course
deviation, and glideslope data. The pilot’s HSI
allows the desired course and heading to be input to
the autopilot. Course deviation data is supplied to
the HSI by the VOR 1 or VOR 2 systems, the
TACAN, or the INS. Glideslope data is supplied by
the VOR 1 or VOR 2 systems. The HSI displays
warning flags when the VOR, TACAN, INS or glideslope signals are lost or become unreliable.
Figure 3-10. Pilot’s Horizontal Situation Indicator (RD-650B)
3-21
TM 55-1510-221-10
b. Controls/Indicators and Functions (Pilot's
HSI, fig. 3-10).
(1.) Distance display. Provides digital displays of DME/TACAN or INS waypoint distance.
TACAN distance is displayed in 1/10 mile increments. INS distance to waypoint is displayed in
whole mile increments. The display will show dashes
when the distance input data is invalid or absent.
(2.) Rotating heading (azimuth) dial. Displays gyro stabilized magnetic compass information
on a dial which rotates with the aircraft throughout
360 degrees. The azimuth ring is graduated in 5
degree increments.
(3.) Lubber line. Fixed heading marks
located at the fore (upper) and aft (lower) position.
(4.) HDG flag. Indicates loss of reliable
heading information.
(5.) Heading bug. The notched 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, the heading bug maintains its position on
the heading dial. The difference between the bug
and the fore (upper) lubber line index is the amount
of heading select error applied to the flight director
computer. In the heading mode the ADI will display
the proper bank commands to turn to and maintain
this selected heading.
(6.) Course display. Provides a digital
readout of selected magnetic course.
(7.) Course pointer. The yellow course
pointer is positioned on the heading dial by the
remote course knob, to a magnetic bearing that coincides with the selected course being flown. The
course pointer rotates with the heading dial to provide a continuous readout of course error to the
computer.
(8.) Bearing pointer. Indicates ADF or
NAV relative bearing as selected by the bearing
pointer source switch.
(9.) ADF annunciator. When illuminated,
indicates ADF bearing information is being displayed.
(10.) Bearing pointer source switch. T h e
bearing pointer source switch, located on the pilot’s
HSI, provides for selecting between ADF or NAV
bearing information as presented by the bearing
pointer. Each push of the select switch alternates
selection of ADF or NAV. Upon power-up or following long-term power interruption, NAV is displayed.
3-22
(1 I.) NAV annunciator. When illuminated,
indicates NAV bearing information is being displayed.
(12.) NAV flag. Indicates loss of VOR,
TACAN or INS information, or unreliable navigation signal.
(13.) Compass synchronization annunciator. The compass synchronization annunciator consists of a dot and X symbol display. When the compass system is in the slaved mode, the display will
oscillate between the dot and X symbol, indicating
the heading dial is synchronized with a gyro stabilized magnetic heading.
(14.) Course deviation dots. In VOR or
TACAN operation, each dot represents 5 degree
deviation from the centerline (± 10 degrees). In ILS
operation, each dot represents 1 degree deviation
from the centerline. In INS operation, each dot represents 3.75 nautical miles deviation from centerline.
(15.) Aircraft symbol. The fixed miniature
aircraft symbol corresponds to the longitudinal axis
of the aircraft and lubber line markings. The symbol
shows aircraft position and heading with respect to
a radial course and the rotating heading (azimuth)
dial.
(16.) Course deviation bar. The course deviation bar represents the centerline of the selected
VOR, TACAN, INS or localizer course. The miniature aircraft symbol pictorially shows actual aircraft
position in relation to this selected course.
(I 7.) VERT flag. Covers glide slope pointer
when not receiving glide slope information.
(18.) Glide slope 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 degree displacement.
(19.) 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 radial or
towards the INS waypoint.
(20.) Navigation source annunciators. Five
different annunciators display navigation data
sources. They are: TAC for TACAN, GPS (not
used), INS, NV2 for VOR 2, NV1 for VOR 1. WPT
indicates arrival at INS waypoint.
(21.) Course knob (located on the pedestal).
Positions the course pointer.
TM 55-1510-221-10
(22.) Heading knob (located on the pedestal). Positions the heading bug to a preselected heading.
PILOT SELECT annunciator will illuminate to notify the copilot that both pilots
have selected the same receiver.
(23.) Pilot’s COURSE INDICATOR selector
switch (fig. 2-29). Selects desired source of data for
display on pilot’s HSI and input to autopilot flight
computer.
(I.) Compass synchronization annunciator. The compass synchronization annunciator consists of a dot and X symbol display. When the compass system is in the slaved mode, the display will
oscillate between the dot and X symbol, indicating
the heading dial is synchronized with a gyro stabilized magnetic heading.
(a.) VOR 1. Selects data from VOR
1 system.
(b.)
VOR 2. Selects data from VOR
2 system.
(c.) TACAN.
TACAN system.
Selects data from
(d.) INS. Selects data from INS.
c. Controls/Indicators and Functions (Copilot’s HSI. fig 3-11).
NOTE
If both the pilot and copilot COURSE
INDICATOR select switches are in the
same position, except INS, the pilot has
sole control of course select functions.
The copilot can only monitor deviation
displays from the selected system. A
(2.) VERT flag. Indicates that the information displayed by the glideslope pointer is invalid
and should not be used.
(3.) Rotating heading (azimuth) dial. Displays gyro stabilized magnetic compass information
on a dial which rotates with the aircraft throughout
360 degrees. The azimuth ring is graduated in 5
degree increments.
(4.) Azimuth marks. Fixed azimuth marks
are at 45° bearings throughout 360 degrees of the
compass card for quick reference.
(5.) Course deviation bar. The course deviation bar represents the centerline of the selected
VOR, TACAN, INS or localizer course. The miniature aircraft symbol pictorially shows actual aircraft
position in relation to this selected course.
Figure 3-11. Copilot’s Horizontal Situation Indicator (RD-550)
3-23
TM 55-1510-221-10
(6.) Digital COURSE counter. Provides a
digital readout of selected magnetic course.
(7.) HDG flag. Indicates loss of reliable
heading information.
(8.) Lubber line marks. Fixed heading
marks located at the fore (upper) and aft (lower)
position.
(9.) 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 or TACAN radial or
INS or localizer course. The course pointer rotates
with the heading dial to provide a continuous readout of course error to the computer.
(10.) DIST display. Provides digital display
of station distance.
(Il.) Heading bug. The notched 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.) Bearing pointer. The bearing pointer
provides magnetic bearing to a selected TACAN or
VOR station or INS waypoint.
(13.) Glideslope 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 degree displacement.
(14.) NAV flag. Indicates that information
derived from the selected navigational source (VOR,
TACAN or INS) is invalid and should not be used.
(15.) To-from pointers. 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 or TACAN
radial or toward INS waypoint.
of the aircraft and lubber line markings. The symbol
shows aircraft position and heading with respect to
a radio course and the rotating heading (azimuth)
dial.
(19.) Course knob. Positions the course
indicator.
COURSE INDICATOR
(20.) Copilot's
switch fig. 2-29). Selects desired source of data for
display on copilot’s HSI.
(a.) VOR 1. Selects data from VOR
1 system.
(b.) VOR 2. Selects data from VOR
2 system.
(c.) TACAN. S e l e c t s d a t a f r o m
TACAN system.
(d.) INS. Selects data from INS.
3-19. PILOT’S ATTITUDE DIRECTOR INDICATOR.
a. Description. The pilot’s attitude director
indicator (ADI) (fig. 3-12) combines the attitude
sphere display with computed steering information
to provide the commands required to intercept and
maintain a desired flight path. It also contains an
eyelid display, expanded localizer, glide slope, radio
altitude display, rate-of-turn indicator, mode annunciators, go-around and decision height annunciators,
and inclinometer. Any warning flag in view indicates that portion of information is unreliable.
b. Controls/Indicators and Functions.
(1.) 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.
(2.) Roll attitude index. Displays actual
roll attitude through a movable index and fixed
scale reference marks at 0, 10, 20, 30, 45, 60 and 90
degrees.
(16.) Heading knob. Positions the heading
bug to a preselected compass heading.
(3.) GA (go-around) annunciator. Illuminates when go-around mode has been selected.
(17.) Course deviation dots. I n V O R ,
TACAN or INS operation, each dot represents a 5
degree deviation from the centerline (° 10 degrees).
In ILS operation, each dot represents 1 degree deviation from the centerline. In INS operation, each
dot represents a 3.75 nautical miles deviation from
centerline.
(4.) SPD annunciator. Illuminates when
airspeed is being held by the flight director, in the
IAS mode.
(18.) Aircraft symbol. The fixed miniature
aircraft symbol corresponds to the longitudinal axis
3-24
(5.) ALT annunciator. Illuminates when
altitude is being held by the flight director.
(6.) HDG annunciator. Illuminates when
heading is being held by the flight director, in the
NAV ARM, BC ARM mode.
TM 55-1510-221-10
Figure 3-12. Pilot’s Attitude Director Indicator
(7.) NAV annunciator. Illuminates when
navigation is being controlled by the flight director,
in the NAV CAP, VOR APR mode.
(8.) LOC annunciator. Illuminates whenever the flight director is controlling a localizer
approach, in the NAV CAP mode.
(9.) APR annunciator. Illuminates whenever the flight director is controlling a approach, in
the NAV CAP, VOR APR mode.
(10.) GS annunciator. Illuminates whenever
the flight director is in GS CAP mode, and glide
slope has been captured.
(11.) BC annunciator. Illuminates whenever the flight director is in BC CAP mode, and has
captured the back course approach heading.
(12.) VRT annunciator. Illuminates when
vertical speed is being held by the flight director, in
the VS mode.
(13.) DH annunciator. Illuminates when
aircraft descends below selected decision height as
set on the radio altimeter indicator.
(14.) Eyelid display. Surrounds the attitude
sphere and provides positive attitude identification
by means of a blue eyelid which always shows the
relative position of the sky, and a brown eyelid
which always shows the relative position of the
ground. The eyelids maintain the proper ground-sky
relationship, regardless of sphere position.
(15.) Speed command display. The pointer
indicates relative airspeed provided by the angle-ofattack/speed command system.
(16.) Flight director command cue. Displays
computed commands to capture and maintain a
desired flight path. Always fly the symbolic miniature aircraft to the flight director cue. The cue will
bias from view should a failure occur in either the
pitch or roll channel.
(17.) Radio altitude display. Radio altitude
is digital displayed. The range capability of the display is from -20 to 2500 feet AGL. The display resolution between 200 and 2500 feet is in 10 foot increments. The display resolution below 200 feet is 5
feet. The display will be blank at altitudes over 2500
feet AGL. Dashes are displayed whenever invalid
radio altitude is being received.
(18.) DH SET control knob. Sets decision
height from 0 to 990 feet. Decision height displays
in the DH window on lower left corner of ADI. The
brightness of the digital radio altitude and decision
height display is controlled by the dimming knob
3-25
TM 55-1510-221-10
which is concentric with the DH SET knob. The
dimming knob also dims the distance and course
display on the pilot’s HSI, and the altitude alert display.
(19.) Expanded localizer. Raw localizer displacement data from the navigation receiver (HSI
display) is amplified approximately 7 1/2 times to
permit the expanded localizer pointer to be used as
a sensitive reference indicator of the aircraft’s position, with respect to the center of the localizer. It is
normally used for assessment only, since the pointer
is very sensitive and difficult to fly throughout the
entire approach. During final approach, the pointer
serves as an indicator of the Category II window.
Full scale deflection of the expanded localizer
pointer is equal to 1/4 degree of beam signal. The
expanded localizer is displayed by the localizer
pointer only when a valid localizer signal is available.
(20.) Inclinometer. Gives the pilot a conventional display of aircraft slip or skid, and is used
as an aid to coordinated maneuvers.
(21.) Rate of turn. Rate of turn is displayed
by the pointer at the bottom of the ADI. The marks
at the extreme left and right sides of the scale represent a standard rate turn.
satisfy the commands of the selected flight director
mode.
(26.) Glide slope scale and pointer. Displays
aircraft deviation from glide slope beam center only
when tuned to a ILS frequency and a valid glide
slope is present. The aircraft is below glide path if
pointer is displaced upward. The glide slope dot represents approximately 0.4 degree deviation from the
beam centerline.
3-20. COPILOT’S GYRO HORIZON INDICATOR.
a. Description. The copilot’s gyro horizon
indicator (fig. 3-13) is a flight aid which indicates
the aircraft’s attitude. The attitude given is in relationship to an artificial horizon. There are no front
panel fuses or circuit breakers provided for the copilot’s gyro horizon indicator.
b. Indicators and Functions.
(1.) Bank angle scale. Indicates aircraft
bank angle from zero to 90 degrees with marks at
10, 20, 30, 45, 60, and 90 degrees.
(2.) Bank angle index. Reference indicating zero-degree bank.
(22.) Attitude (ATT) test switch. W h e n
depressed, the sphere will show an approximate attitude change of 20 degrees of right bank at 10
degrees pitch-up. The ATT warning flag will appear.
In addition, all mode annunciator lights except DH
will illuminate.
(3.)
bank angle.
(23.) Radio altitude (RA) test switch. Pressing the RA test button causes the following displays
on the radio altitude readout: all digits display 8’s
then dashes, and then the preprogrammed test altitude as set in the radio altimeter R/T unit, until the
test button is released at which time the actual altitude is displayed. The DH display during the test
displays all 8’s with the altitude display and then
displays the current set altitude for the remainder of
the test. RA test is inhibited as a function of APR
CAP.
power.
(24.) Decision height (DH) display. The digital DH display, displays decision height range from
0 to 990 feet in 10 foot increments. The decision
height is set by the knob in the lower right corner of
the ADI.
(25.) Symbolic miniature 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. The symbolic aircraft is flown to align
the command cue to the aircraft symbol in order to
3-26
Bank angle pointer. Indicates aircraft
(4.) Horizon line. Affixed to sphere,
remains parallel to the earth’s horizon at all times.
(5.) G flag. Presence announces loss of
(6.) Sphere. Indicates orientation with
earth’s axis at all times.
(7.) Inclinometer. Assists the copilot in
making coordinated turns.
(8.) Miniature aircraft. Indicates attitude
of aircraft with respect to the earth’s horizon.
3-21. TURN AND SLIP INDICATORS.
a. Description. The pilot and copilot have
identical turn and slip indicators (fig. 3-14) protected by the circuit breaker placarded TURN &
SLIP on the overhead circuit breaker panel (fig.
2-26).
b. Controls/Indicators and Functions.
(1.) Two-minute turn marks. Fixed markers indicate two-minute turn rate when covered by
turn rate indicator.
TM 55-1510-221-10
Figure 3-13. Copilot’s Gyro Horizon Indicator (GH-14B)
Figure 3-14. Turn and Slip Indicator (329T-1)
3-27
TM 55-1510-221-10
(2.) Turn rate indicator. Deflects to indicate rate of turn.
(a.) No.1. Selects compass system
No. 1 for display.
(3.) Inclinometer. Indicates lateral acceleration (side slip) of aircraft.
(b.) No.2. Selects compass system
No. 2 for display.
3-22. GYROMAGNETIC COMPASS SYSTEMS.
(3.) GYRO SLAVE-FREE switch. Selects
system mode of operation.
a. Description. Two identical compass systems provide accurate directional information for
the aircraft at all latitudes of the earth. As a heading
reference, two modes of operation are used: directional gyro (FREE) mode, or slaved (SLAVE) mode.
In polar regions of the earth where magnetic heading
references are not reliable, the system is operated in
the FREE mode. In this mode, the system furnishes
an inertial heading reference, with latitude corrections introduced manually. In areas where magnetic
heading references are reliable, the system is operated in the SLAVE mode. In this mode, the directional gyro is slaved to the magnetic flux detector,
which supplies long-term magnetic reference to correct the apparent drift of the gyro. Magnetic heading
information from both systems is applied to various
aircraft systems through pilot and copilot COMPASS No.1 - No.2 switches. There are no circuit
breakers for the gyromagnetic compass systems. The
circuits are protected by the 2-ampere F2 and F6
fuses on the No. 1 junction box.
(a.) SLAVE. Selects slaved mode.
Compass flux valve connects to azimuth card.
(b.) FREE. Selects free mode. Flux
valve is not connected to azimuth card.
(4.) INCREASE-DECREASE switch. Provides manual fast synchronization of the system.
(a.) INCREASE. Causes gyro heading output to increase (move in clockwise direction).
(b.) DECREASE. Causes gyro heading output to decrease (move in counter-clockwise
direction).
d. Normal Operation.
(1.) Alignment procedure:
b. Vertical Gyro A vertical gyro provides lineof-sight stabilization to the weather radar and roll
and pitch information to the autopilot. A FAST
ERECT switch at the top of the pilot’s instrument
panel (figure 2-29) provides a means for fast erection of the gyros. Pressing and holding the FAST
ERECT switch will erect the gyro to within 1.0° of
pitch and roll within 60 seconds of power application, and erect to within 0.5° within 2 minutes. Normal operation of the vertical gyro system will not
require use of the fast erect switch. The circuit is
protected by the 3-ampere F22 fuse in the No. 1
junction box.
c.
(2.)
Controls and Functions.
(1.) Pilot’s COMPASS No.1-No.2 switch.
Selects desired source for magnetic heading information to display on pilot’s HSI and copilot’s RMI.
(3.)
1.
Gyro compass slave-free switch SLAVE.
2.
Gyrocompass
increase-decrease
switch - Hold switch momentarily
in the direction desired, and then
release. This will place system in
fast erect mode. The gyro will
then erect at approximately 30
degrees per minute. While in the
fast erect mode, the HEADING
flag (HSI) will be in view. When
the HEADING flag retracts from
view, the heading displayed will
be the magnetic heading.
To determine magnetic heading:
1.
Gyrocompass slave-free switch SLAVE.
2.
RMI rotating heading dial (compass card) - Read heading.
To determine directional gyro head-
ing:
(a.)
No. 1 for display.
No. 1. Selects compass system
1.
Gyrocompass slave-free switch FREE.
(b.)
No.2 for display.
No.2. Selects compass system
2.
Gyrocompass
increase-decrease
switch - Hold until the RMI compass card aligns with the magnetic
heading, then release.
3.
Read heading. The heading will
agree with the appropriate HSI.
COMPASS
No. I-No.2
(2.) Copilot’s
switch. Selects desired source for magnetic heading
information to display on copilot’s HSI and pilot’s
RMI and INS.
3-28
TM 55-1510-221-10
e. Shutdown Procedure. Both compass systems are shut down when the INVERTER No.1 or
INVERTER No.2 switch is turned off. (If either
inverter is on, both compass sets will be energized.)
maneuver to fly toward the preselected altitude. Any
of the following pitch modes may be engaged: Pitch
Hold, Airspeed Hold or Vertical Speed Hold. Upon
initiation of altitude preselect capture, the previously selected pitch mode is automatically reset.
3-23. ALTITUDE SELECT CONTROLLER
The Altitude Select Controller (fig. 3-15) provides a means for setting the desired altitude reference for the altitude alerting and altitude preselect
system.
(1.) Altitude Alert. As the aircraft reaches
a point 1000 feet from the selected altitude, a signal
is generated to light the warning light on the altimeter. This light remains on until the aircraft is 250
feet from the selected altitude. If the aircraft now
deviates by 250 feet or more from the selected altitude, the light is again energized. The light remains
on until the aircraft returns to within 250 feet or
deviates more than 1000 feet from the selected altitude.
(2.) Altitude preselect The altitude is
selected by turning the selector knob until the altitude display reads the desired value. No further
action is taken on the controller. To initiate altitude
preselect, the ALTSEL button is selected on the
flight director controller. The pilot must initiate a
3-24. RADIO ALTIMETER INDICATOR.
Description. The radio altimeter indicator
(fig. 3-16) displays radio altitude information from
2500 feet to touchdown with an expanded linear
scale under 500 feet.
b. Controls/Indicators and Functions.
(1.) DH annunciator. Light illuminates to
alert that aircraft is at or below selected DH.
(2.) Decision height bug. Manually set by
knob to establish DH.
(3.) Failure warning flag (not shown).
When visible, indicates that system information may
be unreliable.
(4.) Altitude pointer. Points to dial reading
for current radio altitude from 0 to 2500 feet.
(5.) Decision height set knob. Used to
manually set DH.
Figure 3-15. Altitude Select Controller (AL-800).
3-29
TM 55-1510-221-10
Figure 3-16. Radio Altimeter Indicator (RA-315)
(6.) TEST pushbutton. Pressed to check
indicator R/T unit and flag operation.
Operating the test button causes the flag to come
into view and altitude pointer to indicate approximately 100 feet. Release of button causes pointer to
return to existing altitude and flag to retract.
3-25. VOR/LOC NAVIGATION SYSTEM.
a. Description. The aircraft is equipped with
two VOR systems, controlled by a dual NAV 1 NAV 2 control panel located on the pedestal (fig.
2-7). Either VOR can direct input signals to the attitude director indicator. Controls are shown on figure
3-1. Each VOR system includes independent
receiver units for VOR/LOC and glideslope (GS).
Each VOR receiver provides a VOR input to a
respective RMI, HSI, and the flight director computer. Each glideslope receiver sends GS flag and
pointer deviation information to the HSI and flight
director computer. VOR/LOC indicators may be
used for navigation during manual control of the aircraft, or the autopilot may be coupled to the VOR
system, accepting VOR inputs to the autopilot computer. The pilot’s unit (VOR 1) is a navigation radio
system which receives and interprets VHF omnidirectional radio range (VOR) and localizer (LOC)
signals, glideslope signals (GS), and marker beacon
3-30
signals. It has a maximum range of 120 nautical
miles line-of-sight. The system operates in a VOR/
LOC frequency range of 108.00 to 117.95 megahertz, in a glideslope frequency range of 329.15 to
335.00 megahertz, and at a marker beacon frequency of 75 megahertz. VOR 2 is similar to VOR
1 except VOR 2 cannot receive or interpret marker
beacon signals. Each VOR system provides course
deviation and glide path data, which can be
switched either to the copilot’s HSI or to the autopilot flight computer and pilot’s HSI, or both. The
audio outputs of VOR 1 and VOR 2 systems are
supplied to the NAV control on the audio control
panels. VOR 1 bearing data is supplied to the singleneedle pointer on both Radio Magnetic Indicators.
VOR 2 bearing data is supplied to the double-needle
pointer on both Radio Magnetic Indicators. VOR 1
uses a marker beacon antenna located on the underside of the forward fuselage (fig. 2-1). VOR 1 and
VOR 2 both use the same glideslope antenna,
located inside the radome. Both VOR’s are protected by separate 2-ampere circuit breakers, located
respectively on the number 1 and number 2 avionics
bus. The circuit breakers are placarded VOR No.1
and VOR No.2 are located on the overhead circuit
breaker panel (fig. 2-26).
TM 55-1510-221-10
Figure 3-17. NAV 1 - NAV 2 Control Panel (VIR-30AGM, VIR-30AG
b. Controls/Indicators and Functions (NAV 1
Control Panel, fig. 3-17).
(2.) Frequency control. Selects operating
frequency of VOR 2 receiver.
(1.) Frequency indicator. Displays selected
frequency of VOR 1 receiver.
(3.) NAV-TEST switch. Activates test of
VOR 2 navigation systems. If the system is functioning properly the following indications will be presented:
(2.) Frequency control. Selects operating
frequency of VOR 1 receiver.
(3.) NAV-TEST pushbutton. Activates test
of VOR 1 navigation system. If the system is functioning properly, the following indications are presented:
(a.) RMI. Single needle indicates 0°.
(b.) HSI. Indicates lateral deviation
to the right and glideslope deviation down, if flag is
tuned to ILS frequency. It tuned to NAV frequency,
indicates 0° and the G/S flag is in view,
(4.) OFF/VOL control. Activates VOR 2
receiver. Permits monitoring VOR 2 audio and
adjusts volume of signals received.
(a.) RMI. Single needle indicates 0°.
(b.) HSI. Indicates lateral deviation
to the right and glideslope deviation down, if NAV
is tuned to ILS frequency. If tuned to NAV frequency, indicates 0° and G/S flag is in view.
d.
Controls and Functions, Instrument Panel.
(4.) VOL-OFF control. Activates VOR 1
receiver. Permits monitoring VOR 1 audio and
adjusts volume of signals received.
(1.) Pilots COURSE INDICATOR switch.
Selects VOR receiver to control pilot’s HSI.
c. Controls/Indicators and Functions (NAV 2
Control Panel, Fig. 3-17).
HSI.
(1.) Frequency indicator. Displays selected
frequency of VOR 2 receiver.
HSI.
(a.) VOR 1. VOR 1 controls pilot’s
(b.) VOR 2. VOR 2 controls pilot’s
3-31
TM 55-1510-221-10
(2.) Copilot's
COURSE INDICATOR
switch. Selects VOR receiver to control copilot’s
HSI.
(a.)
VOR 1. VOR 1 controls copilot’s
(b.)
VOR 2. VOR 2 controls copilot’s
HSI.
2. Pilot’s, copilot’s COURSE INDICATOR switches (instrument
panel) - Select VOR source.
3. Check glideslope flags unmasked.
(4.) VOR receiver operation:
1. VOR frequency control knob
(VOR control panel, fig. 3-17) Select desired frequency.
HSI.
(3.) NAV-A switch (audio control panel,
fig. 3-1). Applies VOR audio to respective headsets.
2. Volume control knob (VOR control panel, fig. 3-17) - Full on.
(4.) MKR BCN HI-LO (marker beacon
audio control panel, fig. 3-2). Controls sensitivity of
marker beacon receiver.
e.
VOR Operation.
(1.) Turn-on procedure:
1.
Insure that aircraft DC and AC
power is on.
2. Avionics master power switch
(overhead control panel, fig. 2-12)
- ON.
3. Frequency controls (VOR control
panel) - Set for both receivers.
4. Volume controls (VOR control
panel - Turn clockwise to activate
sets and adjust volume.
5. NAV A, then NAV B audio
switches (audio control panel, fig.
3-1) - ON. Confirm proper signal,
then OFF.
6. RMI and HSI - Confirm proper
indications.
(2.) Normal Operation.
1.
Pilot/copilot
course
indicator
switches (instrument panel) Select VOR source.
2.
To determine course to station on
pilot’s HSI: TO-FROM pointer
reads TO (up) position.
3.
To determine bearing from station on pilot’s HSI: TO-FROM
pointer reads FROM position
(down).
4.
To determine course to station on
RMI: Select VOR, verify single
needle points course to station.
(3.) Localizer (LOC) operation:
1.
3-32
VOR frequency knob (NAV
panel) - Select frequency.
3. NAV A, audio switch - ON.
Adjust audio.
(5.) Shutdown procedure: Volume control
(VOR control panel, fig. 3-17) - OFF.
3-26. MARKER BEACON RECEIVER.
a. Description. The marker beacon receiver is
located inside the No.1 VOR receiver. The marker
beacon receiver obtains power through the VOR
receiver. The marker beacon provides visual and
aural indication of the aircraft’s position over a 75
MHz marker beacon ground transmitter. Upon
entering the range of marker beacon signals, blue,
amber, or white annunciator lights will illuminate,
and corresponding aural signals will indicate aircraft
passage over the (0) outer, (M) middle, (A) inner or
airway marker beacons. Range is vertical to 50,000
feet. Volume, and sensitivity controls are located on
the marker beacon audio control panel (fig. 3-1).
b. Controls/Indicators and Functions.
(1.) Marker beacon sensitivity switch
(marker beacon audio control panel, fig. 3-2). Controls sensitivity of marker beacon receiver.
(a.) HI position. Enables high sensitivity operation of marker beacon receiver.
(b.) LO position. Enables low sensitivity operation of marker beacon receiver.
(2.) "0" indicator. Illuminates when aircraft passes over an outer marker beacon.
(3.) "M" indicator. Illuminates when aircraft passes over a middle marker beacon.
(4.) "A" indicator. Illuminates when aircraft passes over an inner or airway marker beacon.
c. Marker Beacon Operation.
1.
Marker beacon volume control (marker
beacon audio control panel, fig. 3-2) As required.
TM 55-1510-221-10
2. Marker beacon HI-LO switch (marker
beacon audio control panel, fig. 3-2) As required.
indicator
3. Marker
lights
beacon
(instrument panel, fig. 2-29) - Confirm
beacon indication.
3-27. AUTOMATIC DIRECTION FINDER (DF-203).
a. Description. The Automatic Direction
Finder (ADF) (fig. 3-18), is a radio navigation system which provides a visual indication of aircraft
bearing, relative to a selected ground radio station.
It may also be used to home to a selected station,
find aircraft position, or monitor conventional
medium frequency AM radio transmissions. The
system is designed to provide reliable reception of a
400-watt radio station at a range of 65 nautical
miles throughout a 360-degree turn of the aircraft. It
operates in a frequency range of 190 to 1750 kilohertz. Bearing indications are displayed visually on
the RMI's and aural signals are applied to the audio
control panels. The ADF system consists of a
receiver, located on the forward side of the aft cabin
bulkhead inside the pressure vessel; a control unit,
located on the pedestal extension; a non-directional
sense antenna, installed in the aircraft dorsal tin; a
directional loop antenna, located on the underside
of the fuselage; and a quadrangle error corrector,
installed on the loop antenna (to compensate for the
deflection of arriving radio signals by the wings and
fuselage of the aircraft). The system is protected by
a l-ampere ADF, a 5-ampere RADIO RELAY, and
a 35-ampere AVIONICS BUS FEEDER No.2 circuit
breaker located on the overhead circuit breaker
panel (fig. 2-26).
The only warning that the crewmember
has for an unreliable ADF signal or loss
of the ADF receiver is the loss of the
ADF audio signal in the crewmember’s
headset. This signal should be monitored
during all phases of the approach in IMC
conditions.
NOTE
Keying the HF radio set while operating
the ADF radio set will cause a momentarily unreliable ADF signal.
Figure 3-18. ADF Control Panel (DF-203)
3-33
TM 55-1510-221-10
b. Controls and Functions (ADF Control
Panel, fig. 3-18).
(I.) L-LOOP-R control. Operative only
when the function switch is in the LOOP or ADF
position. Center position removes rotation signals
from the loop antenna and the ADF pointer on the
RMl's. L (left) or R (right) of center applies rotation
signals to loop antenna and ADF pointer on RMI's
for 360-degree rotation left or right. Center position
holds antenna position.
NOTE
Range and voice filtering are cancelled
when both FILTER R and FILTER V are
ON. Range and voice audio will be heard.
e. ADF Normal Operation.
(1.)
To operate the set as ADF.
1.
Mode selector - ADF.
2.
BFO-OFF switch - BFO.
(2.) BFO-OFF. At BFO (on) setting, permits line tuning with Beat Frequency Oscillator
(BFO). Also provides audio tone when receiving
unmodulated CW. OFF position turns BFO off.
3.
Range switch - Select frequency
range.
4.
(3.) Tuning meter. Indicates relative
strength of received signals.
Audio control panel (fig. 3-l)
NAV B switch - On and adjusted.
5.
Gain control - As required.
6.
TUNE control - Rotate for maximum reading on tuning meter
and zero BFO beat.
7.
BFO-OFF switch - OFF.
8.
Double needle switch (RMI, fig.
3-9) - ADF.
9.
Double needle pointer (RMI, fig.
3-9) - Read course to station.
(4.) TUNE control. Selects operating frequency.
(5.) Range switch. Selects operating frequency band.
(6.) FREQUENCY indicator. Indicates
selected frequency.
(7.) Mode selector. S e l e c t s o p e r a t i n g
mode.
(a.) OFF. Turns set off.
(2.) To operate set for sense antenna
receiving only.
(b.) ADF. Permits automatic direction finding or homing operation.
1.
Mode selector (ADF control
panel, fig. 3-18 - ANT.
2.
(c.) ANT. Permits reception using
sense antenna.
Range switch - Select operating
range.
3.
(d.) LOOP. Permits audio-null homing and manual direction finding operations.
Tune control - Rotate for maximum reading on tuning meter.
4.
Gain control - As required.
(8.) GAIN control. Adjusts volume of
received signal.
(3.)
To operate set for aural-null direction
finding.
1.
Mode selector (ADF control
panel, fig. 3-18) - ANT.
2.
BFO-OFF switch - BFO.
3.
Range switch - Select operating
range.
d. Controls and Functions (Marker Beacon
Audio Panel, fig. 3-2).
4.
Tune control - Tune desired station.
(1.) FILTER V-OFF switch.
Selects
whether voice filter will be used with ADF audio.
5.
Gain control - Adjust for minimum audio output.
(2.) FILTER R-OFF switch. S e l e c t s
whether range filter will be used with ADF audio.
6.
Double needle switches (RMI, fig.
3-9) - As required.
c. Controls and Functions (Audio Control
Panel, jig. 3-1).
(1.) NAV-B switch. O N position applies
ADF audio to respective headset.
3-34
TM 55-1510-221-10
7.
BFO-OFF switch - OFF.
8.
Mode selector switch - LOOP.
9.
Loop switch - L or R. Turn left or
right until a null is reached (minimum sound in headsets).
10.
Double needle on RMI (fig. 3-9) Read course to station,
NOTE
The true null and direction to the radio
station may be indicated by either end of
the single needle. This ambiguity must be
solved to determine proper direction to
the station.
(4.) Shutdown procedure: Mode selector
switch (ADF control panel, fig. 3-18) - OFF.
3-28. TACAN SYSTEMS.
a. Description. Two Tactical Air Navigation
(TACAN) systems are provided. One is dedicated to
the INS and is used only for position updating, and
provides only DME information to the INS.The
other is used in conjunction with other avionics systems, including the flight director system and the
autopilot. For normal navigation TACAN is a radio
navigation system which provides aircraft distance
and bearing information relative to a TACAN
ground station. Both systems operate in the L band
frequency range of 962 to 1213 MHz. Their range,
though limited to line-of-sight, is designed to provide reliable reception of a TACAN ground station
at a distance of 170 nautical miles at an aircraft altitude of 20,000 feet. The normal time required for
the systems to lock on to a selected ground station
signal is three seconds. Both systems are protected
by a 2-ampere circuit breaker, placarded TACAN,
located on the overhead circuit breaker panel (fig.
2-26).
b. TACAN System (Non-INS dedicated). The
TACAN system (Non-INS dedicated) consists of a
range unit (which includes the system transmitter)
and a bearing unit, both located in the right nose
avionics compartment; a distance indicator (fig.
3-19) located on the instrument panel; a control
unit (fig. 3-20) located in the pedestal extension;
and an antenna, located on the top of the fuselage.
The TACAN system (non INS dedicated) operates
in conjunction with TACAN and VORTAC ground
stations to provide distance, ground speed, time-tostation, and bearing-to- station data. It operates in
the L band frequency range on one of 252 preselected frequencies, 126 X mode and 126 Y mode channels. Course deviation to TACAN stations are dis-
played on the HSI’s. Distance, time-to-station, and
ground speed are displayed on the TACAN digital
display (fig. 3-19). The ground speed and time-tostation are accurate only if the aircraft is flying
directly toward the ground station at a sufficient distance that the slant range and ground range are
nearly equal. The (Non-INS dedicated) TACAN system may be connected to and used with the autopilot system. When employed as the primary means of
navigation, aircraft flight may be controlled manually or by the autopilot. Indications of aircraft heading and bearing to ground stations are displayed on
the horizontal situation indicators. Magnetic bearing
to a station is displayed by the RMI and pilot’s HSI
bearing pointer. TACAN distance, ground speed,
and time-to-station are all displayed on the TACAN
indicator located on the copilot’s instrument panel
(fig. 2-29). TACAN distance is displayed on the
HSI’s. The TACAN control panel (fig. 3-20) enables
selection of the TACAN frequency (channel) to be
used, and provides for self-test of TACAN circuits.
At the present time, most TACAN and VORTAC
stations are operated in the X mode. When Y mode
stations are operational, air navigation charts will
designate the Y mode stations. The small (outer)
control provides system power ON-OFF and station
identifier tone, volume and control.
c. INS TACAN System. The INS TACAN system is coupled directly to INS circuits. It is dedicated only to updating the INS, is activated when
the INS is operational, and is controlled only by the
INS. The INS TACAN consists of a range unit and
a distance indicator, both located on the INS equipment rack and both identical to counterparts in the
(non-INS dedicated) TACAN and antenna, located
on the underside of the fuselage (fig. 2-1). No controls are required or provided for the INS TACAN
system.
d. Controls/Indicators and Functions.
(1.) TACAN control panel.
1.
TEST SWITCH. Activates system
self-test.
2.
CHANNEL INDICATOR. Displays selected TACAN channel.
3.
X-Y SWITCH. Selects X or Y
mode for TACAN channels.
4.
CONTROL.
VOL
TACAN volume.
5.
OUTER CHANNEL SELECTOR
KNOB. For manual selection of
tens and hundreds part of channel
number.
6.
INNER CHANNEL SELECTOR
KNOB. For manual selection of
units part of channel number.
Adjusts
3-35
TM 55-1510-221-10
Figure 3-19. TACAN Control Panel (AN/ARN-136)
Figure 3-20. TACAN Distance Indicator (SANS-706)
3-36
TM 55-1510-221-10
7. ON-OFF SWITCH. Activates or
deactivates system.
(2.) TACAN distance indicator.
3.
Course indicator selector switch (instrument panel, fig. 2-29) - TACAN.
4.
TACAN X-Y switch - As required.
NM INDICATOR. Displays slant
range distance in nautical miles
from aircraft to selected TACAN
ground station.
5.
TACAN channel selector knobs - Select
desired channel.
6.
Wait 5 seconds for signal acquisition
and lock-on.
INDICATOR.
KT
ground speed in knots.
7.
If bearing lock-on is not obtained, perform an inflight self-test to insure correct operation of the system. Anytime
a course indicator NAV or VOR LOC
flag is in view, bearing, course deviation, and TO-FROM information may
be inaccurate and should be disregarded.
8.
Insure that audio station identification
signal is correct for the ground station
selected.
9.
RMI single-needle pointer and pilot
HSI bearing pointer -Read bearing to
station.
4. Self-test procedure: Course knob
on pilot’s HSI - Set to 180°, press
and hold TEST switch.
10.
HSI course control knob - Set desired
course.
5.
11.
HSI course deviation bar - Read deviation from selected course. Course
arrow will show wind correction angle
when the course deviation bar is centered and the aircraft is tracking the
selected course.
12.
TACAN indicator and pilot’s HSI Read distance to station.
13.
To determine course TO or course
FROM a TACAN station, rotate course
knob (pilot’s HSI) until course deviation bar is centered and the TO-FROM
pointer reads TO or FROM.
14.
To use TACAN during pilot-controlled
flight, control aircraft by manual controls, responding to information displayed on the flight director, RMI,
HSI, TACAN, and other instruments.
15.
To use TACAN with the autopilot,
engage autopilot and monitor autopilot
performance on flight director, RMI,
HSI and TACAN indicators. Verify
adherence to preset heading and
course, and confirm the execution of
displayed steering commands.
1.
2.
Displays
3. MIN INDICATOR. Displays
time to TACAN station in minutes.
e.
(Non-INS dedicated) TACAN Operation.
(I.) Turn-On procedure:
1.
Power switch (TACAN control
panel, fig. 3-20) - ON.
2.
Volume control - As required.
3. Course indicator switches (instrument panel, fig. 2-29) -TACAN.
Pilot’s HSI course deviation bar Centered, with course knob set to
180 ±2 degrees, and TO-FROM
indicator indicating TO.
6. RMI bearing pointers (fig. 3-9)
and pilot’s HSI bearing pointer Point to 180 degrees.
7.
HSI course knob - Increase the
selected course. The course deviation bar on a 180 ± 2 degrees TO
indication and the bearing pointers on each RMI indicator read
180 degrees. Using the course
knob, increase the selected course,
the course deviation bar will
move left. Decrease the selected
course, the course bar will center,
then move to the right of center.
Full scale deflection will be 10
± 1°.
f. Normal Operating Procedure:
1.
RMI single needle switch (fig. 3-9) VOR-TACAN.
2.
RMI selector switch (instrument panel,
fig. 2-29) -TACAN.
3-37
TM 55-1510-221-10
NOTE
The TACAN ground speed reading will be
accurate only when the aircraft is on a
course directly to or from the TACAN
station.
g. Shutdown procedure: TACAN power switch
(fig. 3-20) - OFF.
3-29. AUTOMATIC FLIGHT CONTROL SYSTEM.
The RC-12H aircraft is certified with
wingtip pods installed. Should the pods
be removed, the autopilot system must be
replaced with a standard C-12D autopilot.
Effected wiring must also be changed.
a. Description. The Automatic Flight Control
System is a completely integrated autopilot/flight
director/air data system which has a full complement of horizontal and vertical flight guidance
modes. These include all radio guidance modes, and
air data oriented vertical modes.
When engaged and coupled to the flight director
(FD) commands, the system will control the aircraft
using the same commands displayed on the attitude
director indicator. When engaged and uncoupled
from the flight director commands, manual pitch
and roll commands may be inserted using the pitch
wheel and turn knob.
When the autopilot is coupled, the flight director
instruments act as a means to monitor the performance of the autopilot. When the autopilot is not
engaged, the same modes of operation are available
for flight director only. The pilot maneuvers the aircraft to satisfy the Flight Director commands, as
does the autopilot when it is engaged.
b. Air Data Computer. A digital air data computer located in the forward avionics compartment
provides the altitude information for the pilot’s
altimeter indicator, altitude alerter, and transponder. The computer also provides altitude and airspeed hold function data to the flight control computers. The air data computer receives 28 VDC
power through, and is protected by, a 2 amp circuit
breaker placarded AIR DATA - ENCDR located in
the AVIONICS section of the overhead circuit
breaker panel. All air data computer functions are
automatic in nature and require no flight crew
action.
3-38
c. Flight Director Mode Selector. The Flight
Director Mode Selector (fig.3-22), located on the
pedestal, provides for selection of all modes (except
go-around which is initiated by a remote switch
located on the left power lever for the flight director.
The top row of split light annunciated pushbuttons
contains the lateral modes and the bottom row contains the vertical modes. The mode buttons will illuminate when manually selected, or automatically
selected through other modes.
The split light pushbutton annunciators, illuminate amber for armed conditions and green for captured. When more than one lateral or vertical mode
is selected, the flight director system automatically
arms and captures the submode. Mode annunciations are also presented on remote annunciator
blocks, located above the pilot’s Attitude Director
Indicator (ADI), and on the pilot’s ADI. Operating
modes and annunciation events of the Flight Director system are detailed in figure 3-21.
d. Controls/Indicators and Functions (FD
Mode Selector, jig. 3-22):
(1.) Heading Mode Switch (HDG).
Engages heading mode. Commands aircraft to
acquire the heading indicated by a heading marker
on the pilot’s HSI.
(2.) Navigation Mode Switch (NAV).
Engages the navigation mode selected.
(3.) VOR Approach Mode Switch (APR).
Engages approach mode. Commands aircraft to
intercept and track ILS inbound course.
(4.) Back Course Mode Switch (BC).
Engages backcourse mode. Commands aircraft to
intercept back course ILS.
(5.)
VNAV Mode Switch. Not used.
(6.) Standby Mode Switch (SBY). Engages
standby mode.
(7.) Indicated Airspeed Hold Mode Switch
(IAS). Engages indicated airspeed hold mode.
(8.) Vertical Speed Hold Mode Switch
(VS). Engages vertical speed hold mode.
(9.) Altitude Preselect Mode Switch (ALTSEL). Engages altitude preselect mode.
(10.) Altitude Hold Mode Switch (ALT).
Engages altitude hold mode. Commands aircraft to
maintain pressure altitude.
e.
Autopilot Modes of Operation.
(1.) Heading Select Mode (HDG). In the
HDG mode the flight director computer provides
TM 55-1510-221-10
fig 3-12
Figure 3-21. Flight Director Modes and Annunciators
3-39
TM 55-1510-221-10
nate. At capture, a command is generated to capture
and track the VOR beam. The course error signal is
gain programmed as a function of airspeed. Crosswind washout is included which maintains the aircraft on beam center in the presence of crosswind.
The intercept angle is used in determining the capture point to ensure smooth and comfortable performance during bracketing.
When passing over the station, an overstation
sensor detects station passage removing the VOR
deviation signal from the command until it is no
longer erratic. While over the station, course
changes may be made by selecting a new course on
the HSI.
If the NAV receiver is not valid prior to the capture point, the lateral beam sensor will not trip and
the system will remain in the HDG mode. After capture, if the NAV receiver, compass data or vertical
gyro go invalid, the ADI command cue will bias out
of view. Also, the NAV CAP annunciators will
extinguish if the NAV receiver becomes invalid.
Figure 3-22. Flight Director Mode Selector (MS500)
inputs to the command cue to command a turn to
the heading indicated by the heading bug on the
HSI. The heading select signal is gain programmed
as a function of airspeed. When HDG is selected, it
overrides the NAV, BC APR and VOR APR modes.
In the event of a loss of valid signal from the gyro
or compass, the command cue on the AD1 is biased
out of view. automatically select
(2.) Navigation Mode (NAV). The Navigation Mode represents a family of modes for various
navigation systems including VOR, Localizer,
TACAN and INS.
(a.) VOR Mode. The VOR Mode is
selected by selecting either VOR 1 or VOR 2 on the
Course Indicator select switch on the pilot’s instrument panel, and then depressing the NAV button on
mode selector with the navigation receiver tuned to
a VOR frequency. (A VOR indicator, NV1 or NV2
on the pilot’s HSI, will illuminate. VOR NAV information will display on the pilot’s HSI and RMI).
Prior to VOR capture, the command cue receives a
heading select command as described above and the
HDG mode switch is illuminated along with the
NAV ARM annunciators. Upon VOR capture the
system automatically: switches to the VOR mode;
HDG and NAV ARM annunciators extinguish;
NAV capture (NAV CAP) annunciators will illumi-
3-40
(3.) Localizer Mode. The Localizer Mode
is selected by depressing the NAV button on the
mode selector with the navigation receiver tuned to
a LOC frequency. Mode selection and annunciation
in the LOC mode is similar to the VOR mode. The
localizer deviation signal is gain programmed as a
function of radio altitude, time and airspeed. If the
radio altimeter is invalid, gain programming is a
function of glide slope capture, time and airspeed.
Other valid logic is the same as the VOR mode.
(4.) VOR Approach Mode. T h e V O R
Approach Mode is selected by depressing the APR
button on the mode selector with the navigation
receiver tuned to a VOR frequency and less than 10
DME miles from the station. The mode operates
identically to the VOR mode with the gains optimized for a VOR approach.
(5.) Back Course Mode. The back course
mode is selected by pressing the BC button on the
mode selector. Back course operates the same as the
LOC mode with the deviation and course signals
reversed to make a back course approach on the
localizer. When BC is selected, and outside the lateral beam sensor trip point, BC ARM and HDG
annunciators will illuminate. At the capture point,
BC CAP will be annunciated with BC ARM and
HDG annunciators extinguished. When BC is
selected, the glideslope circuits are locked out.
(6.) Localizer Approach Mode (APR). The
approach mode is used to make an ILS approach.
Pressing the APR button with a ILS frequency
tuned, arms both the NAV and APR modes to capture the localizer and glide slope respectively. No
alternate NAV source can be selected. Operating
TM 55-1510-221-10
LOC mode is the same as described above except, if
the radio altimeter is invalid in APR mode, gain
programming is a function of glide slope capture,
time, and airspeed.
With the APR mode armed, the pitch axis can
be in any one of the other pitch modes except goaround. When reaching the vertical beam sensor trip
point, the system automatically switches to the glide
slope mode. The pitch mode and APR ARM annunciators extinguish and APR CAP annunciator illuminates on the controller. At capture, a command is
generated to make a gradual interception of the glide
slope beam. Capture can be made from above or
below the beam. The glide slope gain is programmed
as a function of radio altitude, time and airspeed.
The APR CAP annunciator on the Mode Selector
will extinguish if the GS receiver becomes invalid
after capture.
Glide slope capture is interlocked so that the
localizer must be captured prior to glide slope capture. If the glide slope receiver is not valid prior to
capture, the vertical beam sensor will not trip and
the system will remain in the pitch mode. After capture, if the NAV receiver, GS receiver, compass data
or vertical gyro becomes invalid, ADI command cue
will bias out of view. If the radio altimeter is not
valid, the glide slope gain programming will be a
function of GS capture, time, airspeed, and the middle marker.
(7.) Pitch Hold Mode. Whenever a roll
mode is selected without a pitch mode, the ADI
command cue will display a pitch attitude hold command. The pitch attitude can be changed by pressing
the CWS button on the control wheel and maneuvering the aircraft. The command cue will be synchronized to zero while the button is depressed.
Upon release of the button, the pitch command will
be such as to maintain the new pitch attitude. In the
pitch hold mode, the ADI command cue will be
biased out of view if the VG is not valid.
(8.) TACAN Mode. The TACAN mode is
selected by selecting “TACAN“ on the course indicator selector switch, located on the pilot’s instrument
panel. A TACAN annunciator, placarded TAC,
located on the pilot’s HSI, will illuminate. TACAN
navigation information will display on the pilot’s
HSI and RMI.
NOTE
The TACAN receiver must be tuned to a
valid TACAN frequency. TACAN functions are identical to VOR using TACAN
information rather then VOR signals. The
ARM/CAP annunciation is the same as in
VOR mode.
(9.) Altitude Hold Mode (ALT). The Altitude Hold Mode is selected by depressing the ALT
button on the mode selector. When ALT mode is
selected, it overrides the APR CAP, GA, IAS, VS,
and ALTSEL CAP modes. In the ALT mode the
pitch command is proportional to the altitude error
provided by the air data computer. The altitude
error signal is gain programmed as a function of airspeed. Depressing and holding the CWS button
allows the pilot to maneuver the aircraft to a new
Altitude Hold reference without disengaging the
mode. Once engaged in the Altitude Hold Mode, the
mode will be reset if the air data computer is not
valid and the ADI command cue will bias out of
view if the VG is not valid.
NOTE
If the Baro setting on the altimeter is
changed, a command is generated to fly
the aircraft back to the original altitude
reference.
(10.) Indicated Airspeed Hold Mode (IAS).
The Indicated Airspeed Hold Mode is selected by
depressing the IAS button on the mode selector.
When IAS is selected, it overrides the APR CAP,
GA, ALT, VS, and ALTSEL CAP modes. In the IAS
mode the pitch command is proportional to airspeed
error provided by the air data computer. Depressing
and holding the CWS button allows the pilot to
maneuver the aircraft to a new Airspeed Hold reference without disengaging the mode. Once engaged in
the IAS mode, the mode will be reset if the air data
computer is not valid. The ADI command cue will
bias out of view if the VG is not valid.
(11.) Vertical Speed Hold Mode (VS). The
Vertical Speed Hold Mode is selected by depressing
the VS button on the mode selector. When VS is
selected, it overrides the APR CAP, GA, ALT, ALTSEL CAP, and IAS modes. In the VS mode, the
pitch command is proportional to VS error provided
by the air data computer. Depressing and holding
the CWS button allows the pilot to maneuver the
aircraft to a new Vertical Speed Hold reference without disengaging the mode. Once engaged in the VS
mode, the mode will be reset if the air data computer is not valid. The ADI command cue will bias
out of view if the VG is not valid.
(12.) Altitude Preselect Mode (ALTSEL).
The Altitude Preselect Mode is selected by pressing
the ALTSEL button on the mode selector. The
desired altitude is selected on the altitude preselect
controller. Pitch hold, VS or IAS may be selected as
a mode to fly to the selected altitude. When outside
the altitude bracket trip point, the ALTSEL ARM
annunciator along with the selected pitch mode is
3-41
TM 55-1510-221-10
illuminated on the mode selector. When reaching
the bracket altitude, the system automatically
switches to the ALTSEL CAP mode and the previously selected pitch mode is cancelled. When the
altitude is reached, the ALTSEL CAP mode is automatically cancelled and the Flight Director switches
to the ALT hold mode. If the air data computer is
not valid, the altitude preselect mode cannot be
selected. The ADI command cue will bias out of
view if the VG is not valid.
a 7-degree visual pitch up attitude command. Selecting GA disconnects the autopilot. However, the yaw
damper remains on.
(13.) Standby Mode (SBY). The Standby
Mode is selected by depressing the SBY button on
the mode selector. This resets all the other flight
director modes and biases the ADI command cue
from view. While depressed, SBY acts as a lamp test
causing all mode annunciators to illuminate and the
flight director warning flag on the ADI to come in
view. When the button is released, the mode annunciator lights extinguish and the flight director warning flag retracts from view.
(1.) Description. The autopilot controller
(fig. 3-23), provides the means of engaging the
autopilot and yaw damper as well as manually controlling the autopilot through the turn knob and
pitch wheel. The autopilot system limits are shown
in Table 3-1.
Once go-around is selected any roll mode can be
selected. The wings level roll command will cancel.
The go-around mode is cancelled by selecting
another pitch mode, or CWS.
f. Autopilot Controller.
(2.) Controls/Indicators and Functions
(Autopilot Controller, fig. 3-23):
(14.) Go-Around Mode. The Go-Around
Mode is selected by depressing the remote goaround switch. When selected all other modes are
reset, and the remote go-around (GA) and yaw damp
(YD ENG) annunciators will be illuminated. The
ADI command cue receives a wings level command
(zero command when roll is zero). The command
cue also receives the go-around command which is
Figure 3-23. Autopilot Con troller
3-42
1.
PITCH WHEEL. Movement of
the pitch wheel will cancel only
ALT HOLD, and ALTSEL CAP.
With vertical modes of VS or IAS
selected on the mode selector,
rotation of the pitch wheel will
change the respective displayed
vertical mode reference. VS or
IAS modes may be cancelled by
pressing the mode button on the
TM 55-1510-221-10
mode selector. If VS or IAS are
not selected, the pitch wheel
works as described above. The
pitch wheel is always disabled
during a coupled glide slope.
2.
3.
4.
5.
6.
7.
BANK LIMIT PUSHBUTTON/
ANNUNCIATOR
SWITCH.
Selection of the Bank Limit mode
on the autopilot controller provides a lower maximum bank
angle while in the Heading Select
mode. LOW will illuminate on
the Bank Limit switch. The lower
bank limit is inhibited and LOW
is extinguished during NAV mode
captures. If Heading Select is
again engaged, Bank Limit will
again be illuminated. Pressing
Bank Limit when illuminated will
return autopilot to normal bank
limits.
SOFT RIDE PUSHBUTTON/
ANNUNCIATOR SWITCH. Soft
ride reduces autopilot gains while
still maintaining stability in rough
air. This mode may be used with
any Flight Director mode
selected.
TURN KNOB. Rotation of the
turn knob out of detent results in
a roll command. The roll angle is
proportional to and in the direction of the turn knob rotation. the
turn knob must be in detent (center position) before the autopilot
can be engaged. Rotation of the
turn knob cancels any other previously selected lateral mode.
YAW DAMP SWITCH. When
the autopilot is not engaged, the
yaw damper may be utilized by
depressing the YD ENGAGE
pushbutton.
AP ENGAGE PUSHBUTTON/
ANNUNCIATOR SWITCH. The
AP ENGAGE switch is used to
engage the autopilot. Engaging the
autopilot automatically engages
the yaw damper. The autopilot
may be engaged with the airplane
in any reasonable attitude.
ELEV TRIM ANNUNCIATORS.
The elevator trim annunciator
indicates UP or DN when a
sustained signal is being applied
to the elevator servo. The
not be
annunciator
should
illuminated when engaging the
autopilot.
(3.) Autopilot Disengagement. The autopilot is normally disengaged by momentarily depressing the control wheel AP DISC switch. The autopilot
may however be disengaged by any of the following:
1.
Actuation of the control wheel AP
DISC button. Disengagement is
confirmed by 5 flashes of the AP
ENG annunciator.
2.
Pressing the respective vertical
gyro FAST ERECT button.
3.
Actuation of respective compass
INCREASE-DECREASE switch.
4.
Selection of Go-Around mode.
Disengagement is confirmed by
the AP ENG annunciator flashing
5 times and illumination of the
GA and YD ENG annunciators.
5.
Pulling the autopilot CONTROL
& AFCS DIRECT circuit breaker.
6.
Pressing
the
autopilot AP
ENGAGE pushbutton.
7.
Actuation of the manual electric
trim.
8.
Any of the following malfunctions
will cause the autopilot to automatically disengage:
a.
Vertical gyro failure.
b. Directional gyro failure.
c.
Autopilot power or circuit
failure.
d. Torque limiter failure.
NOTE
Disengaging under any of the previous
four conditions will illuminate the AP
DISC annunciator and the MASTER
WARNING light. Pressing the control
wheel AP DISC switch will extinguish the
AP DISC annunciator.
(4.) Pitch Sync & Control Wheel Steering
(CWS). The CWS push button located on the control
wheel (fig. 2-17) allows the pilot to manually change
aircraft attitude, altitude, vertical speed and/or airspeed without disengaging the autopilot. After completing the manual maneuver, the CWS pushbutton
is released, and the autopilot will automatically
3-43
TM 55-1510-221-10
Table 3-1, Autopilot System Limits
(Sheet 1 of 2)
MODE
CONTROL OR SENSOR
Engage Limit
Unlimited
Engage Limit
Roil Up to ±90 deg
Pitch Up to ±30 deg
Turn Knob
Roil Angie Limit
Roil Rate Limit
±30 deg
±5 deg/sec
Pitch Wheel
Pitch Angle Limit
±15 deg Pitch
Heading Hold
Roil Angie Limit
Less than 6 deg and no
roll mode selected
Heading
Select
Heading SEL
Knob on HSI or
Remote slew knob on
console
Roil Angle Limit
Roil Rate Limit
±25 deg
±3.5 deg/sec
VOR
Course Knob,
NAV Receiver
and TACAN
Receiver
Yaw
Damper
Yaw Control
Autopilot
Engage
Basic
Autopilot
Capture
Beam Angie
intercept
(HDG SEL)
Roil Angie Limit
Course Cut Limit
at Capture
Capture Point
ON Course
Roil Angie Limit
Crosswind
Correction
Over Station
Course Change
Roll Angie Limit
LOC or
APR or
BC
Course Knob
and
NAV Receiver
LOC Capture
Beam Intercept
Roil Angle Limit
Roil Rate Limit
Capture Point
NAV On-Course
Roll Angle Limit
Crosswind
Correction Limit
Gain Programming
3-44
VALUE
PARAMETER
up to ±90 deg
±25 deg
±45 deg Course
Function of beam,
beam rate, course error.
Max trip point is 175
micro-amps. Min trip point
is 30 microamps
±13 deg of Roil
Up to ±45 deg
Course Error
UP to ±90 deg
±17 deg
Up to ±90 deg
±25 deg
±5 deg/sec
Function of Beam, Beam
Rate and Course Error.
Max Trip Point is 175
microamps. Min Trip point
is 60 micoramps.
±17 deg of roll
±30 deg of course
error
f (time and TAS) starts at
1200 ft radio altitude, gain
reduction = 1 to 5
TM 55-1510-221-10
Table 3-1, Autopilot System Limits
(Sheet 2 of 2)
MODE
LOC or
APR or
BC (cont)
CONTROL OR SENSOR
GS Receiver and
Air Data
Computer
VALUE
PARAMETER
Glideslops Capture
Beam Capture
Pitch Command
Limit
Glideslope Damping
Pitch Rate Limit
Gain Programming
Function of beam
and beam rate. Trip
point is 30 microamps
±10 deg
Vertical velocity
f (TAS)
f (time and TAS)
starts at 1200 ft radio
altitude, gain reduction =
1 to .33
f (Radio Alt) starts at 250
ft. gain reduction = .33 at
250 ft to 0 at 0 ft.
GA
Control Switch
on Throttles
Fixed Pitch-Up
Command, Wings Level
7 deg Pitch Up
Pitch
Sync
CWS Switch on
Wheel
Pitch Attitude
Command
±20 deg max
ALT Hold
Air Data
Computer
ALT Hold Engage Range
0 to 50,000 ft
ALT Hold Engage Error
Pitch Limit
Pitch Rate Limit
±20 ft
±20 deg
f (TAS)
VERT Speed Engage
Range
VERT Speed Hold Engage
Error
Pitch Limit
Pitch Rate Limit
0 to ±6,000 ft/min.
VS Hold
IAS Hold
Air Data Computer
Air Data
Computer
ALT Preselect Air Data Computer
±30 ft/min
±20 deg
f (TAS)
IAS Engage Range
IAS Hold Engage Error
Pitch Limit
Pitch Rate Limit
80 to 450 knots
±5 knots
±20 deg
f (TAS)
Preselect Capture Range
Maximum Vertical
Speed for Capture
Maximum Gravitational
Force During Capture
Maneuver
Pitch Limit
Pitch Rate Limit
0 to 50,000 ft
±4,000 ft/min
±20g
±20 deg
f (TAS)
3-45
TM 55-1510-221-10
resynchronize to the vertical mode. Example: with
IAS mode selected, the pilot may depress the CWS
pushbutton and manually change airspeed. Once
trimmed at the new airspeed the CWS pushbutton is
released, and the autopilot will hold this airspeed. If
a large pitch attitude change is made, the pilot
should trim the aircraft normally before releasing
the CWS button.
NOTE
Either pilot’s CWS button will permit
changing of the autopilot regardless of
which pilot has control of the autopilot.
However, use of the CWS will cancel the
other pilot’s flight director GA mode.
3-30. INERTIAL NAVIGATION SYSTEM.
a. Description. The Inertial Navigation System (INS) is a self-contained navigation and attitude
reference system. It is aided by (but not dependent
upon) data obtained from its own TACAN system,
the GPS, the aircraft encoding altimeter and the
gyromagnetic compass system. The position and
attitude information computed by the INS is supplied to the automatic flight control system, weather
radar system, horizontal situation indicator, and
radio magnetic indicators. In conjunction with other
aircraft equipment, the INS permits operation under
Instrument meteorological Conditions (IMC). The
INS provides a visual display of present position
data in Universal Transverse Mercator (UTM) coordinates or conventional geographic (latitudelongitude) coordinates during all phases of flight.
When approaching the point selected for a leg
switch, an ALERT light will illuminate informing
the pilot of an imminent automatic leg switch or the
need to manually insert course change data. The
INS may be manually updated for precise aircraft
present position accuracy by flying over a reference
point of known coordinates. The INS may be
updated automatically by the TACAN system or the
GPS. Altitude information is automatically inserted
into the INS computer by an encoding altimeter
whenever the INS is operational.
The Mode Selector Unit (MSU) (fig. 3-24) controls system activation and selects operating modes.
The Control Display Unit (CDU) (fig. 3-25) provides controls and indicators for entering data into
the INS and displaying navigation and system status
information.
The INS system is protected by the 10-ampere
INS AC POWER and the 5-ampere INS HTR AC
POWER circuit breakers on the mission AC/DC
power cabinet, by the 5-ampere INS CONTROL circuit breaker on the overhead circuit breaker panel
3-46
and by the 20-ampere circuit breaker on the front of
the INS battery unit.
6. Controls/Indicators and Functions (INS
Mode Selector Unit, fig. 3-24).
(1.) READY NAV lamp. Illuminates to
indicate INS high accuracy alignment has been
attained. If attained during ALIGN mode, light
remains illuminated until NAV mode is selected.
Light illuminates momentarily during alignment, if
alignment is accomplished while in NAV mode.
(2.) BAT lamp. Illuminates to indicate
INS shutdown due to low battery unit voltage.
(3.) Mode select knob. Controls INS activation and selects operating modes.
(a.) OFF. Deactivates INS.
(b.) STBY. Moving to STBY from
OFF mode: Starts fast warmup of system to operating conditions; activates computer so information
may be inserted; all INS controlled warning flags
will indicate warning. Moving to STBY from any
other mode: INS operates as if in attitude reference
mode.
(c.) ALIGN (ground use only,
parked). Moving to ALIGN from OFF mode: Leveling starts after fast warmup heaters are off. Moving
to ALIGN from STBY: Alignment starts if fast warmup heaters are off. Moving to ALIGN from NAV
mode: INS is not downmoded, but will allow automatic shutdown if overtemperature is detected.
(d.) NAV. Activates normal navigation mode after automatic alignment is completed;
must be selected before moving aircraft. Moving to
NAV from STBY mode causes INS to automatically
sequence through STBY and ALIGN to NAV mode,
if present position is inserted and aircraft is parked.
NAV mode is used to shorten time in STBY and to
bypass battery test, if stored heading is valid.
(e.) ATT. Activates attitude reference mode. Used to provide only INS attitude signals. Shuts down computer and CDU leaving only
BAT and WARN lights operative. Once selected,
INS alignment is lost.
c. Controls/Indicators and Functions, (INS
Control Display Unit, fig. 3-25).
(1.) HOLD key. Used with other CDU
controls to stop present position display from changing, in order to update position and to display
recorded malfunction codes. Lights when pressed
first time; goes out when pressed second time or
when inserted data is accepted by computer. When
TM 55-1510-221-10
Figure 3-24. INS Mode Selector Unit
pressed second time, allows displays to resume showing
changing current present position.
(2.) ROLL LIM key. Allows selection of Roll
Limited steering mode. Press to select mode, key lights.
Roll steering output is limited to 10 degrees. Press
second time to exit mode, key light extinguishes. Roll
steering output returns to normal limit of 25 degrees.
(3.) Data display, left and right. Composed
of lights which illuminate to display numbers, decimal
points, degree symbols, left and right directions, and
latitude or longitude directions.
(4.) INSERT / ADVANCE / HI PREC key.
Allows insertion of loaded data into computer.
Enters displayed data into INS. When pressed
before pressing any numerical key, alternates display
of normal and high precision data.
(5.) ALERT lamp. Illuminates amber to
alert pilot 1.3 minutes before impending automatic
course leg change. Extinguishes when switched to
new leg, if AUTO-MAN switch is set to AUTO.
Flashes on and off when passing waypoint, if
AUTO-MAN switch is set to MAN. Light will
extinguish if AUTO is selected or if a course change is
inserted.
(6.) BAT lamp. Illuminates amber to indicate
loss of 115 VAC power and INS operation on INS
battery power.
(7.) WARN /amp. Lights red to alert pilot
INS self-test circuits have detected a system fault.
Illumination may be caused by continuous or
intermittent condition. Intermittent conditions light
WARN light until reset by TEST switch. If continuous
condition does not degrade attitude operation, light goes
out when mode selector is set to ATT.
(8.) Keyboard. Consists of 10 keys for
entering load data into data and FROM-TO displays.
“N”, “S”, "E”. and “W” (on keys 2,8,6 and 4) indicate
direction of latitude and longitude. TAC and DISP (on
keys “7” and “9” ) enable loading and display of
TACAN station data. MV/P and DISP (on keys “3” and
“9”) are associated with loading and display of magnetic
variation and magnetic heading. Pattern steering
parameters and DISP (on keys “5” and “9”) are
associated with loading and display of UTM
coefficients and waypoint move parameters.
(9.) CLEAR key. When pressed, illuminates
and erases data loaded into data displays or FROM-TO
display. Used to cancel erroneous data. After clearing,
data loading can be resumed.
3-47
TM 55-1510-221-10
Figure 3-25. INS Control Display Unit
3-48 Change 4
TM 55-1510-221-10
(10.) WYPT CHG key. W h e n p r e s s e d ,
enables numbers in FROM-TO display to be
changed. If INSERT/ADVANCE key is pressed,
computer will use navigation leg defined by new
number in all navigation computations. If INSERT/
ADVANCE key is not pressed, computer will continue using original numbers in all navigation computations; but distance/time information, based on
new leg, may be called up and read in data displays
(in case of waypoints). When not in TACAN mix
mode, TACAN station number is inserted to display
DIS/TIME information.
(11.) AUTO-MAN TEST switch. This is a
dual purpose control. When the knob is pressed
inward, the TEST switch function is engaged. When
the knob is rotated to either the AUTO or MAN setting, the control serves as a selector between those
modes.
(a.) AUTO. Selects automatic leg
switching mode. Computer switches from one leg to
the next whenever waypoint in TO side of the
FROM-TO display is reached.
(b.) MAN. Selects manual leg switching mode. Pilot must make waypoint changes manually.
(c.) TEST. When pressed, performs
test of INS lights and displays, remote lights and
indicators controlled by INS, and computer input/
output operations.
Used with other controls to activate display of
numerical codes denoting specific malfunctions and
resets malfunction warning circuits.
During alignment, activates the HSI test. Continued pressing of switch provides constant INS outputs to drive cockpit displays in a predetermined
fashion.
NOTE
The INS can provide test signals to the
Horizontal Situation Indicator (HSI) and
connected displays. Pressing TEST switch
during STBY, ALIGN, or NAV modes
will cause all digits on connected digital
displays to indicate “S’s“ and illuminates
the HSI “WAYPOINT” and ALERT
lights. Additional HSI test signals are provided when INS is in ALIGN and the
data selector is at any position other than
DSRTK/STS. Under those conditions,
pressing TEST switch causes the HSI to
indicate heading, drift angle, and track
angle error - all at “0°“ or “30°“. At the
same time, cross track deviation is indicated at “3.75“ nautical miles (one dot)
right or left and INS-controlled HSI flags
are retracted from view.
Output test signal are also supplied to the
autopilot when INS steering is selected. Rotating
AUTO-MAN switch to AUTO and pressing TEST
during align furnishes a 15° left bank steering command. A 15° right bank steering command is furnished when the AUTO/MAN switch is set to MAN.
(12.) FROM-TO display. Display numbers
defining waypoints of navigation leg being flown or,
in the case of a flashing display, displays TACAN
station being used. Waypoint numbers automatically
change each time a waypoint is reached. Unless
flight plan changes during flight, the automatic leg
switching sequence will always be 1, 2, 2 3, 3 4....8
9, 9 1, 1 2, etc.
(13.) Data selector. Selects data to be displayed in data displays or entered into INS. The
rotary selector has 10 positions. Five positions (L/L
POS, L/L WY PT, UTM POS, UTM WY PT and
DSRTK/STS) also allow data to be loaded into data
display then inserted into computer memory.
(a.) TK/GS. Displays aircraft track
angle in left display and ground speed in right display.
(b.) HDG/DA. Displays aircraft true
heading in left display and drift angle in right display.
(c.) XTK/TKE. Displays cross track
distance in left display and track angle in right display.
(d.) L/L POS. Displays or enters
present aircraft position latitude in left data display
and longitude in right data display. Both displays
indicate degrees and minutes to nearest tenth of a
minute. This position also enables the insertion of
present position coordinates during alignment and
present position updates.
(e.) L/L WY PT. Displays or enters
waypoint and TACAN station data, if used in conjunction with the waypoint/TACAN selector. This
position will also cause display of inertial present
position data when the HOLD key is illuminated.
(f.) DIS/TIME. Displays distance
from aircraft to TACAN station or any waypoint, or
between any two waypoints in left display. Displays
time to TACAN station or any waypoint, or between
any two waypoints, in right display.
(g.) WIND. Displays wind direction
in left display and wind speed in right display, when
true airspeed is greater than the air data system
lower limit (115 to 400 KIAS).
3-49
TM 55-1510-221-10
(h.) DSRTK/STS. Displays
desired
track angle to nearest degree in the left data display,
and INS system status in right data display.
(i.) UTM POS. Displays or enters
aircraft position in Universal Transverse Mercator
(UTM) coordinates, with northing data in kilometers in left display and easting data in kilometers in
right display. The extra precision display shows
meters.
prior to engine starting. In this event, the
engines must not be started until after the
INS is placed in the NAV mode.
1.
Applicable circuit breakers Check depressed.
2. Mode selector switch (MSU, fig.
3-23) - ALIGN. Confirm following:
a.
(j.) UTM WY PT. Displays or enters
waypoint and TACAN station data in UTM coordinates. Also enables loading and display of spheroid
coefficients if GRID and DISP keys are pressed
simultaneously.
b. INSERT/ADVANCE pushbutton light (CDU) illuminates.
c. BAT light (MSU, fig. 3-23)
illuminates for approximately 12 seconds at alignment state “8”, then extinguishes.
(14.) Dim knob. Controls intensity of CDU
key lights and displays.
(15.) Waypoint/DME selector. Thumbwheel
switch, used to select waypoints for which data is to
be inserted or displayed. Waypoint station “0“ is for
display only and cannot be loaded with usable data.
d.
INS - Normal Operating Procedures.
NOTE
The following data will be required prior
to operating the INS: latitude and longitude (Geographical) or Universal Transverse Mercator (UTM) coordinates of aircraft during INS alignment. This
information is necessary to program the
INS computer during alignment procedure.
NOTE
When inserting data into INS computer,
always start at left and work to right. The
first digit inserted will appear in right
position of applicable display. It will step
to left as each subsequent digit is entered.
The degree sign, decimal point, and colon
(if applicable) will appear automatically.
(1.) Preflight Procedure.
Insure that cooling air is available to navigation unit before turning the INS on.
NOTE
Aircraft must be connected to a ground
power unit if INS alignment is performed
3-50
FROM-TO display (CDU,
fig. 3-24) indicates “1 2“.
3. Dim knob (CDU) - Adjust for
optimum brightness of CDU displays.
4. AUTO-MAN
TEST
switch
(CDU) - AUTO.
5.
Data selector (CDU) - L/L POS or
UTM POS, as desired. Observe
coordinates of last present position prior to INS shutdown
appear in data displays.
NOTE
Aircraft must not be towed or taxied during INS alignment. Movement of this type
during alignment causes large navigation
errors. If aircraft is moved during alignment, restart alignment by setting mode
selector switch to STBY, then back to
ALIGN and reinserting present position.
NOTE
Passenger or cargo loading in the aircraft
could cause the type of motion which
affects the accuracy of alignment. Any
activity which causes the aircraft to
change attitude shall be avoided during
the alignment period.
switch
6. AUTO-MAN
TEST
(CDU) - Press and hold for test.
Confirm following on CDU:
7. Left and right data displays indicate “88°88.8 N/S” and “88°88.8
E/W“ respectively.
TM 55-1510-221-10
FROM-TO
“8.8“.
9.
The following pushbuttons and
lights illuminate: ROLL LIM,
INSERT/ADVANCE,
HOLD,
WYPT CHG, ALERT, BAT (on
WARN,
CDU and
MSU),
READY NAV and WYPT on
pilot’s HSI.
10.
11.
display
indicates
8.
switch
AUTO-MAN
TEST
(CDU) - Release. Confirm data
displays indicate coordinates in
computer memory.
2.
3. INSERT/ADVANCE pushbutton - Press. Observe
pushbutton light remains
illuminated.
4.
If UTM coordinates are to be
used, verify that appropriate grid
coefficients have been loaded.
(2.) Insert present position:
NOTE
Prior to pressing INSERT/ADVANCE
pushbutton, any incorrectly loaded data
can be corrected by pressing the CLEAR
pushbutton and reloading correct data.
NOTE
While parked aircraft is undergoing alignment, encoding altimeter will supply the
field elevation (aircraft pressure altitude)
into INS.
NOTE
Once present position has been inserted
and computer has advanced to alignment
state “7“, present position cannot be reinserted without downmoding to STBY and
restarting alignment.
NOTE
If longitude and latitude coordinates are
being used, skip following step (a) -2- and
proceed with step (b) -2-.
(a.) Insert UTM coordinates of aircraft present position:
1.
Data selector - UTM POS.
Observe that prior to initial
load, INSERT/ADVANCE
pushbutton light illuminates.
To load zone and easting
values - Press keys in
sequence, starting with “E“.
Example: Zone 16, 425 km
East = E16 425. Observe
that zone and easting in kilometers appear in right data
display as keys are pressed.
To load northing data Press keys in sequence, starting with “N“ or “S“ to indicate north or south hemisphere. Example: 4749 km
North = N 4749. Observe
northing kilometers appear
in left data display as keys
are pressed.
5. INSERT/ADVANCE pushbutton - Press. Observe that
the pushbutton light remains
illuminated.
6. INSERT/ADVANCE pushbutton -‘Press. Observe extra
precision display for present
position northing and easting, to the nearest meter,
appears in left and right data
displays, respectively.
7.
To load extra precision easting data - Press keys in
sequence, starting with “E”.
Example: 297 m East = E
297. Observe that easting
meters appear in right data
display as keys are pressed.
8. INSERT/ADVANCE pushbutton - Press. Observe
pushbutton light remains
illuminated.
9. To load extra precision
northing data - Press keys in
sequence, starting with “N“
“S”. Example: 901 m
North = N 901. Observe
that northing meters appear
in left data displays as keys
are pressed.
3-51
TM 55-1510-221-10
NOTE
with “N” or “S“ to indicate
north or south. Example:
42°54.0' North = N 4 2 5 4
0. Observe that latitude
appears in left data display
as keys are pressed.
Extra precision values are always added
to normal values regardless of which key
(N/S) is pressed to initiate the entry. The
normal entry establishes the hemisphere.
10. INSERT/ADVANCE pushbutton - Press. Observe latitude and longitude data is
displayed in UTM and
INSERT/ ADVANCE pushbutton light extinguishes.
NOTE
The computer will convert coordinates in
the overlap area; however display values
will reference appropriate zone.
NOTE
The “W“ key may be used to initiate easting entries; however computer will always
interpret such entries as an “E“ input. “E“
will be displayed in normal UTM display.
3. INSERT/ADVANCE pushbutton - P r e s s . O b s e r v e
pushbutton light remains
illuminated.
4. To load longitude data Press keys in sequence, starting with “W” or “E” to indicate west or east. Example:
87°54.9’ West = W 8 7 5 4
9. Observe that longitude
appears in right data display
as keys are pressed.
5. INSERT/ADVANCE pushbutton - Press. Observe
pushbutton
light
extinguishes.
6.
NOTE
a.
Extra precision values are always added
to normal values. As an example, South
4,476.995 m will display “4476S“ in normal display and “995“ in extra precision
display. There is no rounding between the
two displays.
Left data display indicates desired track
angle
in
computer
memory.
b.
Right data display indicates --84, -74, -64,
or --54, depending on
which alignment state
the
computer
has
reached.
(b.) To insert geographic coordinates
of aircraft present position:
NOTE
NOTE
Prior to pressing INSERT/ADVANCE
pushbutton, any incorrectly loaded data
can be corrected by pressing the CLEAR
pushbutton and loading correct data.
3-52
Data selector - DSRTK/STS.
Confirm:
1.
Data selector - L/L POS.
Observe that, prior to initial
load,
the
INSERT/
ADVANCE pushbutton light
is illuminated.
2.
To load latitude data - Press
keys in sequence, starting
After present position has been inserted
and computer has advanced to state “7“,
present position cannot be reinserted
without downmoding to STBY and
restarting alignment.
7.
Data selector (CDU) DSRTK/STS. Observe lefthand data display indicates
the desired track in computer memory and right data
display
indicates
status
"-194“.
TM 55-1510-221-10
NOTE
NOTE
If fourth digit from right is blank, a valid
heading has not been stored. Proceed with
normal preflight procedure.
Values for various spheroids are listed in
table 3-1.
8.
4.
Monitor data display for system alignment state “9“ to
alignment state “8“. Observe
right data display will be “--184“.
9. Monitor data display for
malfunction codes. Observe
if the 26V 400 Hz power is
off, “.03184“ will appear in
the right data display and
WARN light illuminates. If
magnetic compass system is
off, “.03184“ will appear in
right data display and
WARN light is extinguished.
10. If there are malfunction
codes, proceed to ABNORMAL PROCEDURES in
this chapter.
If values are correct, return CDU
to normal display mode by
momentarily setting data selector
to any position except UTM WY
PT. If values are to be changed,
continue with following steps:
(4.) Abbreviated INS Interface Test - As
required.
NOTE
Assuming a level aircraft, attitude indicators will become level during alignment
state “8“ and remain level in all modes
until INS is shut down. Warning indicators for INS attitude signals from the INS
are valid while attitude sphere display is
level.
NOTE
NOTE
To achieve best accuracy, engine start and
heavy loading activity should be delayed
until entry into NAV mode.
NOTE
Waypoint data and TACAN station data
may be loaded any time after turn-on.
(3.)
Verify UTM Grid Coefficients:
1.
Data selector (CDU) - UTM WY
PT.
2.
Keys “5” and “9“ - Press simultaneously. Observe FROM-TO display is blank. Earth flatness coefficient appears in left display. The
relative earth radius, in meters,
appears in right display.
NOTE
These values are retained from turn-on to
turn-on unless changed by operator.
3. Verify that values correspond to
those required for spheroid being
used.
The INS can provide test signals to the
Horizontal Situation Indicator (HSI) and
connected displays. Pressing TEST switch
during STBY, ALIGN, or NAV mode
causes all digits on connected digital displays to indicate “8’s,” and lights the
WYPT on the pilot’s HSI and the ALERT
light. Additional HSI test signals are provided when INS is in ALIGN and data
selector is at any position other than
DSRTK/STS. Under those conditions,
pressing TEST switch causes HSI to indicate heading, drift angle, and track angle
error - all at “0°" or “30°.“ At the same
time, cross track deviation is indicated at
“3.75“ nautical miles (one dot) right or
left and INS-controlled HSI flags are
retracted from view.
NOTE
Output test signals are supplied to the
autopilot when INS steering is selected.
Rotating AUTO/MAN switch to AUTO
and pressing TEST during alignment furnishes a 15° left bank steering command.
A 15° right bank steering command is furnished when AUTO/MAN switch is set to
MAN.
3-53
TM 55-1510-221-10
NOTE
(5.) To program destinations or TACAN
coordinates:
The quick test procedure may be performed any time after alignment State “8“
is reached and prior to entry into NAV.
NOTE
1.
Mode selector (MSU) - ALIGN.
Observe CDU displays are illuminated.
2. Data selector (CDU) - DSRTK/
STS. Monitor right data display
until state “8“ (or lower) is
reached. Observe right data display is ---N4, where “N“ is not
“9”.
3. AUTO-MAN switch (CDU)MAN.
4.
Data selector - Set to any position
except DSRTK/STS.
5. INS - Couple to flight director
and autopilot, as applicable. After
performing the preceeding step,
observe:
a.
All lights on MSU - Check
illuminated.
b. All lights on CDU - Check
illuminated. All “8’s“ displayed.
6.
c.
HSI - All angles 30°. Crosstrack deviation bar one dot
right. All INS flags retracted.
d.
Flight Director/Autopilot - A
15° steering command is
issued.
e.
Mission Control Panel - INS
UPDATE and NO INS
UPDATE annunciator illuminated.
CDU TEST switch - Hold
depressed, and rotate AUTOMAN switch to AUTO. Observe
all indications are as in step -6except a 15° left steering command is issued. On HSI, all angles
are “0°" and cross-track deviation
bar is one dot left.
7. CDU TEST switch - Release. If
desired, decouple INS. Observe
that operation returns to normal.
3-54
If latitude and longitude (Geographic)
coordinates are being used, skip following
procedure (5)(a.) and execute next procedure (5)(b.). Enter all of the data for a
given destination or TACAN before starting to enter data for another.
(a.)
Insertion of UTM waypoint coor-
dinates:
1.
Data selector - UTM WY
PT. Data displays will indicate last coordinates inserted
into related waypoint.
2.
Thumbwheel - Set to
waypoint n u m b e r t o b e
loaded.
NOTE
UTM data may be loaded in any order
and, until final entry, a value may be
reloaded.
3.
To load zone and easting Press keys in sequence, starting with “E”. Example: Zone
16, 425 km East = E16 425.
Observe that zone and easting in kilometers appear in
the right data display as keys
are pressed.
4.
INSERT/ADVANCE pushbutton - Press. Observe
pushbutton light is illuminated.
5.
To load northing press keys
in sequence, starting with
“N” or “S” to indicate north
or south hemisphere. Example: 4749 km North = N
4749. Observe that northing
kilometers appear in the left
data display as keys are
pressed.
6.
INSERT/ADVANCE pushbutton - Press. Observe
pushbutton light remains
illuminated.
7 . INSERT/ADVANCE pushbutton - Press. Observe that
TM 55-1510-221-10
TABLE 3-2. VARIOUS VALUES FOR UTM GRID COEFFICIENTS
SPHEROID
FLATNESS
COEFFICIENT
International
Clark 1866
Clark 1880
Everest
Bessel
Modified Everest
Australian National
Airy
Modified Airy
29700
29498
29346
30080
29915
30080
29825
29932
29932
RELATIVE
RADIUS
8388
8206
8249
7276
7397
7304
8160
7563
7340
m
m
m
m
m
m
m
m
m
SOURCE: Universal Transverse Metcator
Grid Technical Manual,
TM 5-241-8, Headquarters,
Department of the Army,
30 April 1973, page 4.
Flatness Coefficient: 100 (I/f)
Relative Radius: a-6,3700.000
an extra precision display
related to resident value of
northing and easting, to the
nearest meter, appears in left
and right data displays,
respectively.
8.
To load extra precision easting value - Press keys in
sequence starting with “E“.
Example: 297 m East = E
297. Observe that easting
meters appear in the right
data display as keys are
pressed.
9. INSERT/ADVANCE pushbutton - Press. Observe
pushbutton light remains
illuminated.
10. To load extra precision
northing value - Press keys
in sequence, starting with
“N” or “S“. Example: 901 m
North = N 901. Observe
that northing meters appear
in left data display as keys
are pressed. The value is
always added to the normal
value regardless of which
key (N/S) is pressed to ini-
tiate the entry. The normal
entry establishes the hemisphere.
11.
INSERT/ADVANCE pushbutton - Press. Within 3 seconds computer converts
input into latitude and longitude for storage in memory.
The stored value is again
converted to UTM for disINSERT/
The
play.
ADVANCE pushbutton light
extinguishes.
Conversion
routines may cause displays
to change by up to 10 m.
NOTE
The computer will convert coordinates in
overlap area; however, data display values
will reference appropriate zone.
NOTE
The “W“ key may be used to initiate easting entries; however, the computer will
always interpret such entries as an “E“
input. “E“ will be displayed in normal
UTM data display.
3-55
TM 55-1510-221-10
pushbutton
guishes.
NOTE
The extra precision values are always
added to normal values. As an example,
South 4,476.995 m will display “4476 (S)“
in the normal display and “995” in extra
precision display. In other words, there is
no rounding between the two displays.
12. Repeat steps 2 through 11
for each waypoint to be
loaded.
(b.)
If desired to insert extra precision coordinate data -Press
INSERT/ADVANCE pushbutton. Observe that arcseconds for loaded latitude
and longitude, to nearest
tenth of a second, appear in
left and right data displays,
respectively.
8.
To load related arc-second
values for latitude - Press
keys in sequence, starting
with “N“. Example: 35.8“
North = N 358.
INSERT/ADVANCE pushbutton - Press. Observe
light
extinpushbutton
guishes.
9.
Insertion of geographic waypoint
coordinates:
1.
Data selector - L/L WY PT.
Data displays indicate last
coordinates inserted into the
selected waypoint.
2 . T h u m b w h e e l - Set to
waypoint n u m b e r t o b e
loaded.
3.
To load latitude - Press keys
in sequence, starting with
“N“ or “S” to indicate north
or south. Example: 42° 54.0'
N o r t h = N 4 2 5 4 0.
INSERT/
that
Observe
ADVANCE pushbutton light
illuminates when first key is
pressed, and latitude appears
in left data display as keys
are pressed.
4. INSERT/ADVANCE pushbutton - Press. Observe
extinlight
pushbutton
guishes.
5. To load longitude - Press
keyboard keys in sequence,
starting with “W“ or “E“
indicating west o r e a s t .
Example: 87°54.9’ West =
W 8 7 5 4 9. Observe that
INSERT/ADVANCE pushilluminates
button
light
when first key is pressed,
and longitude appears in display as keys are pressed.
6. INSERT/ADVANCE pushbutton - Press. Observe
3-56
extin-
7.
NOTE
A load cycle may be terminated prior to
insertion of all four values by moving
data selector or thumbwheel.
light
10.
11.
12.
To load related arc-second
values for longitude - Press
keys in sequence, starting
with “E“. Example: 20.1“
East - E 201.
INSERT/ADVANCE pushbutton - Press. Observe
light
extinpushbutton
guishes.
Repeat steps -2- through -11for each waypoint to be
loaded.
NOTE
e x a m p l e , i f INSERT/
In
above
ADVANCE pushbutton was pressed, the
following normal display would appear:
“42°54.5 (N)“ and 87°54.3(W). The extra
precision values are added to normal values and normal data displays are not
rounded off.
NOTE
The normal geographic coordinates must
always be loaded prior to extra precision
values.
NOTE
The directions “N“ or “S” and “E“ or
“W” are established during normal coordinate entry. Either key may be used to
initiate entry during extra precision loads
and values will be added to the extra precision value without affecting direction.
TM 55-1510-221-10
NOTE
It is characteristic of the computer display
routine to add “0.2“ arc-seconds to any
display of “59.9” arc-seconds. The value
in computer is as loaded by operator.
(6.) To insert TACAN coordinates:
(a.) Insertion of UTM TACAN station data:
NOTE
Prior to pressing the INSERT/ADVANCE
pushbutton, any incorrectly loaded data
can be corrected by pressing CLEAR
pushbutton and loading correct data.
1.
Data selector - UTM WY
PT.
2. Keys “7“ and “9” - Press
simultaneously. Observe that
number of TACAN station
being used for navigation
flashes on and off in
“FROM-TO“ display and
data displays indicate coordinates of station selected by
thumbwheel.
3. Thumbwheel - Set to number of station to be loaded.
Confirm thumbwheel is in
detent.
4.
Station “0” c a n n o t b e
loaded. Observe that if number of station to be loaded is
same as number of the
TACAN station currently
being used, number in
“FROM-TO“ display will be
set to “0“ when TACAN data
is loaded.
5. To load zone and easting Press keys in sequence, starting with “E“. Example: Zone
16, 425 km East = E16 425.
Observe that zone and easting in kilometers appear in
the right display as keys are
pressed.
6. INSERT/ADVANCE pushbutton - Press. Observe that
pushbutton light is illuminated.
7. To load northing - Press
keys in sequence, starting
with “N“ or “S” to indicate
north or south hemisphere.
Example: 4749 km North =
N 4749. Observe that northing kilometers appear in left
data display as keys are
pressed.
8. INSERT/ADVANCE pushbutton - Press. Observe
pushbutton light remains
illuminated.
9. INSERT/ADVANCE pushbutton - Press. Observe that
display
precision
extra
related to the resident value
of northing and easting, to
nearest meter, appears in left
and right data displays,
respectively.
NOTE
UTM data may be loaded in any order.
Until final fourth entry, actuation of
INSERT/ADVANCE pushbutton without
a prior data entry will cause normal and
extra precision UTM data to be alternately displayed.
10.
To load extra precision easting value - Press keys in
sequence, starting with “E”.
Example: 297 m East = E
297. Observe that easting
meters appear in right data
display as keys are pressed.
11.
INSERT/ADVANCE pushbutton - Press. Observe
pushbutton remains illuminated.
12.
To load extra precision
northing value - Press keys
in sequence, starting with
“N” or “S”. Example: 901 m
North = N 901. Observe
that northing meters appear
in left data display as keys
are pressed. The value is
always added to -ormal
value regardless of which
key (N/S) is pressed to initiate entry. The normal entry
establishes hemisphere.
3-57
TM 55-1510-221-10
13.
INSERT/ADVANCE pushbutton - Press. - 10erve that
during the next 1 to 3 seconds, the computer converts
input into latitude and longitude for storage in memory.
The stored value is again
converted back to UTM for
display to operator. The
INSERT/ADVANCE pushbutton light extinguishes.
The conversion routines
may cause data displays to
change by up to 10 m.
NOTE
Altitude inputs are limited to 15,000 feet.
17.
INSERT/ADVANCE pushbutton - Press. Observe
pushbutton
light
extinguishes.
18.
INSERT/ADVANCE pushbutton - Press. Observe that
left data display indicates
last
previously
inserted
channel number, and right
display is blank.
19.
To indicate following load is
channel number - Press keys
"2"
"8"
Observe
INSERT/ADVANCE pushbutton light illuminates.
20.
To load channel number Press keys in sequence.
Example:
109 =
109.
Observe number appears in
left data display as keys are
pressed.
21.
INSERT/ADVANCE pushbutton - Press. Observe
pushbutton
light
extinguishes.
NOTE
The computer will convert coordinates in
overlap area; however, data display values
will reference appropriate zone.
NOTE
The “W” key may be used to initiate easting entries; however, the computer will
always interpret such entries as an “E”
input. “E“ will be displayed in normal
UTM data display.
NOTE
The extra precision values are always
added to normal values. As an example,
South 4,476.995 m will display “4476 S“
in normal display and “995” in extra precision display. In other words, there is no
rounding between the two displays.
NOTE
Any number will be accepted by INS;
however, only stations with a channel
number within range of “1” through
“126“ will be used for TACAN mixing.
NOTE
14.
15.
16.
3-58
INSERT/ADVANCE pushbutton - Press. Observe right
data display indicates last
previously inserted altitude,
and left data display is
blank.
To indicate the following
load is altitude - Press keys
“4”
or
“6“.
Observe
INSERT/ADVANCE pushbutton light illuminates.
To load altitude in feet Press keys in sequence.
Example: 1230 ft = 1230.
Observe
numbers
that
appear in right data display
as keys are pressed.
Channel number has an implied “X“ suffix.
NOTE
Degree symbol (°)should be disregarded
when reading altitude and data display.
22. INSERT/ADVANCE pushbutton - Press. Observe station northing, zone, and
easting reappear.
23.
Repeat steps -1- through -22for each TACAN station.
24. To return INS to normal
mode, momentarily set data
selector to UTM POS.
TM 55-1510-221-10
INSERT/
Observe
that
ADVANCE pushbutton light
illuminates when first key is
longitude
and
pressed,
appears in data display as
keys are pressed.
(b.) Insertion of geographic TACAN
station data:
NOTE
Prior to pressing INSERT/ADVANCE
pushbutton, any incorrectly loaded data
can be corrected by pressing CLEAR
pushbutton and loading correct data.
1.
Data selector - L/L WY PT.
7.
8.
NOTE
If number of station to be loaded is same
as number of TACAN station currently
being used, number in “FROM-TO” display will be set to “0” when TACAN data
is loaded.
2. Keys “7” and “9“ - Press
simultaneously. Observe that
number of TACAN station
being used for navigation
flashes on and off in
“FROM-TO“ display. Data
displays indicate coordinates
of station selected via thumbwheel.
3. Thumbwheel - Set to number of station being loaded.
(Insure thumbwheel is in
detent.)
NOTE
Station “0” cannot be loaded.
4.
To load latitude - Press keys
in sequence, starting with
“N” or “S” to indicate north
or south. Example: 42° 54.0’
North = 4 2 5 4 0. Observe
INSERT/ADVANCE
that
pushbutton light illuminates
when first key is pressed.
5. INSERT/ADVANCE pushbutton - Press Observe pushbutton light extinguishes.
6. To load longitude - Press
keys in sequence, starting
with “W” or ,“E” indicating
west or east. Example: 87°
54.9’West = W 8 7 5 4 9.
INSERT/ADVANCE pushbutton - Press. Observe
extinlight
pushbutton
guishes.
INSERT/ADVANCE pushbutton - Press. Observe that
the arc-seconds related to
loaded latitude and longitude, to nearest tenth of a
second, appear in left and
right data display, respectively.
9.
If extra precision coordinate
data is to be inserted -Press
keys in sequence, starting
with “N“, to load related
arc-second values for latitude. Example: 35.8“ North
= N 358.
10.
INSERT/ADVANCE pushbutton - Press. Observe
extinlight
pushbutton
guishes.
11.
To load related arc-second
values for longitude - Press
keys in sequence, starting
with “E”. Example: 20.1”
East - E 201.
12.
INSERT/ADVANCE pushbutton - Press. Observe
extinlight
pushbutton
guishes.
NOTE
e x a m p l e , i f INSERT/
above
In
ADVANCE pushbutton were pressed, the
following normal display would appear:
“42° 54.5 N” and “87° 54.3 W”. The extra
precision values are added to normal values and normal displays are not rounded
off.
NOTE
The normal geographic coordinates must
always be loaded prior to extra precision
values.
3-59
TM 55-1510-221-10
NOTE
19. To load channel number Press keys in sequence.
Example: 109 = 109. Numbers appear in left data display as keys are pressed.
The directions “N“ or “S” and “E” or
“W“ are established during normal coordinate entry. Either key may be used to
initiate entry during extra precision loads
and the values will be added to extra precision value without affecting direction.
20. INSERT/ADVANCE pushbutton - Press. Observe
light
extinpushbutton
guishes.
NOTE
It is characteristic of the computer display
routine to add 0.2 arc-seconds to any display of 59.9 arc-seconds. The value in
computer is as loaded by operator.
13.
14.
15.
INSERT/ADVANCE pushbutton - Press. Observe that
right data display indicates
last previously inserted altitude, and left data display is
blank.
To indicate the following
load is altitude - Press key
“4“
or
“6“.
Observe
INSERT/ADVANCE pushbutton light illuminates.
NOTE
Any number will be accepted by the INS;
however only stations with a channel
number within range of 1 through 126
will be used for TACAN mixing.
NOTE
The channel number has an implied “X“
suffix.
NOTE
Decimal points and degree symbols
should be disregarded when reading altitude and channel number displays.
To load altitude first - Press
keys in sequence. Example:
1230 ft = 1230. Numbers
appear in right data display
as keys are pressed.
NOTE
Altitude inputs are limited to 15,000 feet.
16.
17.
18.
3-60
INSERT/ADVANCE pushbutton - Press. Observe
pushbutton
light
extinguishes.
INSERT/ADVANCE pushbutton - Press. Observe that
left data display indicates
last
previously
inserted
channel number, and right
data display is blank.
To indicate the following
load is channel number Press key “2” or “8“.
Observe
INSERT/
ADVANCE pushbutton light
illuminates.
(7.)
21.
INSERT/ADVANCE pushbutton - Press. Observe station latitude and longitude
reappear.
22.
Repeat steps -3- through -19for each TACAN station.
23.
To return INS to normal display modes, momentarily set
data selector to L/L POS.
Designating fly-to destinations:
1.
Data selector - L/L WY PT or
UTM WY PT, as required.
2.
Waypoint thumbwheel
destination
number.
number of destination
a p p e a r s i n “TO”
FROM-TO display.
3.
Data
selector HDG/DA.
Observe present aircraft heading
appears, t o n e a r e s t t e n t h o f
degree, in left data display; also
drift angle, to nearest degree,
appears in right data display.
- Select
Observe
waypoint
part
of
TM 55-1510-221-10
NOTE
6.
Navigation information is now available
from the INS for display on the pilot and
co-pilot RMI’s and on the pilot and copilot HSI’s, as determined by the COURSE
INDICATOR and RMI select switches.
(8.)
To fly selected INS course:
1.
Pilot’s COURSE INDICATOR
switch - INS.
2. Pilot’s RMI select switch - INS.
3. Horizontal Situation Indicators
(pilot’s and/or copilot’s HSI) Steer toward indicators.
4. CDU ALERT light - Monitor.
Observe illumination approx 1.3
minutes before reaching point for
automatic leg switch. Indicator
flashes on and off after passing a
waypoint, if AUTO-MAN switch
is in MAN.
(9.) Aided TACAN operation:
NOTE
If high accuracy alignment is required,
wait for the READY NAV light before
selecting NAV mode.
1.
Mode selector - NAV.
2.
Data selector - DSRTK/STS.
3. Key “4“ - Press. Observe right
data display is “000004“ and
INSERT/ADVANCE pushbutton
light is illuminated.
4. INSERT/ADVANCE pushbutton
- Press. Observe right data display
INSERT/
is
"1 -XX4”
and
ADVANCE pushbutton light is
extinguished.
7. To monitor station selection Observe FROM-TO data display.
Observe only the number of stations eligible for mixing will be
displayed. A “0“ indicates that
none of the 9 stations are eligible
for selection.
8. Monitor “INS UPDATE“ annunciator.
NOTE
Mixing will not be annunciated if: (a)
TACAN control is inappropriately set; (b)
TACAN station data loaded in error; (c)
aircraft look-down angle is greater than
30”; (d) horizontal ground distance is less
than two times the altitude. When 2 minutes elapse without an update, the “NO
INS UPDATE“ annunciator will illuminate.
9. To return INS display to normal
- Set data selector to any position
except WYPT or DIS/TIME.
10.
5. Data selector - L/L WY PT or
UTM WY PT.
To monitor program of mix - Set
data selector to DSRTK/STS.
(Observe Accuracy Index (AI) will
decrement to “0“.)
NOTE
To insure favorable geometry during the
update process, the following TACAN station criteria should be observed:
11.
One station must be at least 15
nm off course.
12.
For optimum single TACAN
updating, update should continue
until aircraft has passed the station.
13.
For optimum dual TACAN
updating, u s e o n e “off-track“
TACAN station and one "ontrack“ station.
14.
For optimum multi-TACAN station updating, the stations should
NOTE
Every 30 seconds, the INS will select next
eligible TACAN station in sequence for
updating. To be eligible, TACAN station
range must be between 5 and 150 nm and
channel between 1 and 126.
Keys “7“ and “9“- Press simultaneously. Observe channel number
of the TACAN station being used
for navigation flashes on and off.
Data displays indicate coordinates of station selected via thumbwheel.
3-61
TM 55-1510-221-10
be evenly distributed in azimuth
around the aircraft.
to nearest tenth of a degree. Release keys 3 and 9.
Left display reverts to true heading.
15. Waypoint thumbwheel - Set to
number of first TACAN station to
be used. Observe selected station
number is displayed on the “TO“
side of FROM-TO data display.
(e.) Ground speed: Data selector TK/GS. Observe ground speed appears in right data
display to nearest knot.
16. Update GPS.
(10.) Switching from aided to unaided inertial operation.
1.
Data selector - DSRTK/STS.
(g.) Drift angle: Data selector HDG/DA. Observe drift angle appears in right data
display to nearest degree.
2.
Key “5” - Press. Observe
INSERT/ADVANCE pushbutton
light illuminates; 000005 appears
in right data display.
(h.) Wind speed and direction: Data
selector - WIND. Wind direction appears in left data
display to nearest degree and wind speed appears in
right display to nearest knot.
3. INSERT/ADVANCE pushbutton
Observe
INSERT/
- Press.
light
ADVANCE pushbutton
extinguishes. Data display returns
to normal with “5“ appearing in
first digit of right display.
(i.) Desired track angle: Data selector - DSRTK/STS. Observe desired track angle in
right data display to nearest degree.
NOTE
Benefits of previous aiding are maintained but no additional automatic
updates will be made.
(11.)
(f.) Ground track angle: Data selector - TK/GS. Observe ground track angle appears in
left data display to nearest tenth of a degree.
To obtain readouts from INS:
NOTE
The computer is assumed to be in the
NAV mode for all data displays.
(j.) Track angle error: Data selector
- XTK/TKE. Observe track angle error appears in
right data display to nearest degree.
(k.) Cross track distance: Data selector - XTK/TKE. Observe cross track distance
appears in left data display to nearest nautical mile.
(l.) Distance and time to next
waypoint: Data selector - DIS/TIME. Observe distance to next waypoint, shown in “TO“ side of
FROM-TO display, appears in left data display to
nearest nautical mile. Observe time to reach next
waypoint at present ground speed appears in right
data display to nearest tenth of a minute.
(m.) Extra precision geographic present position display:
(a.) System status: Data selector DSRTK/STS. Observe numbers indicating system
status appear in right data display.
(b.) Geographic present position:
Data selector - L/L POS. Observe latitude and longitude of present position appear in left and right data
displays, respectively. Both displays are to tenth of
a minute.
1.
Data selector - L/L POS.
Latitude and longitude of
present position, to nearest
tenth of a minute, appears in
left and right data displays,
respectively.
2.
INSERT/ADVANCE pushbutton - Press. Observe arcseconds related to present
position latitude and longitude, to nearest tenth of a
second, appear in left and
right data displays, respectively.
(c.) UTM position: Data selector UTM POS. Observe northing and zone with easting
of present position appear in left and right displays,
respectively. Both displays are in kilometers.
(d.) True heading and MAG heading:
Data selector - HDG/DA. Observe aircraft true
heading appears in left data display to nearest tenth
of a degree. Press and hold keys 3 and 9 simultaneously. MAG heading appears in left data display
3-62
(n.) Geographic present inertial position display.
1.
Data selector - L/L WY PT.
TM 55-1510-221-10
2. HOLD pushbutton - Press.
Observe HOLD pushbutton
light illuminates, latitude
and longitude of present
inertial position to a tenth of
degree appear in left and
right data displays, respectively.
ma1 operation and HOLD
light
extinpushbutton
guishes.
(p.) Distance and time to waypoint
other than next waypoint:
1.
NOTE
While HOLD pushbutton light is extinguished, TACAN updates are inhibited.
3. INSERT/ADVANCE pushbutton - Press. Observe arcsecond related to present
inertial position latitude and
longitude, to nearest tenth of
a second, appears in left and
right data displays, respectively.
4. HOLD pushbutton - Press.
Observe INS returns to normal operation and HOLD
light
extinpushbutton
guishes.
(o.) UTM present inertial position
2. Key “0“ - Press. Observe
“FROM“ side of FROM-TO
data display changes to “0“.
3.
Do not press INSERT/ADVANCE pushbutton. This would cause an immediate
flight plan change.
4.
Data selector - DIS/TIME.
Observe distance to desired
waypoint appears in left data
display to nearest nautical
mile. Time to reach desired
waypoint at present groundspeed appears in right data
display to nearest tenth of a
minute.
5.
CLEAR pushbutton - Press.
Observe INS returns to noroperation.
Observe
mal
INSERT/ADVANCE
and
WYPT CHG pushbutton
lights extinguish. Waypoints
defining current navigation
leg appear in FROM-TO display.
Data selector - UTM WY
PT.
2. HOLD pushbutton - Press.
Observe HOLD pushbutton
light illuminates. Northing
and zone with easting of the
present inertial position in
kilometers appear in left and
right data displays, respectively.
NOTE
While HOLD pushbutton light is illuminated, TACAN updates are inhibited,
3. INSERT/ADVANCE pushbutton - Press. Observe extra
precision values related to
position
present inertial
northing and easting, to
nearest meter, appear in left
and right data displays,
respectively.
4. HOLD pushbutton - Press.
Observe INS returns to nor-
corresponding to
Key
desired waypoint - Press.
Observe “TO“ side of
FROM-TO data display
changes to desired waypoint
number.
NOTE
display:
1.
WYPT CHG pushbutton Press. Observe WYPT CHG
INSERT/ADVANCE
and
pushbutton light illuminates.
(q.) Distance and time between any
two waypoints:
1. WYPT CHG pushbutton Press. Observe WYPT CHG
INSERT/ADVANCE
and
pushbutton lights illuminate.
corresponding to
2. Keys
desired waypoints - Press in
sequence. Observe desired
waypoint numbers appear in
3-63
TM 55-1510-221-10
FROM-TO data display as
keys are pressed.
A D V A N C E a n d WYPT
CHG pushbutton lights illuminate. Station
number
flashing discontinues.
NOTE
Do not press INSERT/ADVANCE pushbutton. This would cause an immediate
flight plan change.
3.
Data selector - DIS/TIME.
Observe distance between
desired waypoints appears in
left data display to nearest
nautical mile. Time to travel
between desired waypoints
at present ground speed
appears in right data display
to nearest tenth of a minute.
6.
NOTE
If wrong key is pressed, press CLEAR;
displays will revert to that indicated in
step 2.
7. INSERT/ADVANCE pushbutton - Press. Observe
INSERT/ADVANCE
and
WYPT CHG pushbutton
lights extinguish. The loaded
digit will appear in right
position of FROM-TO display and will be flashing on
and off. Distance to that station to nearest nautical mile
appears in left data display.
The right display continues
to display time to next
waypoint.
4. CLEAR pushbutton - Press.
Observe INS returns to normal
operation.
Observe
WYPT CHG and INSERT/
ADVANCE pushbutton light
extinguishes.
Waypoints
defining current navigation
leg appear in FROM-TO
data display.
(r.)
Distance to any TACAN station:
1.
Data selector - DIS/TIME.
Observe distance to next
waypoint to nearest nautical
mile is in left data display.
Time to next waypoint to
nearest tenth of a minute is
in right data display.
2. Keys “7“ and “9“ - Press
simultaneously.
Observe
number of TACAN station
being used for navigation
flashes on and off in
FROM-TO display. Distance
to TACAN station to nearest
nautical mile is in left data
display.
T i m e t o next
waypoint is in right data display.
3. If in aided TACAN operation - Monitor display.
Observe station number is
selected every 30 seconds.
4.
If not in aided TACAN
operation - Perform steps 5
through 7.
5. WYPT CHG pushbutton Press. Observe INSERT/
3-64
Key
indicating
desired
TACAN station number Press. Observe number will
appear in left digit location
of FROM-TO data display.
8.
Data
selector
WIND,
momentarily. Returns INS
to normal display mode.
NOTE
If in aided TACAN operation and if the
desired station is not being selected, exit
aided operation per procedure: “Switching From Aided to Unaided Inertial
Operation“, perform steps 1 thru 8, and
then return to aided operation per procedure: “Aided TACAN Operation“.
(s.)
Coordinates of any waypoint:
1.
Data selector - L/L WY PT
or UTM WY PT.
2. Waypoint thumbwheel - Set
desired waypoint. Observe
following:
a.
L/L WY PT: latitude
and
longitude of
desired waypoint, to a
tenth of a minute,
appear in left and right
TM 55-1510-221-10
tion, to tenth of minute, appears in left and
right data displays,
respectively.
data displays respectively.
b. UTM WY PT: Northing and zone with easting
desired
of
waypoint, to a kilometer, appear in left and
right data displays
respectively.
3. INSERT/ADVANCE pushbutton - Press. Observe the
following:
a.
b.
L/L WY PT: The arcseconds
related to
desired waypoint latitude and longitude, to
a tenth of an arcsecond appear, in left
and
right
displays
respectively.
b. UTM WY PT: Northing and zone with easting of desired TACAN
station, to a kilometer,
appear in left and right
data displays, respectively.
3. INSERT/ADVANCE
PUSHBUTTON - Press.
Observe the following:
UTM WY PT: The
extra precision display
related to
desired
waypoint northing and
easting, in
meters,
appear in left and right
data displays respectively.
a.
L/L WY PT: The arcseconds
related to
desired TACAN station, to tenth of an arcsecond, appear in left
and right data displays
respectively.
b.
UTM WY PT: The
extra precision display
desired
related to
TACAN station northing and easting, in
meters, appear in left
and right data displays
respectively.
NOTE
L/L WY PT: A coordinate is the addition
of values for degrees, whole minutes and
seconds.
Example: W 87° 54’ 58.6“ = 87°54.9W
and 58.6.
UTM WY PT: A coordinate is the addition of the values for kilometers and
meters.
Example: S 2,474,706m = 2474S and
706.
(t.) TACAN station data:
Keys “7” and “9” - Press
simultaneously.
2. Waypoint thumbwheel - Set
to desires TACAN station.
Observe number of TACAN
station being used for navigation flashes on and off.
NOTE
Direction is indicated in normal data displays.
4.
5. Example: W 87° 54’ 58.6“
will be displayed as “87°54.
9W” and 58.6”.
6.
UTM WY PT: A coordinate
is the addition of values for
kilometers and meters.
7.
Example: S 2,474,706 m will
be displayed as “2474 S”
and “706“.
1.
a. L/L WY PT: Latitude
longitude of
and
desired TACAN sta-
L/L WY PT: A coordinate is
the addition of values for
degrees, whole minutes, and
seconds.
8. INSERT/ADVANCE pushbutton - Press. Observe
TACAN station altitude, in
feet, will appear in right data
display; degree symbol and
decimal points should be
3-65
TM 55-1510-221-10
disregarded. Left data display is blank.
9. INSERT/ADVANCE pushbutton - Press. Observe
TACAN station channel
number, in whole numbers,
will appear in left data display; degree symbol and decimal point should be disregarded. Right data display is
blank.
NOTE
If INSERT/ADVANCE pushbutton is
pressed, the normal coordinates indicated
in step -3- will be displayed.
NOTE
Waypoint thumbwheel may be moved at
any time and normal coordinates for new
TACAN station will be displayed.
10. Data selector - Momentarily
to any position other than
L/L WY PT, UTM WY PT
or DIS/TIME. (Returns INS
to normal operation.)
(u.) Magnetic heading.
1.
Data selector - HDG/DA.
Observe true heading to
nearest tenth degree appears
in right data display.
2.
Keys “3“ and “9” - Press
simultaneously a n d h o l d .
Observe magnetic heading to
nearest tenth of a degree
appears in left data display.
Drift angle continues to be
displayed in right data display.
Keys “3“ and “9” - Release.
Observe left data display
reverts to true heading.
3.
(12.) INS updating:
(a.) Normal geographic present position check and update:
1.
3-66
Data selector - L/L POS.
Observe latitude and longi-
tude of present position
appear in left and right data
displays, respectively.
2.
Illuminated HOLD pushbutton - Press. Observe latitude
and longitude in data displays freeze at values present
when HOLD pushbutton is
pressed.
NOTE
While HOLD pushbutton light is illuminated, TACAN, GPS and data link
updates are inhibited.
3. Keys - Press in sequence to
load latitude of position reference, starting with “N“ or
“S” to indicate north or
south. Example: 42°54.0'
north = N 4 2 5 4 0.
Observe
INSERT/
ADVANCE pushbutton light
illuminates when first key is
pressed, and latitude appears
in left data display as keys
are pressed.
4. INSERT/ADVANCE pushbutton - Press. Observe
INSERT/ADVANCE pushbutton light remains illuminated, and previous value of
latitude reappears.
5.
Keys - Press in sequence to
load longitude of position
reference, starting with “W”
or “E“ to indicate west or
east. Example: 87°54.9’ west
= W 8 7 5 4 9. Observe longitude appears in right data
display as keys are pressed.
6. INSERT/ADVANCE pushbutton - Press. Observe
INSERT/ADVANCE
and
HOLD pushbutton lights
remain illuminated. North
position error and east position error, in tenth of a nautical mile, will appear in left
and right data displays,
respectively.
TM 55-1510-221-10
NOTE
If WARN light illuminates, proceed to
step 7; otherwise proceed to step 8.
7.
8.
Data selector - DSRTK/STS.
Observe action code “02”
and malfunction code “49“.
This indicates that the radial
error between the loaded
position and the INS position exceeds 33 nautical
miles. Operator must evaluate possibility that either
INS is in error or reference
point position is incorrect. It
is possible to force INS to
accept updated position by
setting data selector to L/L
POS and proceeding to step
8).
If displayed values are
within
tolerance,
press
HOLD pushbutton to return
INS to normal operation. If
one or both values are out of
tolerance, proceed to step 9.
9. Key “2” - Press. Observe left
data display is “00000 N”;
INSERT/ADVANCE
and
HOLD pushbutton lights are
illuminated.
10. INSERT/ADVANCE pushbutton - Press. Observe
INSERT/ADVANCE
and
HOLD pushbutton lights
extinguish. Present position
appears in data displays. Present position check and
update is complete.
NOTE
Within 30 seconds, computer will process
correction and revised present position
will appear in data display. If AI prior to
position update is 1 or greater, computer
will accept over 95 percent of correction
shown in difference display. If AI is “0“,
amount of correction accepted will be less
and is a function of time in NAV mode
and number of updates which have been
made.
(b.) Extra precision geographic present position check and update:
1. Data selector - DSRTK/STS.
2. Key “2” - Press. Observe
INSERT/ADVANCE pushbutton light illuminates,
“000002” appears in right
data display.
3. INSERT/ADVANCE pushbutton - Press. Observe right
data display is “1-XX2“,
INSERT/ADVANCE pushbutton light is extinguished,
and any TACAN, GPS or
data link updating is discontinued.
4. Data selector - L/L POS.
Observe latitude and longitude of present position
appears in left and right data
displays, respectively.
5. HOLD pushbutton - Press
(when aircraft passes over
known position reference.)
Observe HOLD pushbutton
light illuminates. Latitude
and longitude in data displays freeze at values present
when HOLD pushbutton
was pressed.
6. Load latitude by pressing
keys in sequence, starting
with “N“ or “S” to indicate
north or south. Example:
42°54.0’ North = N 4 2 5 4
0. Observe latitude appears
in left data display as keys
are pressed.
7. INSERT/ADVANCE pushb u t t o n - Press Observe
INSERT/ADVANCE
and
HOLD pushbuttons remain
illuminated.
8. Load longitude by pressing
keys in sequence, starting
with “W” on “E” indicating
w e s t o r east. Example:
87°54.9’ West = W 8 7 5 4
9.
9. INSERT/ADVANCE pushbutton - Press. Observe
INSERT/ADVANCE
and
HOLD pushbutton lights
remain illuminated.
3-67
TM 55-1510-221-10
10.
INSERT/ADVANCE pushbutton - Press. Observe arcseconds related to present
position latitude and longitude, to nearest tenth of a
second, appear in left and
right data displays, respectively.
11.
Load related arc-second values for latitude in sequence,
starting with “N“. Example:
35.8° North = N 358.
12.
INSERT/ADVANCE pushbutton - Press. Observe
INSERT/ADVANCE
and
HOLD pushbuttons remain
illuminated.)
13.
Load related arc-second values
for
longitude in
sequence, starting with “E“.
Example: 20.1° East = E
201.
(c.)
UTM present position check and
update:
NOTE
UTM data may be loaded in any order
and, until final entry, a value may be
reloaded.
1.
Data selector - UTM POS.
Observe UTM coordinates
of present position appear in
data displays.
2.
HOLD pushbutton - Press
(when aircraft passes over
known position reference.)
Observe HOLD pushbutton
light illuminates. Coordinates in data display freeze
at values present when
HOLD pushbutton
was
pressed.
NOTE
NOTE
Extra precision values are added to normal values and normal displays are not
rounded off.
While HOLD pushbutton light is illuminated, TACAN, GPS and data link
updates are inhibited.
3.
Load zone and easting by
pressing keys in sequence,
starting with “E”. Example:
Zone 16, 425 km East =
E16 425. Observe zone and
easting in kilometers appear
in right data display as keys
are pressed.
4.
INSERT/ADVANCE pushbutton - Press. Observe
INSERT/ADVANCE pushbutton light remains illuminated.
5.
Load northing by pressing
keys in sequence, starting
with “N” or “S” to indicate
north or south hemisphere.
Example: North 4749 km =
N 4749. Observe northing
kilometers appear in left
data display as keys are
pressed.
6.
INSERT/ADVANCE pushbutton - Press. Observe
INSERT/ADVANCE pushbutton light remains illuminated.
NOTE
Normal latitude-longitude coordinates
must always be loaded prior to extra precision values.
NOTE
Directions “N” or “S“ and “E” or “W”
are established during normal coordinate
entry. Either key may be used to initiate
entry during extra precision loads and
values will be added to extra precision
values without affecting direction.
NOTE
It is characteristic of the computer display
routine to add 0.2 arc-seconds to any display of 59.9 arc-seconds. Value in computer is loaded by operator.
14. Proceed to step 6 in procedure: “Normal Geographic
Present Position Check and
Update.“
3-68
TM 55-1510-221-10
7. INSERT/ADVANCE pushbutton - Press. Observe extra
precision display related to
present position northing
and easting, to nearest
meter, appears in left and
right data displays, respectively.
8. Load extra precision easting
value by pressing keys in
sequence, starting with “E“.
Example: 297 m East = E
297. Observe easting meters
appear in right data display
as keys are pressed.
9. INSERT/ADVANCE pushbutton - Press. Observe
INSERT/ADVANCE pushbutton light remains illuminated.
10. Load extra precision northing value by pressing keys in
sequence, starting with “N“
or “S”. Example: 901 m
N o r t h = N 901. Observe
Northing meters appear in
left data display as keys are
pressed. The value is always
a d d e d t o n o r m a l value
regardless of which key
(N/S) is pressed to initiate
entry. Normal entry establishes the hemisphere.
NOTE
The “W” key may be used to initiate easting entries; however, the computer will
always interpret such entries as an “E“
input.
NOTE
The extra precision values are always
added to normal values.
NOTE
Any data inserted when HOLD pushbutton light is not illuminated will be
rejected by computer.
11.
INSERT/ADVANCE pushbutton - Press. Observe
and
INSERT/ADVANCE
HOLD pushbutton lights
remain illuminated. North
position error and east position error in kilometers will
appear in left and right data
displays, respectively.
12.
If WARN light illuminates,
proceed to step 13; otherwise proceed to step 9 in
procedure: “Extra Precision
Geographic Present Position
Check and Update.“
13.
Data selector - DSRTK/STS.
Observe action code “02“
and malfunction code “49“.
This indicates radial error
between loaded position and
INS position exceeds 62
kilometers. Operator must
evaluate possibility that INS
is in error or reference point
position is incorrect. It is
possible to force INS to
accept updated position by
setting data selector to UTM
POS and proceeding to step
10 of procedure: “Extra Precision Geographic Present
Check
Position
14.
If updating is to be rejected
- Press HOLD pushbutton.
and
HOLD
Observe
INSERT/ADVANCE pushbutton lights extinguish. INS
returns to normal operation.
(d.) Position update eradication:
NOTE
This procedure is not considered common. Its use is limited to those times
where an operational error has resulted in
an erroneous position fix.
1.
Data selector - DSRTK/STS.
2. Key “1“ - Press. Observe
INSERT/ADVANCE pushbutton light illuminates,
000001 appears in right data
display.
3. INSERT/ADVANCE pushbutton - Press. Observe
INSERT/ADVANCE pushbutton light extinguishes.
Within 30 seconds, data displays return to normal with
“0“ (normal inertial mode)
3-69
TM 55-1510-221-10
in last digit of right display.
AI will be set to approximately three times the number of hours in NAV.
(13.) Flight course changes.
(a.)
Manual flight plan change inser-
tion:
NOTE
The INS may be shut down, downmoded
to STBY or ALIGN mode, or operated in
the navigation mode after landing. The
determining factor in choosing course of
action is expected length of time before
the next takeoff.
1.
WYPT CHG pushbutton Press. Observe WYPT CHG
INSERT/ADVANCE
and
pushbuttons illuminate.
2. Select new FROM and TO
waypoints by pressing corresponding keys
3. WYPT CHG pushbutton Press.
Observe
new
waypoint numbers appear in
FROM-TO data displays as
keys are pressed.
NOTE
Selecting zero as FROM waypoint will
cause desired track to be defined by computed present position (inertial present
position plus fixes) and TO waypoint.
4. INSERT/ADVANCE pushbutton - Press. Observe
WYPT CHG and INSERT/
ADVANCE
pushbuttons
extinguish.
NOTE
Waypoint zero always contains ramp
coordinates if no manual flight plan
changes are made. When a manual flight
plan change is made, present position at
instant of insertion is stored in waypoint
zero.
(14.) After landing procedures:
If INS will be unattended for an extended
period, it should be shut down.
Do not leave INS operating unless aircraft
or ground power and cooling air are available to system.
3-70
NOTE
Do not tow or taxi aircraft during INS
alignment. Movement during alignment
requires restarting alignment.
(a.) Transient stops.
NOTE
Action to be taken during a transient stop
depends upon time available and on
availability of accurate parking coordinates (latitude and longitude.)
1. Realignment - INS operating.
(Recommended if sufficient time and accurate
parking coordinates are available.)
NOTE
INS can be downmoded to perform a
realignment and azimuth gyro calibration.
Alignment to produce an alignment state
number of “5” can be accomplished in
approximately 17 minutes. During the 17
minute period, an automatic azimuth
gyro recalibration is determined on basis
of difference between inertial present
position before downmoding and inserted
present position. To obtain further refinement of azimuth gyro drift rate, calculated on basis of newly computed error
data, INS can be left in alignment mode
for a longer period, allowing the alignment state number to attain some value
lower than “5“.
2. Data selector - STBY, then
to ALIGN.
3. Present position coordinates
- Insert, according to procedure: “Geographic Present
Position
Insertion” or
“ U T M P r e s e n t Position
TM 55-1510-221-10
4. Realignment - INS shutdown. Perform complete
alignment procedures.
5. Position update. Recommended if time is not available for realignment.
NOTE
Perform position update using parking
coordinates in accordance with procedure: “Insertion of Geographic Waypoint
Coordinates.“ If parking coordinates are
not available, proceed as follows:
Continue operation in NAV, if INS accuracy appears acceptable.
Perform position update using best estimate of parking coordinates.
6. Downmoding to standby:
Mode selector - STBY.
NOTE
INS can be downmoded to standby operation which will maintain navigation unit
at operating temperature with gyro wheels
running. INS is downmoded to standby as
follows:
Do not leave INS operating unless aircraft
or ground power and cooling air are available to system.
7. Shutdown: Mode selector OFF.
NOTE
e. Abnormal Procedures:
(1.) General. INS contains self-testing features which provide one or more warning indications when a failure occurs. The WARN light on th
CDU provides a master warning for most malfunctions occuring in the navigation unit. Malfunctions
in the MSU or CDU will normally be obvious
because of abnormal indications of displays and
lights. A battery unit malfunction will shut down
INS when battery power is used.
(2.) Automatic INS shutdown.
(a.) Overtemperature. An overtemperature in navigation unit will cause INS to shut
down (indicated by blank CDU displays) when
mode selector is at STBY or ALIGN during ground
operation. The WARN light on CDU will illuminate
and will not extinguish until mode selector is rotated
to OFF. The cooling system should be checked and
corrected if faulty. If cooling system is satisfactory,
navigation unit should be replaced.
(b.) Low battery charge. A low battery unit charge will cause INS to shut down when
INS is operating on battery unit power. Both WARN
light on CDU and BAT light on MSU will illuminate and not extinguish until the mode selector is set
to OFF. The battery unit should be replaced when
this failure occurs.
(c.) Interpretation of failure indications. It is important to be able to correctly interpret
failure indications in order to take effective action.
Failure indications are listed below under two main
categories: WARN light illuminated, and WARN
light extinguished. Under each of these categories
are listed other indications which will give the operator sufficient information to take action.
1. WARN light illuminated.
Take the following action:
a. If action codes 01, 02,
03, 04, 05 are displayed
- See table 3-2.
b. No action or malfunction codes displayed computer
Indicated
failure.
Improper displays Indicates NU computer
failure.
2. WARN light extinguished. If
CDU displays are blank,
incorrect or frozen - CDU
failure is indicated.
c.
INS will retain inertial present position
data computed at time INS is downmoded. This value is compared with present position inserted for next alignment
and difference is used to determine azimuth gyro drift rate.
3-71
TM 55-1510-221-10
NOTE
It is not possible to load displays from the
keyboard. A temporary failure of a
numerical key may prevent data loading.
If a number cannot be loaded into latitude or longitude displays, after pressing/
wiggling the key several times, the cause
may be the momentary hang-up of
another key. To identify the faulty key,
rotate the data selector to DSRTK/STS.
The right digit on right display will indicate suspect key. Press and release suspect
key several times. To test whether the
keyboard problem is corrected, try pressing any other numerical key. Its number
should now appear as the right digit. If
this test is successful, press the CLEAR
key and return data selector to original
data loading position. Otherwise, a CDU
failure is indicated.
minated.
(d.)
data selector is being
rotated) problem is
normally in the navigation unit.
b.
If displays respond normally to the data selector, the problem is normally absence of 115V
AC power to INS.
(e.) For corrective action: Check to
assure proper settings of following switches and circuit breakers essential to INS operation:
1. Overhead circuit breaker
panel (fig. 2-26) - Circuit
breakers in:
a.
AVIONICS MASTER
CONTR
b. INS CONTROL
c.
CDU BAT indicator light is illu-
AVIONICS MASTER
PWR No.1
d. AVIONICS MASTER
PWR No.2
Operation on battery is an indication that
there may may be no aircraft power to
blower motor with resultant loss of cooling. The INS can operate only a limited
time (normally 15 minutes) on battery
power before a low voltage shutdown will
occur. Then, immediate corrective action
must be taken.
2. Overhead control panel (fig.
2-12): INVERTER No. 1 or
INVERTER No.2 switch ON (either).
3. Mission control panel (fig.
4-1):
a.
b. Bus cross tie switch ON/AUTO.
NOTE
CDU BAT indicator will illuminate for
12 seconds in alignment State “8“ (about
5 minutes after turn-on). This is normal
and indicates a battery test is in progress.
No corrective action is required.
NOTE
During ground operation, it is recommended that operation on battery power
not exceed 5 minutes.
1.
T o d e t e r m i n e corrective
action: (Monitor CDU displays while rotating the
CDU selector switch.)
a.
3-72
If displays are frozen
(do not change while
#1 INV or #2 INV
switch - ON.
4.
Mission AC/DC Power Cabinet (fig. 2-2): INS AC PWR
circuit breaker - In.
NOTE
CDU BAT indicator should extinguish
after above corrective action. If it remains
illuminated, INS will eventually shut
down when battery voltage drops below
approximately 19VDC. Flight crew
should prepare for shutdown.
(3.) Malfunction indications and procedures: Table 3-3 details the procedure for a Malfunction Code Check. Table 3-4 lists a number of malfunction indications which occur under given modes
of operation. Follow procedure given. Table 3-5
details action codes and recommended action. Table
TM 55-1510-221-10
3-6 lists failed test symptoms by malfunction codes
and lists codes for recommended actions.
Table 3-3. Malfunction Code Check
step
Indication
Control
Operation
Indicator or
Display
Indication
1
2
Data Selector
Rotate to DSRTK/STS
WARN light
RH data display
3
TEST switch
Press and release
RH data display
4
Repeat step 3 repeatedly, recording all malfunction codes until second and third digits again indicate
an action code or go blank. Refer to Table 3-4 for action codes and recommended action and to Table
3-5 for malfunction code definition.
If WARN light extinguishes and two digits go blank, failure was intermittent and has been cleared. If digits do not go blank, perform action according to displayed recommended action code.
5
Lights
Action code second and third
digits
Lowest number malfunction code
which has occurred since this
procedure was performed replaces
action code.
3-73
TM 55-1510-221-10
Table 3-4. Malfunction Indications and Procedures
Mode of
Operation
STBY or ALIGN
Malfunction
Indication
WARN on, CDU blank
(DIM control clockwise),
MSU BAT off
Procedure
1. Rotate MSU mode
selector OFF.
Probable Cause
Automatic
shutdown
caused by
over-temperature.
2. Check aircraft
cooling system and
correct if faulty.
3. Realign INS.
STBY, ALIGN, NAV
WARN on, MSU BAT on,
CDU blank.
1. Rotate MSU mode
selector OFF.
2. Insure all switches
and circuit breakers
applicable to INS
operation are set
properly.
3. If in flight, rotate
MSU mode selector
OFF.
4. If on ground, replace
battery unit. Battery
unit test may be bypassed by rotating
mode selector to
OFF, then to NAV
and reloading
position coordinates.
When INS advances
to alignment State 7
(PI=7) rotate mode
selector to ALIGN.
STBY, ALIGN, or NAV
3-74
WARN on, CDU is
operating
Perform Malfunction
Code Check as
described in Table 3-2.
Loss of INS
power and low
battery
Unit (BU).
Navigation
failure or
interfacing
system problem.
TM 55-1510-221-10
Table 3-5. Action Codes and Recommended Actions
Code
Recommended Action
01
02
Shut down INS.
Watch for degradation (NAV), During ground operation, downmode to STBY and restart
alignment.
INS may be used for navigation. One or more analog outputs are not functioning properly. Check
26 VAC circuit breakers, HSI and autopilot.
Downmode to STBY and restart alignment (ground operation only).
Correct problem in interfacing system (could be INS). Will not seriously affect performance.
03
04
05
3-75
TM 55-1510-221-10
Table 3-6. Recommended Codes
Malf
Code
10
11
12*
13
14
15
16
17
16
20*
22*
23*
24*
25*
26*
27*
31
32
33
34
35
36
37
38
40
42
44
45
47
49
51*
54*
57
59
60
62
63
Failed Test
Invalid heading
GR/CS program pin connected in error
Canned altitude profile in use (input altitude invalid)
Y velocity change
X velocity change
Torque limited
Invalid pitch and roll
Invalid magnetic heading
Excessive saturation time
Bearing to waypoint
Bearing to waypoint
Drift angle
Steering converter
True heading converter
XTK converter
Tick mark to fast
Ground speed
Memory parity
Azimuth stabilization loop
Inner roll stabilization loop
Pitch stabilization loop
Accelerometer loop
Z platform overtemperature
XY platform overtemperature
Heading error
Drift angle !gt!45°
Azimuth gyro drift ratye
Gyro scale factor or loaded altitude
15-second loop
Fix measurement too large
Excessive wind
Incomplete conversion from UTM to L/L
XY platform rotation rate
600 millisecond loop
X or Y sample and hold change
XY platform rotation rate
CDU self-checks
*Failed test does not illuminate WARN light on CDU.
3-76
Modes of
Operation
ALIGN
ALIGN
ALIGN, NAV
NAV
NAV
ALIGN, NAV
ALIGN, NAV
ALIGN, NAV
ALIGN
ALIGN, NAV
ALIGN, NAV
ALIGN, NAV
ALIGN, NAV
ALIGN, NAV
ALIGN, NAV
STBY
NAV
STBY, ALIGN,
ALIGN, NAV
ALIGN, NAV
ALIGN, NAV
ALIGN, NAV
NAV
NAV
ALIGN
NAV
ALIGN
ALIGN
NAV
NAV
ALIGN, NAV
STBY, ALIGN,
ALIGN
STBY, ALIGN,
ALIGN
NAV
STBY, ALIGN,
Recommended
Action Code
NAV
NAV
NAV
NAV
04
01
05
02
02
02
05
05
04
03
03
03
03
03
03
01
02
02
01
01
01
01
01
01
04
02
02
04
02
02
05
05
02
02
04
02
02
TM 55-1510-221-10
3-30A. GLOBAL POSITIONING SYSTEM
(AN/ASN-149, (B)3).
a . Description. Complete provisions are installed for a global positioning system (GPS). The
GPS is used to provide updated position information to the inertial navigation system. The GPS system consists of a control/display unit, receiver, observer headset and GPS key panel, antenna
electronics unit, and an antenna.
ed by the data select switch. If more than one page
is available a double arrow is displayed in the lower
right comer of the display. Pressing the slew key will
access the next page. Repeated pressing of the slew
key will return the display to the first page after the
last page has been accessed.
(i.) Data selector switch. For all data
selector switch positions there are two modes of displayed data:
(1.) Control/display unit (CDU). The control/display unit (fig. 3-25A), located on the electronics rack in the cabin, accomplishes all display
and control functions necessary for the operation of
the GPS receiver.
(2.) Observer Headset and GPS Key Panel.
The observer headset and GPS key panel (fig. 325B), located on the electronics rack in the cabin,
contains a headset connector and GPS key and loading controls.
1.
Destination mode (active waypoint as
destination)
2. Waypoint (WP) examine mode (any
waypoint)
Pressing the WP key switches the CDU between
the two modes.
The 10 position data select switch is used to select the type of information to be displayed on the
CDU:
GPS Controls, Indicators, and Functions.
POS.
Position data is displayed.
(1.) GPS control/display unit (fig. 3-25A).
MSN.
Mission data is displayed.
(a.) Line selection keys. Four line selection keys, located to the left of the CDU display
screen, are used to initiate and terminate data entries, and to select various system options.
OPT.
Option data is displayed. Six pages
of information pertaining to the
GPS receiver are made available
when the OPT position is selected.
(b.) Display screen. System information is shown on the cathode ray tube display screen.
The display screen can show four lines of text with
13 alphanumeric characters on each.
STAT.
Status data is displayed.
VAR-DTM.
Magnetic variation and map datum
data is displayed.
ERR.
Error data is displayed.
WIND.
Wind data is displayed.
DIS-TG.
Distance and time to go data is displayed.
(d.) Display brightness control. A
control knob placarded BRT is provided to control
the brightness of the cathode ray tube display screen.
Clockwise rotation of the control increases brightness.
TRK-GS.
Track and ground speed data is displayed.
DTK-VA.
Desired track and vertical angle
data is displayed.
(e.) Data entry keys (0 through 9).
The data entry keys are used to enter alphanumeric
data.
(j.) Waypoint key. The waypoint
key, placarded WP, is used to enter and examine
waypoint data.
(f.) USE LTR key. The use letter
key, placarded USE LTR, is used to select alphabetic prompt in free format data entry. The USE LTR
key terminates alphabetic entry when pressed.
(k.) Mark key. The MARK key is
used for MARK and FREEZE functions.
b.
(c.) Mode selector switch. The fourposition mode selector switch, placarded PULL
OFF, INIT, NAV, and PULL TEST, is used to select
the operating mode of the GPS system.
(g.) Clear key. The clear key, placarded CLR, is used to clear erroneous data entry
and message displays.
(h.) Slew key. The slew key is used to
access additional pages within a data display select-
(2.) Observer Headset and GPS Key Panel
Controls, Indicators, and Functions fig. 3-25B).
(a.) Interphone hot microphone, normal, key radio switch. This switch, placarded INTPH HOT MIC-NORM-KEY RADIO allows selection of hot microphone intercom, normal, and key
radio positions.
Change 2
3-76A
TM 55-1510-221-10
Figure 3-25A. GPS Control/Display Unit (CDU)
3-76B
Change 2
TM 55-1510-221-10
Figure 3-25B. Observer Headset and GPS Key Panel
Change 2
3-76C
TM 55-1510-221-10
(b.) Observer headset connector. Alows connection of observer headset.
(c.) GPS Zeroize switch. Actuating
he guarded switch, placarded ZEROIZE GPS, will
declassify the GPS receiver.
(d.) GPS key connector. Connects
start. A quick start uses stored position, time, and
recent ephemeris information. A cold start is used
only when the GPS is unable to perform a normal
startup.
(a.) GPS Normal Start.
1.
GPS key.
(e.) Load switch. This pushbutton
switch initiates loading process.
(f.) Load status indicator light. Illuninates to indicate load status.
(g.) Dust cap. Covers GPS key conlector when not in use.
(h.) Observer microphone connector.
Connects observer microphone.
(i.) GPS time of day switch. T h i s
pushbutton switch is used to transmit GPS time of
day to Have Quick II radios.
c.
GPS System Modes of Operation.
(1.) Off Mode. W h e n t h e P U L L O F F
node has been selected, power is removed from the
system, except panel lighting.
Mode selector switch - INIT.
When built-in-test is complete the display will show
data corresponding to the
data selector switch position.
NOTE
Data display will not be illuminated for
about 30 seconds after GPS has been
turned on. Ensure that the display brightness control has been set to the full clockwise position to receive the INIT display,
then adjust as desired.
NOTE
If the GPS has been OFF for more than
30 seconds when INIT mode was selected,
the set will perform the initial built-in-test
which takes approximately 30 seconds.
2.
Data selector switch - POS.
If ENTER POS message is
displayed press line select
key 3 next to message. Position must be entered.
3.
(2.) Initialize Mode. When the INIT (initialize) mode has been selected, position and time,
estimates can be entered via the keypad. Waypoint
data may be entered and examined, and option selections made. No navigation functions can be performed.
Displayed position - Check.
Verify or enter new updated
position and altitude as required.
4.
(3.) Navigation Mode. Selection of the
NAV (navigation) mode enables normal GPS functions (satellite tracking and navigation), including
data transfer to and from other aircraft systems.
Data selector switch - TRKGS. Verify correct track and
groundspeed are displayed.
If not valid, enter correct
values.
5.
Slew key-Press. Enter current time, year, and day of
year on page 2.
NOTE
Critical memory and other circuits which
cannot be turned off remain powered by
batteries in the receiver.
(4.) Test Mode. Selection of the PULL
TEST mode initiates a full command test of GPS
user equipment for line replaceable unit (LRU) fault
identification and isolation.
d. GPS Normal Operation.
(1.) GPS Start Procedures. The GPS must
be initialized prior to being used for navigation.
There are three types of start: normal, quick, and
cold. A position estimate, time estimate, and almanac (or ephemeris) data are required for normal
3-76D
Change 2
NOTE
Prior to next step, ensure all required initialization data has been entered correctly, as they cannot be changed after selection of NAV mode.
6.
Mode selector switch - NAV.
GPS will begin to search fol
satellite signals.
TM 55-1510-221-10
memory is used, and the internal low power time source
(LPTS) is used to initialize
time.
NOTE
If COLD alternates with the figure of
merit display, the GPS is performing a
cold start.
7.
2.
As the GPS is acquiring satellites, position, time, and
velocity estimates can be
checked to ensure that they
are within startup error limits. If so, monitor STAT
page 1. If not, a normal start
is required.
3.
After SAT 4 is achieved with
good EPE (Estimated Position Error), and FM (Figure
of Merit) of FM3 or below,
check position, velocity, and
time.
4.
GPS is now ready for normal navigation.
Data selector switch - STAT.
NOTE
The number of satellites (SAT) being acquired and tracked can be observed. Estimated position error (EPE) and figure of
merit (FM) can be monitored. The GPS
will be ready for use when SAT 3 or SAT
4 is displayed on STAT page 1.
8. Select page 2 of STAT.
Check almanac age (ALM).
If greater than 5000 hours,
force a cold start.
(c.) GPS Cold Start.
9. While the GPS is acquiring
satellites, periodically check
STAT page 1 for SAT 3 or
SAT 4 message. Figure of
merit (FM) is another indication of a converging position fix and can be directly
monitored from page 1 of
any data selection, where
FM alternates with the system map datum and other
alerts.
10.
SAT 3 or SAT 4 should be
displayed within five minutes. If not, check that position,
time,
track, and
groundspeed have been entered correctly. Also check
that satellites are available.
If all information is correct
and satellites are available,
force a cold start.
(b.) GPS Quick Start.
1. Mode selector switch - Set to
NAV directly from OFF. After power-on test has been
completed, the GPS uses velocity estimates from the aircraft’s sensors (if available).
If velocity is not available
from the aircraft, zero velocity is assumed. If position
and time are not available
from the aircraft, the position estimate from GPS
1.
Mode selector switch - INIT.
2. Data selector switch - OPT.
e.
3.
Slew key - Select page 4.
4.
Enter 04 on line 1.
5.
Line select key 2 - Press next
to COLD START to initiate.
6.
Line select key 3 - Press next
to COLD START to clear
cold start message and resume normal display.
7.
Mode selector switch - NAV.
CHAALS Use of GPS and INS.
(1.) CHAALS Concept. CHAALS (Coherent High Accuracy Airborne Location System), is an
emitter location system that provides timely, high
accuracy locations required for targeting and to support emitter associations and battlefield situation assessment. CHAALS provides this capability through
coherent processing of differential doppler (DD) and
time difference of arrival (TDOA) information received at a ground facility from the aircraft.
CHAALS receivers aboard the aircraft will receive and digitize emitter signals. The data will be
transmitted over the data link to the GR/CS integrated processing facility (IPF). There, CHAALS
processors will perform the required computations
to produce accurate emitter locations. The precise
navigation required will be provided by the inertial
navigation system (INS) and the global positioning
system (GPS). GPS also provides the primary means
Change 2
3-76E
TM 55-1510-221-10
of time synchronizing the CHAALS receivers (signal
conditioners or SC’s) aboard the aircraft. A backup
for the GPS will be provided by the data link. The
resultant emitter reports will be sent to GR/CS by
CHAALS.
(2.) GPS (and INS) Involvement. The accurate and timely navigation (position and velocity)
is provided by integrating an INS with a GPS, and
integrating both (through a series of intermediaries)
with a CHAALS ground based navigation processor
(NP). The SC, data link, and CHAALS HSSP (High
Speed Signal Processor) from the communication
link. The critical airborne interfaces for CHAALS
navigation and time synchronization include the following:
1.
INS to GPS:
a.
3-76F
Acceleration
Change 2
b. Velocity
c.
Position
d. Altitude
2.
INS to CHAALS: Same as INS to GPS
3. GPS to CHAALS:
a.
Time mark pulse (time synchroniza,
tion)
b.
Navigation data block (position, veloci
ty, and time)
c.
Error state vector data block (9 element ESV, time)
d.
TM/covariance data block (time, TM
time, covariance)
e.
Status data block (status including
DOP’s and FOMN)
TM 55-1510-221-10
Section IV.
TRANSPONDER AND RADAR
3-31. WEATHER RADAR SET (AN/APN-215).
a. Description. The weather radar set fig.
3-25) provides a visual The weather radar set (fig.
3-25), provides a visual 120° around the nose of the
aircraft, extending to a distance of 240 nautical
miles. The presentation on the screen shows the
location of potentially dangerous areas, such as
thunderstorms and hailstorms, in terms of distance
and azimuth with respect to the aircraft. The radar
is also capable of ground mapping operations. The
weather radar set is protected by a 5-ampere
RADAR circuit breaker located on the overhead circuit breaker panel (fig. 2-26).
b. Controls/Indicators and Functions.
(1.) GAIN control. Used to adjust radar
receiver gain in the MAP mode only.
(2.) STAB OFF switch. Push type on/off
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 a maximum of 15 degrees up
or down from horizontal attitude of aircraft.
(5.) 60° switch. Push type on/off switch.
When activated, reduces antenna scan from 120° to
60 degrees.
(6.) TRACK switches. Momentary action
type switches. When activated, a yellow track line
extending from the apex of the display through top
range mark appears and moves either left or right,
depending on the switch pressed. The track line
position will be displayed in degrees in the upper
left comer 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 type on/off
switch. When activated, the last image presented
before pressing the switch is displayed and held. The
word HOLD will flash on and off in the upper left
comer 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.
(a.) OFF. Turns set off.
(b.) STBY. Places set in standby
mode. This position also initiates a 90-second
warm-up 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 opera-
tion.
(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 or bottom mode is reached, the opposite
switch must be pressed to further change the mode.
(10.) NAV switch. If pressed with the INS
operating and the weather radar operating in a
weather depiction mode, the screen will display INS
waypoints that are located within the range displayed and within the degree of coverage left or right
of the present heading of the aircraft.
(11.) BRT control. Used to adjust screen
brightness.
c.
Weather Radar - Normal Operation.
Do not operate the weather radar set
while personnel or combustible materials
are within 18 feet of the antenna reflector
When the weather radar set is operating,
high-power radio frequency energy is
emitted from the antenna reflector, which
can have harmful effects on the human
body and can ignite combustible materials.
3-77
TM 55-1510-221-10
Do not operate the weather radar set in a
confined space where the nearest metal
wall is 50 feet or less from the antenna
reflector. Scanning such surfaces may
damage receiver crystals.
(1.) Turn-on procedure: Function switch TEST or ON, as required. (Information will appear
after time delay period has elapsed.)
2.
MODE switches - Press and
release as required to select WX.
3.
BRT control - As required.
4.
TILT control - Adjust until
weather pattern is displayed.
Include the areas above and
below the rainfall areas to obtain
a complete display.
5.
MODE switches - Press and
release to select WXA. Areas of
intense rainfall will appear as
flashing red. These areas must be
avoided.
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.
(2.) Initial adjustment operating procedure:
1.
BRT control - As required.
2. MODE switches - Press and
release as required.
3. RANGE switches - Press and
release as required.
4.
TILT control - Move up or down
to observe targets above or below
aircraft level. The echo display
will change in shape and location
only.
NOTE
Refer to FM 1-30 for weather observation, interpretation and application.
(5.)
1.
(3.) Test procedure:
1.
Function switch - TEST.
(4.) Weather observation operating procedure:
1.
3-78
3.
Function switch - ON.
BRT control - As required.
4. GAIN control - As required to
present usable display.
BRT control - As required.
4. Screen - Verify proper display.
(The test display consists of two
green bands, two yellow bands,
and a red band on a 120-degree
scan. The word TEST will be displayed in the upper right comer.
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.)
Function switch - ON.
2. MODE switches - Press and
release as required to select MAP.
2. RANGE switches- Press and
release as required to obtain 80mile display.
3.
Ground mapping operating procedure:
(6.) Standby procedure: Function switch STBY.
(7.) Shutdown procedure: Function switch
- OFF.
(8.) Weather radar emergency operation.
Not applicable.
3-32. TRANSPONDER SET (APX-100).
a. Description. The transponder system
receives, decodes, and responds to interrogations
from Air Traffic Control (ATC) radar to allow aircraft identification, altitude reporting, position
tracking, and emergency tracking. The system
receives a radar frequency of 1030 MHz and transmits preset coded reply pulses on a radar frequency
of 1090 MHz at a minimum peak power of 200
watts. The range of the system is limited to line-ofsight.
TM 55-1510-221-10
Figure 3-26. Weather Radar Control-Indicator (AN/APN-215)
3-79
TM 55-1510-221-10
The transponder system consists of a combined
receiver/transmitter/ control panel (fig. 3-27) located
on the pedestal extension; a pair of remote switches,
one on each control wheel; and two antennas,
located on the underside and top of the fuselage.
The system is protected by the 3-ampere TRANSPONDER and the 35-ampere AVIONICS MASTER PWR No. 1 circuit breakers on the overhead
circuit breaker panel (fig. 2-26).
(5.) MASTER CONTROL. Selects system
operating mode.
(a.) OFF. Deactivates system.
(b.) STBY.
warm-up (standby) mode.
(2.) TEST-MON indicator. Illuminates to
indicate system malfunction or interrogation by a
ground station.
(3.) ANT switch. Selects desired antenna
for signal input.
(c.) NORM. Activates normal oper(d.) EMER. Transmits emergency
reply code.
(6.) STATUS ANT indicator. Illuminates
to indicate the BIT or MON fault is caused by high
VSWR in antenna.
(7.) STATUS KIT indicator. Illuminates to
indicate the BIT or MON fault is caused by external
computer.
(b.) DIV. S e l e c t s d i v e r s e ( b o t h )
(8.) STATUS ALT indicator. Illuminates
to indicate the BIT or MON fault is caused by the
altitude digitizer.
(c.) BOT. Selects lower antenna.
(9.) IDENT-MIC-OUT switch. Selects
source of aircraft indentification signal.
(a.)
TOP. Selects upper antenna,
antennas.
(4.) RAD TEST-OUT switch. E n a b l e s
reply to TEST mode interrogations from test set.
(a.) IDENT. Activates transmission
of identification pulse (1P).
Figure 3-27. Transponder Control Panel (AN/APX - 100)
3-80
system
ating mode.
b. Controls/Indicators and Functions.
(1.) TEST-GO indicator. Illuminates to
indicate successful completion of built-in-test (BIT).
Activates
TM 55-1510-221-10
(b.) MIC. Enables either control
wheel POS IDENT switch to activate transmission
of ident signal from transponder.
(c.) OUT. Disallows outgoing signal.
(10.) MODE 4 reply indicator light. Illuminates to indicate a reply has been made to a valid
Mode 4 interrogation.
(11.) MODE 4 AUDIO OUT switch. Selects
monitor mode for mode 4 operation.
(a.) AUDIO. Enables sound and sight
monitoring of mode 4 operation.
(b.) LIGHT. E n a b l e s m o n i t o r i n g
REPLY indicator for mode 4 operation.
(c.) OUT.
Deactivates
monitor
c.
Transponder - Normal Operation.
(1.) Turn-on procedure: MASTER switch STBY. Depending on the type of receiver installed,
the TEST/MON NO GO indicator may illuminate.
Disregard this signal.
(2.) Test procedure:
NOTE
Make no checks with the master switch in
EMER, or with M-3/A codes 7600 or
7700 without first obtaining authorization
from the interrogating station(s).
1.
2.
mode.
(12.) MODE 3/A code selectors. Select
desired reply codes for mode 3/A operation.
Allow set two minutes to warm
up.
Select codes assigned for use in
modes 1 and 3/A by depressing
and releasing the pushbutton for
each switch until the desired
number appears in the proper
window.
(13.) MODE 1 code selectors. Select desired
reply codes for mode 1 operation.
3.
Lamp indicators - Operate pressto-test feature.
(14.) MODE 4 TEST-ON-OUT switch.
Selects test mode of Mode 4 operation.
4.
M-1 switch - Hold in TEST.
Observe that no indicator lights
illuminate.
5.
M-1 switch - Return to ON.
6.
Repeat steps 4 and 5 for the M-2,
M-3/A and M-C mode switches.
7.
MASTER control - NORM.
tion.
8.
(15.) MODE 4 code control. Selects preset
mode 4 code.
MODE 4 code control - A. Set a
code in the external computer.
9.
MODE 4 AUDIO OUT switch OUT.
(a.) TEST. Activates built-in-test of
mode 4 operation.
(b.)
ON. Activates mode 4 operation.
(c.) OUT. Disables mode 4 opera-
(16.) M-C, M-3A, M-2, and M-1 switches.
Select test or reply mode of respective codes.
(3.) Modes 1, 2, 3/A, and/or 4 operating
procedure.
(a.) TEST. Activates self-test of
selected code. Transponder can also reply
(b.) ON. Activates normal operation.
(c.) OUT. Deactivates operation of
selected code.
(17.) MODE 2 code selectors. Select desired
reply codes for Mode 2 operation. The cover over
mode select switches must be slid forward to display
the selected mode 2 code.
(18.) POS IDENT pushbutton (control
wheels, jig. 2-17). When pressed, activates transponder identification reply.
NOTE
If the external security computer is not
installed, a NO GO light will illuminate
any time the Mode 4 switch is moved out
of the OFF position.
1.
MASTER control - NORM.
2. M-1, M-2, M-3/A, and/or MODE
4 ON-OUT switches - ON. Actuate only those switches corresponding to the required codes.
The remaining switches should be
left in the OUT position.
3-81
TM 55-1510-221-10
3.
MODE 1 code selectors - Set (if
applicable).
4.
MODE 3/A code selectors - Set (if
applicable).
5.
MODE 4 code control - Set (if
required).
6.
MODE 4 REPLY indicator Monitor to determine when transponder set is replying to a SIF
interrogation.
7.
MODE 4 AUDIO OUT switch Set (as required to monitor Mode
4 interrogations and replies).
8.
MODE 4 audio and/or indicator Listen and/or observe (for Mode 4
interrogations and replies).
9.
IDENT-MIC-OUT switch - Press
to IDENT momentarily.
10.
MODE 4 TEST-ON-OUT switch
- TEST.
11.
Observe that the TEST GO indicator light illuminates.
12.
MODE 4 TEST-ON-OUT switch
- ON.
13.
ANT switch - BOT.
14.
Repeat steps 4, 5, and 6. Observe
that the TEST GO indicator illuminates.
15.
TOP-DIV-BOT-ANT switch TOP.
16.
Repeat step 14.
17.
TOP-DIV-BOT-ANT switch DIV.
18.
Repeat step 14.
19.
When possible, obtain the cooperation of an interrogating station
to exercise the TEST mode. Execute the following steps:
a.
RAD TEST-OUT switch RAD TEST.
identification-position replies while operating in
code Modes 1, 2, and/or 3/A, in response to ground
station interrogations. This type of operation is initiated by the operator as follows:
1.
Modes 1, 2, and/or 3/A - ON, as
required.
2. IDENT-OUT-MIC switch - Press
momentarily to IDENT, when
directed.
NOTE
Holding circuits within the transponder
receiver-transmitter will transmit identification-position signals for 15 to 30 seconds. This is normally sufficient time for
ground control to identify the aircraft’s
position. During the 15 to 30 second
period, it is normal procedure to acknowledge via the aircraft communications set
that identification- position signals are
being generated.
NOTE
Set any of the M1, M2, M3/A, M-C, or
MODE 4 switches to OUT to inhibit
transmission of replies in undesired
modes.
NOTE
With the IDENT-OUT-MIC switch set to
the MIC position, the POS IDENT button must be depressed to transmit identification pulses.
(5.) Shutdown procedure:
1.
To retain Mode 4 code in external
computer during a temporary
shutdown:
a.
MODE 4 CODE switch Rotate to HOLD.
b. Wait 15 seconds.
c.
MASTER control - OFF.
2.
To zeroize the Mode 4 code in the
external computer turn MODE 4
CODE switch to ZERO.
3.
RAD TEST-OUT switch OUT.
MASTER control - OFF. This will
automatically zeroize the external
computer unless codes have been
retained (step 1 above).
(4.) Transponder set identification-position
operating procedure: The transponder set can make
d. Transponder - Emergency Operation. Not
applicable.
b. Obtain verification from
interrogating station that a
TEST MODE reply was
received.
c.
3-82
TM 55-1510-221-10
Figure 3-28. Pilot’s Altimeter Indicator (BA-141)
3-33. PILOT’S ALTIMETER INDICATOR.
The pilot’s altimeter, on the upper left side of
the instrument panel (fig, 3-28), is a servoed unit
under control of the Air Data Computer and is part
of the Flight Director/Autopilot system. Altitude is
displayed by a 10,000 foot counter, a 1000 foot
counter, a 100 foot counter, and a single needle
pointer (coupled with the 100 foot counter) which
indicates hundreds of feet on a circular scale in 20
foot increments. Below an altitude of 10,000 feet, a
diagonal striped symbol will appear on the 10,000
foot counter. The barometric pressure knob allows
ground supplied pressure values to be adjusted and
displayed in inches Hg or millibars. If AC power to
the altimeter is lost, a warning OFF flag will appear
in the upper counter drum display window to indicate power loss, unreliable altimeter readings, and
possible loss of encoder transmissions to ground stations. Circuits are protected by a 3-ampere fuse in a
junction box.
When the BARO knob is adjusted to ground
supplied instructions, the updated altitude pressure
is routed to the Air Data Computer. The ADC
recomputes all data on hand, sends corrected altitude pressure information to the Flight Director and
autopilot, servo commands to correct the display on
the pilot’s altimeter, and supplies altitude informa-
tion to the transponder (for transmission to .
a.
Controls/Indicators and Functions.
(1.) ALT alert annunciator. Illuminates
when aircraft is within 1000 feet of preselected altitude during capture maneuver and extinguishes
when aircraft is within 250 feet of preselected altitude. After capture, light will illuminate if aircraft
departs more than 250 feet from the selected altitude.
(2.) Altitude counter drums. Indicates aircraft altitude in tens of thousands, thousands, and
hundreds of feet above sea level.
(3.) IN HG Indicator. Indicates local barometric pressure in inches of mercury. Adjusted by
use of BARO knob.
(4.) Needle indicator. Indicates aircraft
altitude in hundreds of feet with subdivisions at 20
foot increments.
(5.) BARO Knob. Used to manually set
barometric pressure displayed in the MB and IN
HG windows.
3-83
TM 55-1510-221-10
(6.) MB Indicator. Indicates local barometric pressure in millibars. Adjusted by use of
BARO knob.
(7.) Below 10,000 feet symbol. Presence
indicates aircraft altitude is below 10,000 feet.
(8.) OFF Flag. Presence indicates loss of
power to instrument and unreliable readings.
3-34. COPILOT’S ENCODING ALTIMETER.
Description. The copilot’s altimeter (fig.
3-29):’ provides an indication of present aircraft
pressure altitude above sea level. It also supplies
information to the INS and GPS. The air data computer supplies altitude information to the transponder.
b. Controls/Indicators and Functions.
NOTE
If the OFF flag is visible, either DC
power is off, the fuse has blown, or there
is an altimeter encoder failure. Since the
OFF flag monitors only the encoder input
to the altimeter and not transponder condition, the altitude reporting function
may be inoperative without the OFF flag
showing, in the case of transponder failure or improper control settings. It is also
possible to get a good Mode C test on the
transponder control with the OFF flag
showing. If the OFF flag remains visible,
radio contract should be made with a
ground radar site to determine if the altitude reporting function is operative.
b.
Pilot’s Altimeter - Normal Operation.
(1.) Turn-on procedure: Servoed altimeter
will operate when transponder is operating with
M-C switch set to center position.
(2.) Operating procedure:
1.
2.
Barometric set knob - Set desired
altimeter setting in IN. HG. window.
OFF flag - Check not visible.
3. Needle indicator - Check operation.
(1.) ALT alert indicator. Not used.
(2.) Needle indicator. Indicates aircraft
altitude in hundreds of feet with subdivisions at 20foot intervals.
(3.) MILLIBARS window. Indicates local
barometric pressure in millibars. Adjusted by use of
set knob.
(4.) IN HG window. Indicates local barometric pressure in inches of mercury. Adjusted by
use of set knob.
(5.) BARO knob. Used to manually set
barometric pressure displayed in the MB an IN HG
windows.
(6.) Drum indicator. Indicates aircraft altitude in ten-thousands, thousands, and hundreds of
feet above sea level.
(7.) Test button. Used to test altimeter
operation.
c.
Encoding Altimeter - Normal Operation.
(1.) Turn-on procedure: Encoding altimeter will operate when transponder is operating with
M-C switch set to center position.
(2.) Operating procedure:
1.
Barometric set knob - Set desired
altimeter setting in IN. HG. window.
2.
CODE OFF flag - Check not visible.
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 also nullities use of the altimeter for
IFR flight.
c. Pilot’s Altimeter - Emergency Operation.
Disregard pilot’s altimeter and utilize copilot’s
altimeter.
3-84
3. Needle indicator - Check operation.
4. TEST button a.
Push - Reading decreases by
500 feet.
b.
Release - Returns to original
reading.
TM 55-1510-221-10
1. Altitude alert annunciator
2. Altitude pointer
3. MILLIBARS barometric pressure counter
4. IN HG barometric pressure counter
5. BARO knob
6. TEST button
7. Counter drum display
AP 012930
Figure 3-29. Copilot’s Encoding Altimeter
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 also nullifies use of the altimeter for
IFR flight.
d. Encoding Altimeter - Emergency Operation.
Altimeter circuit breaker - Pull (if encoder fault
occurs).
3-85/(3-86 blank)
TM 55-1510-221-10
CHAPTER 4
Mission Equipment
Section I.
MISSION AVIONICS
4-1. MISSION AVIONICS COVERAGE.
Complete provisions only are installed for the
GPS, CHAALS and AQL mission systems. Equipment descriptions and operating instructions are to
be obtained from appropriate vendor and Army
Technical manuals.
4-2. MISSION CONTROL PANEL.
top section contains the mission caution/advisory
annunciator panel, see Table 2-8. The center section
contains one DC volt/ammeter, two digital AC volt/
frequency meters, two AC digital load meters, one
antenna steering synchro control, and the antenna
steering mode selector switch. The bottom section
contains the mission equipment control switches
and the mission equipment circuit breakers.
The mission control panel (fig 4-l), mounted on
the copilot’s sidewall, consists of three sections. The
Section II.
AIRCRAFT SURVIVABILITY EQUIPMENT
4-3. M-130 FLARE AND CHAFF DISPENSING
SYSTEM.
a. Description. The M- 130 flare and chaff dispensing system provides effective survival countermeasures against radar guided weapons systems and
infrared seeking missile threats. The system consists
of two dispenser assemblies with payload module
assemblies, a dispenser control panel, a flare dispense switch, two control wheel mounted chaff dispensing switches, an electronic module assembly,
and associated wiring. The flare and chaff dispensing system is protected by a 5-ampere circuit
breaker, placarded M130 POWER located on the
mission control panel (fig. 4-l).
Right engine nacelle dispenser is for chaff
only.
(1.) Dispenser assemblies. Two interchangeable dispenser assemblies are mounted on the
aircraft. One is located in the aft portion of the right
nacelle and the other is mounted on the right side of
the fuselage. On this aircraft the dispenser in the
nacelle will be used for chaff only while the dispenser mounted on the fuselage can be used for
either flares or chaff. The selector switch (placarded
C-F) on the dispenser can be set for either chaff or
flares. The unit also contains the sensor for the flare
detector. The dispenser assembly breech plate has
the electrical contact pins which fire the impulse cartridges. The unit also contains the sequencing mechanism.
(2.) Payload module assemblies. A removable payload module assembly is provided for each
dispenser assembly. Each payload module has 30
chambers which will accept either flares or chaffs.
Flares or chaffs are loaded into the rear-end (studded end) of the payload module, and secured in
place by a retaining plate.
(3.) Electronic module assembly (EM).
The electronic module assembly contains the programmer, the flare detector and a safety switch. The
unit is located behind the pilot’s seat.
(a.) Flare detector. The flare detector
is provided to insure that a flare is burning when it
is ejected from the dispenser payload module. If the
initial flare fails to ignite, the detector automatically
fires another flare within 75 milliseconds. If the second flare fails to ignite, the detector will fire a third
flare. If the third flare ignition is not detected, the
detector will not fire another flare until the system
is activated again by pressing the FLARE DISPENSE switch.
(b.) Programmer. The programmer is
used for the chaff mode only. It has four switches
for setting count and interval of salvo and burst.
(c.) Safety switch. The safety switch
(with safety pin and red flag) prevents firing of chaff
4-1
TM 55-1510-221-10
Figure 4- 1. Mission Control Panel
4-2
TM 55-1510-221-10
or flares when the safety pin is inserted. The safety
pin shall be removed only while the aircraft is in
flight or during test of the system.
(4.) Flare dispenser switch. A single pushbutton switch (fig. 4-2) placarded FLARE DISPENSE, located on the control pedestal, will fire a
flare from the dispenser payload module each time
it is pressed. If the FLARE DISPENSE switch is
held down, it will dispense a flare every 2.3 seconds.
(5.) Control wheel mounted chaff dispense
switches. Two pushbutton switches placarded
CHAFF DISP, one located on top left portion of the
pilot’s control wheel and the other located on the
top right portion of the copilot’s control wheel, activates the chaff dispensing system when pressed.
(6.) Wing mounted safety switch. A wing
mounted safety switch (with safety pin and red flag)
located on top of the right wing, just aft of the
nacelle, prevents the firing of chaff or flares when
the pin is inserted. This safety pin shall be inserted
while the aircraft is on the ground and removed
prior to flight or during system test.
(7.) Dispenser control panel (DCP). The
flare dispenser control panel (fig. 4-3) is mounted in
the control pedestal. Control functions are as follows:
(a.) RIPPLE FIRE switch. A guarded
switch placarded RIPPLE FIRE fires all remaining
flares when moved to the up position. It is used in
the event of an inflight emergency to dispense all
flares from the dispenser payload module.
(b.) FLARE counter setting knob.
Facilitates setting FLARE counter to the number of
flares in the payload module before flight.
(c.) FLARE counter. Indicates the
number of flares remaining in the dispenser payload
module.
(d.) ARM light. An amber press to
test indicator light placarded ARM illuminates when
the ARM-SAFE switch is in the ARM position,
when the safety pins are removed from the electronic module and the wing safety switch. Clockwise
rotation will dim the indicator light.
(e.) CHAFF counter. Indicates the
number of chaffs remaining in the payload module.
(f) CHAFF counter setting knob.
Facilitates setting CHAFF counter to the number of
chaffs in the payload module before flight.
SELECTOR
(g.) MAN-PGRM
SWITCH. Selects manual or programmed chaff dispense.
AP006537
Figure 4-2. Flare Dispense Switch
4-3
TM 55-1510-221-10
(h.) ARM-SAFE switch. When in the
SAFE position, power is removed from the M-130
system. When in the ARM position, power is
applied to the M-130 system.
1. MAN. Bypasses the programmer and fires one chaff each time one of the
chaff dispense switches is pressed.
2 . P G R M . Chaff is fired in
accordance with the preset chaff program as set into
the electronic module (count and interval of bursts
and salvo).
(i.) Ripple Fire Switch Cover. Prevents accidental switch activation.
(8.) Ammunition for dispenser. Ammunition for the system consists of countermeasure chaff
Ml, countermeasure flares M206, and impulse cartridges M796.
(a.) Countermeasure chaff M1. These
units consist of a plastic case 8 inches in length and
0.97 inches square. The base of the chaff case is
flanged to provide one-way assembly into the dispenser payload module. The chaff consists of aluminum coated fiberglass strands.
(b) Countermeasure flare M206.
These units consist of an aluminum case 8 inches in
length and 0.97 inches square. The base of the flare
is flanged to provide one-way assembly into the payload module. The flare material consists of a magnesium and teflon composition. A preformed packing
is required in the base of the flare unit prior to
inserting the impulse cartridge.
(c.) Impulse cartridge M796. This
cartridge fits into the base of either the flare or chaff
and is electrically initiated to eject flares or chaff
from the dispenser payload module.
b. Normal Operation.
NOTE
If aircraft is to be flown with flare dispenser assembly removed, fairing should
be removed from fuselage.
(1.) General. At the present time surfaceto-air intermediate range guided missiles that are
launched against the aircraft must be visually
detected by the aircraft crew. Crew members must
insure visual coverage over the ground area where a
missile attack is possible. The aircraft radar warning
system will only alert the pilot and copilot when the
aircraft is being tracked by radar-guided anti-aircraft
weapons systems. It will not indicate the tiring of
weapons against the aircraft.
4-4
(2.) Crew responsibilities. The pilot or designated crew member is responsible for removing
the safety pin from the right wing before flight, and
for replacing it immediately after flight. After the
aircraft is airborne the pilot is responsible for
removing the safety pin from the electronic module
and moving the ARM-SAFE switch on the dispenser
control panel to ARM. Before landing, he is responsible for re-inserting the safety pin in the electronic
module and moving the ARM-SAFE switch to
SAFE. While airborne the pilot and copilot are
responsible for scanning the terrain for missile
threats. When either pilot recognizes a missile
launch he will press the FLARE DISPENSE button
to eject flares.
(3.) Conditions for firing. The dispenser
system should not be fired unless a missile launch is
observed or radar guided weapons systems is
detected and locked on. If a system malfunction is
suspected, aircraft commander may authorize
attempts to dispense flares or chaff as a test in a
non-hostile area.
Aircraft must be in flight to dispense
flares.
(a.) Firing procedure.
1. Flares. Upon observing a
missile launch the pilot or copilot (whoever sights
the launch first) will fire a flare. If more than one
missile launch is observed, the firing sequence
should be continued until the aircraft has cleared the
threat area.
2_ Chaff Upon receiving an
alert from the aircraft radar warning system, the
pilot or copilot will fire the chaff and initiate an evasive maneuver. The number of burst/salvo and number of salvo/program and their intervals as established by training doctrine will be set into the
programmer prior to take-off (refer to TM 9-1095
206- 13 & P for information on setting programmer).
If desired, the operator may override the programmed operational mode and fire chaff countermeasures manually by moving the dispenser function selector switch to MANUAL and pressing a
dispenser switch.
(b.) Firing responsibility. When the
pilot or copilot observes a missile launch or radar
warning indication, he fires flares or chaff and
assumes command of the dispenser system, and fires
succeeding flares as required. He will advise the
other pilot that a missile launch has been observed
TM 55-1510-221-10
1. RIPPLE FIRE switch
2. FLARE counter setting knob
3. FLARE counter indicator
4. ARM light
5. CHAFF counter indicator
6. CHAFF counter setting knob
7. MANUAL-PROGRAM switch
8. ARM-SAFE switch
9 RIPPLE FIRE switch cover
AP 006768
Figure 4-3. Flare Dispenser Control Panel
4-5
TM 55-1510-221-10
or a radar warning signal has been received, and
announce that flares or chaff have been fired.
9. On DCP, set ARM-SAFE switch to ARM.
ARM lamp will illuminate.
NOTE
4-4. SYSTEM DAILY PREFLIGHT/RE-ARM TEST.
The following test procedures shall be conducted
prior to the first flight of each day and prior to each
re-arming of the dispensers. The first dispenser
tested shall be the one used to dispense flares and
the second one shall be the one used to dispense
chaff. Notify AVUM if any improper indications
occur during the tests.
When the test set is installed on the dispenser assembly and 28 volts DC aircraft
power has been applied, the sequencer
switch inside of dispenser assembly resets,
making an audible sound as it rotates.
There will be no such sound if the
sequencer switch has been previously
reset or if switch is in position 12 or 24.
NOTE
On test set, TS PWR ON lamp (clear)
illuminates and remains illuminated
throughout the test sequence until aircraft
power to test set (via test set power cable)
is disconnected or shut off.
Assure payload module is not connected
to dispenser assembly at any time during
the following test procedure.
1.
On flare dispenser assembly, assure the C-F
selector switch is in F (flare) position.
2.
Obtain M-91 test set and assure that TEST
SEQUENCE switch is in START/HOME
position.
3.
4.
5.
Connect base plate of test set to
dispenser assembly. Secure both
studs uniformly hand tight, using
hexagonal wrench provided in test
ing case.
breech of
mounting
5/32 inch
set carry-
Obtain test set power cable from the M-91
test set carrying case and connect cable
between exterior connection J 1 (28V DC) on
aircraft and aircraft power + 28V DC (J1)
of test set.
Remove safety pin from EM and in the top
skin of the right wing.
10.
11.
Set mission chaff program on EM.
Perform the following operations on the
M-91 test set:
a.
Press to test the remaining three lamps
on test set. Each lamp will illuminate.
NOTE
Replace any lamp that does not illuminate when pressed. If none of the indicating lamps illuminate, return test set to
AVUM.
b. Rotate TEST SEQUENCE switch
clockwise to the next position, TS
RESET. No visual indication will
occur.
c. Rotate TEST SEQUENCE switch
clockwise to SV SELF TEST position.
STRAY VOLTAGE lamp (red) will
illuminate.
On DCP, assure that RIPPLE FIRE
switch guard is in down position.
6. Provide aircraft power to DCP by setting
M- 130 POWER circuit breaker to ON position.
7.
On DCP, press ARM lamp. Lamp will illuminate. Release ARM lamp. Lamp will
extinguish.
8.
On DCP, set FLARE counter to 30 CHAFF
COUNTER to 30 and MAN-PGRM switch
to MAN position.
4-6
d. Rotate TEST SEQUENCE switch
clockwise to next position, TS RESET.
STRAY VOLTAGE lamp (red) will
extinguish.
e.
Rotate TEST SEQUENCE switch
clockwise to next position, STRAY
VOLT. STRAY VOLTAGE lamp (red)
should not illuminate.
f. Rotate TEST SEQUENCE switch
clockwise to next position, SYS NOT
RESET. SYS NOT RESET lamp
(amber) should not illuminate. If lamp
TM 55-1510-221-10
illuminates, press and release MANUAL SYSTEM RESET switch and SYS
NOT RESET lamp should then extinguish.
c.
d. Rotate TEST SEQUENCE switch
counterclockwise to START/HOME
position.
NOTE
When the MANUAL SYSTEM RESET
switch is pressed and released, and 28
volts DC power has been applied, the
sequencer switch inside the dispenser
assembly resets, making an audible sound
as it rotates. If the sequencer switch has
been previously reset or if the switch is in
position 12 or 24, there will be no such
sound.
g.
Rotate TEST SEQUENCE switch
clockwise to next position, DISP
COMP.
12. Press FLARE DISP switch once. For each
depressing, the FLARE counter on DCP
should count down in groups of three.
13.
On DCP, raise RIPPLE FIRE switch guard
and set toggle switch to up position until
FLARE counter counts down to 00. Return
switch guard to down position. On DCP,
reset FLARE counter back to 30. DISPENSER COMPLETE lamp (green) on test
set will illuminate.
14. Perform the following operations on the
M-91 test set:
a.
Rotate TEST SEQUENCE switch
counter-clockwise to SYS NOT RESET
position. SYS NOT RESET lamp
(amber) will illuminate. DISPENSER
COMPLETE lamp (green) will remain
illuminated.
b. Press and release MANUAL SYSTEM
RESET switch. SYS NOT RESET lamp
(amber) will extinguish.
NOTE
When the MANUAL SYSTEM RESET
switch is pressed and released, and 28
volts DC power has been applied, the
sequencer switch inside the dispenser
assembly resets, making an audible sound
as it rotates. If the sequencer switch has
been previously reset or if the switch is in
position 12 or 24, there will be no such
sound.
Rotate TEST SEQUENCE switch
counterclockwise to STRAY VOLT
position. STRAY VOLTAGE lamp
(red) should not illuminate.
NOTE
When the TEST SEQUENCE switch is
turned to the START/HOME position,
the DISPENSER COMPLETE lamp will
extinguish, the STRAY VOLTAGE lamp
will illuminate and then will extinguish
when passing through the TS RESET
position.
15.
On CHAFF dispenser assembly, assure that
C-F selector switch is in C (chaff) position.
16. Remove M-91 test set from first dispenser
assembly.
17.
Connect M-91 test set to breech assembly of
second dispenser assembly. Secure both
mounting studs uniformly hand tight using
ball hexagonal key screwdriver provided in
test set carrying case.
NOTE
When the test set is installed on the dispenser assembly and 28 volts DC aircraft
power has been applied, the sequence
switch inside the dispenser assembly
resets, making an audible sound as it
rotates. There will be no such sound if the
sequencer switch has been previously
reset or if switch is in position 12 or 24.
NOTE
On test set, TS PWR ON lamp (clear)
illuminates and remains illuminated
through the test sequence until aircraft
power to test set (via test set power cable)
is disconnected or shut off.
18. Perform the following operations on the
M-91 test set:
a.
Press to test all four lamps on test set.
Each lamp will illuminate.
4-7
TM 55-1510-221-10
CHAFF counter should decrease in accordance with the program set on the EM.
NOTE
Replace any lamp that does not illuminate when pressed. If none of the indicating lamps illuminate, return test set to
AVUM.
22.
Repeatedly press other CHAFF DISPENSE
switch until CHAFF counter on DCP reads
00.
23.
On test set, observe DISPENSE COMPLETE lamp (green) is illuminated and then
perform the following operations:
b. Rotate TEST SEQUENCE switch
clockwise to TS RESET position. No
visual indication will occur.
c.
Rotate TEST SEQUENCE switch
clockwise to SV SELF TEST position.
STRAY VOLTAGE lamp (red) will
illuminate.
a.
Rotate TEST SEQUENCE switch
counter-clockwise to SYS NOT RESET
position. SYS NOT RESET lamp
(amber) will illuminate.
b.
Press and release MANUAL SYSTEM
RESET switch. SYS NOT RESET lamp
(amber) will extinguish.
d. Rotate TEST SEQUENCE switch
clockwise to next position, TS RESET.
STRAY VOLTAGE lamp (red) will
extinguish.
e.
NOTE
Rotate TEST SEQUENCE switch
clockwise to next position, STRAY
VOLT. STRAY VOLTAGE lamp (red)
should not illuminate.
When the MANUAL SYSTEM RESET
switch is pressed and released, and 28
volts DC power has been applied, the
sequencer switch inside the dispenser
assembly resets, making an audible sound
as it rotates. If the sequencer switch has
been previously reset or if the switch is in
position 12 or 24, there will be no such
sound.
f. Rotate TEST SEQUENCE switch
clockwise to next position, SYS NOT
RESET. SYS NOT RESET lamp
(amber) should not illuminate. If lamp
illuminates, press and release MANUAL SYSTEM RESET switch and SYS
NOT RESET lamp should then extinguish.
c.
NOTE
d. Rotate TEST SEQUENCE switch
counter-clockwise to START/HOME
position.
When the MANUAL SYSTEM RESET
switch is pressed and released, and 28
volts DC power has been applied, the
sequencer switch inside the dispenser
assembly resets, making an audible sound
as it rotates. If the sequencer switch has
been previously reset or if the switch is in
position 12 or 24, there will be no such
sound.
g. Rotate TEST SEQUENCE switch
clockwise to next position, DISP
COMPL.
19.
20.
21.
4-8
Press pilot CHAFF DISP switch once, Press
copilot CHAFF DISP switch once. On DCP,
for each depressing, the CHAFF counter
should count down by an increment of one.
On DCP, set MAN-PGRM switch to PGRM
position.
Press any one of CHAFF DISP switches in
aircraft. On DCP, the number shown on
Rotate TEST SEQUENCE switch
counter-clockwise to STRAY VOLT
position. STRAY VOLTAGE lamp
(red) should not illuminate.
NOTE
When the TEST SEQUENCE switch is
turned to the OFF position, the DISPENSER COMPLETE lamp will extinguish, the STRAY VOLTAGE lamp will
illuminate and then will extinguish when
the OFF position is reached.
24. Install safety pins.
25.
Disconnect test set power cable.
26.
Remove M-91 test set from dispenser assembly and restore in carrying case along with
the power cable and hexagonal wrench.
27. On DCP, set ARM-SAFE switch to SAFE
position.
28.
On DCP, reset CHAFF counter to 30.
TM 55-1510-221-10
29.
Disconnect aircraft power by pulling the 5
ampere M130 POWER circuit breaker
located on the mission control panel (fig.
4-l).
30.
Proceed immediately to ammunition loading
procedures.
4-5. AMMUNITION.
a.
Ammunition Loading Procedure.
The system must have been tested to
assure that there is no stray voltage and
all aircraft power must be removed from
the system prior to unloading the payload
module.
7. On the dispenser control panel, assure
ARM-SAFE switch is in SAFE position.
8. On the electronic module and right
wing assure safety pins and flag assemblies are installed.
Only one shipping container is to be
opened at a time. If a shipping container
has been opened and only partially emptied, the remaining contents will be
secured in the container with an appropriate type of packaging material or filler to
adequately prevent jostling. All munitions
in storage must be in their original shipping containers.
1.
Place payload module assembly on
work bench in approved safe area so
that the retaining plate is facing up.
2.
Remove retaining plate by unscrewing
two retaining bolts.
3.
Insert one flare (or chaff) at a time into
each chamber of payload module.
4. Remove plastic dust cap from each
chaff or flare.
9. Slide payload module assembly into
dispenser assembly and secure two stud
bolts, hand tight, using 5/32 inch hexagonal wrench.
b. Ammunition Unloading Procedure.
All aircraft power to the dispenser system
must be turned off prior to removal of
payload module from dispenser assembly.
Safety pin flag shall be installed in the
electronic module prior to landing and
the safety pin flag shall be installed in the
wing-mounted safety switch immediately
after landing.
1.
2.
Prior to insertion of an impulse cartridge,
be sure there is preformed packing in the
flare cartridge. (There will be no preformed packing in chaff cartridges.) Reinstall any preformed packing that is inadvertently removed with dustcap. The
loading of impulse cartridges into a flare
or chaff shall be accomplished one at a
time.
On dispenser control panel, assure
ARM-SAFE switch is in SAFE position.
Assure safety pin and flag are inserted
into electronic module and in the wing
mounted safety switch.
If there is an indication that a misfire
occurred, notify emergency ordnance disposal personnel for disposition and disposal.
Insert one impulse cartridge into each
flare (or chaff).
Remove module from dispenser assembly by unscrewing two stud bolts with
a 5/32 inch hexagonal wrench and sliding out of dispenser assembly.
6 . Install retainer plate assembly by
screwing to two retainer bolts into payload module.
4. Remove retaining plate from payload
module by unscrewing two retaining
bolts.
5.
3.
4-9
TM 55-1510-221-10
5. Remove expended and unexpended
impulse cartridges and flares (or chaff)
from payload module.
6. Repack unexpended items in original
containers and return to stores.
a. Radar Signal Detecting Set Control Panel
Functions (AN/APR-39(V)1) (fig. 4-4).
(1.)
PWR switch. Turns set on or off.
(2.)
SELF TEST switch. Initiates self test.
(3.) DSCRM switch. Turns discriminate
function on or off.
NOTE
It is not unusual for the case of a chaff
cartridge to crack when fired. It does not
effect performance of the item and should
not be reported as a malfunction.
(4.) AUDIO control. Adjusts audio level.
b. Radar Signal Detecting Set Indicator Functions (fig. 4-5).
(1.) MA indicator. Illuminates to indicate
the presence of an MA threat.
4-6. RADAR SIGNAL DETECTING SET (AN/APR39(V)1).
The radar signal detecting set (control panel, fig.
4-3) indicates the relative position of search radar
stations. Audio warning signals are applied to the
pilot’s and copilot’s headsets. The radar signal
detecting set is protected by the 7.5-ampere circuit
breaker placarded APR39, located on the mission
control panel (fig. 4-l). The associated antennas are
shown in figure 2-1. For operating instructions, refer
to TM I l-5841-283-20. Pattern # 1 self test, shall be
as shown in figure 4-4.
(2.) Display. Indicates relative position of
search radar stations.
(3.) BRIL control. Adjusts brilliance,
(4.) DA Y-NIGHT control. Rotate to adjust
intensity of display.
4-7. RADAR WARNING RECEIVER (AN/APR-44()
(V3).
The radar warning receiver (fig. 4-6) indicates
the presence of certain types of search radar signals.
1. POWER switch
2. SELF TEST switch
3. DESCRIMINATE function switch
4. AUDIO level control
AP 003891
Figure 4-4. Radar Signal Detecting Set Control Panel (AN/APR-39(V)1)
4-10
TM 55-1510-221-10
1.
2.
3.
4.
MA indicator
Display
BRIL control
DAY-NIGHT control
AP 005715
Figure 4-5. Radar Signal Detecting Set Indicator
1. Radar warning indicator
2. VOLUME control
3. POWER switch
AP 003892
Figure 4-6. Radar Warning Receiver Control Panel (AN/APR-44() (V3)
4-11
TM 55-1510-221-10
The radar warning receiver is protected by the 5-ampere circuit breaker placarded APR44, located on
the mission control panel (fig. 4-1). For operating
instructions, refer to TM 11-5841-291-12.
4-12
a. Radar warning indicator. Illuminates to
indicate the presence of an AI or SAM threat.
b.
VOLUME control. Adjusts volume.
c.
POWER switch. Turns set on and off.
TM 55-1510-221-10
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. Limits concerning maneuvers, weight, and
center of gravity are also covered in this chapter.
5-4. MINIMUM CREW REQUIREMENTS.
The minimum crew required for aircraft operation
is two pilots. 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.
Instruments which display operating limitations
are illustrated in figure 5-1. The operating limitations
are color coded on the instrument faces. Color coding of
each instrument is explained in the illustration.
5-6.
INSTRUMENT MARKING COLOR CODES.
Operating limitations and ranges are illustrated by
the colored markings which appear on the dial faces of
engine, flight, and utility system instruments. Red
markings indicate the limit above or below which
continued operation is likely to cause damage or shorten
life. The green markings indicate the safe or normal
range of operation. The yellow markings indicate the
range when special attention should be given to the
operation covered by the instrument. Operation is
permissible in the yellow range, but should be avoided.
White markings on the airspeed indicator denotes flap
operating range.
The blue marking on the airspeed indicator denotes best
rate of climb with one engine inoperative, at maximum
gross weight, maximum forward c.g., sea level standard
day conditions.
5-7. PROPELLER LIMITATIONS.
The maximum propeller overspeed limit is 2200
RPM. Propeller speeds above 2000 RPM indicate
failure of the primary governor. Propeller speeds above
2080 RPM indicate failure of both primary and
secondary governors. Torque is limited to 81% for
sustained operation above 2000 RPM.
5-8. STARTER LIMITATIONS.
The starters in this aircraft are limited to an
operating period of 30 seconds ON, then 5 minutes
OFF, for two starter operations. After two starter
operations the starter shall be operated for 30 seconds
ON, then 30 minutes OFF.
5-1
TM 55-1510-221-10
TORQUE
49% MAXIMUM BELOW 1600 RPM
2O-100% NORMAL OPERATING RANGE
100% MAXIMUM
123% TRANSIENT (5 SECONDS)
TURBINE TACHOMETER (N1 SPEED)
52% MINIMUM LOW IDLE
88% MAXIMUM REVERSE (ONE MINUTE)
101.5% MAXIMUM
100.1% MAXIMUM BELOW -48 C
102.6% TRANSIENT (10 SECONDS)
TURBINE GAS TEMPERATURE
400-75OºC NORMAL OPERATING RANGE
66OºC MAXIMUM LOW IDLE
750°C MAXIMUM CONTINUOUS
750°C MAXIMUM REVERSE (1 MINUTE)
850°C MAXIMUM TRANSIENT
1000ºC MAXIMUM STARTING (5 SECONDS)
PROPELLER TACHOMETER
1600-2000 RPM NORMAL OPERATING RANGE
1900 RPM MAXIMUM REVERSE (1 MINUTE)
2000 RPM MAXIMUM
2200 RPM TRANSIENT (5 SECONDS)
OIL TEMPERATURE AND PRESSURE
OIL TEMPERATURE SCALE
10-99ºC NORMAL OPERATING RANGE
0-99ºC CRUISE CLIMB AND RECOMMENDED SPEED
-40-99ºC STARTING. LOW IDLE. HIGH IDLE
99ºC MAXIMUM
104 C TRANSIENT (5 MINUTES)
OIL PRESSURE SCALE
60 PSI MINIMUM
60 TO 85 PSI. 49% TORQUE MAXIMUM
85-135 PSI NORMAL OPERATING ABOVE 21.000 FEET
85-105 PSI CAUTION RANGE BELOW 21.000 FEET
105-135 PSI NORMAL OPERATING BELOW 21.000 FEET
200 PSI MAXIMUM STARTING WITH COLD OIL
NOTE: + 10 PSI FLUCTUATIONS ARE ACCEPTABLE
Figure 5-1 Instrument Markings (Sheet 1 of 3)
5-2
Change 4
TM 55-1510-221-10
243 KIAS MAXIMUM (Vmo) (.47 MACH)
NOTE
MAXIMUM ALLOWABLE AIRSPEED (RED)
STRIPED) POINTER IS SELF ADJUSTING
WITH ALTITUDE
91 KIAS MINIMUM SINGLE-ENGINE
CONTROL SPEED (Vmca)
127 KIAS ONE-ENGINE INOPERATIVE
BEST RATE-OF-CLIMB (Vyse)
78-153 KIAS FULL FLAP OPERATING RANGE
198 KIAS MAXIMUM APPROACH FLAP
EXTENSION SPEED
PNEUMATIC PRESSURE
12-20 PSI NORMAL OPERATING RANGE
20 PSI MAXIMUM
PROPELLER DEICER AMMETER
14-18 AMPERES NORMAL OPERATION
R
B
G
APO12879
Figure 5-1. Instrument Markings (Sheet 2 of 3)
5-3
TM 55-1510-221-10
FUEL QUANTITY
O-265 LBS NO TAKEOFF RANGE
CABIN ALTIMETER AND DIFFERENTIAL PRESSURE
O-6.1 PSI NORMAL RANGE
6.1 PSI MAXIMUM
FLAP POSITION INDICATOR
40% TAKEOFF AND APPROACH
R
Y
G
APOO4768.3
Figure 5-1. Instrument Markings (Sheet 3 of 3)
5-4
TM 55-1510-221-10
5-9. AUTOPILOT LIMITATIONS.
WARNING
The RC-12H aircraft is certified with wingtip
pods installed. Should the pods be removed,
the autopilot system must be replaced with a
standard C-12D autopilot. Affected wiring must
also be changed.
a. An autopilot preflight check must be. conducted
and found satisfactory prior to each flight on which the
autopilot is to be used.
b. A pilot must be seated at the controls with the
seat belt fastened when the autopilot is in operation.
c. Operation of the autopilot and yaw damper is
prohibited during takeoff and landing, and below 200
feet above terrain. Maximum speed for autopilot
operation is 243 KIAS/O.47 Mach.
d. During a coupled ILS approach do not operate
the propellers in the 1750 to 1850 RPM range.
5-10. FUEL SYSTEM LIMITS.
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.
a. Operating Limits. Operation with FUEL
PRESS light on is limited to IO hours. Log FUEL
PRESS light on time on DA Form 2408-13. One
standby boost pump may be inoperative for takeoff.
(Crossfeed fuel will not be available from the side with
the inoperative standby boost pump.) Operation on
aviation gasoline is time limited to 150 hours between
engine overhaul and altitude limited to 20,000 feet with
one standby boost pump inoperative. Crossfeed
capability is required for climb, when using aviation
gasoline above 20,000 feet.
b. Fuel Management. Auxiliary tanks will not be
filled for flight unless the main tanks are full. Maximum
allowable fuel imbalance is 400 lbs. Do not take off if
fuel quantity gages indicate in yellow arc (less than 265
lbs. of fuel in each main tank). Crossfeed only during
single engine operation.
c. Fuel System Anti-Icing. Icing inhibitor
conforming to MIL-I-27686 will be added to
commercial fuel, not containing an icing inhibitor,
during fueling operations, regardless of ambient
temperatures. The additive provides anti-icing
protection and also functions as a biocide to kill
microbial growth in aircraft fuel systems.
5-11. BRAKE DEICE LIMITATIONS.
The following limitations apply to the brake deice
system:
a. The brake deice system shall not be operated at
ambient temperatures above 15°C.
b. The brake deice system shall not be operated
longer than 10 minutes (one timer cycle) with the
landing gear retracted. If operation does not
automatically terminate approximately 10 minutes after
gear retraction, turn the brake deice switch OFF.
c. Maintain 85% N1 o r h i g h e r d u r i n g
simultaneous operation of the brake deice and surface
deice systems. If adequate pneumatic pressure cannot
be provided for simultaneous operation of the brake
deice and surface deice systems, turn OFF the brake
deice system.
d. In order to maintain an adequate supply of
systems pneumatic bleed air, the brake deice system
shall be turned OFF during single engine operation.
5-12. PITOT HEAT LIMITATIONS.
Pitot heat should not be used for more than 15
minutes while the aircraft is on the ground.
Change 4 5-5
TM 55-1510-221-10
Section ill. POWER LlMlTS
5-13. ENGINE LIMITATIONS.
Observe the following limitations (table 5-1) during
operation of this aircraft equipped with two Pratt and
Whitney of Canada, Ltd. PT6A-41 engines. Each
column is a separate limitation. The limits presented do
not
necessarily occur simultaneously. Whenever
operating limits are exceeded, pilot should record the
value and duration of the condition encountered in the
aircraft log. Operation of the engines is monitored by
instruments, with the operating limits marked on the
face of each instrument.
CAUTION
5-15. POWER DEFlNlTlONS FOR ENGINE
OPERATIONS.
The following definitions describe the engine
power ratings.
a. Takeoff Power. The maximum power available, limited to periods of five minutes duration.
b. Maximum Continuous Power. The highest
power rating not limited by time. Use of this rating is
intended for emergency situations at the discretion of
the pilot.
5-16. GENERATOR LlMlTS.
Engine operation using only the engine driven
fuel pump without boost pump fuel pressure is
limited to 10 cumulative hours. All time in this
category shall be entered on DA Form 2408-13
for the attention of maintenance personnel.
CAUTION
Use of aviation gasoline is time-limited to 150
hours of operation during any TimeBetween-Overhaul (TBO) period. It may be
used in any quantity with primary or alternate
fuel.
5-14. OVERTEMPERATURE AND OVERSPEED
LIMlTATlONS.
a. Whenever the limiting temperatures are
exceeded and cannot be controlled by retarding the
power levers, the engine will be shut down and a
landing made as soon as possible.
b. During engine operation, the temperatures,
speeds and time limits listed in the Engine Operating
Speeds Limitations chart (table 5-1) must be observed.
When these limits are exceeded, the incident will be
entered as an engine discrepancy in the appropriate
maintenance forms. It is particularly important to record
the amount and duration of over temperature and/or
overspeed.
c. Continuous engine operation above 725ºC will
reduce engine life.
5-6
Maximum generator load is limited to 100% for
flight and variable during ground operations. Observe
the limits shown in Table 5-2 during ground operation.
TM 55-1510-221-10
Table 5-1. Operating Limits
OPERATING
CONDITION
GAS
GENERATOR
RPM N1 (10)
RPM %
SHP
TORQUE
PERCENT
(1)
MAXIMUM
OBSERVED
TGT ºC
TAKEOFF (3)
850
100%
750
38,100
MAX CONT
850
100% (4)
750
MAX CLIMB
850
100% (4)
MAX CRUISE
HIGH IDLE
LOW IDLE
STARTING
TRANSIENT
MAX REVERSE(9)
-
123% (7)
-
PROP
RPM
N2
OIL
PRESS
PSI (2)
OIL
TEMP
ºC
101.5
2000
10 to 99
38,100
101.5
2000
725
38,100
101.5
2000
105 to
135
105 to
135
105 to
135
-
-
60(min)
105 to
135
-40 to 99
-40 to 99
-4O(min)
0 to 104 (3)
0 to 99
660 (6)
1000(7)
850
750
(5)
19,500
52(min)
38,500(8)
102.6(8) 2200(7)
1900
88
10 to 99
0 to 99
NOTES:
(1) Torque limit applies within range of 1600-2000 propeller RPM (N 2). Below 1600 RPM, torque is limited to 49%.
(2) Normal takeoff and maximum continuous operation oil pressure at gas generator speeds above 72% with oil
temperature between 60 and 71°C is 105 to 135 PSIG up to 21,000 feet. Above 21,000 feet, the minimum oil pressure
is 85 PSIG. Plus or minus 10 PSIG fluctuations are acceptable. Oil pressure between 60 and 85 PSIG should be
tolerated only for the completion of the flight at power setting not to exceed 49% torque. Oil pressure below 60 PSIG
are unsafe and require that either the engine be shut down or a landing be made as soon as possible, using the
minimum power required to sustain flight. During extremely cold starts, oil pressure may reach 200 PSI.
(3) These values are time limited to 5 minutes.
(4) Cruise torque values vary with altitude and temperature.
(5) At approximately 70% N1 .
(6) High TGT at ground idle may be corrected by reducing accessory load and/or increasing N 1 RPM.
(7) These values are time limited to 5 seconds.
(8) These values are time limited to 10 seconds.
(9) This operation is time limited to 1 minute.
(10) For every 5.6ºC below -48ºC ambient temperature, reduce maximum allowable N1 by 1.6%.
TABLE 5-2. GENERATOR LIMITS
GENERATOR LOAD
0 to 50%
50 to 80%
80 to 100%
MINIMUM GAS GENERATOR RPM - N1
WITHOUT AIR CONDITIONING
53%
60%
63%
*WITH AIR CONDITIONING
60%
65%
70%
*Right engine only
5-7
TM 55-1510-221-10
Section IV. LOADING LIMITS
5-17. CENTER OF GRAVITY LIMITATIONS.
Center of gravity limits and instructions for
computation of the center of gravity are contained
in Chapter 6. The center of gravity range will remain
within limits, providing the aircraft loading is
accomplished according to instructions in Chapter 6.
5-18. WEIGHT LIMITATIONS.
Maximum takeoff gross weight is 15,000
pounds. Maximum landing weight is 15,000 pounds.
Maximum ramp weight is 15,090 pounds. Maximum zero fuel weight is 11,500 pounds.
Section V.
The ability to sustain loss of engine power
and successfully stop, continue the takeoff, or climb before or after gear retraction is not assured for all conditions.
Thorough mission planning must be
accomplished prior to takeoff by analysis
of maximum takeoff weight permitted by
takeoff distance, accelerate-stop, positive
one-engine-inoperative climb at lift off,
accelerate-go, takeoff climb gradient, and
climb performance. These data will
describe performance capabilities for critical mission decisions.
AIRSPEED LIMITS, MAXIMUM AND MINIMUM
5-19. AIRSPEED LIMITATIONS.
5-23. WING FLAP EXTENSION SPEEDS.
Airspeed indicator readings contained in procedures, text, and illustrations throughout this Operator’s Manual are given as indicated airspeed (IAS).
Airspeed indicator markings (fig. 5-1) and placarded
airspeeds, located on the cockpit overhead control
panel (fig. 2- 12), are also indicated airspeeds.
The airspeed limit for APPROACH extension
(40%) of the wing flaps is 198 KIAS. The airspeed
limit for full DOWN extension (100%) of the wing
flaps is 153 KIAS. If wing flaps are extended above
these speeds, the flaps or their operating mechanisms may be damaged.
5-20. MAXIMUM ALLOWABLE AIRSPEED.
The maximum allowable airspeed is 243 KIAS/
0.42 Mach.
5-24. MINIMUM SINGLE-ENGINE CONTROL AIRSPEED (Vmc ).
5-21. LANDING GEAR EXTENSION SPEED.
Chapter 7, Section X describes minimum singleengine control airspeeds. The minimum singleengine control airspeed (Vmc ) at sea level standard
conditions is 91 KIAS.
The airspeed limit for extending the landing gear
and for flight with the landing gear extended is 180
KIAS.
5-22. LANDING GEAR RETRACTION SPEED.
5-25. MAXIMUM
SPEED.
The airspeed limit for retracting the landing gear
is 162 KIAS.
The maximum design maneuvering speed is 168
KIAS.
5-8
DESIGN
MANEUVERING
TM 55-1510-221-10
FLIGHT ENVELOPE CHART
Figure 5-2. Flight Envelope
Change
4
5-9
TM 55-1510-221-10
Section VI. MANUEVERING LIMITS
5-26. MANEUVERS.
For abrupt maneuvers above 168 KIAS. refer to Flight
Envelope (tip. 5-2).
4. Any maneuver which results in a positive load
factor of 3.06G's or a negative load factor of
1.224G’s with wing flaps up, or a positive load
factor of 2.0G’s or a negative 1224G’s with
wing flaps down.
a. The following maneuvers are prohibited.
1. Spins.
b. Recommended turbulent air penetration airspeed is
158 KIAS.
3. Aerobatics of any kind.
5-27. BANK AND PITCH LIMITS.
3. Abrupt maneuvers above 168 KIAS.
a. Bank limits are 60° left or right.
b. Pitch limits are 30° above or below the horizon.
Section VII. ENVIRONMENTAL RESTRICTIONS
5-28. ALTITUDE LIMITATIONS.
The maximum altitude that the aircraft may be operated
at is 31,000. When operating with inoperative yaw damp,
the altitude limit is 17,000 feet.
5-29. TEMPERATURE LIMITS.
a. The aircraft shall not be operated when the ambient
temperatures are warmer than ISA 37°C at SL to 25,000
feet. ISA 31°C above 25,000 feet.
b . Engine ice vanes shall be retracted at 1°C and above.
5-30. FLIGHT UNDER IMC (INSTRUMENT
METEOROLOGICAL CONDITIONS).
This aircraft is qualified for operation in instrument
meteorological conditions.
5-30A. ICING LlMlTATlONS (TYPICAL).
While in icing conditions, if there is an
unexplained 30% increase of torque needed
to maintain airspeed in level flight, a
cumulative total of two or more inches of ice
accumulation on the wing, an unexplained
decrease of 15 knots IAS, or an unexplained
deviation between pilot’s and copilot’s
airspeed indicators, the icing environment
should be exited as soon as practicable. Ice
5-10
Change 5
accumulation on the pitot tube assemblies
could cause a complete loss of airspeed
indication.
The following conditions indicate a possible
accumulation of ice on the pitot tube assemblies and
unprotected airplane surfaces. If any of these conditions
are observed, the icing environment should be exited as
soon as practicable.
1. Total ice accumulation of two inches or more on the
wing surfaces. Determination of ice thickness can be
accomplished by summing the estimated ice thickness on
the wing prior to each pneumatic boot deice cycle (e.g. four
cycles of minimum recommended 1/2-inch accumulation.
2. A 30 percent increase in torque per engine required
to maintain a desired airspeed in level flight (not to exceed
85 percent torque) when operating at recommended
holding speed.
3. A decrease in indicated airspeed of 15 knots after
entering the icing condition (not slower than 1.4 power off
stall speed) if maintaining original power setting in level
flight. This can be determined by comparing pre-icing
condition entry speed to the indicated speed after a surface
and antenna deice cycle is completed.
4. Any variations from normal indicated airspeed
between the pilot’s and copilot’s airspeed indicators.
TM 55-1510-221-10
5-30B. ICING LIMITATIONS (SEVERE).
WARNING
Severe icing may result from environmental
conditions outside of those for which the
airplane is certificated. Flight in freezing
rain, freezing drizzle, or mixed icing
conditions (supercooled liquid water and ice
crystals) may result in a build-up on
protective surfaces exceeding the capability of
the ice protection system, or may result in ice
forming aft of these protected surfaces. This
ice may not shed using ice protection systems,
and may seriously degrade the performance
and controllability of the airplane.
a. During flight, severe icing conditions that exceed
those for which the airplane is certficated shall be
determined by the following visual cues. If one or more of
these visual cues exists. immediately request priority
handling from air traffic control to facilitate a route or an
altitude change to exit the icing conditions:
(1) Unusually extensive ice accreted on the
airframe in areas not normally observed to collect ice.
(2) Accumulation of ice on the upper (or lower. as
appropriate) surface of the wing aft of the protected area.
5-31. CROSSWIND LIMITATION.
The maximum crosswind component is 25 knots at 90°.
The maximum angle of bank in a slip during landing is 8°.
Landing the aircraft in a crab will impose side loads on the
landing gear and should he recorded on the DA Form 2408
13-1. Refer to Chapter 8 for crosswind landing technique.
5-32. OXYGEN REQUIREMENTS.
A minimum ten minute supply of supplemental oxygen
shall be available during flight at or above an altitude of
25,000 feet based on the highest total aircraft oxygen flow
rates.
In addition to the supply required by the information
in the above paragraph. sufficient oxygen will be carried
for each flight. assuming a decompression will occur at the
altitude or point of flight that is most critical from the
standpoint of oxygen need. and that after decompression
the aircraft will descend. in accordance with the emergency
procedures. to a flight altitude that Will allow successful
termination of the flight. Following the decompression, the
cabin pressure altitude is considered to be the same as the
flight altitude.
An oxygen system data/duration table may be found in
Chapter 2.
5-33. CABIN PRESSURE LIMITS.
(3) Accumulation of ice on the propeller spinner
farther aft than normally observed.
b. Since the autopilot may mask tactile cues that
indicate adverse changes in handling characteristics. use of
the autopilot is prohibited when any of the visual cues
specified above exist. or when unusual lateral trim
requirements or autopilot trim warnings are encountered
while the airplane is in icing conditions.
NOTE
All icing detection lights must be operative prior
to flight into icing conditions at night. This
supersedes any relief provided by the master
minimum equipment list (MMEL) or equivalent.
Maximum cabin differential is 6.2 PSI.
5-34. CRACKED CABIN WINDOW / WINDSHIELD.
If a crack occurs in an outer cabin window. the aircraft
is limited to an altitude of 25,000 feet, and maximum cabin
pressure differential is limited to 4.6 PSI. Maximum
operating time with a crack in an outer cabin window is 20
hours. If an external windshield crack is noted. no action is
required in flight. If an external crack occurs in either
cabin window or the windshield, refer to Chapter 9.
Emergency Procedures.
Change 5
5-10.1/(5-10.2 blank)
TM 55-1510-221-10
Section VIII. OTHER LIMITATIONS
5-35. MAXIMUM DESIGN SINK RATE.
The maximum design sink rate is 500 feet per
minute.
5-36. INSTRUMENT LANDING SYSTEM LIMITS.
During ILS approach do not operate the propellers in the 1750 to 1850 RPM range.
5-37. INTENTIONAL ENGINE OUT SPEED.
Inflight engine cuts below the safe one-engine
inoperative speed (Vsse - 109 KIAS) are prohibited.
assumed to be the same as the OAT. The minimum
oil temperature graph (fig. 5-3) is provided for use as
a guide in preflight planning, based on known or
forecast operating conditions, to allow the operator
to become aware of operating temperatures where
icing at the fuel control could occur. If the plot
should indicate that oil temperatures versus OAT
are such that ice formation could occur during takeoff or in flight, anti-icing additive per MIL-I-27686
should be mixed with the fuel at refueling to insure
safe operation. In the event that authorized fuels
(Prist) are not available the limitation of this chart
apply.
5-38. LANDING ON UNPREPARED RUNWAY.
Except in an emergency, propellers should
be moved out of reverse above 40 knots
to minimize propeller blade erosion, and
during crosswind to minimize stress
imposed on propeller, engine and airframe. Care must be exercised when
reversing on runways with loose sand or
dust on the surface. Plying gravel will
damage propeller blades and dust may
impair the pilot’s forward visibility at low
aircraft speeds.
5-39. MINIMUM OIL TEMPERATURE REQUIRED
FOR FLIGHT.
Engine oil is used to heat the fuel upon entering
the fuel control. Since no temperature measurement
is available for the fuel at this point, it must be
Anti-icing additive must be properly
blended with the fuel to avoid deterioration of the fuel cell. The additive concentration by volume shall be a minimum of
0.060% and a maximum of 0.15%.
Approved procedure for adding anti-icing
concentrate is contained in Chapter 2,
Section XII.
JP-4 fuel per MIL-T-5624 has anti-icing
additive per MIL-I-27686 blended in the
fuel at the refinery and no further treatment is necessary. Some fuel suppliers
blend anti-icing additive in their storage
tanks. Prior to refueling, check with the
fuel supplier to determine if fuel has been
blended. To assure proper concentration
by volume of fuel on board, blend only
enough additive for the unblended fuel.
5-11
TM 55-1510-221-10
MINIMUM OIL TEMPERATURE REQUIRED FOR
OPERATION WITHOUT ANTI-lClNG ADDlTlVE
0
-60
-50
-40
-30
-20
FUEL TEMPERATURE (OAT) ~ ºC
-10
0
Figure 5-3. Minimum Oil Temperature
Section IX.
REQUIRED EQUIPMENT FOR VARIOUS CONDITIONS OF FLIGHT
5-40. REQUIRED EQUIPMENT LISTING.
a. A Required Equipment for Various Conditions of Flight listing (fig. 5-4) is provided to enable
the pilot to indentify those systems/components
required for flight. For the sake of brevity, the listing does not include obviously required items such
as wings, rudders, flaps, engines, landing gear, etc.
Also the list does not include items which do not
affect the airworthiness of the aircraft such as galley
equipment, entertainment systems, passenger convenience items, etc. However, it is important to note
that ALL ITEMS WHICH ARE RELATED TO
THE AIRWORTHINESS OF THE AIRCRAFT
AND NOT INCLUDED ON THE LIST ARE
AUTOMATICALLY REQUIRED TO BE OPERATIVE.
b. It is the final responsibility of the pilot to
determine whether the lack or inoperative status of
a piece of equipment on his aircraft will limit the
conditions under which he may operate the aircraft.
5-12
(-) Indicates item may be inoperative for the
specified flight condition,
(*) Refers to remarks and/or exceptions column
for explicit information or reference.
Numbered items contained in ( ) are
required for flights by AR 95-l.
c. The pilot is responsible for exercising the
necessary operational control to assure that no aircraft is flown with multiple items inoperative, without first determining that any interface or interrelationship between inoperative systems or components
will not result in a degradation in the level of safety
and/or cause an undue increase in crew workload.
d. The exposure to additional failures during
continued operation with inoperative systems or
components must also be considered in determining
that an acceptable level of safety is being maintained. The REL may not deviate from requirements
of the Operators Manual limitations section, emergency procedures or safety of flight messages.
TM 55-1510-221-10
SYSTEM and/or COMPONENT
ELECTRICAL POWER
1. AC Volts/Frequency Meter
2. Battery
3. Battery Charge Monitor System and Annuncia
tor
4. DC Generator
5. DC Generator Annunciator
6. DC Load Meter
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
2
2
2
2
2
2
2
2
2
2
2
2
2
2
2
7. lnverter
8. lnverter Annunciator
2
1
2
1
2
1
2
1
2
1
1. Bleed Air Fail Annunciators
2
2
2
2
2
2. Altitude Warning Annunciator (cabin)
1
1
1
1
1
3. Cabin Rate of Climb Indicator
4. Differential Pressure/Cabin Altitude Indicator
5. Duct Overtemp Annunciator
6. Outflow Valve
7. Pressurization Controller
8. Safety Valve
9. Bleed Air Shutoff Valve
1
1
1
1
1
1
2
1
1
1
1
1
1
2
1
1
1
1
1
1
2
1
1
1
1
1
1
2
1
1
1
1
1
1
2
2
2
2
2
2
1. Flap Position Indicator
1
1
1
1
1
2. Flap System
1
1
1
1
1
3. Stall Warning Horn
1
1
1
1
1
4. Trim Tab Position Indicator
(Rudder, Aileron, Elevator)
3
3
3
3
3
5. Yaw Damp System
1
1
1
1
1
One may be inoperative provided
corresponding loadmeter is moni
tored.
May be inoperative provided both
inverters are operative.
ENVIRONMENTAL
Provided bleed air is not used from
side of failed light.
May be inoperative provided airplane remains unpressurized.
FIRE PROTECTION
1. Engine Fire Detector System and Annunciator
FLIGHT CONTROLS
May be inoperative provided that
the flap travel is visually inspected
prior to takeoff.
May be inoperative provided that
the trim tabs are checked in the
neutral position prior to each take off
and checked for full range of operation.
May be inoperative for flight at and
below 17,000 feet.
Figure 5-4. Required Equipment Listing (1 of 3)
5-13
TM 55-1510-221-10
Figure 54. Required Equipment Listing (2 of 3)
5-14
TM 55-1510-221-10
Figure 5-4. Required Equipment Listing (3 of 3)
5-15/(5-16 blank)
TM 55-1510-221-10
CHAPTER 6
WEIGHT/BALANCE AND LOADING
Section I. GENERAL.
6-1. EXTENT OF COVERAGE.
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 RC-12H aircraft are in Class 1.
Additional directives governing weight and balance
Section II.
of Class 1 aircraft forms and records are contained
in DA PAM 738-751 and TM 55-1510-342-23.
6-3. AIRCRAFT COMPARTMENT AND STATIONS.
The aircraft is separated into two compartments
associated with loading. These compartments are the
cockpit and the cabin. Figure 6-l shows the general
description of aircraft compartments.
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.
6-5. CHARTS AND FORMS.
The standard system of weight and balance control requires 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 (fig. 6-2).
b. Form F - Weight and Balance Clearance
Form F, DD Form 365-4 (Tactical), fig. 6-3).
6-6. RESPONSIBILITY.
The aircraft manufacturer inserts all aircraft
identifying data on the title page of the booklet entitled WEIGHT AND BALANCE DATA and on the
various charts and forms. All charts, including one
sample Weight and Balance Clearance Form F, if
applicable, are completed at time of delivery. This
record is the basic weight and balance data of the
aircraft at delivery. All subsequent changes in weight
and balance are compiled by the weight and balance
technician.
6-7. WEIGHT DEFINITIONS.
Weight definitions are as follows:
a. Basic Weight, The basic weight of an aircraft is that weight which includes all fixed operating equipment and unusable fuel and engine oil. It
is only necessary to add variable or expendable load
items for various missions. The basic weight of an
aircraft varies with structural modifications and
changes in fixed operating equipment. The term
basic weight, when qualified with a word indicating
the type of missions such as Basic Weight for Combat, Basic Weight for Ferry, etc., may be used in
conjunction with directives stating what the equipment will be for these missions. For example, extra
fuel tanks and various items of equipment installed
for long range ferry flight, which are not normally
carried on combat missions, will be included in
Basic Weight for Ferry but not in Basic Weight for
Combat.
b. Operating Weight. The operating weight is
the basic weight of the aircraft, including the crew
and all equipment required for the mission, but not
including fuel or payload.
c. Gross Weight. The gross weight is the total
weight of an aircraft contents, and fuel.
6-1
TM 55-1510-221-10
Figure 6-1. Aircraft Compartments and Stations
6-2
TM 55-1510-221-10
1.) The takeoff gross weight is the operating weight plus the variable and expendable load
items which vary with the mission.
(2.) The landing gross weight is the takeoff gross weight minus the expended load items.
6-8. BALANCE DEFINITIONS.
Balance definitions are as follows:
a. Reference Datum. The reference datum is
an imaginary vertical plane at, or forward of, the
nose of the aircraft from which all horizontal distances are measured for balance purposes. Diagrams
of each aircraft show this reference datum as fuselage station zero.
b. Arm. Arm, for balance purposes, is the horizontal distance in inches from the reference datum
to the center of gravity of the item. Arm may be
determined from the Aircraft Compartment and Station Diagram (fig. 6-l).
c. M o m e n t . Moment is the product of a
weight multiplied by its arm. Moment divided by a
constant is generally used to simplify balance calculations by reducing the number of digits. For this
aircraft, inches and moment/100 have been used.
d. Average Arm. Average arm is the arm
obtained by adding the weights and the moments of
a number of items and dividing the total moment by
the total weight.
e. Basic Moment. Basic moment is the sum of
the moments of all items making up the basic
weight. When using data from an actual weighing of
an aircraft, the basic moment is the total moment of
the basic aircraft with respect to the reference
datum.
f
Center of Gravity (CG). Center of gravity is
the point about which an aircraft would balance if
suspended. Its distance from the reference datum is
found by dividing the total moment by the total
weight of the aircraft.
g. 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.
last weight and moment/100 entry is considered the
current weight and balance status of the basic aircraft (fig. 6-2).
6-10. WEIGHT AND BALANCE CLEARANCE
FORM F, DD FORM 365-4 (TACTICAL).
Form F (fig. 6-3) is a summary of the actual disposition of load in the aircraft. It records the balance status of the aircraft step by step. It serves as
a work sheet to record weight and balance calculations and any corrections that must be made to
insure that the aircraft will be within weight and CG
limits. It is necessary to complete a Form F prior to
flight when an aircraft is loaded in a manner for
which no previous valid Form F is available. A copy
must remain in the aircraft for the duration of the
flight. Form F (Tactical) is completed as follows:
1.
Insert necessary identifying information at
top of form. In blank spaces of
LIMITATIONS table, enter gross weights
for takeoff and landing obtained from the
WEIGHT LIMITATIONS paragraph in
Chapter 5.
2.
Ref 1. Enter aircraft basic weight and index
or mom/100 figure. Obtain this information
from last entry on Chart C (fig. 6-2).
3.
Ref 2. Leave blank (oil is included in basic
weight).
4.
Using
Ref
3.
letter
compartment
shown in
designations as
Aircraft
Compartment and Station Diagram (fig.
6-1) enter number, weight, and mom/100
figures of crew at their takeoff positions. Use
actual crew weights if available. Enter total
of each compartment in WEIGHT and
MOM/100 columns. To determine MOM/
100 of crew use Table 6-1, Occupants Useful
Load, Weights and Moments.
NOTE
The maximum baggage compartment
weight is 410 pounds. Also the floor loading limit of 100 Lbs/Sq Ft shall not be
exceeded.
6-9. CHART C - BASIC WEIGHT AND BALANCE
RECORD.
5. Ref 4. Enter sum of weights and sum of
mom/100 figures for Ref 1 through Ref 3 to
obtain OPERATING WEIGHT and corresponding mom/ 100 figure.
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
6. Ref 5. Enter the item description (flare/
chaff), total amount, weight, and MOM/100
of all expendable stores.
6-3
TM 55-1510-221-10
Figure 6-2. Chart C - Basic Weight and Balance Record, DD Form 365-3
7.
Ref 6. Not applicable to RC-12H aircraft.
9.
Ref 8. Not applicable to RC-12H aircraft.
8.
Ref 7. Enter the number of gallons, weight
and sum of mom/100 figures for takeoff
fuel. The weight of fuel used in warm up and
taxiing should not be included. If JP-4 fuel
at a density of 6.5 Lb/Gal is being used, use
Fuel Moments Table (Table 6-2) to determine fuel weight and mom/ 100 figures. If
other than JP-4 is being used, use the following procedures to obtain accurate fuel weight
and mom/100 figures from the Fuel Moment
Chart and to record information on Form F.
10.
Ref 9. Enter sum of weights and moments
for reference 4 through reference 8 opposite
TAKEOFF CONDITION (uncorrected).
11.
Ref 10. Enter TAKEOFF C.G. (uncorrected)
as determined from weight and moment
values of reference 9.
a.
Assume 6.7 Lb/Gal for JP-5 fuel or 6.0
Lb/Gal for AVGAS.
b.
Multiply the total number of gallons of
fuel in the aircraft times the computed
fuel density figure and thereby
determine actual fuel weight.
c.
Use Fuel Moment Table (Table 6-2) to
determine mom/ 100 figure.
d. Enter the weight and corresponding
mom/100 figure in the corresponding
columns of Ref 7. Also, enter a figure
for the total fuel gallons known to be in
the aircraft.
6-4
Check WEIGHT figure opposite Ref 10 against
GROSS WT. TAKEOFF. Check mom/100 figure
opposite Ref 10 with Center-of-Gravity Limits
Table (Table 6-4) to ascertain that CG is within
allowable limits.
12.
Ref 11. If changes in amount or distribution
of loads are required, indicate necessary
adjustments by proper entries in CORRECTIONS table. Enter a brief description of
adjustment made in column marked ITEM.
Add all weight and moment decreases and
insert totals in space opposite TOTAL
WEIGHT REMOVED. Add all weight and
moment increase and insert totals in space
opposite TOTAL WEIGHT ADDED. Subtract smaller from larger of two totals and
enter differences (with applicable + or sign) opposite NET DIFFERENCE. Transfer
these NET DIFFERENCE figures to spaces
opposite Ref 11.
TM 55-1510-221-10
Figure 6-3. Weight and Balance Clearance, DD Form 365-4 (Tactical)
6-5
TM 55-1510-221-10
Table 6-1. Occupants Useful Load, Weights and
Moments
Fuel.
a.
USEFUL LOAD WEIGHTS
AND MOMENTS
OCCUPANTS
13.
WEIGHT
CREW
F.S.129
MOM/100
80
90
100
110
120
130
140
150
160
170
180
190
200
210
220
230
240
250
103
116
129
142
155
168
181
194
206
219
232
245
258
271
284
297
310
323
Ref 12. Enter sum of the difference between
Ref 10 and Ref 11. Recheck against GROSS
WT. TAKEOFF in LIMITATIONS table to
assure that this figure does not exceed
allowable limit.
Record figures in LIMITATIONS table as follows: In FORWARD and AFT space for PERMISSIBLE C.G. TAKEOFF enter an inches figure
obtained as follows: Match weight figure recorded in
Ref 12 with a corresponding figure in the GROSS
WEIGHT-POUNDS on Center-of-Gravity Limits
Table (Table 6-4). Determine the FORWARD limit
in inches figure for the weight matched and record
in the FORWARD space stated. Enter AFT C.G.
LIMIT in inches in AFT space.
14. Ref 13. By referring to Center-of-Gravity
Limits Table (Table 6-4), determine takeoff
C.G. position. Enter this figure in space provided opposite TAKEOFF C.G. in inches.
Insure that this position is within the FWD
and AFT C.G. limit in LIMITATIONS
table.
15. Ref 14. Less expendables.
6-6
Enter estimated weight of fuel to be
expended. Subtract this figure from
weight of fuel on board (reference 7).
This figure represents the landing fuel
weight. Use the Fuel Moments Table
(Table 6-2) to determine landing fuel
moment. Subtract the landing fuel
moment from the total fuel moment on
board (reference 7). This figure represents the moment of fuel expended.
Enter in reference 14.
NOTE
Do not consider reserve fuel as expended
when determining ESTIMATED LANDING CONDITIONS.
Flare/Chaff cartridges.
b.
Determine total weight of flares and/or
chaff cartridges t h a t h a v e b e e n
expended. Use the Aircraft Survivability Equipment table (Table 6-3) to
determine landing expendables
moment. Enter figures in WEIGHT
and MOM/ 100 columns.
c.
Ref 15. Enter differences in weights
and MOM/100 figures between reference 12 and totals of reference 14.
Record figures in LIMITATIONS table as follows: In FORWARD and AFT space for PERMISSIBLE C.G. LANDING, enter a figure obtained as follows: Match weight figure recorded in reference 15
with a corresponding figure i n t h e G R O S S
WEIGHT-POUNDS on Center-of-Gravity Limits
Table (Table 6-4). Determine the FORWARD limit
in inches figure for the weight matched and record
in the FORWARD space stated. Within the AFT
space for PERMISSIBLE C.G. LANDING, record
the AFT C.G. limit in inches. Check data against
PERMISSIBLE C.G. LANDING in LIMITATIONS
table.
d. Ref 16. Refer to Center-of-Gravity
Limits Table (Table 6-4) to determine
landing C.G. position. Enter this figure
in space provided opposite ESTIMATED LANDING C.G. in inches.
Check Ref 16 against PERMISSIBLE C.G.
LANDING in LIMITATIONS Table.
Necessary signatures must appear at bottom of
form.
TM 55-1510-221-10
F U E L
D E N S I T Y / W E I G H T
v s
T E M P E R A T U R E
MODEL: UC-12B
DATE: 14 MAY 1979
DATA BASIS: FLIGHT TEST
CONFIGURATION:
ENGINE: PT6A-41
PROPELLER: T10178
FUEL GRADE: JP-5
FUEL DENSITY: 6.8 LB/GAL
EXAMPLE:
FUEL
TEMPERATURE: . . . . . . . . . . . . . . . .
FUEL GRADE: . . . . . . . . . . . . . . . . . . . . .
SPECIFIC WEIGHT . . . . . . . . . . . . . . . .
FUEL QUANTITY: . . . . . . . . . . . . . . . . . .
FUELWEIGHT: . . . . . . . . . . . . . . . . . . .
28°C
JP-5
= 6.7 LB/US GAL
130 US GAL
(6.7 X 130) = 871 LBS
AP 004484
Figure 6-4. Density Variation of Aviation Fuel.
6-7
TM 55-1510-221-10
Table 6-2. Fuel Center-of-Gravity Moments
USEFUL LOAD WEIGHTS AND MOMENTS
USABLE FUEL
6-8
TM 55-1510-221-10
Table 6-3. Survivability Equipment Weights and Moments
Item
NACELLE DISPENSER
Weight
(lb)
Dispenser (Empty)
Chaff Cartridges (30)
10
9
TOTAL : (Dispenser/30 chaff cartridges)
19
Moment/100
I
I
21
19
40
FUSELAGE DISPENSER
Dispenser (Empty)
Chaff and/or Flare Cartridges
TOTAL : (Dispenser/30 chaff and/or flare
cartridges)
10
9
49
26
19
55
Change 4 6-9
TM 55- 1510-221-10
Table 6-4. Center-of-Gravity Moments (sheet 1 of 3)
CENTER OF GRAVITY MOMENT TABLE - MOMENT/100
6-10
TM 55-1510-221-10
Table 6-4. Center-of-Gravity Moments (sheet 2 of 3)
CENTER OF GRAVITY MOMENT TABLE - MOMENT 100
6-11
TM 55-1510-221-10
Table 6-4. Center-of-Gravity Moments (sheet 3 of 3)
CENTER OF GRAVITY MOMENT TABLE - MOMENT/ 100 (CONT’D)
6-12
TM 55-1510-221-10
Table 6-5. C.G. Limits (Landing Gear Down) - Restricted Category
*CENTER OF GRAVITY LIMITS (LANDING GEAR DOWN) RESTRICTED CATEGORY
NOTES:
-The moment/100 for retraction of the alighting gear is - 60.4. Loadings based on wheels-down condition which fall within
the limiting moments in the table, will be satisfactory for flight with alighting gear retracted.
6-13
TM 55-1510-221-10
Section III. FUEL/OIL
6-11. FUEL LOAD.
6-12. FUEL AND OIL DATA.
Fuel loading imposes a restriction on the
amount of load which can be carried. The required
fuel must first be determined, then that weight subtracted from the total weight of crew and fuel.
Weight up to and including the remaining allowable
capacity can be subtracted directly from the weight
of crew and fuel. As the fuel load is increased, the
loading capacity is reduced.
a. Fuel Moment Table. This table (Table 6-2)
shows fuel moment/ 100 given US gallons or pounds
for JP-4 and JP-5.
b. Oil Data. Total oil weight is 62 pounds and
is included in the basic weight of the aircraft. Servicing information is provided in Section XII of Chapter 2.
Section IV. CENTER OF GRAVITY
6-13. CENTER OF GRAVITY LIMITATIONS.
Removal of mission gear may result in
exceeding the forward center-of-gravity
limit.
Center of gravity limitations are expressed in
ARM inches which refers to a positive measurement
from the aircraft’s reference datum. The forward CG
limit at 11,279 Lbs. or less is 181.0 ARM inches.
The forward-sloping CB limit line from 11,279 Lbs.
to 13,500 Lbs., and straight up to 15,000 Lbs., is
fuselage station 188.3. At 15,000 Lbs. or less, the aft
CG limit is 195.1 ARM inches. The Center of Gravity Limitations Table (Table 6-4) is designed to
establish forward and aft CG limitations.
Section IV. Cargo Loading
6-14. LOAD PLANNING. The basic factors to be
considered in any loading situation are as follows:
a. Cargo shall be arranged to permit access to
all emergency equipment and exits during flight.
b. Floorboard structural capacity shall be considered in the loading of heavy or sharp-edged containers and equipment. Shorings shall be used to distribute highly condensed weights evenly over the
cargo areas.
c. All cargo shall be adequately secured to
prevent damage to the aircraft, other cargo, or the
item itself.
6-15. LOADING PROCEDURE.
NOTE
The cabin door is weight limited to a
maximum of 300 pounds to prevent possible structural damage.
6-14
Loading of cargo is accomplished through the
cabin door (21.5 in. X 50.0 in.) or the cargo door
(52.0 in. X 52.0 in).
6-16. SECURING LOADS.
All cargo shall be secured with restraints strong
enough to withstand the maximum force exerted in
any direction. The maximum force can be determined by multiplying the weight of the cargo item
by the applicable load factor. These established load
factors (the ratio between the total force and the
weight of the cargo item) are 1.5 to the side and
rear, 3.0 up, 6.6 down, and 9.0 forward.
TM 55-1510-221-10
CHAPTER 7
PERFORMANCE
TABLE OF CONTENTS
Introduction to Performance ..........................................................................................................................
Takeoff Flight Path ..........................................................................................................................................
Comments Pertinent to the use of Performance Graphs ..............................................................................
Table of Contents ............................................................................................................................................
Airspeed Calibration - Normal System ........................................................................................................
Altimeter Correction - Normal System ........................................................................................................
Airspeed Calibration - Alternate System ......................................................................................................
Altimeter Correction - Alternate System ......................................................................................................
Free Air Temperature Correction ................................................................................................................
ISA Conversion ..............................................................................................................................................
Fahrenheit to Celsius Temperature Conversion ..........................................................................................
Minimum Takeoff Power at 2000 RPM ......................................................................................................
Maximum Takeoff Weight Permitted by Enroute Climb Requirement ......................................................
Takeoff Weight to Achieve Positive One-Engine-Inoperative Climb
at Lift-Off - Flaps 0% ....................................................................................................................................
Takeoff Weight to Achieve Positive One-Engine-Inoperative Climb
at Lift-Off - Flaps 40% ..................................................................................................................................
Wind Components ........................................................................................................................................
Takeoff Distance - Flaps 0% ........................................................................................................................
Accelerate-Stop - Flaps 0% ............................................................................................................................
Accelerate-Go Distance Over 50 FT Obstacle - Flaps 0% ..........................................................................
Takeoff Climb Gradient - One-Engine-Inoperative - Flaps 0% ..................................................................
Takeoff Distance - Flaps 40% ......................................................................................................................
Accelerate-Stop - Flaps 40% ..........................................................................................................................
Accelerate-Go Distance Over 50 FT Obstacle - Flaps 40% ........................................................................
Takeoff Climb gradient - One Engine Inoperative Flaps 40% ......................................................................................................................................................
Climb - Two-Engine - Flaps 0% ..................................................................................................................
Climb - Two-Engine - Flaps 40% ................................................................................................................
Climb - One Engine Inoperative ..................................................................................................................
Service Ceiling One-engine-Inoperative ........................................................................................................
Time, Fuel, and Distance to Cruise Climb ..................................................................................................
Maximum Cruise Power 1900 RPM ISA-30°C ..........................................................................................
Maximum Cruise Power 1900 RPM ISA-20°C ..........................................................................................
Maximum Cruise Power 1900 RPM ISA-1OºC ..........................................................................................
Maximum Cruise Power 1900 RPM ISA ..................................................................................................
Maximum Cruise Power 1900 RPM ISA+10ºC ........................................................................................
Maximum Cruise Power 1900 RPM ISA+20ºC ........................................................................................
Maximum Cruise Power 1900 RPM ISA+30ºC ........................................................................................
Maximum Cruise Power 1900 RPM ISA+37ºC ........................................................................................
Maximum Cruise Speeds 1900 RPM ..........................................................................................................
Maximum Cruise Power 1900 RPM ........................................................................................................
Fuel Flow At Maximum Cruise Power 1900 RPM ..................................................................................
Range Profile - Maximum Cruise Power 1900 RPM ................................................................................
Maximum Range Power 1700 RPM ISA-30°C ..........................................................................................
Maximum Range Power 1700 RPM ISA-20°C ..........................................................................................
Maximum Range Power 1700 RPM ISA-10ºC ..........................................................................................
7-l
7-3
7-6
7-8
7-10
7-11
7-12
7-13
7-14
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
7-47
7-48
7-49
7-1
TM 55-1510-221-10
Maximum Endurance Power 1700 RPM ISA+20ºC ................................................................................
Maximum Endurance Power 1700 RPM ISA+30ºC ................................................................................
Maximum Endurance Power 1700 RPM ISA+37ºC ................................................................................
Range Profile - Long Range Power 1700 RPM ........................................................................................
Range Profile - 542 Gallons Usable Fuel ....................................................................................................
Endurance Profile - 542 Gallons Usable Fuel ............................................................................................
Time, Fuel, And Distance to Descend ........................................................................................................
Climb - Balked Landing ................................................................................................................................
Normal Landing Distance Without Propeller Reversing
- Flaps 100% ..................................................................................................................................................
Landing Distance Without Propeller Reversing - Flaps 0% ......................................................................
Maximum Range Power 1700 RPM ISA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Maximum Range Power 1700 RPM ISA+ 10°C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Maximum Range Power 1700 RPM ISA+20ºC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Maximum Range Power 1700 RPM ISA+30ºC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Maximum Range Power 1700 RPM ISA+37ºC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Maximum Endurance Power 1700 RPM ISA-30°C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Maximum Endurance Power 1700 RPM ISA-20°C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Maximum Endurance Power 1700 RPM ISA-1OºC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Maximum Endurance Power 1700 RPM ISA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Maximum Endurance Power 1700 RPM ISA+ 10°C .,.........................................................................,....
7-2
7-60
7-61
7-62
7-63
7-64
7-65
7-66
7-67
7-68
7-69
7-50
7-51
7-52
7-53
7-54
7-55
7-56
7-57
7-58
7-59
TM 55-1510-221-10
CHAPTER 7
PERFORMANCE
7-l. INTRODUCTION TO PERFORMANCE.
The graphs in this Section present performance
information for takeoff, climb, cruise, and landing at
various parameters of weight, altitude, and temperature.
The following example presents calculations for
a proposed flight from Denver to Reno using the
conditions listed below:
7-2. CONDITIONS.
At Stapleton International (DEN):
Free Air Temperature . . . . . . . . . . . . . . . . . . . . . . . . . . 28°C (82°F)
Field Elevation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5333 feet 1
Altimeter Setting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30.02 in. Hg
Wind . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 030º at 13 knots
Runway 35R Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12,000 feet1
Route of trip:
DEN - J116 - EKR - J173 - SLC - J154 BAM - J32 - RN0
Cruise Altitude:
26,000 feet
At Cannon International (RNO):
Free Air Temperature . . . . . . . . . . . . . . . . . . . . . . . . . . 32°C (90°F)
Field Elevation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4412 feet1
Altimeter Setting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29.60 in. Hg
Wind . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 200º at 15 knots
Runway 25 Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6101 feet1
1 Source: NOAA Standard Instrument Departures for Western United States, 9 APR
1987.
2
Source: NOAA Enroute High Altitude -U.S.
Chart H-l, 9 APR 1987.
3
MEA on NOAA Enroute Low Altitude -U.S.
Chart L-8, 9 JUN 1983.
4 Includes distance between airport and
VORTAC, per NOAA Airport/Facility
Directory (Southwest US.), 9 APR 1987.
in. Hg. Always subtract the reported altimeter setting FROM 29.92 in. Hg, then multiply the answer
by 1000 to find the difference in feet between field
elevation and pressure altitude.
Pressure Altitude at DEN:
29.92 in. Hg - 30.02 in. Hg = -0.10
-0.10 x 1000 feet = -100 feet
The pressure altitude at DEN is 100 feet
below field elevation.
Pressure altitude at DEN = 5333 - 100 = 5233 feet.
Pressure altitude at RNO:
29.92 in. Hg - 29.60 in. Hg = 0.32
0.32 x 1000 feet = 320 feet
The pressure altitude at RN0 is 320 feet
above field elevation.
Pressure altitude at RN0 = 4412 + 320 = 4732 feet.
7-4. PERFORMANCE EXAMPLE.
Maximum takeoff weight (from LIMITATIONS
Section) = 15,000 pounds
7-5. MAXIMUM TAKEOFF WEIGHT PERMITTED
BY ENROUTE CLIMB REQUIREMENT.
Enter the graph at 5233 feet take-off field pressure altitude to 28°C takeoff FAT:
Maximum Allowable Takeoff Weight 14,200 pounds
The maximum takeoff weight permitted by the
Enroute Climb Requirement graph is the only operating limitation required to meet applicable FAR
requirements. Information has been presented, however, to determine the takeoff weight, field requirements, and takeoff flight path assuming an engine
failure occurs during the take-off procedure. The following illustrates the use of these charts.
7-3. PRESSURE ALTITUDE.
7-6. TAKEOFF WEIGHT TO ACHIEVE POSITIVE
ONE-ENGINE-INOPERATIVE CLIMB AT LIFTOFF
(Flaps 0%).
To determine the approximate pressure altitude
at origin and destination airports, add 1000 feet to
field elevation for each 1.00 in. Hg that the reported
altimeter setting value is below 29.92 in. Hg, and
subtract 1000 feet for each 1.00 in. Hg above 29.92
Enter the graph at 5233 feet to 28ºC, to determine the maximum weight at which the accelerate-go procedure should be attempted.
Maximum Accelerate-Go Weight . . . . 13,480 pounds
7 - 3
TM 55-1510-221-10
7-7. ACCELERATE-STOP (FLAPS 0%).
Enter the Accelerate-Stop graph at 28ºC, 5233
feet pressure altitude, 13,480 pounds, and 10 knots
head wind component:
Accelerate-Stop Distance . . . . . . . . . . . . . . . . . . . . . . . . 5050 feet
Takeoff Decision Speed (VR ) . . . . . . . . . . . . . . . . . . . . 98 knots
7-8. TAKEOFF DISTANCE (FLAPS 0%).
Enter the graph at 28ºC, 5233 feet pressure altitude, 13,480 pounds, and IO knots head wind component:
Ground Roll . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3420 feet
Total Distance Over 50-foot Obstacle . . . . 5100 feet
Takeoff Speed:
At Rotation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100 knots
At 50 Feet . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 16 knots
Horizontal distance used to climb from 50 feet to 100 feet (100 - 50) (1000 ÷ 51) = 981 feet
Total Distance = 5500 + 981- 6531 feet
Results are illustrated below:
7-12. FLIGHT PLANNING.
Calculations for flight time, block speed, and
fuel requirements for a proposed flight are detailed
below using the same conditions presented on page
4-3, and a takeoff weight of 12,000 pounds.
Enter ISA CONVERSION graph at the conditions indicated:
DEN-BVL
Pressure Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26,000 feet
FAT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -20°C
ISA Condition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ISA + 17°C
The following example assumes the aircraft is
loaded so that takeoff weight is 10,000 pounds.
BLV-RN0
Pressure Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26,000 feet
FAT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -10°C
IAS Condition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ISA + 27°C
7-10. ACCELERATE-GO DISTANCE OVER 50FOOT OBSTACLE (FLAPS 0%).
Enter the TIME, FUEL, AND DISTANCE TO
CLIMB Graph at 28°C to 5233 feet, and to 12,000
pounds, and enter at -10°C to 26,000 feet, and to 12,
000 pounds, and read:
7-9. TAKEOFF FLIGHT PATH EXAMPLE.
Enter the graph at 28°C, 5233 feet pressure altitude, 10,000 pounds, and 10 knots head wind component:
Total Distance Over 50-foot Obstacle . . . . 5550 feet
Speed at Rotation (VR) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 93 knots
Speed at 35 Feet Above Runway (Climb Speed) Knots
Time to Climb . . . . . . . . . . 30.0-3.1 = 26.9=27 minutes
Fuel Used to Climb . . . . . . . . . . . . . . . . 379.3-60.0=319.3=
320 pounds
Distance Traveled . . . . . . 85.7-8.7=77 nautical miles
7-11. TAKEOFF CLIMB GRADIENT - ONE
ENGINE INOPERATIVE (FLAPS 0%).
Time to Descend . . . . . . 17.7-3.1= 14.6 = 15 minutes
Fuel Used to Descend . . . . 198.5-50.0=149 pounds
Distance Traveled . . . . . . . . . . . . . . . . . . . . . . . . 86.5- 15.3 =71.2=
71 nautical miles
Enter the graph at 28°C, 5233 feet pressure altitude, and 10,000 pounds:
Climb Gradient . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.1%
Climb Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 knots
A 5.1% climb gradient is 51 feet of vertical
height per 1000 feet of horizontal distance.
NOTE
The graphs for take-off climb gradient
assume a zero-wind condition. Climbing
into a head wind will result in higher
angles of climb, and hence better obstacle
clearance capabilities.
Calculations of the horizontal distance to clear
an obstacle 100 feet above the runway surface:
7-4
Enter the TIME, FUEL, AND DISTANCE TO
DESCEND Graph at 26,000 feet, and enter again at
4732 feet, and read:
An estimated average cruise weight of 11,200
pounds was used for this example.
Enter the tables for MAXIMUM ENDURANCE
POWER 1700 RPM for ISA + 10°C ISA + 20°C,
and ISA + 30°C and read the cruise speeds for 26,
000 feet at 12,000 pounds and 11,000 pounds:
Cruise True Airspeed (ISA + 17°C) . . . . . . 169 knots
Cruise True Airspeed (ISA + 27°C) . . . . . . 172 knots
Enter
the
*MAXIMUM
ENDURANCE
POWER 1700 RPM Tables for ISA + 10°C ISA +
2OºC, and ISA + 30°C at 12,000 pounds and 11,000
pounds and interpolate the recommended torque
settings for ISA + 17°C and ISA + 27°C at 11,200
pounds.
ISA + 17°C . . . . . . . . . . . . . . . . . . . . . . . . 40% torque per engine
TM 55-1510-221-10
Figure 7-1. Takeoff Flight Path
ISA + 27°C . . . . . . . . . . . . . . . . . . . . . . . . 41% torque per engine
*MAXIMUM
ENDURANCE
Enter
the
POWER 1700 RPM Tables for ISA + 10°C ISA +
20°C and ISA + 30°C at 12,000 pounds and 11,000
pounds at 26,000 feet, and interpolate the fuel flows
for ISA + 17°C and ISA + 27°C at 11,200 pounds.
ISA + 17°C
Fuel Flow Per Engine . . . . . . . . . . . . . . . . . . . . . . 198.75
Total Fuel Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 397.50
ISA + 27°C
Fuel Flow Per Engine . . . . . . . . . . . . . . . . . . . . . . . . 203.5
Total Fuel Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 407
Lbs/hr
Lbs/hr
Lbs/hr
Lbs/hr
NOTE
For flight planning, enter these charts at
the forcasted ISA condition; for enroute
power settings and fuel flows, enter at the
actual indicated FAT.
Time and fuel used were calculated at MAXIMUM ENDURANCE POWER 1700 RPM as follows:
Time =
Distance Ground Speed
Distance x Total Fuel Flow
Fuel Used =
Ground Speed
Results are as follows:
7-13. RESERVE FUEL.
Reserve Fuel is calculated as 45 minutes at Maximum Range Power 1700 RPM. Use planned cruise
altitude (26,000 feet), forecasted ISA condition (ISA
+ 27°C) and estimated weight at end of planned
trip (10,309 pounds). (Since the lowest weight column in the tables is 11,000 pounds, assume weight
at the end of the planned trip to be 11,000 pounds,
and use that fuel flow value for this example.)
Enter the tables for MAXIMUM RANGE
POWER 1700 RPM for ISA + 20°C and ISA +
30°C at 11,000 Lbs and 26,000 feet, and read the
total fuel flows:
ISA + 20°C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 478 Lbs/hr
ISA + 30°C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 448 Lbs/hr
Then interpolate for the fuel flow at ISA + 27°C
as follows:
Change in Fuel Flow = 478 - 448 = 30 Lbs/hr
Change in Temperature = (ISA + 20°C) - (ISA
+ 30°C) = 10°C
Rate of Change in Fuel Flow = Change in Fuel
Flow ÷ Change in Temperature
7-5
TM 55-1510-221-10
Rate of Change in Fuel Flow = (30 Lbs/hr) ÷
(10°C)
Rate of Change in Fuel Flow = 3.0 Lbs/hr
decrease per 1°C increase
Temperature increase from ISA + 20°C to ISA
+ 27°C = 7°C
Total Change in Fuel Flow = 7 x 3.0 Lbs/hr =
21.0 Lbs/hr
Total Fuel Flow = (ISA + 20°C Fuel Flow) +
(Total Change in Fuel Flow)
Enter the NORMAL LANDING DISTANCE
WITHOUT PROPELLER REVERSING - FLAPS
100% Graph at 32°C, 4732 feet, 10,309 pounds, and
10 knots head wind component:
Ground Roll . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1800 feet
Total Distance Over 50-foot Obstacle . . . . 2510 feet
Approach Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99 knots
Enter the CLIMB - BALKED LANDING Graph
at 32°C, 4732 feet, and 10,309 pounds:
Rate of Climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1270 ft/min
Climb Gradient . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10.2%
Total Fuel Flow = (478) - (21) = 457 Lbs/hr
Reserve Fuel = 45 minutes x Total Fuel Flow
Reserve Fuel = (0.75) x (457 Lbs/hr) = 342.75
= 343 lbs.
Total Fuel Requirement = 1781 + 343 = 2124
pounds
7-14. ZERO FUEL WEIGHT LIMITATION.
For this example, the following conditions were
assumed:
Ramp Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12,090 pounds
Weight of Usable Fuel Onboard . . . . . . 2124 pounds
Zero Fuel Weight = Ramp Weight - Weight of
Usable Fuel Onboard
Zero Fuel Weight = (12,090) - (2124) = 9966
pounds
Maximum zero fuel weight limitation (from
LIMITATIONS section) = 11,500 pounds.
Maximum Zero Fuel Weight Limitation has not
been exceeded.
Anytime the Zero Fuel Weight exceeds the Maximum Zero Fuel Weight Limit, the excess must be
off-loaded from PAYLOAD. If desired, additional
FUEL ONLY may then be added until the ramp
weight equals the Maximum Ramp Weight Limit of
15.090 Lbs.
7-15. LANDING INFORMATION.
The estimated Landing Weight is determined by
subtracting the fuel required for the trip from the
Ramp Weight:
Ramp Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12,090 Lbs
Fuel Required for Total Trip . . . . . . . . . . 1781 pounds
Landing Weight (12,090 - 1781) . . . . 10,309 pounds
7-6
7-16. COMMENTS PERTINENT TO THE USE OF
PERFORMANCE GRAPHS.
a . In addition to presenting the answer for a
particular set of conditions the example on a graph
also presents the order in ‘which the various scales
on the graph should be used. For instance, if the
first item in the example is FAT, then enter the
graph at the existing FAT.
b. The reference lines indicate where to begin
following the guidelines. Always project to the reference line first, then follow the guidelines to the next
known item by maintaining the same PROPORTIONAL DISTANCE between the guide line above
and the guide line below the projected line. For
instance, if the projected line intersects the reference
line in the ratio of 30% down/70% up between the
guidelines, then maintain this same 30%/70% relationship between the guide lines and follow them to
the answer or next known item.
c. The associated conditions define the specific conditions from which performance parameters
have been determined. They are not intended to be
used as instructions; however, performance values
determined from charts can only be achieved if the
specified conditions exist.
d. The full amount of usable fuel is available
for all approved flight conditions.
e. Indicated airspeeds (IAS) were obtained
using the Airspeed Calibration - Normal System
graph.
f: Notes have been provided on various
graphs and tables to approximate performance with
ice vanes extended. The effect will vary, depending
upon airspeed, temperature, altitude, and ambient
conditions. At lower altitudes, where operation on
the torque limit is possible, the effect of ice vane
extension will be less, depending upon how much
power can be recovered after the ice vanes have
been extended.
TM 55-1510-221-10
AIRSPEED CALIBRATION - NORMAL SYSTEM
Figure 7-2. Airspeed Correction - Normal System
7-7
TM 55-1510-221-10
ALTIMETER CORRECTION - NORMAL SYSTEM
Figure 7-3. Altimeter Correction - Normal System
7-8
Figure 7-4. Airspeed Calibration - Alternate System
TM 55-1510-221-10
7-9
TM 55-1510-221-10
7-10
Figure 7-5. Altimeter Correction - Alternate System
TM 55-1510-221-10
INDICATED OUTSIDE AIR TEMPERATURE CORRECTION
Figure 7-6. Free Air Temperature Correction
7-11
TM 55-1510-221-10
ISA CONVERSION
PRESSURE ALTITUDE vs FREE AIR TEMPERATURE
Figure 7-7. ISA Conversion
7-12
Figure 7-8, Fahrenheit to Celsius Temperature Conversion
TM 55-1510-221-10
7-13
TM 55-1510-221-10
MINIMUM TAKEOFF POWER AT 2000 RPM
(65 KNOTS)
NOTES:
1. TORQUE INCREASES APPROXIMATELY 1% FROM 0 TO 65 KNOTS.
2. THE PERCENT TORQUE INDICATED IN THIS FIGURE IS THE MINIMUM VALUE AT 65 KNOTS AT
WHICH TAKEOFF PRESENTED IN THIS SECTION CAN BE REALIZED. ANY EXCESS POWER WHICH
CAN BE DEVELOPED WITHOUT EXCEEDING ENGINE LIMITATIONS SHOULD BE UTILIZED.
3. FOR OPERATION WITH ICE VANES EXTENDED, INCREASE FIELD PRESSURE ALTITUDE 1000 FEET
BEFORE ENTERING GRAPH.
Figure 7-9. Minimum Takeoff Power at 2000 RPM
7-14
TM 55-1510-221-10
MAXIMUM TAKEOFF WEIGHT PERMITTED BY ENROUTE
CLIMB REQUIREMENT
ASSOCIATED CONDITIONS:
EXAMPLE:
POWER ....................... MAXIMUM CONTINUOUS
INOPERATIVE PROPELLERS ........ FEATHERED
LANDING GEAR ...................... UP
FLAPS . . . . . . . . . . . . . . . . . . . . . . . 0%
PRESSURE ALTITUDE . . . . . . 5233 FT
FAT . . . . . . 28°C
MAXIMUM TAKEOFF
WEIGHT
14.200 LBS
NOTE: ONE-ENGINE-INOPERATIVE PERFORMANCE WEIGHT LIMIT IS FOR
RATE-OF-CLIMB CAPABILITIES AT 5000 FT PRESSURE ALTITUDE.
REFER TO THE l CLIMB - ONE ENGINE INOPERATIVE’ GRAPH
FOR ACTUAL CLIMB CAPABILITIES APPLICABLE TO THE
PARTICULAR TEMPERATURE AND ALTITUDE BEING CONSIDERED.
Figure 7-10. Maximum Takeoff Weight Permitted by Enroute Climb Requirement
7-15
TM 55-1510-221-10
TAKEOFF WEIGHT - FLAPS 0%
TO ACHIEVE POSITIVE ONE-ENGINE-INOPERATIVE CLIMB AT LIFT-OFF
ASSOCIATED CONDITIONS:
EXAMPLE:
POWER
.
.
.
TAKEOFF
FLAPS
0%
LANDING GEAR ................ DOWN
INOPERATIVE PROPELLER .... FEATHERED
PRESSURE ALTITUDE . . . . 5233 FT
FAT
...... 28°C
TAKEOFF WEIGHT ........... 13,480 LBS
Figure 7-11. Takeoff Weight - Flaps 0%, To Achieve Positive One-Engine Climb at Lift-Off
7-16
TM 55-1510-221-10
TAKEOFF WEIGHT - FLAPS 40%
TO ACHIEVE POSITIVE ONE-ENGINE-INOPERATIVE CLIMB AT LIFT-OFF
Figure 7-12. Takeoff Weight - Flaps 40%, To Achieve Positive One-Engine Climb at Lift-Off Flaps 40%
7-17
TM 55-1510-221-10
WIND COMPONENTS
Demonstrated Crosswind is 17 kts
EXAMPLE:
WIND SPEED
ANGLE BETWEEN WIND DIRECTION AND FLIGHT PATH
HEADWIND COMPONENT
CROSSWIND COMPONENT
Figure 7-13. Wind Components
7-18
Change 4
20 KTS
50%
13 KTS
15 KTS
Figure 7-14. Takeoff Distance - Flaps 0%
7-19
TM 55-1510-221-10
Change 4
Change 4
TM 55-1510-221-10
7-20
Figure 7-15. Accelerate-Stop - Flaps 0%
TM 55-1510-221-10
Figure 7-16. Accelerate-Go Distance Over 50-FT Obstacle - Flaps 0%
7-21
TM 55-1510-221-10
7-22
Figure 7-17. Takeoff Climb Gradient - One-Engine-Inoperative - Flaps 0%
Figure 7-18. Takeoff Distance - Flaps 40%
TM 55-1510-221-10
7-23
TM 55-1510-221-10
7-24
Figure 7-19. Accelerate-Stop - Flaps 40%
TM 55-1510-221-10
Figure 7-20. Accelerate-Go Distance Over 50-FT Obstacle - Flaps 40%
7-25
Figure 7-21. Takeoff Climb gradient - One-Engine-Inoperative - Flaps 40%
TM 55-1510-221-10
7-26
Figure 7-22. Climb Two-Engine - Flaps 0%
TM 55-1510-221-10
7-27
TM 55-1510-221-10
7-28
Figure 7-23. Climb Two-Engine - Flaps 40%
Figure 7-24. Climb One-Engine Inoperative
TM 55-1510-221-10
7-29
TM 55-1510-221-10
Figure 7-25. Service Ceiling One-engine-Inoperative
7-30
Figure 7-26. Time, Fuel, and Distance to Cruise Climb
TM 55-1510-221-10
7-31
TM 55-1510-221-10
Figure 7-27. Maximum Cruise Power 1900 RPM ISA-30°C (Sheet 1 of 2)
7-32
TM 55-1510-221-10
Figure 7-27. Maximum Cruise Power 1900 RPM ISA-30°C (Sheet 2 of 2)
7-33
TM 55-1510-221-10
Figure 7-28. Maximum Cruise Power 1900 RPM ISA-20°C (Sheet 1 of 2)
7-34
TM 55-1510-221-10
Figure 7-28. Maximum Cruise Power 1900 RPM ISA-20°C (Sheet 2 of 2)
7-35
TM 55-1510-221-10
Figure 7-29. Maximum Cruise Power 1900 RPM ISA- 10°C (Sheet 1 of 2)
7-36
TM 55-1510-221-10
Figure 7-29. Maximum Cruise Power 1900 RPM ISA- 10°C (Sheet 2 of 2)
7-37
TM 55-1510-221-10
Figure 7-30. Maximum Cruise Power 1900 RPM ISA (Sheet 1 of 2)
7-38
TM 55-1510-221-10
Figure 7-30. Maximum Cruise Power 1900 RPM ISA (Sheet 2 of 2)
7-39
TM 55-1510-221-10
Figure 7-37. Maximum Cruise Power 1900 RPM ISA+ 10°C (Sheer 1 of 2)
7-40
TM 55-1510-221-10
Figure 7-31. Maximum Cruise Power 1900 RPM ISA+ 10°C (Sheet 2 of 2)
7-41
TM 55-1510-221-10
Figure 7-32. Maximum Cruise Power 1900 RPM ISA+20°C (Sheet 1 of 2)
7-42
TM 55-1510-221-10
Figure 7-32. Maximum Cruise Power 1900 RPM ISA+20°C (Sheet 2 of 2)
7-43
TM 55-1510-221-10
Figure 7-33. Maximum Cruise Power 7900 RPM ISA+30°C (Sheet 1 of 2)
7-44
TM 55-1510-221-10
Figure 7-33. Maximum Cruise Power 1900 RPM ISA+30°C (Sheet 2 of 2)
7-45
TM 55-1510-221-10
Figure 7-34. Maximum Cruise Power 1900 RPM ISA+37°C (Sheet 1 of 2)
7-46
TM 55-1510-221-10
Figure 7-34. Maximum Cruise Power 1900 RPM ISA+37°C (Sheet 2 of 2)
7-47
TM 55-1510-221-10
Figure 7-35. Maximum Cruise Speeds 1900 RPM
7-48
TM 55-1510-221-10
Figure 7-36. Maximum Cruise Power 1900 RPM
7-49
TM 55-1510-221-10
Figure 7-37. Fuel Flow At Maximum Cruise Power 1900 RPM
7-50
Figure 7-38. Range Profile - Maximum Cruise Power 1900 RPM
TM 55-1510-221-10
7-51
TM 55-1510-221-10
Figure 7-39. Maximum Range Power 1700 RPM ISA-30°C (Sheet 1 of 2)
7-52
TM 55-1510-221-10
Figure 7-39. Maximum Range Power 7700 RPM ISA-30°C (Sheet 2 of 2)
7-53
TM 55-1510-221-10
Figure 7-40. Maximum Range Power 1700 RPM ISA-20°C (Sheet 1 of 2)
7-54
TM 55-1510-221-10
Figure 7-40. Maximum Range Power 1700 RPM ISA-20°C (Sheet 2 of 2)
7-55
TM 55-1510-221-10
Figure 7-41. Maximum Range Power 1700 RPM ISA- 10°C (Sheet 1 of 2)
7-56
TM 55-1510-221-10
MAXIMUM RANGE POWER
1700 RPM
ISA-10°C
Figure 7-41. Maximum Range Power 1700 RPM ISA- 10°C (Sheet 2 of 2)
7-57
TM 55-1510-221-10
Figure 7-42. Maximum Range Power 1700 RPM ISA (Sheet 1 of 2)
7-58
TM 55-1510-221-10
MAXIMUM RANGE POWER
1700 RPM
ISA
Figure 7-42. Maximum Range Power 1700 RPM ISA (Sheet 2 of 2)
7-59
TM 55-1510-221-10
MAXIMUM RANGE POWER
1700 RPM
ISA +10 °C
Figure 7-43. Maximum Range Power 1700 RPM ISA+10°C (Sheet 1 of 2)
7-60
TM 55-1510-221-10
MAXIMUM RANGE POWER
1700 RPM
ISA+10 °C
Figure 7-43. Maximum Range Power 1700 RPM ISA+ 10°C (Sheet 2 of 2)
7-61
TM 55-1510-221-10
MAXIMUM RANGE POWER
1700 RPM
ISA +20 °C
Figure 7-44. Maximum Range Power 1700 RPM ISA+20°C (Sheet 1 of 2)
7-62
TM 55-1510-221-10
MAXIMUM RANGE POWER
1700 RPM
ISA +20 °C
Figure 7-44. Maximum Range Power 1700 RPM ISA+20°C (Sheet 2 of 2)
7-63
TM 55-1510-221-10
MAXIMUM RANGE POWER
1700 RPM
ISA+30 °C
Figure 7-45. Maximum Range Power 1700 RPM ISA+3O°C (Sheet 1 of 2)
7-64
TM 55-1510-221-10
MAXIMUM RANGE POWER
1700 RPM
ISA +30 °C
Figure 7-45. Maximum Range Power 1700 RPM ISA+3O°C (Sheet 2 of 2)
7-65
TM 55-1510-221-10
MAXIMUM RANGE POWER
1700 RPM
ISA+37 °C
Figure 7-46. Maximum Range Power 1700 RPM ISA+37°C (Sheet 1 of 2)
7-66
TM 55-1510-221-10
MAXIMUM RANGE POWER
1700 RPM
ISA +37 °C
Figure 7-46. Maximum Range Power 1700 RPM ISA+37°C (Sheet 2 of 2)
7-67
TM 55-1510-221-10
Figure 7-47. Maximum Endurance Power 1700 RPM ISA-30°C (Sheet 1 of 2)
7-68
TM 55-1510-221-10
Figure 7-47. Maximum Endurance Power 1700 RPM ISA-30°C (Sheet 2 of 2)
7-69
TM 55-1510-221-10
Figure 7-48. Maximum Endurance Power 1700 RPM ISA-20°C (Sheet 1 of 2)
7-70
TM 55-1510-221-10
Figure 7-48. Maximum Endurance Power 1700 RPM ISA-20°C (Sheet 2 of 2)
7-71
TM 55-1510-221-10
Figure 7-49. Maximum Endurance Power 1700 RPM ISA-10° (Sheet 1 of 2)
7-72
TM 55-1510-221-10
Figure 7-49. Maximum Endurance Power 1700 RPM ISA-10°C (Sheet 2 of 2)
7-73
TM 55-1510-221-10
Figure 7-50. Maximum Endurance Power 1700 RPM ISA (Sheet 1 of 2)
7-74
TM 55-1510-221-10
Figure 7-50. Maximum Endurance Power 1700 RPM ISA (Sheet 2 of 2)
7-75
TM 55-1510-221-10
Figure 7-51. Maximum Endurance Power 1700 RPM ISA+ 10°C (Sheet 1 of 2)
7-76
TM 55-1510-221-10
Figure 7-51. Maximum Endurance Power 1700 RPM ISA+ 10°C (Sheet 2 of 2)
7-77
TM 55-1510-221-10
Figure 7-52, Maximum Endurance Power 1700 RPM ISA+20°C (Sheet 1 of 2)
7-78
TM 55-1510-221-10
Figure 7-52. Maximum Endurance Power 1700 RPM ISA+20°C (Sheet 2 of 2)
7-79
TM 55-1510-221-10
Figure 7-53. Maximum Endurance Power 1700 RPM ISA+3O°C (Sheet I of 2)
7-80
TM 55-1510-221-10
Figure 7-53. Maximum Endurance Power 1700 RPM ISA+30°C (Sheet 2 of 2)
7-81
TM 55-1510-221-10
Figure 7-54. Maximum Endurance Power 1700 RPM ISA+37°C (Sheet 1 of 2)
7-82
TM 55-1510-221-10
Figure 7-54. Maximum Endurance Power 1700 RPM ISA+37°C (Sheet 2 of 2)
7-83
TM 55-1510-221-10
7-84
Figure 7-55. Range Profile - Long Range Power 1700 RPM
TM 55-1510-221-10
RANGE PROFILE - 542 GALLONS USABLE FUEL
STANDARD DAY (ISA)
ASSOCIATED CONDITIONS:
EXAMPLE:
WEIGHT ...............
FUEL ...................
FUEL DENSITY .......
WIND ..................
PRESSURE ALTITUDE ........................ 20.000 FT
NOTE:
15,090 LBS BEFORE ENGINE START
AVIATION KEROSENE
6.7 LBS/GAL
ZERO
RANGE AT:
MAXIMUM CRUISE POWER .............. 1059 NM
MAXIMUM ENDURANCE POWER ........ 1079 NM
LONG RANGE POWER .................... 1107 NM
RANGE ALLOWS FOR TAXI AND RUNUP; INCLUDES CRUISE CLIMB AND DESCENT;
AND ALLOWS FOR 45 MINUTES RESERVE FUEL AT LONG RANGE POWER.
Figure 7-56. Range Profile - 542 Gallons Useable Fuel
7-85
TM 55-1510-221-10
7-86
Figure 7-57. Endurance Profile - 542 Gallons Useable Fuel
TM 55-1510-221-10
TIME, FUEL, AND DISTANCE TO DESCEND
EXAMPLE:
ASSOCIATED CONDITIONS:
INITIAL ALTITUDE. . . 26,000 FT
FINAL ALTITUDE . . . . 4732 FT
TIME TO DESCEND . . (17.7-3.1) 15 MIN
FUEL TO DESCEND . . (198.5-50.0) 149 LBS
DISTANCE TO DESCEND . . (86.5-15.3) 71 NM
DESCENT SPEED: VMO/MMO
POWER . . . AS REQUIRED TO DESCEND
AT 1500 FT/MIN
LANDING GEAR.. UP
FLAPS
.
.
0%
Figure 7-58. Time, Fuel, And Distance to Descend
7-87
TM 55-1510-221-10
7-88
Figure 7-59. Climb - Balked Landing
TM 55-1510-221-10
Figure 7-60. Normal Landing Without propeller Reversing - Flaps 100%
7-89
TM 55-1510-221-10
LANDING DISTANCE WITHOUT PROPELLER REVERSING - FLAPS 0%
EXAMPLE:
ASSOCIATED CONDITIONS:
POWER RETARDED TO MAINTAIN APPROPRIATE
DESCENT RATE ON FINAL APPROACH. ABOVE
13.500 LBS USE 500 FT/MIN: AT OR BELOW
13,500 LBS USE 600 FT/MIN:
FLAPS
0%
RUNWAY PAVED, LEVEL, DRY SURFACE
APPROACH SPEED.. IAS AS TABULATED
BRAKING
MAXIMUM
FLAPS - 100% LANDING
DISTANCE OVER
50-FT OBSTACLE.. 2510 FT
LANDING WEIGHT 10,309 LBS
FLAPS-UP LANDING
DISTANCE OVER
50-FT OBSTACLE. _. 3410 FT
APPROACH SPEED 111 KTS
NOTES: 1. LANDING WITH FLAPS FULL DOWN (100%) IS NORMAL PROCEDURE. USE THE GRAPH BELOW WHEN IT
IS NECESSARY TO LAND WITH FLAPS UP (0%).
2. TO DETERMINE FLAPS-UP LANDING DISTANCE, READ FROM THE “NORMAL LANDING DISTANCE WITHOUT PROPELLER REVERSING - FLAPS 100%” GRAPH THE LANDING DISTANCE APPROPRIATE TO
FAT, ALTITUDE, WIND, AND 50-FT OBSTACLE, THEN ENTER THE GRAPH BELOW WITH DERIVED VALUE
AND READ THE FLAPS-UP LANDING DISTANCE.
Figure 7-61. Landing Distance Without Propeller Reversing - Flaps 0%
7-90
TM 55-1510-221-10
CHAPTER 8
Normal Procedures
Section I.
MISSION PLANNING
8-l. MISSION PLANNING.
8-4. PERFORMANCE.
Mission planning begins when the mission is
assigned and extends to the preflight check of the
aircraft. It includes, but is not limited to, checks of
operating limits and restrictions; weight, balance,
and loading; performance: publications; flight plan;
and crew and passenger briefings. The pilot in command shall insure compliance with the contents of
this manual that are applicable to the mission.
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 variation 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.
8-2. OPERATING LIMITS AND RESTRICTIONS.
8-5. FLIGHT PLAN.
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.
8-3. WEIGHT, BALANCE, AND LOADING.
The aircraft must be loaded and weight and balance verified per Chapter 6. WEIGHT, BALANCE,
AND LOADING.
Section II.
A flight plan must be completed and filed per
AR 95-1, DOD FLIP, and local regulations.
8-6. CREW BRIEFINGS.
A crew briefing must be conducted for a thorough understanding of individual and team responsibilities. The briefing should include, but not be
limited to, copilot, crew chief, and ground crew
responsibilities and the coordination necessary to
complete the mission most efficiently. A review of
visual signals is desirable when ground guides do not
have a direct voice communications link with the
crew. Refer to Section VI for crew briefings.
OPERATlNG PROCEDURES AND MANEUVERS
8-7. OPERATING PROCEDURES AND MANEUVERS.
This section deals with normal procedures and
includes all steps necessary for safe and efficient
operation of the aircraft from the time a preflight
begins until the flight is completed and the aircraft
is parked and secured. Unique feel, characteristics,
and reaction of the aircraft during various phases of
operation and the techniques and procedures used
for taxiing, takeoff. climb. etc., are described,
including precautions to be observed. Only the
duties of the minimum crew necessary for the actual
operation of the aircraft are included. For operation
of avionics equipment. refer to the operating handbooks that accompany the aircraft loose tools.
8-8. ADDITIONAL DATA.
Additional crew duties are covered as necessary
in Section VI, CREW DUTIES. Mission equipment
checks are contained in Chapter 4, MISSION
EQUIPMENT. 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 III,
FLIGHT CHARACTERISTICS, and are repeated in
this section only when required for emphasis.
Checks that must be made 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.
8-1
TM 55-1519-221-10
8-9. CHECKLIST.
Normal procedures are given primarily in checklist form and are amplified as necessary in accompanying paragraph form when a detailed description of
a procedure or maneuver is required. A condensed
version of the amplified checklist, omitting all
explanatory text, is contained in the Operator’s and
Crewmember’s Checklist, TM 55-1510-221-CL. To
provide for easier cross referencing, the procedural
steps are numbered to coincide with the corresponding numbered steps in TM 55-1510-221-CL.
If high or gusty winds are present, and the
flight controls are unlocked, control surfaces may be damaged by buffeting.
* 3.
Flight controls - Unlock and check.
* 4.
Parking brake - Set.
8-10. USE OF CHECKLIST.
Although a good working knowledge of all aircraft procedures is desirable, it is not mandatory
that they be committed to memory. The pilot is
responsible for the initiation and accomplishment of
all required checks. Checklist items will be called
out orally and the action verified using the pilot’s
checklist (-CL). The copilot will normally read the
checklist and perform such duties as indicated, as
well as those directed by the pilot. ‘As required”
will not be used as a response; instead the actual
position or setting of the unit or item, such as “ON”
or “UP“ or “APPROACH’ will be stated. Upon
completion of each checklist, the copilot will advise
the pilot that the checklist called for has been completed.
The elevator trim system must not be
forced past the limits which are indicated
on the elevator trim tab position indicator.
5.
Elevator trim - Set to “0” (neutral).
Do not cycle landing gear handle on the
ground.
* 6.
Gear - DN.
* 7.
ICC vane pull handles - In.
8-11. CHECKS.
* 8.
Keylock switch - ON.
Items which apply only to night or only to
instrument flying shall have an "N" or “I” respectively, immediately preceding the check to which it
is pertinent. The symbol “O’ shall be used to indicate “if installed.’ Those duties which are the
responsibility of the copilot, at the command of the
pilot, will be indicated by a circle “O” around the
step number, i.e., ➃ Circuit breakers - In. The star
symbol " ★ " indicates an operational check contained
in the performance section of the condensed checklist. The asterisk symbol "*" indicates that performance of the step is mandatory for all thru-flights.
The asterisk applies only to checks performed prior
to takeoff. Placarded items appear in upper case.
* 9.
Battery switch - ON.
8-12. BEFORE EXTERIOR CHECK.
* 1.
Publications - Check DA Forms 2408-12,
-13, -14, and -18, DD Form 365-4, locally
required forms and publications, and availability of operator’s manual (-10) and checklist (-CL).
* 2.
Oxygen system - Check that oxygen quantity
is sufficient for the entire mission.
8-2
Change 4
10.
Ice vane switches - RETRACT.
11.
Lighting systems - Check as required, to
include navigation lights, recognition lights,
landing/taxi light, wing ice lights, beacons,
emergency lights, and interior lights, then
OFF.
NOTE
The emergency lights override switch
should be placed in the TEST position
and the emergency lights (5) checked for
illumination and intensity. A dim light
indicates a weak battery pack. At the
completion of the check, the switch must
be cycled from the TEST position to the
OFF/RESET position and then placed in
AUTO.
12.
Fuel gages - Check fuel quantity and
gage operation.
TM 55-1510-221-10
★ 13. Pitot tubes (2), stall warning vane, heated
fuel fuel vents (2) - Check.
1. Stall warning heat switch - ON.
2. Pitot heat switches (2) - ON. Check cover
removed.
3. Fuel vent heat switches (2) - ON.
4. Left wing heated fuel vent - Check by feel for
heat and condition.
5. Stall warning vane - Check by feel for heat
and condition.
6. Right wing heated fuel vent - Check by feel
for heat and condition.
7. Stall warning heat switch - OFF.
8. Pitot heat switches (2) - OFF.
9. Heated fuel vent switches (2) - OFF.
14.
Battery switch - As required.
Change 4
8-2.1/(8-2.2 blank)
TM 55-1510-221-10
1. Area 1 - Left wing, landing gear, engine, nacelle and propeller
2. Area 2 - Nose section
3. Area 3 - Right wing, landing gear, engine, nacelle and propeller
4. Area 4 - Fuselage, right side
5. Area 5 - Empennage
6. Area 6 - Fuselage, left side
AP 012082
Figure 8-1. Exterior Inspection
8-3
TM 55-1510-221-10
15. Mission equipment and circuit breakers Check and set.
16. Toilet - Check condition.
17. Emergency equipment - Check that all
required emergency equipment is available
and that fire extinguishers and first aid kits
have current inspection dates.
Parachutes - Check secure and for current
inspection and repack dates.
8-13.
EXTERIOR CHECK.
j. Outboard wing fuel vent - Check free of
obstructions.
k. Outboard deice boot - Check for secure
bonding, cracks, loose patches, stall strips, and general
condition.
1. Stall warning vane - Check free.
m. Monopole antenna - Check general
condition.
* n. Tiedown - Release.
8-14. FUEL SAMPLE.
NOTE
Fuel and oil quantity check may be
performed prior to EXTERIOR CHECK.
During warm weather open fuel cap slowly to
prevent being sprayed by fuel under pressure
due to thermal expansion.
* 1. Fuel sample - Check collective fuel sample from
a l l d r a i n s for possible contamination.
Thru-flight check is only required if aircraft has
been refueled.
8-15.
i. Main tank fuel and cap - Check fuel level
visually, condition of seal, and cap tight and properly
installed.
LEFT WING, AREA I.
1. Left wing area - Check as follows (fig. 8-l):
* a. General condition - Check for skin damage
such as buckling, splitting, distortion, dents, or fuel
leaks.
b. Flaps - C h e c k f o r f u l l r e t r a c t i o n
(approximately 0.25 inch play) and skin damage such as
buckling, splitting, distortion, or dents.
o. Inboard dipole antenna set - Check for
security and cracks at mounting points. Check bonding
secure, boots free of cuts and cracks.
p. Wing ice light - Check condition.
q. AC GPU access door - Secure.
r. Recessed and heated fuel vents - Check free of
obstructions.
s. Inverter inlet and exhaust louvers - Check
free of obstructions.
8-16. LEFT MAIN LANDING GEAR.
1. Left main landing gear - Check as follows:
* a. Tires - Check for cuts, bruises, wear, proper
inflation and wheel condition.
b. Brake assembly - Check brake lines for
damage or signs of leakage, brake linings for wear (0.25
inch maximum, between housing and lining carrier),
brake deice assembly and bleed air hose for condition
and security.
c. Fuel sump drains (3) - Check for leaks.
d. Controls and moveable trim tab - Check
security and moveable trim tab position.
* c. Shock strut - Check for signs of leakage,
minimum strut extension (5.56 inches), and that left and
right strut extension is approximately equal.
NOTE
All static wicks (27) must be installed for
optimum radio performance.
e. Static wicks (4) - Check security and
condition.
f. Wing pod, navigation lights, deice boots and
antennas - Check condition.
g. Recognition light - Check condition.
h. Outboard antenna set - Check condition.
8-4
Change 4
d. Torque knee - Check condition.
e. Safety switch - Check condition, wire, and
security.
★ f. Fire extinguisher pressure - Check pressure
within limits.
g. Wheel well, doors, and linkage - Check for
signs of leaks, broken wires, security, and general
condition.
TM 55-1510-221-10
h.
Fuel sump drains (forward) - Check for
leaks.
8-18. CENTER SECTION LEFT SIDE.
1.
8-17. LEFT ENGINE AND PROPELLER.
1.
Left engine - Check as follows:
A cold oil check is unreliable. Oil should
be checked within 10 minutes after stopping engine. If more than 10 minutes
have elapsed, motor engine for 30 seconds, then recheck. If more than 10 hours
have elapsed, run engine for 2 minutes,
then recheck. Add oil as required. Do not
overfill.
* a.
b.
Engine compartment, left side - Check
for fuel and oil leaks, security of oil
cap, door locking pin, and general condition.
NOTE
Secure front cowling latches first.
c.
Left cowl locks - Locked.
d.
Left exhaust stack - Check for cracks
and free of obstructions.
* e.
Propeller blades and spinner - Check
blade condition, boots, security of spinner and free propeller rotation.
* f.
Engine air inlets and ice vane - Check
free of obstruction and ice vane
retracted.
g.
h.
Bypass door - Check condition.
i.
Right exhaust stack - Check for cracks
and free of obstructions.
j.
Engine compartment, right side Check for fuel and oil leaks, ice vane
linkage, door locking pin, and general
condition. Lock compartment access
door.
b.
Auxiliary tank fuel sump drain - Check
for leaks.
C.
Inboard deice boot - Check for secure
bonding, cracks, loose patches, and
general condition.
* d.
Auxiliary tank fuel gage and cap Check fuel level visually, condition of
seal, and cap tight and properly
installed.
e.
Monopole antenna - Check condition.
8-19. FUSELAGE UNDERSIDE.
1.
Engine oil - Check oil level no more
than 2 quarts low, cap secure, locking
tab aft, and access door locked.
Center section - Check as follows:
a. Heat exchanger inlet and outlet - Check
for cracks and free of obstruction.
Fuselage underside - Check as follows:
* a. General condition - Check for skin
damage such as buckling, splitting, distortion, dents, or fuel leaks.
b.
Antennas - Check security, and general
condition.
8-20. NOSE SECTION, AREA 2.
1.
Nose section - Check as follows:
a.
Outside air temperature probe - Check
condition.
b.
Avionics door, left side - Check secure.
c.
Air conditioner exhaust - Check free of
obstructions.
d.
Wheel well - Check for signs of leaks,
broken wires and general condition.
e.
Doors and linkage - Check condition,
security, and alignment.
f.
Nose gear turning stop - Check condition.
* g.
Tire - Check for cuts, bruises, wear,
appearance of proper inflation, and
wheel condition.
* h.
Shock strut - Check for signs of leakage
and 3.0 inches minimum extension.
Right cowl locks - Locked.
i.
Torque knee - Check condition.
j.
Shimmy damper and linkage - Check
for security and condition.
k.
Landing and taxi lights - Check for
security and condition.
8-5
TM 55-1510-221-10
l.
Pitot tubes - Check covers removed,
alignment, security, and free of
obstructions.
m.
n.
Radome - Check condition.
Antenna pod - Check general condition.
Monopole antennas (2) - Check general
condition.
o.
p.
q.
r.
ping engine. If more than 10 minutes
have elapsed, motor engine for 30 seconds, then recheck. If more than 10 hours
have elapsed, run engine for 2 minutes,
then recheck. Add oil as required. Do not
overfill.
* a.
Engine oil - Check oil level, oil cap
secure (locking tab aft), and access
door locked.
b.
Engine compartment, left side - Check
for fuel and oil leaks, security of oil
cap, door locking pins, and general
condition.
c.
Left cowl locks - Locked.
d.
Left exhaust stack - Check for cracks
and free of obstructions.
* e.
Inboard deice boot - Check for secure
bonding, cracks, loose patches, and
general condition.
Battery access panel - Secure.
Propeller blades and spinner - Check
blade condition, boots, security of spinner, and free propeller rotation.
* f.
Engine air inlets and ice vane - Check
free of obstruction and ice vane
retracted.
Battery vents - Check free of obstruction.
g.
h.
Bypass door - Check condition.
i.
Right exhaust stack - Check for cracks
and free of obstructions.
j.
Engine compartment, right side Check for fuel and oil leaks, ice vane
linkage, door locking pins, and general
condition. Lock compartment access
door.
Windshields and wipers - Check windshield for cracks and cleanliness and
wipers for contact with glass surface.
Air conditioner inlet - Check free of
obstructions.
Avionics door, right side - Check
secure.
8-21. CENTER SECTION, RIGHT SIDE.
1.
Center section - Check as follows:
a.
b.
c.
* d.
e.
f.
g.
Auxiliary tank fuel gage and cap Check fuel level visually, condition of
seal, and cap tight and properly
installed (locking tab aft).
Battery compartment drain - Check
free of obstruction.
Battery ram air intake - Check free of
obstruction.
INS TAS temperature probe - Check
condition and free of obstructions.
h.
Auxiliary tank fuel sump drain - Check
for leaks.
i.
Heat exchanger inlet and outlet - Check
for cracks and free of obstructions.
j.
Monopole antenna - Check condition.
8-22. RIGHT ENGINE AND PROPELLER.
8-23. RIGHT MAIN LANDING GEAR.
1.
Right main landing gear - Check as follows:
a.
Fuel sump drains (forward) - Check for
leaks.
* b.
Tires - Check for cuts, bruises, wear,
proper inflation and wheel condition.
c.
Brake assembly - Check brake lines for
damage or signs of leakage, brake linings for wear (0.25 inch maximum,
between piston housing and lining carrier), brake deice assembly and bleed
air hose for condition and security.
* d.
Shock strut - Check for signs of leakage, minimum strut extension (5.50
inches), and that left and right strut
extension is approximately equal.
1. Right engine and propeller - Check as follows:
A cold oil check is unreliable. Oil should
be checked within 10 minutes after stop-
8-6
Right cowl locks - Locked.
TM 55-1510-221-10
e. Torque knee - Check condition.
f. Safety switch - Check condition, wire, and
security.
★ g. Fire extinguisher pressure - Check
pressure within limits.
h. Wheel well, doors, and linkage - Check
for signs of leaks, broken wires, security,
and general condition.
8-24. RIGHT WING, AREA 3.
1. Right wing - Check as follows:
* r. General condition - Check for skin
damage such as buckling, splitting,
distortion, dents, or fuel leaks.
8-25. FUSELAGE RIGHT SIDE, AREA 4.
1. Fuselage right side - Check as follows:
* a. General condition - Check for skin
damage such as buckling, splitting,
distortion or dents.
b. Emergency light - Check condition.
C. Flare/chaff dispenser - Check number of
flares in payload module and for security.
a. Recessed and heated fuel vents - Check
free of obstructions.
d. Beacon - Check condition.
b. Inverter inlet and exhaust louvers - Check
free of obstructions.
e. Underside fuselage antennas - Cheek
general condition.
c. DC GPU access door - Secure.
f. Towel bar antennas (2) - Check general
d. Inboard dipole antenna set - Check for
security and cracks at mounting points,
bonding secure, free of cuts and cracks.
g. P-band antenna - Check general condition.
e. Wing ice light - Check condition.
f. Outboard deice boot - Check for secure
bonding, cracks, loose patches, stall strips,
and general condition.
* g. Tiedown - Release.
* h. Main tank fuel and cap - Check fuel level
visually, condition of seal, and cap tight
and properly installed.
i. Outboard wing fuel vent - Check free of
obstructions.
j. Outboard antenna set - Check condition.
k. Recognition light - Check condition.
1. Wing pod navigation lights, deice boots
and antennas - Check condition.
m. Static wicks (4) - Check security and
condition.
n. Controls - Check security and condition of
ground adjustable tab.
o. Fuel sump drains (3) - Check for leaks.
p. Flaps - Check for full retraction
(approximately 0.25 inch play) and skin
damage, such as buckling, splitting,
distortion, or dents.
q . Chaff dispenser - Check number of chaffs
in payload module and for security.
h. Tailcone access door - Check secure.
i. Oxygen filler door - Check secure.
j. Static ports - Check clear of obstructions.
k. APR 44 antennas (2) - Check.
l. Emergency locator transmitter antenna Check condition.
8-26. EMPENNAGE, AREA 5.
1. Empennage - Check as follows:
a. Vertical stabilizer, rudder, and trim tab Check for skin damage, such as buckling,
distortion, or dents, and trim tab rig.
b. Static wicks (19) - Check installed.
c. Antennas - Check security, and general
condition.
d. Deice boots - Check for secure bonding,
cracks, loose patches, and general
condition.
e. Horizontal stabilizer, and elevator - Check
for skin damage, such as buckling,
distortion and dents.
NOTE
Any difference between the indicated position
on the trim tab position indicator and the actual
Change 4 8-7
TM 55-1510-221-10
(1) Safe arm and diaphram plunger Check position (lift door step).
position of the elevator trim tab signifies an
unairworthy condition and must be corrected
prior to the next flight of the aircraft.
(2) Index marks on rotary cam locks (6)
- Check aligned with indicator
windows.
f. Elevator trim tab - Verify “0” (neutral)
position.
b. Cargo door - Check closed and latched by
the following:
WARNING
(1) Upper handle - Check closed and
latched. (Observe through cargo door
latch handle access cover window.)
If the possibility of ice accumulation on the
horizontal stabilizer or elevator exists, takeoff
will not be attempted.
(2) Index marks on rotary cam locks (4)
- Check aligned with indicator
windows.
g . Position and beacon lights - Check
condition.
h.
(3) Lower pin latch handle - Check
closed and latched. (Observe through
cargo door lower latch handle access ,
cover window.)
Rotating boom dipole antenna - Check
condition and positron.
i . Wide band data link antenna pod - Check
for cracks and chips.
(4) Carrier rod - Check indicator aligned
with orange stripe on carrier rod.
(Observe through window aft lower
comer.)
8-27. FUSELAGE, LEFT SIDE, AREA 6.
1.
Fuselage left side - Check as follows:
(5) Battery switch - OFF.
* a. General condition - Check for skin
damage such as buckling, distortion, or
dents.
(6) Cargo door - Check closed and
latched.
(7) Cabin door - Close but leave
unlatched. Check CABIN DOOR
annunciator light illuminated.
b. Static ports - Check clear of obstructions.
c. ELT-ARMED.
(8) Cabin door - Open. Check CABIN
light
annunciator
extinguished
d. APR-44 antennas (2) - Check.
e. P band antenna - Check general condition.
(9) Battery switch - ON. Check CABIN
DOOR annunciator light illuminated.
f. Towel bar antennas (2) - Check general
condition.
(10) Cabin door - Close and latch. Check
CABIN DOOR annunciator light
extinguished.
g. Emergency light - Check condition.
h. Cabin door - Check door seal and general
condition.
i. Fuselage to side - Check general
condition and antennas.
* j. Chocks and tiedowns - Check removed.
8-28. INTERIOR CHECK.
(11) Battery switch - OFF.
3. Emergency exit - Check secure and key
removed.
4. Mission cooling ducts - Check open and free of
obstructions.
1.
Cargo/loose equipment - Check secure.
5. Flare/chaff dispenser preflight test Completed.
2.
Cabin/cargo doors - Test and lock:
6. KKY-28/58 key loaded - As required.
a. Cabin door - Check closed and latched by
the following.
8-8 Change 4
★ 7.
Crew briefing - As required. Refer to Section
VI.
TM 55-1510-221-10
8-29. BEFORE STARTING ENGINES.
★ 1.
Oxygen system - Check as required.
a . Oxygen supply pressure gages - Check.
b . Supply control lever (green) - ON.
c . Diluter control lever - 100% OXYGEN.
d . Emergency control lever (red) - Set to
TEST MASK position while holding mask
directly away from face, then return to
NORMAL.
* 3.
5. Magnetic compass - Check for fluid, heading
and current deviation card.
* 6. Pedestal controls - Set as follows:
CAUTION
Movement of power levers into reverse range
while engines are shut down may result in
bending and damage to control linkages.
a. POWER levers - IDLE.
e. Oxygen masks - Put on and adjust.
b. Propeller levers - HIGH RPM.
f. Emergency pressure control lever - Set to
TEST MASK position and check mask for
leaks, then return lever to NORMAL.
c. CONDITION levers - FUEL CUTOFF.
g. Flow indicator - Check, during inhalation
blinker appears, during exhalation blinker
disappears). Repeat a minimum of 3
times.
2.
c. Crossfeed switch - OFF.
Circuit breakers - Check in.
Overhead control panel switches - Set as
d. Flaps- UP.
e. Friction locks - Check and set.
* 7. Pedestal extension switches - Set as follows:
a. Flare/chaff dispenser control - SAFE.
b. Avionics - As required.
c. Rudder boost switch - ON.
a. Light dimming controls - As required.
b. Cabin temperature mode selector switch OFF.
c. Ice & rain switches - As required.
d. Exterior light switches - As required.
e. Master panel lights switch - As required.
f. Inverter switches - As required.
g. Avionics master power switch - As
required.
h. Environmental switches - As required.
9. Outside air temperature gage - Check, note
current reading.
10. Instrument panel - Check and set as follows:
a. Pilot’s and copilot’s course indicator
switches -As required.
b. Pilot’s and copilot’s RMI switches - As
required.
c. Pilot’s and copilot’s microphone switch As required.
i. Autofeather switches - OFF.
d. Pilot’s and copilot’s compass switches As required.
j. # 1 engine start switches - OFF.
e. Gyro switches - SLAVE.
k. Master switch - As required.
f. Flight instruments - Check instruments for
protective glass, warning flags (12 pilot, 6
copilot), static readings, and heading
correction card.
1. #2 engine start switches - OFF.
* 4.
8. Gear alternate engage and ratchet handles stowed.
Fuel panel switches - Check as follows:
a. Standby fuel pump switches - OFF.
b. Auxiliary transfer override switches AUTO.
g. Radar - OFF.
h. APR-39 and APR-44 - OFF.
i. Engine instruments - Check for protective
glass and static readings.
Change 4 8-9
TM 55-1510-221-10
20. External power advisory annunicator lights -
11. Deleted.
As required. (Aircraft EXTERNAL POWER
12. Mission panel switches and circuit breakers set and OFF.
13. Pressurization controls - Set.
14. Subpanels - Check and set as follows:
a Fire protection test switch - OFF.
b. Landing, taxi, and recognition lights OFF.
c. Landing gear control switch - Recheck
DN.
d. Cabin lights - As Required.
15. Pilot’s static air source - NORMAL.
NOTE
Do not use alternate static source during
takeoff and landing except in an emergency.
Pilot’s instruments will show a variation in
airspeed and altitude.
16. Pilot’s and copilot’s audio control panels - As
required.
17. Deleted.
★ 18. Fuel pumps/crossfeed operation - Check as
a . Fire pull handles - Pull.
b . Standby fuel pump switches - On.
c . Battery switch - ON.
d . #1 fuel pressure and #2 fuel pressure
warning lights - Illuminated
e. Fire pull handles - In.
f. #l fuel press and #2 fuel press warning
annunciator lights - Extinguished.
g. Standby fuel pump switches - Off.
h. #1 fuel pressure and #2 fuel pressure
warning lights - Illuminated.
i. Crossfeed - check. Check system
operation by
activating
switch
momentarily left then right, noting that #1
FUEL PRESS and #2 FUEL PRESS
warning annunciator lights extinguish and
that the FUEL CROSSPEED advisory
annunciator light illuminates as switch is
energized.
19. AC and DC GPU - As required.
8-10 Change 4
and mission EXT DC PWR ON annunciator
lights illuminated.)
21. DC power - Check. (22 VDC minimum for
battery, 28 maximum for GPU starts).
★ 22. Annunciator panels - Test as required.
a MASTER
CAUTION,
MASTER
WARNING, #1 FUEL PRESS, #2 FUEL
PRESS, GEAR DN, L BL AIR FAIL, R
BL AIR FAIL, INST AC, #1 DC GEN. #1
INVERTER, #1 NO FUEL XFR, #2 NO
FUEL XFR, #2 INVERTER, #2 DC GEN,
- Check illuminated.
b. ANNUNCIATOR TEST switch - Press
and hold. Check that the annunciator
panels, FIRE PULL handle lights, marker,
beacon lights, ANT Azimuth indicator,
MASTER CAUTION and MASTER
WARNING lights are illuminated. Release
switch and check that all lights except
those in step (1) are extinguished.
c. MASTER CAUTION and MASTER
WARNING lights - Press. Check that both
lights extinguish.
★ 23. Stall and gear warning system - Check as
follows:
a. STALL WARN TEST switch - TEST.
Check that warning horn sounds.
b. LDG GEAR WARN TEST switch TEST. Check that waming horn sounds
and that the LDG GEAR CONTR handle
warning lights (2) illuminate.
★ 24. Fire Protection system - Check as follows:
a . Fire Detector Test switch - Rotate
counterclockwise to check three DETR
positions. FIRE PULL handles should
Illuminate in each position. Reset
MASTER WARNING in each position.
b . Fire Detector Test switch - Rotate
counterclockwise to check two EXTGH
positions. SQUIB OK light, associated #1
EXTGH DISCH and #2 EXTGH DISCH
annunciator caution light and MASTER
CAUTION LIGHT should illuminate in
each position.
25. INS - Align as required.
8-30. * FIRST ENGINE START (BATTERY
START).
TM 55-1510-221-10
NOTE
7.
Oil pressure - Check (60 PSI minimum).
The engines must not be started until
after the INS is placed into the NAV
mode or OFF as required.
8.
Ignition and engine start switch - OFF, after
50% N1.
Starting procedures are identical for both
engines. When making a battery start, the right
engine should be started first. When making a
ground power unit (GPU) start, the left engine
should be started first due to the GPU receptacle
being located adjacent to the right engine. A crewmember should monitor the outside observer
throughout the engines start.
10.
1.
Avionics master switch - As required.
2.
Exterior light switches - As required.
9. CONDITION lever - HI IDLE. Monitor
TGT as the condition lever is advanced.
Generator switch - RESET, then ON.
8-31. SECOND ENGINE START (BATTERY
START).
First engine generator load 50% or less.
1.
2. Propeller - Clear.
3.
Ignition and engine start switch - ON. Propeller should begin to rotate and associated
#l IGN ON or #2 IGN ON annunciator
light should illuminate. Associated #l FUEL
PRESS or # 2 FUEL PRESS annunciator
light should extinguish.
3. Propeller - Clear.
4.
Ignition and engine start switch - ON. Propeller should begin to rotate and associated
#1 IGN ON or #2 IGN ON annunciator
light should illuminate. Associated #1 FUEL
PRESS or #2 FUEL PRESS light should
extinguish.
If ignition does not occur within 10 seconds after moving condition lever to
LOW IDLE, initiate engine clearing procedure. If for any reason a starting
attempt is discontinued, the entire starting sequence must be repeated after
allowing the engine to come to a complete
stop (5 minute minimum).
If ignition does not occur within 10 seconds after moving condition lever to
LOW IDLE, initiate engine clearing procedure. If for any reason a starting
attempt is discontinued, the entire starting sequence must be repeated after
allowing the engine to come to a complete
stop (5 minute minimum).
4. CONDITION lever (after N 1 RPM passes
12% minimum) - LOW IDLE.
5. CONDITION lever (after N 1 RPM stabilizes, 12% minimum) - LOW IDLE.
Monitor TGT to avoid a hot start. If
there is a rapid rise in TGT, be prepared
to abort the start before limits are
exceeded. During starting, the maximum
allowable TGT is 1000°C for five seconds.
If this limit is exceeded, use ABORT
START procedure and discontinue start.
Enter the peak temperature and duration
on DA Form 2408-13.
6. TGT and N1 - Monitor (TGT 1000°C maximum, N1 52% minimum).
Monitor TGT to avoid a hot start. If
there is a rapid rise in TGT, be prepared
to abort the start before limits are
exceeded. During starting, the maximum
allowable TGT is 1000°C for five seconds.
If this limit is exceeded, use ABORT
START procedure and discontinue start.
Enter the peak temperature and duration
on DA Form 2408-13.
5. TGT and N 1 - Monitor (TGT 1000°C maximum, N1 52% minimum).
6.
Oil pressure - Check (60 PSI minimum).
7.
Ignition and engine start switch - OFF after
50% N1.
8-11
TM 55-1510-221-10
8.
Battery charge light - Check (light should
illuminate approximately 6 seconds after
generator is brought on line. Light should
extinguish within 5 minutes following a normal engine start on battery).
9.
Inverter switches - ON, check INVERTER
annunciator lights extinguished.
10.
Second engine generator - RESET, then ON.
11.
CONDITION levers - As required.
8-32.
ABORT START.
1.
CONDITION lever - FUEL CUTOFF.
2.
Ignition and engine start switch - STARTER
ONLY.
3.
TGT - Monitor for drop in temperature.
4.
Ignition and engine start switch - OFF.
8-33.
CONDITION lever - FUEL CUTOFF.
2.
Ignition and engine start switch - OFF (5
minute minimum).
Do not exceed starter limitation of 30 seconds ON and 5 minutes OFF for two
starting attempts and engine clearing procedure. Allow 30 minutes off before additional starter operation.
4.
Ignition and engine start switch - STARTER
ONLY (15 seconds minimum, 30 seconds
maximum).
INS - As required.
NOTE
The engines must not be started until
after the INS is placed into the NAV
mode or OFF as required.
8-12
3.
Exterior light switches - As required.
4.
Propeller - Clear.
5.
Ignition and engine start switch - ON. Propeller should begin to rotate and associated
#1 IGN ON or #2 IGN ON should illuminate. Associated #1 FUEL PRESS or #2
FUEL PRESS warning annunciator light
should extinguish.
6.
CONDITION lever (after N1, RPM stabilizes, 12% minimum) - LOW IDLE.
Monitor TGT to avoid a hot start. If
there is a rapid rise in TGT, be prepared
to abort the start before limits are
exceeded. During engine start, the maximum allowable TGT is 1000°C for five
seconds. If this limit is exceeded, use
ABORT START procedure and discontinue start. Enter the peak temperature
and duration on DA Form 2408- 13.
7.
TGT and N1 - Monitor (TGT 1000°C maximum, N1 52% minimum).
8.
Oil pressure - Check (60 PSI minimum).
9.
Ignition and engine start switch - OFF after
50% N1.
10.
CONDITION lever - HI IDLE. Monitor
TGT as the condition lever is advanced.
11.
DC GPU - Disconnect as required,
12.
Generator switch (GPU disconnected) RESET, then ON.
Ignition and engine start switch - OFF.
8-34. * FIRST ENGINE START (GPU START).
1.
Avionics master switch - As required.
ENGINE CLEARING.
1.
3.
2.
8-35.
SECOND ENGINE START (GPU START).
1.
Propeller - Clear.
2.
Ignition and engine start switch - ON. Propeller should begin to rotate and associated
#1 IGN ON or #2 IGN ON annunciator
light should illuminate. Associated #1 FUEL
PRESS or #2 FUEL PRESS annunciator
light should extinguish.
TM 55-1510-221-10
CAUTION
If ignition does not occur within 10 seconds
after moving condition lever to LOW IDLE,
initiate engine clearing procedure. If for any
reason a starting attempt is discontinued, the
entire starting sequence must be repeated after
allowing the engine to come to a complete stop
(5 minute minimum).
a. Bleed air valves - OPEN.
b. CONDITION levers - HI IDLE.
c. Brake deice switch - DEICE. Check
BRAKE DEICE ON advisory annunciator
light illuminated.
* 2. Cabin temperature and mode - Set.
CAUTION
3.
CONDITION lever (after N1 RPM passes 12%
minimum) - LOW IDLE.
Verify airflow is present from aft cockpit
eyeball outlets to insure sufficient cooling for
mission equipment.
CAUTION
NOTE
Monitor TGT to avoid a hot start. If there is a
rapid rise in TGT, be prepared to abort the start
before limits are exceeded. During engine start,
the maximum allowable TGT is 1000°C for
five seconds. If this limit is exceeded, use
ABORT START procedure and discontinue
start. Enter the peak temperature and duration
on DA Form 2408-13.
For maximum cooling on the ground, turn the
bleed air valve switches to ENVIRO OFF
position.
4.
TGT and N 1 - Monitor (TGT 1000°C
maximum, N1 52% minimum).
5.
Oil pressure - Check (60 PSI minimum).
6.
Ignition and engine start switch - OFF, after
50% N1
7.
Propeller levers - FEATHER.
8.
GPU - Disconnect. (Check aircraft external
power and mission external power light
extinguished.)
9.
Propeller levers - HIGH RPM.
10. Aircraft inverter switches - ON, check #1
INVERTER and #2 INVERTER annunciator
lights extinguished.
11. Generator switches - RESET, then ON.
12. CONDITION levers - As required.
8-36. BEFORE TAXING.
* 1.
Brake deice - As required. To activate the
brake deice system proceed as follows:
★ * 3. AC/DC power - Check for:
a. AC frequency - 394 to 406 Hz.
b. AC voltage - 104 to 124 VAC.
c. DC load - Check.
d. DC voltage - 27.0 to 28.5 VDC.
WARNING
Do not operate radar in congested areas. Injury
could result to personnel in close proximity to
operating radar.
CAUTION
Do not operate the weather radar in an area
where the nearest effective surface is 50 feet or
less from the antenna reflector. Scanning such
surfaces within 50 feet of the antenna reflector
may damage receiver crystals.
* 4. Avionics master switch - ON.
5. Mission panel - Set and checked as required.
★ 6. Automatic flight control system - Check as
follows:
a. Altitude alert.
8-13
TM 55-1510-221-10
NOTE
Pause a few seconds between each step to
allow time for the proper indications.
(1)
Set alert controller more than 1000
feet above altitude indicated on
pilot’s altimeter. The pilot’s altimeter
alert light should be extinguished.
(2)
Decrease the alert controller to
within 1000 feet of the pilot’s
altimeter setting. The alert light
should illuminate.
(3)
Decrease the controller to less than
250 feet above the pilot’s altimeter
setting. The alert light should
extinguish.
(4)
Increase the controller to 300 ±50
feet above the pilot’s altimeter
indication and check that the alert
light illuminates.
(5)
Set the desired altitude.
b. Autopilot.
(1) Autopilot controller UP TRIM, DN
TRIM annunciators - CHECK not
illuminated.
CAUTION
A steady illumination of UP TRIM or DN
TRIM annunciator indicates that the automatic
synchronization is not functioning and the
autopilot should not be engaged.
(b) Deleted.
(7) Elevator trim follow-up - Check.
(a) Control wheel - Hold aft of mid
travel. Trim wheel should run
nose down after approximately 3
seconds. Trim down annunciator
should
illuminate
after
approximately 8 seconds.
(b) Control wheel - Hold forward of
mid travel. Trim wheel should
run nose up after approximately 3
seconds, trim up annunciator
should
illuminate
after
approximately 8 seconds, and AP
TRIM FAIL annunciator and
MASTER WARNING lights
should illuminate after
approximately 15 seconds.
The elevator trim system must not be forced
beyond the limits which are indicated on the
elevator trim tab indicator.
(8) AP/YD & TRIM DISC Button Depress through second level.
Autopilot and yaw damper should
disengage and ELECT TRIM O F F
annunciator should illuminate. AP
ENG and YD ENG annunciators on
instrument panel should flash 5
times.
(2)
Turn knob - Center.
(9) Elevator trim control switch - OFF,
then ON.
(ELEC TRIM OFF
annunciator should extinguish).
(3)
Elevator trim control switch - ON.
(10) AP - Re-engage.
(4)
Control wheel - Hold to mid travel.
(5)
AP button - Press. AP ENGAGE and
YD ENGAGE annunciators on.
(11) Turn controller - Check that control
wheel follows in each applied
direction, then center.
(6) Deleted.
(a) Deleted.
8-14 Change 4
TM 55-1510-221-10
(12) Pitch wheel - Check that trim responds
to pitch wheel movement. (UP TRIM
and DN TRIM annunciators may
illuminate).
(13) Heading bug - Center and engage HDG.
Check that control follows a turn in each
direction.
(14) Disengage AP by selecting GA. Check
that AP disengages and FD commands
7° nose up, wings level attitude. YD
disengage - Autopilot mode selector STBY.
7. Electric elevator trim - Check.
8-37. TAXIING.
CAUTION
Excessive use of brakes with the increased
weight of this aircraft will increase the
possibility of brake failure and/or brake fire.
Judicious use of the brakes is recommended
with coordinated use of beta range.
Taxi speed can be effectively controlled by the use of
power application and the use of the variable pitch
propellers in beta range. Normal turns may be made
with the steerable nose wheel; however, a turn may be
tightened by using full rudder and inside brake as
necessary. Turns should not be started with brakes
alone, nor should the aircraft be pivoted sharply on one
main gear.
a. Elevator trim switch - ON.
b. Pilot and copilot trim switches - Check
operation.
WARNING
Operation of the electric trim system should
occur only by movement of pairs of switches.
Any movement of the elevator trim wheel
while depressing only one switch element
denotes a system malfunction. The electric
elevator trim control switch must then be
turned OFF and flight conducted by
operating the elevator trim wheel manually.
Do not use autopilot.
1. Brakes - Check.
2. Flight instruments - Check for normal operation.
8-38. ENGINE RUNUP.
1.
*2.
Mission control panel - Set.
Propeller manual feathering - Check as follows:
a. CONDITION lever - LOW IDLE.
b. Left propeller lever - FEATHER Check that
propeller feathers.
(1) Pilot and copilot. Check individual
element for no movement of trim, then
check proper operation of both elements.
c. Left propeller lever - HIGH RPM.
d. Repeat procedure for right propeller.
(2) Check pilot switches override copilot
switches while trimming in opposite
directions, and trim moves in direction
commanded by pilot.
C. Check pilot and copilot trim disconnects
while activating trim.
d. Elevator trim switch - OFF then ON
annunciator
TRIM OFF
(ELECT
extinguishes).
*3.
Autofeather- Check as follows:
a. CONDITION levers - LOW IDLE.
b. AUTOFEATHER switch - Hold to TEST.
#2
AUTOFEATHER and
(#1
AUTOFEATHER advisory annunciator
lights should remain extinguished.)
8.
Avionics - Check and set as required.
until
Advance
c. P O W E R l e v e r s AUTOFEATHER lights are illuminated
(approximately 22% torque).
9.
INS - NAV mode, if on.
d. Left POWER lever - Retard and check:
10. Flaps - Check.
11.
Altimeters - Set and check.
torque
(1) At 16 - 21%
AUTOFEATHER light out.
-
#2
14% torque - Both
(2) At 9 AUTOFEATHER lights out and left
propeller starts to feather.
Change 4
8-15
TM 55-1510-221-10
e. Left POWER lever - Approximately 22%
torque.
f. Repeat steps b and d for right engine.
g. POWER levers - IDLE (both lights out,
neither propeller feathers).
*4. Overspeed governors - Check as follows:
a. POWER levers - Set approximately 1950
RPM (both engines).
b. #1 propeller governor test switch - Hold to
TEST position.
c. #1 propeller RPM 1830 to 1910 - Check.
d. Repeat steps b and c for # 2 engine.
e. POWER levers - Set 1800 RPM.
*5. Primary governors - Check as follows:
a. POWER levers - Set 1800 RPM.
b. Propeller levers - Move aft to detent. Check
that propeller RPM drops to 1600 to 1640
RPM.
c. Propeller levers - HIGH RPM.
c. Surface deice switch - SINGLE CYCLE
AUTO. Check for a drop in pneumatic
pressure and wing deice boot inflation and
after 6 seconds for a second drop in
pneumatic pressure.
d. Surface deice switch - MANUAL. Check that
surface boots inflate, and remain inflated,
then OFF.
e. Antenna deice switch - SINGLE. Check for a
drop in pneumatic pressure and antenna
deice boot inflation.
f. Antenna deice switch - MANUAL. Check that
boots inflate, and remain inflated, then OFF.
g. Engine inlet lip heat switches - ON. Check
that #1 LIP HEAT ON and #2 LIP HEAT ON
advisory annunciator lights are illuminated,
and the #1 LIP HEAT and #2 LIP HEAT
caution annunciator lights are extinguished,
then OFF.
h. RADOME ANTI-ICE switch - ON. Check
that RADOME HEAT annunciator is
illuminated, then OFF.
*10. Pneumatic pressure - Check as follows:
a. Left bleed air valve switch - PNEU &
ENVIRO OFF.
*6. Ice vanes - Check as follows:
b. Pneumatic pressure - Check 12 to 20 PSI.
a. Ice vane switches - EXTEND. Verify torque
drop, TGT increase, and illumination of #1
ICE VANE EXTEND and #2 ICE VANE
EXT annunciators.
b. Ice vane switches - RETRACT. Verify return
to original torque and TGT, and that #1 ICE
VANE EXTEND and #2 ICE VANE EXT
annunciators extinguish.
7. CONDITION levers - HI IDLE.
8. POWER levers - IDLE.
*9. Anti-ice and deice systems - Check as follows:
a. Windshield anti-ice switches - NORMAL and
HI. C h e c k P I L O T a n d C O P I L O T
(individually) for loadmeter rise, then OFF.
b. Propeller switches - INNER and OUTER
(momentarily). Check for loadmeter rise.
8-16
Change 4
c. Right pneumatic and environmental switch PNEU & ENVIRO OFF. Check that L BL
AIR FAIL and R BLAIR FAIL, and L BL AIR
OFF and R BL AIR OFF annunciator lights
are illuminated.
d. Pneumatic pressure - Verify zero.
TM 55-1510-221-10
e.
Left pneumatic and environmental switches OPEN. Check that L BL AIR FAIL and R BL
AIR FAIL, and L BL AIR OFF and R BL AIR
OFF annunciator lights are extinguished.
f.
Pneumatic pressure - Verify 12 to 20 PSI.
g.
Right pneumatic and environmental switches OPEN.
* 11. Pressurization system - Check as follows:
a.
Cabin door
extinguished.
caution
light
-
Check
b.
Storm windows - Check closed.
c.
Bleed air valve switches - Check OPEN.
d. Cabin altitude - Set 500 feet lower than
airfield elevation.
e.
Cabin pressure/dump switch - TEST (hold).
f.
Cabin rate-of-climb gage - Check for descending indication and, when confirmed, release
cabin pressure/dump switch from TEST.
g. Aircraft altitude - Set to planned cruise
altitude plus 500 feet. (If this setting does not
result in a CABIN ALT indication of at least
500 feet over takeoff field pressure altitude,
adjust as required).
12. CONDITION levers - As required
13. ANTI-ICE - As required.
NOTE
If windshield anti-ice is needed prior to takeoff,
use normal setting for a minimum of 15
minutes prior to selecting high temperature to
provide adequate preheating and minimize
effects of thermal shock.
8-39. BEFORE TAKEOFF.
0
1. Autofeather switch - ARM.
2. Bleed air valves - As required.
3. Ice & rain switches - As required.
0
0
0
0
4. Fuel panel - Check fuel quantity and switch
positions.
5. Flight and engine instruments - Check for normal
indications.
6.
Cabin altitude and rate-of-climb controller - Set.
7. Annunciator panels - Check (note indications).
8. Propeller levers - HIGH RPM.
9. Flaps - As required.
11. Avionics - Set.
12. Flight controls - Check
0
0
0
13. Departure briefing - Complete.
8-40. LINE UP.
1.
Transponder - As required.
2.
Engine autoignition switch - ARM.
3. Power stabilized - Check approximately 25%
minimum.
0
4.
CONDITION levers - LOW IDLE.
5.
Lights - As required.
6.
Mission control panel - Set.
8-41. TAKEOFF.
To aid in planning the takeoff and to obtain
maximum aircraft performance, make full use of the
information affecting takeoff shown in Chapter 7.
The data shown is achieved by setting brakes, setting takeoff power, and then releasing brakes. When
runway lengths permit, the normal takeoff may be
modified by starting the takeoff after power has
been stabilized at approximately 25% torque, then
applying power smoothly so as to attain full power
no later than 65 KIAS. This will result in a
smoother takeoff but will significantly increase
takeoff distance.
a. Normal Takeoff. After LINE Up procedures
have been completed, release brakes and smoothly
apply power to within 5% of target. Power should
be applied at a rate that will produce takeoff power
by 40 KIAS. Maintain directional control with
nosewheel steering rudder, and differential power,
while maintaining wings level with ailerons. The
pilot should retain a light hold on the power levers
throughout the takeoff and be ready to initiate
ABORT procedures if required. The copilot should
insure that the AUTOFEATHER advisory lights are
illuminated (if applicable), adjust and maintain power
at the exact takeoff power settings, and monitor all
engine instruments. The pilot will rotate at the
recommended rotation speed (Vr) and establish the
climb attitude (9° to 16°) that will attain best rateof-climb airspeed (Vy) during the initial climb.
Rotation should be at a rate that will allow liftoff
at liftoff airspeed (V1of).
b. Crosswind Takeoff. Position the aileron control
into the wind at the start of the takeoff roll to
maintain a wings level attitude. Under strong
crosswind conditions, leading with upwind power at
the beginning of the takeoff roll will assist in maintaining directional control. As the nosewheel comes
off the ground, the rudder is used as necessary to
10. Trim - Set.
Change 3
8-17
TM 55-1510-221-10
prevent turning (crabbing) into the wind. Rotate in
a positive manner to keep from side-skipping as
weight is lifted from the shock struts. To prevent
damage to the landing gear, in the event that the aircraft were to settle back onto the runway, remain in
“slipping” flight until well clear of the ground, then
crab into the wind to continue a straight flight path.
c.
Minimum Run Takeoff.
Spectacular takeoff performance can be
obtained by lifting off at speeds below
those recommended in Chapter 7. However, control of the aircraft will be lost if
an engine failure occurs immediately following liftoff until a safe speed can be
attained. Except during soft field takeoff,
liftoff below recommended speeds will
not be performed.
Minimum run takeoffs are performed with flaps
extended to 40% although at some conditions, use of
flaps during takeoff may result in the inability to
attain positive single-engine climb if an engine fails
immediately after liftoff.
To compensate for torque effect during the
beginning of the takeoff roll, align the aircraft with
the nose approximately 10° right of centerline. After
LINE UP procedures have been completed, hold
brakes firmly and apply TAKEOFF POWER, allowing for some increase in power as airspeed increases
during the takeoff roll. Copilot action is the same as
for normal takeoff. Release brakes and maintain
directional control and nosewheel steering and rudder. Do not use brakes unless absolutely necessary.
Hold the elevators in a neutral position, maintaining
wings level with ailerons. Allow the aircraft to roll
with its full weight on the wheels until the recommended rotation speed (V,) is reached. At this speed
rotate smoothly and firmly at a rate that will allow
liftoff at liftoff air speed (V1of). When flight is
assured, retract the landing gear.
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 obstacle clearance airspeed (V,). Climb at this speed until clear of
the obstacle. After the obstacle is cleared, lower the
nose slowly and accelerate to best rate-of-climb airspeed (V). Retract flaps after attaining single engine
best rate-of-climb airspeed (V yse).
8-18
NOTE
The best angle-of-climb speed (V,) is very
close to single engine power-off stall
speed. To provide for a margin of safety
in the event of engine failure immediately
after takeoff, the obstacle clearance airspeed value is used in lieu of true Vx, for
maximum angle takeoff climbs. Takeoff
performance data shown in Chapter 7 is
based on the use of obstacle clearance
climb speed.
e. Soft Field Takeoff. If a takeoff must be
made in conditions of mud, snow, tall grass, rough
surface or other conditions of high surface friction,
the following procedure should be used. Set flaps at
TAKEOFF (40%), align the aircraft with the runway,
and with the yoke held firmly aft, begin a slow
steady acceleration, avoiding rapid or transient
accelerations. Continue to hold full aft yoke so as to
transfer the weight of the aircraft from the wheels to
the wings as soon as possible. When the aircraft
rotates, control pitch attitude (nose) so as to lift off
from the soft surface at the slowest possible speed.
When airborne, level off immediately in ground
effect just above the surface, and accelerate to normal lift-off airspeed (V1of) before rotating to climb
attitude and retracting the landing gear. Consider
the effects of snow or mud on gear retraction as
applicable.
8-42. AFTER TAKEOFF.
Immediately after takeoff, the pilot flying
the aircraft should avoid adjusting controls located on the aft portion of the
extended pedestal to preclude inducing
spatial disorientation due to Coriolis illusion.
After the aircraft is positively airborne and flight
is assured, retract the landing gear. Adjust pitch attitude as required to maintain best rate-of-climb airspeed (V,). Retract flaps after attaining best singleengine rate-of-climb airspeed (V,,). The copilot
should continue to maintain power at the computed
setting and to monitor instruments. At single-engine
maneuvering altitude, adjust pitch attitude to obtain
cruise climb airspeed. As cruise climb airspeed, is
attained, adjust power to the climb power setting.
The copilot then activates the yaw damp and checks
that the cabin is pressurizing. Both pilots check the
wings and nacelles for fuel or oil leaks. The procedural steps after takeoff are as follows:
TM 55-1510-221-10
1. Power - Set. Refer to the cruise power graphs contained in Chapter 7. To account for ram air
temperature increase, it is essential that temperature be obtained at stabilized cruise airspeed.
1. Gear-up.
2. Flaps-UP.
3.
Landing lights - OFF.
4.
Climb power - Set.
5.
PROP SYNC Switch - As required
0
YAW DAMP switch - As required.
6.
NOTE
A new engine operated at the torque value
presented in the cruise power charts will show a
TGT margin below the maximum cruise limit.
Maximum cruise power settings for temperature
and altitude (derived from Chapter 7) if
exceeded will adversely affect engine life. With
ice vanes retracted, if cruise torque settings
shown on cruise power charts cannot be obtained without exceeding TGT limits, the engine should be inspected.
Autofeather switch - As required.
Brake deice - As required.
0
9
Windshield Anti-ice - As required.
NOTE
Turn windshield anti-ice on to normal when
passing 10,000 feet AGL or prior to entering the
freezing level (whichever comes first). Leave
on until no longer required during descent for
landing. High temperature may be selected as
required, after a minimum warm-up period of
15 minutes.
10. Cabin pressurization - Check, adjust RATE control
knob so that cabin rate-of-climb equals one-third
aircraft rate-of-climb.
0
0
11. Wings and nacelles - Check.
12. Flare/chaff dispenser safety pin (electronic
module) - Remove.
Chaff function selector switch - As required.
0
14. APR-39 and APR-44 - As required.
NOTE
Ice vanes must be extended when operating in
visible moisture at +5°C or less. Visible moisture is moisture in any form (clouds, ice crystals, snow, rain, sleet, hail, or any combination
of these).
0
0
0
3. Auxiliary fuel gages - Monitor. Insure that fuel is
being transferred from auxiliary tanks. (Chapter 2,
Section IV.)
4. Altimeters - Check. Verify that altimeter setting
complies with transition altitude requirement.
5. Engine instrument indications - Check all engine
instruments for normal indications.
8-43. CLIMB.
a. Cruise Climb. Cruise climb is performed at a
speed which is the best combination of climb, fuel
bum-off, and distance covered Set propellers at
1900 RPM and torque as required. Adhere to the
following airspeed schedule as closely as possible.
SL to 10,000 feet
10,000 to 20,000 feet
20,000 to 31,000 feet
2. Ice & rain switches - As required. Insure that antiice equipment is activated before entering icing
conditions.
140 KIAS
131 KIAS
121 KIAS
b. Climb - Maximum Rate. Maximum rate of
climb performance is obtained by setting propellers
at 2,000 RPM. torque at 100% (or maximum climb
TGT), and maintaining best rate-of-climb airspeed.
Refer to Chapter 7 for best rate-of-climb airspeed
for specific weights.
8-44. CRUISE.
Cruise power settings are entirely dependent upon
the prevailing circumstances and the type of mission
being flown. Refer to Chapter 7 for airspeed, power
settings, and fuel flow information. The following
procedures are applicable to all cruise requirements.
6. Recognition lights - As required.
8-45. DESCENT.
Descent from cruising altitude should normally be
made by letting down at cruise airspeed with
reduced power. Refer to Chapter 7 for power settings and rates of descent.
NOTE
Cabin altitude and rat-f-climb controller
should be adjusted prior to starting descent.
a. Descent - Max Rate (Clean). To obtain the
maximum rate of descent in clean configuration, perform the following:
0
1. Cabin pressurization - Set. Adjust CABIN CONTROLLER dial as required and adjust RATE control knob so that cabin rate of descent equals
one-third aircraft rate of descent.
Change 3
8-19
TM 55-1510-221-10
2. POWER levers - IDLE.
5. Altimeters - Set to current altimeter setting.
3. Propeller levers - HIGH RPM.
6. Flare/chaff dispenser arm-safe switch - SAFE.
4. Flaps - UP.
7. Flare/chaff dispenser safety pin (electronic
module) - Insert.
5. Gear - UP.
6. Airspeed - V
* 8. Crew briefing - Complete.
mo
.
0
7. Ice & rain switches - As requited.
8. Recognition lights - As required.
b. Descent - Max Rate (Landing Configuration).
If required to descend at a low airspeed (e.g., to
conserve airspace or in turbulence), approach flaps
and landing gear may be extended to increase the
rate and angle of descent while maintaining the
slower airspeed. To perform the maximum rate of
descent in landing configuration, perform the following:
0
1. Cabin pressurization - Set. Adjust CABIN CONTROLLER dial as requited and adjust RATE control knob so that cabin descent rate equals
one-third aircraft descent rate.
2. POWER levers - IDLE.
3. Propeller levers - HIGH RPM.
4. Flaps - APPROACH.
5. Gear - DN.
6. Airspeed - 180 KIAS maximum.
0
7. Ice & rain switches - As required.
8. Recognition lights - As required.
8-46. DESCENT-ARRIVAL.
Perform the following checks prior to the final
descent for landing.
Cabin pressurization - Set. Adjust CABIN
CONTROLLER dial as required.
Ice & rain switches - As required.
Windshield anti-ice - As required.
NOTE
Set windshield ANTI-ICE to normal or high as
required well before descent into icing conditions or into warm moist air to aid in defogging.
Turn off windshield anti-ice when descent is
completed to lower altitudes and when heating
is no longer required. This will preclude possible wind screen distortions.
4. Recognition lights - ON.
8-20
Change 3
8-47. BEFORE LANDING.
1. Propeller synchronization switch - OFF.
0
2. Autofeather switch - ARM.
3. Propeller levers - As required.
NOTE
During approach, propellers should be set to
1900 RPM to prevent glideslope interference
(ILS approach), provide better power response
during approach, and minimize attitude change
when advancing propeller levers for landing.
4. Flaps (below 198 KIAS) - APPROACH.
5. Gear - DN.
6. Landing lights - As required.
0
7. Brake deice - As required.
8-48. OBSTACLE CLEARANCE APPROACH AND
MINIMUM RUN LANDING.
When landing over obstacles that require a
steeper than normal approach path, or when greater
precision is required due to restricted runway
lengths, the “Power Approach/Precision Landing”
technique should be employed as follows: Prior to
intercepting the descent path, complete the LANDING check and stabilize airspeed (V ref at 1.2 times
power-off stall speed in landing configuration (V so).
After intercepting the desired approach angle maintain a constant descent by controlling the descent
with power and airspeed with elevator. Transition
smoothly from approach to landing attitude. Touchdown should be made on the main gear with the
nose slightly high, with power as required to control
rate of descent for a smooth touchdown. Immediately after touchdown. allow the nosewheel to make
ground contact and apply full reverse power and
braking, as required. If possible, remove reverse
thrust as the aircraft slows to 40 KIAS to minimize
propeller blade erosion.
NOTE
Using 1.2 x V so for approach airspeed will provide increased performance and more responsive controls however, performance data are not
available for approach at this slower airspeed.
TM 55-1510-221-10
8-49. LANDING.
CAUTION
Except in an emergency, propellers should be
moved out of reverse above 40 knots to
minimize propeller blade erosion, and during
crosswind to minimize stress imposed on
propeller, engine and airframe. Care must be
exercised when reversing on runways with
loose sand and/or dust on the surface. Flying
gravel will damage propeller blades and dust
may impair the pilot’s forward visibility at low
airplane speeds. Performance data charts for
landing computations assume that the runway
is paved, level and dry. Additional runway
must be allowed when these conditions are not
met. Refer to Chapter 7 for landing data. Do
not consider headwind during landing
computations; however, if landing must be
downwind, include the tailwind in landing
distance computations. Plan the final approach
to arrive at 50 feet over the landing area at
approach speed (V ref) plus 1/2 wind gust
speed. Perform the following procedures as the
an-craft nears the runway:
1. Autopilot and yaw damp - Disengaged.
2. Gear down lights - Check three green.
3. Propeller levers - HIGH RPM.
a. Normal. Landing. As the aircraft nears the
runway, flare slightly to break the rate or descent and
reduce power smoothly to idle as the nose. of the aircraft
is rotated to landing attitude. Avoid the tendency to ride
the ground effect cushion while waiting for the aircraft
to slow down to a soft landing. As the aircraft touches
down, gently lower the nosewheel to the runway and
use reversing, brakes, or beta range, as required. If
reversing is used, remove reverse power as the aircraft
slows to 40 KIAS to minimize propeller blade erosion.
b. Crosswind Landing. Refer to Chapter 7 for
recommended Vref speeds. Use the "crab-into-the
wind" method to correct for drift during final approach.
The "crab" is changed to a slip (aileron into wind and
top rudder) to correct for drift during flare and
touchdown. After landing, position ailerons as required
to correct for crosswind effect. For crosswind exceeding
the published limits, a combination "slip and crab"
method at touchdown should be used.
brakes unless absolutely necessary. Every precaution
must be taken to prevent the nose wheel from digging
into the surface.
d. Touch-and-Go Landings. The instructor should
select a point on the runway where all pretakeoff
procedures will have been completed prior to the pilot’s
initial application of power. In selecting this point,
prime consideration shall be given to the required
accelerate-stop distance pre-computed for the runway in
use. The nosewheel should be on the runway and rolling
straight before the touch-and-go procedures are
initiated. After the pilot applies power to within 5
percent of target, the copilot’s (instructor’s) actions are
the same as during a normal takeoff. If touch-and-go
landings are approved for training purposes use the
following procedure:
0
0
0
1
Propeller levers - HIGH RPM.
2
Flaps - As required.
3
Trim - Set.
4.
Power stabilized - Check approximately 25%
minimum.
5.
Takeoff power - Set.
8-50. GO-AROUND.
When a go-around is commenced prior to the
LANDING check, use power as required to climb to, or
maintain, the desired altitude and airspeed. If the
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 insuring that the
aircraft will not touch the ground. Retract the flaps to
APPROACH, adjusting pitch attitude simultaneously to
avoid an altitude loss. Accelerate to best rate-of-climb
airspeed (Vy) retracting flaps fully after attaining the
V ref speed used for the approach. Perform the
following:
1.
Power - As required.
2. Gear - UP.
3. Flaps - UP.
4.
Landing lights - OFF.
5.
Climb power - Set.
c. Soft Field Landing. When landing on a soft
unprepared surface such as mud, tall grass, or snow,
plan a normal power approach with flaps fully extended.
Decelerate to the slowest possible airspeed just prior to
touchdown, using power to control the final rate of
descent to as slow as possible. Do not stall prior to
touchdown as the nose attitude and rate of descent will
become unacceptable. On touchdown apply full back
(aft) elevator and then reduce power slowly. Do not use
8-21
TM 55-1510-221-10
0
6.
Yaw damp - As required.
Brake deice-OFF.
12.
Propeller levers - FEATHER.
13.
CONDITION levers - FUEL CUTOFF.
8-51. AFTER LANDING.
WARNING
Complete the following procedures after the
aircraft has cleared the runway:
Do not turn exterior lights off until propeller’s
rotation has stopped.
CONDITION levers - As required.
0
0
2.
Engine autoignition switch - OFF.
3
Ice & rain switches - OFF.
Flaps - UP.
Transponder - As required.
Radar - As required.
Lights - As required.
0
8.
Mission control panel - Set.
8-52. ENGINE SHUTDOWN.
CAUTION
14.
Exterior lights - OFF.
15.
Master panel lights switch - OFF.
16.
Avionics master switch - Off.
17.
Master switch - OFF. -
18.
Keylock switch - OFF.
19.
Oxygen system - OFF.
8-53. BEFORE LEAVING AIRCRAFT.
1.
Wheels - Chocked.
2.
Parking brake - As required.
NOTE
To prevent sustained loads on rudder shock
links, the aircraft should be parked with the
nose gear centered.
Brakes should be released after chocks are in
place (ramp conditions permitting).
1.
Brake deice - OFF.
3.
Flight controls - Locked.
2.
Parking brake - Set.
4.
Overhead flood lights - Off.
3.
Landing/taxi lights - OFF.
5.
Standby fuel pump switches - OFF.
4.
Cabin temperature mode selector switch OFF.
6.
Transponder - OFF.
5.
Autofeather switch - OFF.
7.
Mode 4 - As required.
6.
Vent and aft vent blower switches AUTO.
8.
KY-28/58 - Zeroize as required.
9.
Emergency exit lock - As required.
10.
Aft cabin light - OFF.
11.
Door light - OFF.
7.
INS - OFF.
8.
Mission equipment - OFF, as required.
9.
Inverter switches - OFF.
CAUTION
10. Battery condition - Check as required.
11. TGT - Check. TGT must be 660°C or below
for one minute prior to shutdown.
CAUTION
Monitor TGT during shutdown. If sustained
combustion is observed, proceed immediately
to ABORT START procedure.
8-22 Change 4
If strong winds are anticipated while the
aircraft is unattended, the propellers shall be
secured to prevent their windmilling with zero
engine oil pressure.
12. Walk-around inspection - Complete. Conduct
a thorough walk-around inspection. Checking
for damage, fluid leaks. and levels. Check that
covers, tiedowns. restraints.
TM 55-1510-221-10
safety pins and chocks are installed as
required.
13.
Aircraft forms - Complete. In addition to
established requirements for reporting any
system defects, unusual and excessive operation such as hard landings, etc., the flight
crew will also make entries on DA Form
2408-13 to indicate when limits in the Operator’s Manual have been exceeded.
14.
Aircraft - Secured. Lock cabin door as
required.
NOTE
A cold oil check is unreliable. Oil should
be checked within 10 minutes after stop
ping engine.
Section III. INSTRUMENT FLIGHT
8-54. GENERAL.
This aircraft is qualified for operation in instrument meteorological conditions. Flight handling,
stability characteristics and range are approximately
the same during instrument flight conditions as
when under visual flight conditions.
8-55. INSTRUMENT FLIGHT PROCEDURES.
Refer to FM 1-5, FM 1-230; FLIP; AR 95-1; FC
1-2 18; or applicable foreign government regulations,
and procedures described in this manual.
8-56. INSTRUMENT TAKEOFF.
Complete the BEFORE TAKEOFF check.
Engage the heading (HDG) mode on the autopilot
computer/control (do not engage autopilot). Set
heading marker (HDG) to runway heading and align
the aircraft with the runway centerline, insuring that
nosewheel is straight before stopping aircraft. Hold
brakes and complete the LINEUP check. Insure that
the roll steering bar is centered. Power application
and copilot duties are identical to those prescribed
for a “visual” takeoff. After the brakes are released,
initial directional control should be accomplished
predominantly with the aid of outside visual references. As the takeoff progresses, the crosscheck
should transition from outside references to the
flight director and airspeed indicator. The rate of
transition is directly proportional to the rate at
which the outside references deteriorate. Approaching rotation speed (V,), the crosscheck should be
totally committed to the instruments so that erroneous sensory inputs can be ignored. At rotation
speed, establish takeoff attitude on the flight director. Maintain this pitch attitude and wings-level attitude until the aircraft becomes airborne. When both
the vertical-velocity indicator and altimeter show
positive climb indications, retract the landing gear.
After the landing gear is retracted, adjust the pitch
attitude as required to attain best rate-of-climb airspeed (V,). Use PITCH-SYNC as required to reposi-
tion the flight director pitch steering bar. Retract
flaps after attaining best single-engine rate-of-climb
speed (V yse), and re-adjust pitch as required. Control
bank attitude to maintain the desired heading. Support flight director indications throughout the
maneuver by crosschecking “raw data“ information
displayed on supporting instruments.
NOTE
Due to possible precession error, the pitch
steering bar may lower slightly during
acceleration, causing the pitch attitude to
appear higher than actual pitch attitude.
To avoid lowering the nose prematurely,
crosscheck the vertical-velocity indicator
and altimeter to insure proper climb performance. The erection system will automatically remove the error after the acceleration ceases.
8-57. INSTRUMENT CLIMB.
Instrument climb procedures are the same as
those for visual climb. Enroute instrument climbs
are normally performed at cruise climb airspeeds.
8-58. INSTRUMENT CRUISE.
There are no unusual flight characteristics during cruise in instrument meteorological conditions.
8-59. INSTRUMENT DESCENT.
When a descent at slower than recommended
speed is desired, slow the aircraft to the desired
speed before initiating the descent. Normal descent
to approach altitude can be made using cruise airspeed. Normally, descent will be made with the aircraft in a cruise configuration, maintaining desired
speed as required.
8-23
TM 55-1510-221-10
8-60. INSTRUMENT APPROACHES.
There are no unusual preparations or control
techniques required for instrument approaches. The
approaches are normally flown at an airspeed of V ref
+20 until transitioning to visual flight.
8-61. AUTOPILOT APPROACHES.
There are no special preparations required for
placing the aircraft under autopilot control. Refer to
Chapter 3 for procedures to be followed for automatic approaches.
NOTE
The ILS localizer and glideslope warning
flags indicate insufficient signal strength
to the receiver. Certain electrical mechanical malfunctions between the receiver
and indicators may result in erroneous
localizer/glideslope information without a
warning flag. It is recommended that ILS
information be crosschecked with other
flight instruments prior to and during
final approach. Utilization of NAV TEST
on VOR control prior to the final
approach fix may detect certain malfunctions not indicated by the warning flags.
Section III. FLIGHT CHARACTERISTICS
8-62. STALLS.
A prestall warning in the form of very light buffeting can be felt when a stall is approached. An
aural warning is provided by a warning horn. The
warning horn starts sounding approximately five to
ten knots above stall speed with the aircraft in any
configuration. If correct stall recovery technique is
used, very little altitude will be lost during the stall
recovery. For the purpose of this section, the term
"power-on" shall mean that both engines and propellers of the aircraft are operating normally and
power is set at approximately 50%. The term
"power-off" shall mean that both engines are operating at idle power. Landing gear position has no
effect on stall speed. During practice, enter poweroff stalls from normal glides. Enter power-on stalls
by smoothly increasing pitch attitude to decrease
airspeed by approximately one knot per second until
stall occurs.
a. Power-On Stalls. The power-on stall attitude is very steep and unless this high-pitch attitude
is maintained, the aircraft will generally “settle“ or
“mush“ instead of stall. It is difficult to stall the aircraft inadvertently in any normal maneuver. A light
buffet precedes most stalls, and the first indication
of approaching stall is generally a decrease in control
effectiveness, accompanied by a “chirping“ tone
from the stall warning horn. The stall itself is characterized by a rolling tendency to the right, if the
aircraft is allowed to yaw. The proper use of rudder
will prevent the tendency to roll. A slight pitching
tendency will develop if the aircraft is held in the
stall, resulting in the nose dropping sharply, then
pitching up toward the horizon; this cycle is
repeated until recovery is made. Control is regained
8-24
very quickly with little altitude loss, providing the
nose is not lowered excessively. Begin recovery with
forward movement of the control wheel and a gradual return to level flight. The roll tendency caused
by yaw is more pronounced in power-on stalls, as is
the pitching tendency; however, both are easily controlled after the initial entry. Power-on stall characteristics are not greatly affected by wing flap position, except that stalling speed is reduced in
proportion to the degree of wing flap extension.
b. Power-Off Stalls. Power-off stalls are characterized by a right rolling tendency, as the stall is
approached. Elevator control is effective to the stop
and the pitch attitude can be maintained with a
deceleration rate of 1 knot/sec. Light to moderate
buffet commences approximately 5 - 8 knots above
the stall and the warning horn will sound and continue to the stall. With wing flaps down, the right
rolling tendency is more pronounced and stalling
speed is much slower than with the wing flaps up.
The Stall Speed Chart (fig. 8-2) shows the indicated
power-off stall speeds with the aircraft in various
configurations. Altitude loss during a full stall will
be approximately 800 feet.
c. Accelerated Stalls. The aircraft gives noticeable stall warning in the form of buffeting when the
stall occurs. The stall warning and buffet can be
demonstrated in turns by applying excessive back
pressure on the control wheel.
8-63. SPINS.
Intentional spins are prohibited. If a spin is
inadvertently entered use the following recovery
procedure:
Figure 8-2. Stall Speed
8-25
TM 55-1510-221-10
NOTE
Spin demonstrations have not been conducted. The recovery technique is based
on the best available information. The
first three actions should be accomplished
as nearly simultaneous as possible.
1.
POWER levers - IDLE.
2. Apply full rudder opposite the direction of
spin rotation.
3. Simultaneously with rudder application,
push the control wheel forward and neutralize ailerons.
4.
When rotation stops, neutralize rudder.
Do not pull out of the resulting dive too
abruptly as this could cause excessive
wing loads and a possible secondary stall.
5.
Pull out of dive by exerting a smooth, steady
back pressure on the control wheel, avoiding
an accelerated stall and excessive aircraft
stresses.
8-64. DIVING.
Maximum diving airspeed (red line) is 243
KIAS or 0.47 Mach. 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.
8-65. MANEUVERING FLIGHT.
Maneuvering speed (V,), 168 KIAS), is the maximum speed that abrupt control movements can be
applied without exceeding the design load factor of
the aircraft. The data is based on 15,000 pounds.
8-66. FLIGHT CONTROLS.
The aircraft is stable under all normal flight conditions. Aileron, elevator, rudder and trim tab controls function effectively throughout all normal flight
conditions. Elevator control forces are relatively
light in the extreme aft CG (center of gravity) condition, progressing to moderately high with CG at the
forward limit. Extending and retracting the landing
gear causes only slight changes in control pressure.
Control pressures, resulting from changes in power
settings or the repositioning of the wing flaps are not
excessive in the landing configuration at the most
forward CG position. The minimum speed at which
the aircraft can be fully trimmed is 89 KIAS (gear
and flaps down, propellers at high RPM, and 15,000
pounds power for a 3° angle of descent. Control
forces produced by changes in speed, power setting,
wing flap position and landing gear position are light
and can be overcome with one hand on the control
wheel. Trim tabs permit the pilot to reduce these
forces to zero. During single engine operation, the
rudder boost system aids in relieving the relatively
high rudder pressures resulting from the large variation in power.
8-67. LEVEL FLIGHT CHARACTERISTICS.
All flight characteristics are conventional
throughout the level flight speed range.
Section V. ADVERSE ENVIRONMENTAL CONDITIONS
8-68. INTRODUCTION.
The purpose of this part is to inform the pilot of
the special precautions and procedures to be followed during the various weather conditions that
8-26
may be encountered in flight. This part is primarily
narrative, only those checklists that cover specific
procedures characteristic of weather operations are
included. The checklist in Section II provides for
adverse environmental operations.
TM 55-1510-221-10
8-69. COLD WEATHER OPERATIONS.
Operation of the surface deice system in
ambient temperatures below -40°C can
cause permanent damage to the deice
boots. Operational diffIculties may be
encountered during extremely cold
weather, unless proper steps are taken
prior to or immediately after flight. All
personnel should understand and be fully
aware of the necessary procedures and
precautions involved.
a. Preparation For Flight. Accumulations of
snow, ice, or frost on aircraft surfaces will adversely
affect takeoff distance, climb performance and stall
speeds to a dangerous degree. Such accumulations
must be removed before flight. In addition to the
normal exterior checks, following the removal of ice,
snow, or frost, inspect wing and empennage surfaces
to verify that these remain sufficiently cleared. Also,
move all control surfaces to confirm full freedom of
movement. Assure that tires are not frozen to wheel
chocks or to the ground. Use ground heaters, antiice solution, or brake deice, to free frozen tires.
When heat is applied to release tires, the temperature should not exceed 71 °C (160°F). Refer to Chapter 2 for anti-icing, deicing, and defrosting treatment.
b. Engine Starting. When starting engines on
ramps covered with ice, propeller levers should be in
the FEATHER position to prevent the tires from
sliding. To prevent exceeding torque limits when
advancing CONDITION levers to HIGH IDLE during the starting procedure, place the power lever in
BETA and the propeller lever in HIGH RPM before
advancing the condition lever to HI IDLE.
Warm-Up and Ground Test. Warm-up proc.
cedures and ground test are the same as those outlined in Section II.
d. Taxiing. Whenever possible, taxiing in
deep snow, light weight dry snow or slush should be
avoided, particularly in colder OAT conditions. If it
is necessary to taxi through snow or slush, do not set
the parking brake when stopped. If possible, do not
park the aircraft in snow or slush deep enough to
reach the brake assemblies. Chocks or sandbags
should be used to prevent the aircraft from rolling
while parked. Before attempting to taxi, activate the
brake deice system, insuring that the bleed air valves
are OPEN and that the condition levers are in HI
IDLE. An outside observer should visually check
wheel rotation to insure brake assemblies have been
deiced. The condition levers may be returned to
LOW IDLE as soon as the brakes are free of ice.
e. Before Takeoff.
(1.) If icing conditions are expected, activate all anti-ice systems before takeoff, allowing sufficient time for the equipment to become effective.
(2.) If the possibility of ice accumulation
on the horizontal stabilizer or elevator exists, takeoff
will not be attempted. If icing conditions are
expected, activate all anti-ice systems before takeoff,
allowing sufficient time for the equipment to
become effective.
f. Takeoff. Takeoff procedures for cold
weather operations are the same as for normal takeoff. Taking off with temperatures at or below freezing, with water, slush or snow on the runway, can
cause ice to accumulate on the landing gear and can
throw ice into the wheel well areas. Such takeoffs
shall be made with brake deice on and with the ice
vanes extended to preclude the possibility of ice
build-up on engine air inlets. Monitor oil temperatures to insure operation within limits. Before flight
into icing conditions, the pilot and copilot WSHLD
ANTI-ICE switches should be set at NORMAL position.
g. During Flight.
(1.) Brake deice. After takeoff from a runway covered with snow or slush, it may be advisable
to leave brake deice ON to dislodge ice accumulated
from the spray of slush or water. Monitor BRAKE
DE-ICE annunciator for automatic termination of
system operation and then turn the switch OFF.
(2.) Flight controls. During flight, trim
tabs and controls should also be exercised periodically to prevent freezing.
(3.) Anti-icing equipment. Insure that antiicing systems are activated before entering icing conditions. Do not activate the surface deice system
until ice has accumulated one-half to one inch. The
propeller deice system operates effectively as an
anti-ice system and it may be operated continuously
in flight. If propeller imbalance due to ice does
occur, it may be relieved by increasing RPM briefly,
then returning to desired setting.
8-27
TM 55-1510-221-10
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 surface
boots and 1/8 to 1/4 inch of ice to accumulate on the antenna boots, then activate the deicer boots to remove the ice.
Repeat this procedure as required.
(4.) Ice vanes. Ice vanes must be extended
when operating in visible moisture or when freedom
from visible moisture cannot be assured, at 5°C
OAT or less. Ice vanes are designed as an anti-ice
system, not a deice system. After the engine air inlet
screens are blocked, lowering the ice vanes will not
rectify the condition. Ice vanes should be retracted
at 15°C OAT and above to assure adequate engine
oil cooling.
(5.) Stall speeds. Stalling airspeeds should
be expected to increase when ice has accumulated
on the aircraft causing distortion of the wing airfoil.
For the same reason, stall warning devices are not
accurate and should not be relied upon. Keep a comfortable margin of airspeed above the normal stall
airspeed. Maintain a minimum of 140 KIAS during
sustained icing conditions to prevent ice accumulation on unprotected surfaces of the wing. In the
event of windshield icing, reduce airspeed to 226
KIAS or below.
h. Descent. Use normal procedures in Section
II. Brake icing should be considered if moisture was
encountered during previous ground operations or
inflight in icing conditions with gear extended.
i. Landing. Landing on an icy runway should
be attempted only when absolutely necessary and
should not be attempted unless the wind is within
10 degrees of runway heading. Application of brakes
without skidding the tires on ice is very difficult,
due to the sensitive brakes. In order not to impair
pilot visibility, reverse thrust should be used with
caution when landing on a runway covered with
snow or standing water. Use the procedures in Section II for normal landing.
j. Engine Shutdown. Use normal procedures
in Section II.
k. Before Leaving the Aircraft. When the aircraft is parked outside on ice or in a fluctuating
freeze-thaw temperature condition the following
procedures should be followed in addition to the
normal procedures in Section II. After wheel chocks
are in place, release the brakes to prevent freezing.
Fill fuel tanks to minimize condensation, remove
any accumulation of dirt and ice from the landing
8-28
gear shock struts, and install protective covers to
guard against possible collection of snow and ice.
8-70. DESERT OPERATION AND HOT WEATHER
OPERATION.
Dust, sand, and high temperatures encountered
during desert operation can sharply reduce the operational life of the aircraft and its equipment. The
abrasive qualities of dust and sand upon turbine
blades and moving parts of the aircraft and the
destructive effect of heat upon the aircraft instruments will necessitate hours of maintenance if basic
preventive measures are not followed. In flight, the
hazards of dust and sand will be difficult to escape,
since dust clouds over a desert may be found at altitudes up to 10,000 feet. During hot weather operations, the principle difficulties encountered are high
turbine gas temperatures (TGT) during engine starting, over-heating of brakes, and longer takeoff and
landing rolls due to the higher density altitudes. In
areas where high humidity is encountered, electrical
equipment (such as communication equipment and
instruments) will be subject to malfunction by corrosion, fungi and moisture absorption by nonmetallic
materials.
a. Preparation For Flight. Check the position
of the aircraft in relation to other aircraft. Propeller
sand blast can damage closely parked aircraft. Check
that the landing gear shock struts are free of dust
and sand. Check instrument panel and general interior for dust and sand accumulation. Open main
entrance door and cockpit vent storm windows to
ventilate the aircraft.
N1 speeds of 70% or higher may be
required to keep oil temperature within
limits.
b. Engine Starting. Use normal procedures in
Section II. Engine starting under conditions of high
ambient temperatures may produce a higher than
normal TGT during the start. The TGT should be
closely monitored when the condition lever is
moved to the LO IDLE position. If overtemperature
tendencies are encountered, the condition lever
should be moved to the IDLE CUTOFF position
periodically during acceleration of gas generator
RPM (N1). Be prepared to abort the start before
temperature limitations are exceeded.
c. Warm-Up Ground Tests. Use normal procedures in Section II. To minimize the possibility of
damage to the engines during dusty/sandy condi-
TM 55-1510-221 -10
tlons. activate ICE VANES if the temperature is below
15°C.
d. Taxing.
Use normal procedures in Section II.
When practical. avoid taxiing over sandy terrain to
minimize propeller damage and engine deterioration that
results from Impingement of sand and gravel. During hot
weather operation. use minimum braking action to prevent
overheating.
e. Takeoff
Use normal procedures in Section II.
Avoid taking off in the wake of another alrcraft of the
runway surface is sandy or dusty.
f. During Flight.
II.
g. Descent.
Use normal procedures in Section
Use normal procedures in Section II.
altitude may be expected. The recommended penetration
speed in severe turbulence is 158 KIAS. Pitch attitude and
constant power settings are vital to proper flight technique.
Establish recommended penetration speed and proper
attitude prior to entering turbulent air to minimize most
difficulties. False Indications by the pressure Instruments
due to barometric pressure variations within the storm
make them unreliable. Maintaining a pre-established
attitude will result in a fairly constant airspeed. Turn
cockpit and cabin lights on to minimize the blinding effects
of lighting. Do not use autopilot altitude hold. Maintaln
constant power settings and pitch attitude regardless of
airspeed or altitude indications. Concentrate on maintaining
a level attitude by reference to the Flight Director/Attitude
Indicator. Maintain original heading. Maker no turns
unless absolutely necessary.
8-72. ICE AND RAIN (TYPICAL).
h. Landing. Use normal procedures in Section II.
WARNING
i. Engine Shutdown. Use normal procedures in
Section II.
While in icing conditions. if there is an
unexplained 30% increase of torque needed
to maintain airspeed in level flight, a
cumulative total of two or more inches of ice
accumulation on the wing, an unexplained
decrease of 15 knots IAS. or an unexplained
deviation between pilot’s and copilot’s
airspeed indicators, the icing environment
should be exited as soon as practicable. Ice
accumulation on the pitot tube assemblies
could cause a complete loss of airspeed
indication.
CAUTION
During hot weather. if fuel tanks are completely
filled. fuel expansion may cause overflow,
thereby creating a fire hazard.
j. Before Leaving Aircraft. Use normal procedures in
Section II. Take extreme care to prevent sand or dust from
entering the fuel and oil system during servicing. During
hot weather. release the brake immediately after installing
wheel chocks to prevent brake disc warpage.
8-71. TURBULENCE AND THUNDERSTORM
OPERATION.
CAUTION
Due to the comparatively light wing loading,
control in severe turbulence and thunderstorms
is critical. Since turbulence imposes heavy
loads on the aircraft structure. make all
necessary changes in aircraft attitude with the
least amount of control pressures to avoid
excessive loads on the aircraft’s structure.
Thunderstorms and areas of severe turbulence should be
avoided. If such areas are to be penetrated. it will be
necessary to counter rapid changes in attitude and accept
major Indicated altitude variations. Penetration should be
of an altitude which provides adequate maneuvering
margins as a loss or gain of several thousand feet of
The following conditions indicate a possible
accumulation of ice on the pitot tube assemblies and
unprotected aIrplane surfaces. if any of these conditions
are observed. the icing environment should be exited as
soon as practicable.
(1) Total ice accumulation of two inches or more on
the wing surfaces. Determination of ice thickness can be
accomplished by summing the estimated ice thickness on
the wing prior to each pneumatic boot deice cycle (e.g. four
cycles of minimum recommended 1/2-inch accumulation.
(2) A 30 percent Increase in torque per engine
required to maintain a desired airspeed in level flight (not
to exceed 85 percent torque) when operating at
recommended holding/loiter speed.
(3) A decrease in Indicated airspeed of 15 knots
after entering the icing condition (not slower than 1.4
power off stall speed) if maintaining original power setting
in Ievel flight. This can be determined by comparing pre-
Change 5
8-28.1
TM 55-1510-221-10
icing condition entry speed to the indicated speed after a
surface and antenna deice cycle is completed.
conditions are inadvertently encountered, turn on the antiicing systems prior to the first sign of ice formation.
(4) Any variations from normal indicated airspeed
between the pilot’s and copilot’s airspeed indicators.
(2) Deicer boots. Do not operate deicer boots
continuously. Allow at least one-half inch of ice on the
boots before activating the deicer boots to remove the ice.
Continued flight in severe icing conditions should not be
attempted. If ice forms on the wing area aft of the deicer
boots, climb or descend to an altitude where conditions are
less severe.
a. Typical Ice. Icing occurs because of supercooled
water vapor such as fog, clouds or rain. The most severe
icing occurs on aircraft surfaces in visible moisture or
precipitation with a true outside air temperature between
-5°C and +1°C; however, under some circumstances.
dangerous icing conditions may be encountered with
temperatures below -10°C. The surface of the aircraft
must be at a temperature of freezing or below for it to stick.
If severe icing conditions are encountered, ascend or
descend to altitudes where these conditions do not prevail.
If flight into icing conditions is unavoidable, proper use of
aircraft anti-icing and deicing systems may minimize the
problems encountered. Approximately 15 minutes prior to
flight into temperature conditions which could produce
frost or icing conditions, the pilot and co-pilot windshield
anti-ice switches should be set at normal or high
temperature position (after preheating) as necessary to
eliminate windshield ice. Stalling airspeeds should be
expected to increase when ice has accumulated on the
aircraft causing distortion of the wing airfoil. For the same
reason. stall warning devices are not accurate and should
not be relied upon. Keep a comfortable margin of airspeed
above the normal stall airspeed with ice on the aircraft.
Maintain a minimum of 140 knots during sustained icing
conditions to prevent ice accumulation on unprotected
surfaces of the wing. In the event of windshield icing,
reduce airspeed to 226 knots or below.
b. Rain. Rain presents no particular problems other
than restricted visibility and occasional incorrect airspeed
indications.
c. Taxiing. Extreme care must be exercised when
taxiing on ice or slippery runways. Excessive use of either
brakes or power may result in an uncontrollable skid.
d. Takeoff. Extreme care must be exercised during
takeoff from ice or slippery runways. Excessive use of
either brakes or power may result in an uncontrollable skid.
e. Climb. Keep aircraft attitude as flat as possible and
climb with higher airspeed than usual, so that the lower
surfaces of the aircraft will not be iced by flight at a high
angle of attack.
g. Landing. Extreme care must be exercised when
landing on ice or slippery runways. Excessive use of either
brakes or power may result in an uncontrollable skid. Ice
accumulation on the aircraft will result in higher stalling
airspeeds due to the change in aerodynamic characteristics
and increased weight of the aircraft due to ice buildup.
Approach and landing airspeeds must be increased
accordingly.
NOTE
When operating on wet or icy runways, refer to
stopping distance factors shown in Chapter 7.
8-72A. ICING (SEVERE).
a. The following weather conditions may be conducive
to severe in-flight icing:
(1) Visible rain at temperatures below zero degrees
Celsius ambient air temperature.
(2) Droplets that splash or splatter on impact at
temperatures below zero degrees Celsius ambient air
temperature.
b. The following procedures for exiting a severe icing
environment are applicable to all flight phases from takeoff
to landing.
(1) Monitor the ambient air temperature. While
severe icing may form at temperatures as cold as -18
degrees Celsius, increased vigilance is warranted at
temperatures around freezing with visible moisture present.
(2) Upon observing the visual cues specified in the
limitations section of the airplane flight manual (Military
Operations Manual) for the identification of severe icing
conditions (reference paragraph 5-30B.), accomplish the
following:
f. Cruise Flight.
(1) Prevention of ice formation. Prevention of ice
formation is far more effective and satisfactory than
attempts to dislodge the ice after it has formed. If icing
8-28.2 Change 5
(a) Immediately request priority handling from
air traffic control to facilitate a route or an altitude change
to exit the severe icing conditions in order to avoid
TM 55-1510-221-10
extended exposure to flight conditions more severe than
those for which the airplane has been certificated.
(b) Avoid abrupt and excessive maneuvering
that may exacerbate control difficulties
(c) Do not engage the autopilot.
(d) If the autopilot is engaged. hold the control
wheel firmly and disengage the autopilot.
(f) Do not cxtend flaps during extended
operation in icing conditions. Operations with flaps
extended can result in a reduced angle-of-attack. with the
possibility of ice forming on the upper surface further aft
on the wing than normal. possibly aft of the protected area.
(g) If the flaps are extended. do not retract
them until the airframe is clear of ice.
(h) Report these weather conditions to air
traffic control.
(e) If an unusual roll response or
uncommanded roll control movement is observed, reduce
the angle-of-attack.
Section VI. CREW DUTIES
* 8-73. CREW BRIEFING.
The following guide should be used in accomplishinp
required passenger briefings. Items that do not pertain to a
specific mission may be omitted.
7. Smoking
8. Oxygen.
9. Refueling.
a. Crew introduction.
10. Weapons and prohibited items.
b. Equipment.
11. Protective masks.
1. Personal, to include ID tags.
2. Professional (medical equipment. etc.).
3. Survival.
c. Flight data.
1. Route.
12. Toilet
e. Emergency procedures.
1. Emergency exits.
2. Emergency equipment
3. Emergency landing/ditching procedures.
2. Altitude.
* 8-74. DEPARTURE BRIEFING.
3. Time enroute.
4. Weather.
d. Normal procedures.
1. Entry and exit of aircraft.
The following is a guide that should be used as
applicable in accomplishing the required crew briefing
prior to takeoff. However. if the crew has operated
together previously and the pilot is certain that the copilot
understands all items of the briefing. he may omit the
briefing by stating “standard briefing.” when the briefing is
called for during the BEFORE TAKEOFF CHECK.
2. Seating and seat position.
a. ATC clearance - Review.
3. Seat belts.
1. Routing
4. Movement in aircraft.
5. Internal communications.
6. Security of equipment.
Change 5
8-29/(8-30 blank)
TM 55-1510-221-10
2. Initial altitude.
b.
Departure procedure - Review.
1.
c.
SID.
the copilot understands all items of the briefing, he
may omit the briefing by stating “standard briefing,“
when the briefing is called for during the
DESCENT-ARRIVAL check.
Weather/altimeter setting.
2. Noise abatement procedure.
a.
3. VFR departure route.
b. Airfield/facilities - Review.
Copilot duties - Review.
2. 1.
Field elevation.
1.
Adjust takeoff power.
2. Runway length.
2.
Monitor engine instruments.
3. Runway condition.
3.
Power check at 65 KIAS.
4.
Call out engine malfunctions.
1. Approach plan/profile.
5.
Tune/ident all nav/com radios.
2. Altitude restrictions.
6.
Make all radio calls.
3. Missed approach.
7.
Adjust transponder and radar as
required.
8.
Complete flight log during flight (note
altitudes and headings).
9.
Note departure time.
c. Approach procedure - Review.
a.
Point.
b. Time.
c.
4.
Intentions.
Decision height or MDA.
5. Lost communications.
d. PPC - Review.
1.
Takeoff power.
2. V
3.
d. Back up approach/frequencies.
e.
V y, (climb to 500’ AGL).
4. V
Copilot duties - Review.
r
1.
Nav/Com set-up.
2.
Monitor altitude and airspeeds.
yse
3. Monitor approach.
4.
8-75. ARRIVAL BRIEFING.
The following is a guide that should be used as
applicable in accomplishing the required crew briefing prior to landing; however, if the crew has operated together previously and the pilot is certain that
f.
Call out visual/field in sight.
Landing performance data - Review.
1.
Approach speed.
2. Runway required.
8-31/(8-32 blank)
TM 55-1510-221-10
CHAPTER 9
EMERGENCY PROCEDURES
Section I. AIRCRAFT SYSTEMS
9-1. AIRCRAFT SYSTEMS.
9-4. AFTER EMERGENCY ACTION.
This section describes the aircraft systems emergencies that may reasonably be expected to occur
and presents the procedures to be followed. Emergency procedures are given in checklist form when
applicable. A condensed version of these procedures
is in the Operator’s and Crewmember’s Checklist,
TM 55-1510-221-CL. Emergency operations of avionics equipment are covered when appropriate in
Chapter 3, Avionics, and are repeated in this section
only as safety of flight is affected.
After a malfunction has occurred, appropriate
emergency actions have been taken, and the aircraft
is on the ground, an entry shall be made in the
remarks section of DA Form 2408-13 describing the
malfunction.
9-2. IMMEDIATE
CHECKS.
ACTION
EMERGENCY
Immediate action emergency items are underlined for your reference and shall be committed to
memory. During an emergency, the checklist will be
called for to verify the memory steps performed and
to assist in completing any additional emergency
procedures.
NOTE
The urgency of certain emergencies
requires immediate action by the pilot.
The most important single consideration
is aircraft control. All procedures are subordinate to this requirement. Reset MASTER CAUTION after each malfunction
to allow systems to respond to subsequent
malfunctions.
9-3. DEFINITION OF LANDING TERMS.
The term LANDING IMMEDIATELY is
defined as executing a landing without delay. (The
primary consideration is to assure the survival of
occupants.) The term LAND AS SOON AS POSSIBLE is defined as executing a landing at the nearest
suitable landing area without delay. The term
LAND AS SOON AS PRACTICABLE is defined as
executing a landing to the nearest suitable airfield.
9-5. EMERGENCY EXITS AND EQUIPMENT.
Emergency exits and equipment are shown in
figure 9-1.
9-6. EMERGENCY ENTRANCE.
Entry may be made through the cabin emergency hatch. The hatch may be released by pulling
on its flush-mounted pull-out handle, placarded
EMERGENCY EXIT - PULL HANDLE TO
RELEASE. The hatch is of the nonhinged plug type
which removes completely from the frame when the
latches are released. After the latches are released,
the hatch may be pushed in.
9-7. ENGINE MALFUNCTION.
a. Flight Characteristics Under Partial Power
Conditions. There are no unusual flight characteristics during single-engine operation as long as airspeed is maintained at or above minimum control
speed (V,,) and above power-off stall speed. The
capability of the aircraft to climb or maintain level
flight depends on configuration, gross weight, altitude, and outside air temperature. Performance and
control will improve by feathering the propeller of
the inoperative engine, retracting the landing gear
and flaps, and establishing the appropriate singleengine best rate-of-climb speed (V yse). Minimum
control speed (V,,) with flaps retracted is approximately 1 knot higher than with flaps at takeoff (40%)
position.
b. Engine Malfunction During And After
Takeoff. The action to be taken in the event of an
engine malfunction during takeoff depends on
whether or not liftoff speed (V 1of) has been attained.
If an engine fails immediately after liftoff, many
variables such as airspeed, runway remaining, air-
9-1
TM 55-1510-221-10
Figure 9-1. Emergency Exits and Equipment
9-2
TM 55-1510-221-10
craft weight, altitude at time of engine failure, and
single-engine performance must be considered in
deciding whether it is safer to land or continue flight.
c. Engine Malfunction Before Liftoff (Abort). If an
engine fails and the aircraft has not accelerated to
recommended liftoff speed (V 1of), retard power levers
immediately to IDLE and stop the aircraft with brakes
and reverse thrust. Perform the following:
1. Power - Maximum controllable.
NOTE
If airspeed is below V yse maintain whatever
airspeed has been attained (between V 1of and
V yse) until sufficient altitude can be obtained
to trade off altitude for airspeed to assist in
acceleration to V yse.
1. Power levers - IDLE.
2. Gear - UP
2. Braking - As required.
3. Flaps - UP.
NOTE
If able to land on remaining runway, check
gear down and use brakes and reverse
thrust as required. If insufficient runway
remains for stopping, perform the following:
3. Condition levers - FUEL CUTOFF.
4. Fire pull handles - Pull.
5. Master switch - OFF.
d. Engine Malfunction After Liftoff. If an engine
fails after becoming airborne, maintain single-engine
best rate-of-climb speed (V yse) or, if air speed is below
V yse maintain whatever airspeed is attained between
liftoff (V 1of) and V yse until sufficient altitude is attained
to trade altitude for-airspeed and accelerate to V yse.
(1.) Engine Malfunction after liftoff (abort),
perform the following and land in a
wingslevel attitude:
1. Power levers - Reduce.
2. Gear - DN.
3. Complete normal landing.
NOTE
4. Landing lights - OFF.
5. Brake deice - OFF.
6. Engine cleanup - Perform.
7. Generator load - 100% max.
NOTE
Holding three to five degrees bank (one-half
ball width) towards the operating engine will
assist in maintaining directional control and
improve aircraft performance.
e. Engine Malfunction During Flight. If an engine
malfunctions during cruise flight, maintain control of the
aircraft while maintaining heading or turn as required.
Add power as required to keep airspeed from decaying
excessively and to maintain altitude. Identify the failed
engine by feel (if holding rudder pressure to keep the
aircraft from yawing; the rudder being pressed indicates
the good engine) and engine instruments, then confirm
identification by retarding the power lever of the
suspected failed engine. Refer to Chapter 7 for
single-engine cruise information. If one engine
malfunctions during flight, perform the following:
1. Autopilot/yaw damp - DISENGAGE.
2. Power - As required.
3. Dead engine - Identified.
If able to land on remaining runway, check
gear down and use brakes and reverse
thrust as required. If insufficient runway
remains for stopping, perform the following:
4. Power lever (dead engine) - IDLE.
5. Propeller lever (dead engine) - FEATHER.
4. Condition levers - FUEL CUTOFF
6. Propeller synchronization switch - Off.
5. Fire pull handles - Pull.
7. Gear - As required.
6. Master switch - OFF.
8. Flaps - As required.
(2.) Engine malfunction after liftoff (flight
continued) perform the following:
9. Generator load - 100% max.
10. Power - Set for single engine cruise.
Change 4 9-3
TM 55-1510-221-10
11. Engine cleanup - Perform.
NOTE
At V yse, speeds, holding three to five degrees
bank (one-half ball width) towards the
operating engine will assist’ in maintainin
directional control and improve aircraft
performance.
f. Engine Malfunction During Final Approach. If
an engine malfunctions during final approach (after
LANDING CHECK) the propeller should not be
manually feathered unless time and altitude permit or
conditions require it. Continue approach using the
following procedure:
1. Power-As required
2. Gear - DN.
g. Engine Malfunction (Second Engine). If the
second engine fails, do not feather the propeller if an
engine restart is to be attempted. Engine restart without
starter assist can not be accomp lished with a feathered
propeller, and the propeller Will not unfeather without
the engine operating 140 KIAS is recommended as the
best all around glid e speed (considering engine restart,
distance covered, transition to landing configuration,
etc.), although it does not necessarily result in the
minimum rate of descent. Perform the following
procedure if the second engine fails during cruise flight.
1. Airspeed - 140 KIAS.
2. Powerlever - IDLE
3. Propeller lever - Do not FEATHER.
4.
Conduct engine restart procedure.
h. Engine Shutdown In Flight. If it becomes
necessary to shut an engine down during flight, perform
the following:
1.
Power lever - IDLE.
2.
Propeller lever - FEATHER.
3.
Condition lever - FUEL CUTOFF.
4.
Engine clean up - Perform.
i. Engine Cleanup. The clean up procedure to be
used after engine malfunction, shutdown, or an
unsuccessful restart is as follows:
9-4
1.
Autoignition switch - OFF.
2.
Autofeather switch - OFF.
3.
Generator switch - OFF.
4.
Propeller synchronization switch - OFF.
Change 4
j. Engine Restart During Flight Using Starter.
Engine restarts may be attempted at all altitudes. If
restart is attempted, perform the following:
1.
Cabin temperature mode selector switch - OFF.
2.
Electrical load - Reduce to minimum.
3.
Fire pull handle - In.
4.
Power lever - IDLE.
5.
Propeller lever - FEATHER.
6.
Condition lever - FUEL CUTOFF.
7. TGT (operative engine) - 700°C or less.
8.
Ignition and engine start switch - ON.
9.
Condition lever - LOW IDLE.
NOTE
If a rise in TGT does not occur within 10
seconds after moving the condition lever to
LOW IDLE, abort the start.
10. TGT - Monitor (1,000°C for 5 seconds
maximum).
11. Oil pressure - Check.
12. Ignition and engine start switch - OFF at 50%
N1.
13. Generator switch - RESET, then ON.
14. Engine cleanup - Perform if engine restart
unsuccessful.
15. Cabin temperature mode selector switch - As
required.
16. Electrical equipment - As required.
17. Autoignition switch - ARM.
18. Propellers - Synchronize.
19. Power - As required.
k. Engine Restart During Flight (Not Using
Starter). A restart without starter assist may be
accomplished provided airspeed is at or above 140
KIAS altitude is below 20,000 feet, and the propeller is
not feathered. If altitude permits, diving the aircraft will
increase N1 and assist in restart. N1 required for airstart
should be at or above 9%. If a start is attempted,
perform following:
1.
Cabin temperature mode selector switch - OFF.
TM 55-1510-221-10
2.
Electrical load - Reduce to minimum.
3.
Generator switch (affected engine) OFF.
4.
Fire pull handle - Check in.
5.
Power lever - IDLE.
6.
Propeller lever - HIGH RPM.
7.
Condition lever - FUEL CUTOFF.
8.
Airspeed - 140 KIAS minimum.
9.
Altitude below 20,000 feet - Check.
10.
Engine autoignition switch - ARM.
11.
Condition lever - LOW IDLE.
NOTE
If N, is below 12%, starting temperatures
tend to be higher than normal. To preclude overtemperature (1000°C or above)
during engine acceleration to idle speed,
periodically move the condition lever into
FUEL CUTOFF position as necessary.
NOTE
If a rise in TGT does not occur within 10
seconds after moving the condition lever
to LOW IDLE, abort the start.
12.
TGT - Monitor (1 ,OOO”C for 5 seconds
maximum).
13.
Oil pressure - Check.
14.
Generator switch - RESET, then ON.
15.
Engine Cleanup - Perform if engine
restart unsuccessful.
16.
Cabin temperature mode selector
switch - As required.
17.
Electrical equipment - As required.
18.
Autoignition switch - ARM.
19.
Propellers - Synchronized.
20.
Power - As required.
l. Maximum Glide. In the event of failure of
both engines, maximum gliding distance can be
obtained by feathering both propellers to reduce
propeller drag and by maintaining the appropriate
airspeed with the gear and flaps up. Figure 9-2 gives
the approximate gliding distances in relation to altitude.
m. Landing With Two Engines Inoperative.
Maizain best glide speed (figure 9-2). If sufficient
altitude remains after reaching a suitable landing
area, a circular pattern will provide best observation
of surface conditions, wind velocity, and direction.
When the condition of the terrain has been noted
and the landing area selected, set up a rectangular
pattern. Extending APPROACH flaps and landing
gear early in the pattern will give an indication of
glide performance sooner and will allow more time
to make adjustments for the added drag. Fly the
base leg as necessary to control point of touchdown.
Plan to overshoot rather than undershoot, then use
flaps as necessary to arrive at the selected landing
point. Keep in mind that, with both propellers feathered the normal tendency is to overshoot due to less
drag. In event a positive gear-down indication cannot be determined, prepare for a gear-up landing;
also, unless the surface of the landing area is hard
and smooth, the landing should be made with the
landing gear up. If landing on rough terrain, land in
a slightly tail-low attitude to keep nacelles from possibly digging in. If possible, land with flaps fully
extended.
9-8. LOW OIL PRESSURE.
In the event of a low oil pressure indication, perform the procedures below as applicable:
1.
Oil pressure below 105 PSI below 21,000
feet or 85 PSI 21,000 feet and above, torque
- 49% maximum.
2.
Oil pressure below 60 PSI - Perform engine
shutdown, or land as soon as practicable
using minimum power to insure safe arrival.
9-9. CHIP DETECTOR WARNING LIGHT ILLUMINATED.
If a L CHIP DETR or a R CHIP DETR warning
light illuminates, and safe single-engine flight can be
maintained; perform engine shutdown.
9-10. DUCT OVERTEMP CAUTION ANNUNCIATOR LIGHT ILLUMINATED.
If a DUCT OVERTEMP caution annunciator
light is illuminated, insure that the cabin floor outlets are open and unobstructed, then perform the
following steps in sequence until the light is extinguished. After completion of steps 1 thru 4, if light
does not extinguish, Allow approximately 30 seconds after each adjustment for the system temperature to stabilize. The overtemperature condition is
considered corrected at any point during the procedure that the light extinguishes.
1.
Cabin air control - In.
9-5
TM 55-1510-221-10
Figure 9-2. Maximum Glide Distance
9-6
TM 55-1510-221-10
2.
Cabin temperature mode selector switch AUTO.
3.
Cabin temperature control rheostat - Full
decrease.
4.
Vent blower switch - HI.
5.
Cabin temperature mode selector switch MAN COOL.
6.
Manual temperature switch - DECREASE
(hold).
7.
Left bleed air valve switch - ENVIRO OFF.
8.
If the light is still illuminated after 30 seconds: Left bleed air valve switch - OPEN.
9.
Right bleed air valve switch - ENVIRO
OFF.
10.
If the light is still illuminated after 30 seconds: Right bleed air valve switch - OPEN.
NOTE
If the overtemperature light has not extinguished after completing the above procedure, the warning system has malfunctioned.
3.
Ice vane - Operate manually.
4.
Airspeed - Resume normal airspeed.
9-12.
ENGINE BLEED AIR SYSTEM FAILURE.
a. Bleed Air Failure Light Illuminated. Steady
illumination of the warning light in flight indicates
a possible ruptured bleed air line aft of the engine
firewall. The light will remain illuminated for the
remainder of flight. Perform the following:
NOTE
L BL AIR FAIL or R BL AIR FAIL lights
illuminate during
may momentarily
simultaneous surface deice and brake
deice operation at low N1 speeds.
1.
Brake deice - OFF.
2.
TGT and torque - Monitor (note readings).
3. Bleed air valve switch - PNEU &
ENVIRO OFF.
NOTE
9-11. ICE VANE FAILURE.
Ice vane failure is indicated by #1 VANE FAIL
or #2 VANE FAIL caution annunciator light illumination. If an ice vane fails to operate electrically,
perform the following:
Brake deice on the affected side, and rudder boost, will not be available with bleed
air valve switch in PNEU & ENVIRO
OFF.
4. Cabin pressurization - Check.
b. Excessive Differential Pressure. If cabin differential pressure exceeds 6.1 PSI, perform the following:
After the ice vanes have been manually
extended, they must be mechanically
retracted. No electrical extension or
retraction shall be attempted as damage
to the electric actuator may result. Linkage in the nacelle area must be reset prior
to operation of the electric system. Do not
reset ice vane control circuit breaker.
Do not retract ice vanes electrically after
manual extension.
1.
Airspeed - 160 KIAS or below.
2.
Ice vane control circuit breaker - Pull.
1.
Cabin altitude and rate-of-climb controller -Select higher setting.
2.
If condition persists: LEFT BLEED
AIR VALVE switch - ENVIRO OFF
(light illuminated).
3.
If condition still persists: RIGHT
BLEED AIR VALVE switch ENVIRO OFF (light illuminated).
4.
If condition still persists - Descend
immediately.
5.
If unable to descend: CABIN PRESS
PRESS
-CABIN
DUMP
switch
DUMP.
6.
Bleed air valve switches - OPEN, if
cabin heating is required.
9-7
TM 55-1510-221-10
9-13. LOSS OF PRESSURIZATION (ABOVE 10,
000 FEET).
If cabin pressurization is lost when operating
above 10,000 feet or the ALT WARN warning
annunciator light illuminates, perform the following:
1.
Crew oxygen masks - 100% and on.
5. Yaw damp - OFF.
6. Brake deice - OFF.
9-17. SINGLE-ENGINE LANDING CHECK.
Perform the following procedure during final
approach to runway.
9-14. CABIN DOOR CAUTION LIGHT ILLUMINATED.
1.
Autopilot/yaw damp - Disengaged.
2.
Gear lights - Check (three green).
Remain clear of cabin door and perform the following:
3. Propeller lever (operative engine) - HIGH
RPM.
1.
Bleed air valve switches - ENVIRO OFF.
2.
Descend below 14,000 feet as soon as practicable.
3. Oxygen - As required.
NOTE
To insure constant reversing characteristics, the propeller control must be in the
HIGH RPM position.
9-15. SINGLE-ENGINE DESCENT/ARRIVAL.
9-18. SINGLE-ENGINE GO-AROUND.
NOTE
Approximately 85% N1 is required to
maintain pressurization schedule.
Perform the following procedure prior to the
final descent for landing.
1.
Cabin controller - Set.
2.
Ice and rain switches - As required.
3. Altimeters - Set.
4.
* 5.
Recognition lights - ON.
The decision to go around must be made as
early as possible. Elevator forces at the start of a goaround are very high and a considerable amount of
rudder control will also 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. Insure that the aircraft does not touch
the ground before retracting the landing gear.
Retract the flaps only as safe airspeed permits
(TAKEOFF until V ref, then UP). Perform singleengine go-around as follows:
Arrival briefing - Complete.
9-16. SINGLE-ENGINE BEFORE LANDING.
1.
Propeller lever - As required.
Once flaps are fully extended, a singleengine go-around may not be possible
when close to the ground under conditions of high gross weights and/or high
density altitude.
NOTE
During approach, propeller should be set
at 1900 RPM to prevent glideslope interference (ILS approach), provide better
power response during approach, and to
minimize attitude change when advancing
propeller levers for landing.
1.
Power - Maximum allowable.
2. Gear - UP.
3.
Flaps - As required.
2. Flaps - APPROACH.
4. Landing lights - OFF.
3. Gear - DN.
5. Power - As required.
4.
6.
9-8
Landing lights - As required.
Yaw damp - As required.
TM 55-1510-221-10
3.
9-19. PROPELLER FAILURE (OVER 2080 RPM).
If an overspeed condition occurs that cannot be
controlled with the propeller lever, or by reducing
power, perform the following:
1.
Power lever (affected engine) - IDLE.
2.
Propeller lever - FEATHER.
3.
Condition lever - As required.
4. Propeller synchronization - OFF.
5.
Engine cleanup - As required.
9-20. FIRE.
The safety of aircraft occupants is the primary
consideration when a tire occurs; therefore, it is
imperative that every effort be made by the flight
crew to put the fire out. On the ground it is essential
that the engines be shut down, crew evacuated, and
fire fighting begun immediately. If the aircraft is airborne when a fire occurs, the most important single
action that can be taken by the pilot is to land safely
as soon as possible.
NOTE
Flight into the sun at high aircraft pitch
attitude may actuate the tire warning system. Lowering the nose and/or changing
headings will confirm a warning system
failure-caused by sun rays.
(3.) Engine fire in fright (identified). If an
engine fire is confirmed in flight, perform the following:
Due to the possibilities of fire warning
system malfunctions, the fire should be
visually identified before the engine is
secured and the extinguisher actuated.
a. Engine Fire. The following procedures shall
be performed in case of engine fire:
(1.) Engine/nacelle fire during start or
ground operations. If engine/nacelle fire is identified
during start or ground operation, perform the following:
1.
Propeller levers - FEATHER.
2. Condition levers - FUEL CUTOFF.
3.
Fire pull handle - Pull.
If fire extinguisher has been used to extinguish an engine fire, do not attempt to
restart, until maintenance personnel have
inspected the aircraft and released it for
flight.
4.
Push to extinguish switch - Push.
5.
Master switch - OFF.
(2.) Engine fire in flight fire pull handle
light illuminated). If an engine fire is suspected in
flight, perform the following:
1.
Power lever - IDLE.
2. If fire pull handle light out is
extinguished: Advance power.
If fire pull handle light is still illuminated: Engine fire in flight procedures (identified) - Perform.
1.
Power lever - IDLE.
2.
Propeller lever - FEATHER.
3.
Condition lever - FUEL CUTOFF.
4.
Fire pull handle - Pull.
5.
Fire extinguisher - Actuate as
required.
6.
Engine cleanup - Perform.
b. Fuselage Fire. If a fuselage fire occurs, perform the following:
The extinguisher agent (Bromochlorodifluoromethane) in the fire extinguisher
can produce toxic effects if inhaled.
1.
Fight the fire.
2.
Land as soon as possible.
c. Wing Fire. There is little that can be done
to control a wing fire except to shut off fuel and
electrical systems that may be contributing to the
fire, or which could aggravate it. Diving and slipping
the aircraft away from the burning wing may help.
If a wing fire occurs, perform the following:
1.
Perform engine shutdown on affected
side.
9-9
TM 55-1510-221-10
Land as soon as possible.
NOTE
d. Electrical Fire. Upon noting the existence
or indications of an electrical fire, turn off all
affected electrical circuits, if known. If electrical fire
source is unknown, perform the following:
Opening storm window (after depressurizing) will facilitate smoke and fume
removal.
2.
1. Crew oxygen - 100%.
2.
Master switch - OFF (visual conditions
only).
3. All nonessential electrical equipment OFF.
NOTE
With loss of DC electrical power, the aircraft will depressurize. All electrical
instruments, with the exception of the
propeller RPM, Nt RPM, and TGT gages
will be inoperative.
4. Battery switch - ON.
5. Generator switches (individually) RESET, then ON.
6.
Circuit breakers - Check for indication
of defective circuit.
As each electrical switch is returned to
ON (note loadmeter reading) and check
for evidence of fire.
7.
Essential electrical equipment - On (individually until fire source is isolated).
8.
Land as soon as practicable.
e. Smoke and Fume Elimination. To eliminate smoke and fumes from the aircraft, perform the
following:
9-10
1.
Crew oxygen - 100% and ON.
2.
Bleed air valve switches - ENVIRO
OFF.
3.
Vent blower switch - AUTO.
4.
Aft vent blower switch - OFF.
5.
Cabin temperature mode selector
switch - OFF.
6.
If smoke and fumes are not eliminated:
Cabin pressure dump switch - CABIN
PRESS DUMP.
7.
Engine oil pressure - Monitor.
9-21. FUEL SYSTEM.
a. Fuel Pressure Warning Annunciator Light
Illuminated. Illumination of the #1 FUEL PRESS or
#2 FUEL PRESS warning light usually indicates failure of the respective engine-driven boost pump. Perform the following:
1.
Standby pump switch - ON.
2. Fuel pressure warning annunciator
light - Check extinguished.
3. If fuel pressure warning light is still
illuminated: Record unboosted time.
b. No Fuel Transfer Caution Light Illuminated. Illumination of a #1 NO FUEL XFR or #2
NO FUEL XFR annunciator light with fuel remaining in the respective auxiliary fuel tank indicates a
failure of that automatic fuel transfer system. Proceed as follows:
1.
AUX TRANSFER switch (affected
side) - OVERRIDE.
2.
Auxiliary fuel quantity - Monitor.
3.
AUX TRANSFER switch (after respective auxiliary fuel has completely transferred) - AUTO.
c. Nacelle Fuel Leak. If nacelle fuel leaks are
evident, perform the following:
1.
Perform engine shutdown.
2.
Fire pull handle - Pull.
3.
Land as soon as practicable.
d. Fuel Crossfeed. Fuel crossfeed is normally
used only during single-engine operation. The fuel
from the dead engine side may be used to supply the
live engine by routing the fuel through the crossfeed
system. During extended flights, this method of fuel
usage will provide a more balanced lateral load condition in the aircraft. For fuel crossfeed, use the following procedure:
1.
AUX TRANSFER switches - AUTO.
TM 55-1510-221-10
NOTE
With the FIRE PULL handle pulled. the
fuel in the auxiliary tank for that side will
not be available (usable) for crossfeed.
this may indicate the fuel vent float valve
in the wet section has stuck. Rocking the
aircraft or changing pitch will probably
unstick it. If not. fuel may be crossfed.
2.
Standby pumps - OFF.
9-22 ELECTRICAL SYSTEM EMERGENCIES.
3.
Crossfeed switch - As required.
a. DC Generator Caution Annunciator Light
Illuminated. Illumination of a #1 DC GEN or #2
DC GEN caution annunciator light indicates failure
of a generator or one of its associated circuits (generator control unit). If one generator system becomes
inoperative. all nonessential electrical equipment
should be used judiciously to avoid overloading the
remaining generator. The use of accessories which
create a very high drain should be avoided. If both
generators are shut off due to either generator system failure or engine failure. all nonessential equip
ment should be turned off to preserve battery power
for extending the landing gear and wing flaps. When
a DC GEN light illuminates, perform the following:
4. Fuel crossfeed advisory annunciator
light - Check illuminated.
NOTE
With the FIRE PULL handle pulled, the
respective WI FUEL PRESS or #2 FUEL
PRESS light will remain illuminated on
the side supplying fuel.
5. Fuel pressure light extinguished Check.
6. Fuel quantity - Monitor.
e. Illumination of the #1 NAC LOW or #2
NAC LOW caution annunciator fight. Illumination
of the #1 NAC LOW or #2 NAC LOW caution
annunciator light indicates that the affected tank has
20 minutes remaining at sea level, normal cruise
power consumption rate. Proceed as follows:
1.
Generator switch - OFF, RESET, then
ON.
2.
Generator switch (no reset) - OFF.
3. Mission control switch - OVERRIDE.
4. Operating loadmeter - 100% maximum.
b. Both DC Generator Warning Annunciator
Lights Illuminated.
Failure of the fuel tank venting system will
prevent the fuel in the wing tanks from gravity
feeding into the nacell tank. Fuel vent system
failure may be indicated by illumination of the
#1 or #2 NAC LOW caution light with greater
than 20 minutes of usable fuel indicated in the
main tank fuel system. The total usable fuel
remaining in the main fuel supply system with
the LOW FUEL caution light illuminated may
be as Iittle as 114 pounds, regardless of the
totaI fuel quantity indicated. Continued fIight
may result in engine flameout due to fuel starvation.
1.
Twenty minutes fuel remaining - Confirm:
2.
Land as soon as possible.
NOTE
1.
All nonessential equipment - OFF.
2.
Land as soon as practicable.
c. Excessive Loadmeter Indication (Over
100%). If either loadmeter indicates over 100%, perform the following:
1.
Battery switch - OFF (monitor loadmeter).
2. Loadmeter over 100% - Nonessential
electrical equipment OFF.
3.
Loadmeter under 100% - BATT switch
ON.
d. Inventer Caution Annunciator Light Illuminated. Illumination of the #1 INVERTER or #2
lNVERTER caution annunciator light indicates failure of the affected inverter. When either inverter
fails, the total aircraft load is automatically switched
to the remaining inverter. When a #1 INVERTER
or #2 INVERTER caution annunciator light illuminates, perform the following:
1.
Affected #1 INVERTER or #2
INVERTER switch - OFF.
If a "NAC LOW" light occurs about the
time the "AUX" tanks go empty and the
fuel gages show the main tanks "FULL".
Change 1
9-11
TM 55-1510-221-10
e. INST AC Warning Annunciator Light Illuminated. Illumination of the INST AC warning light
indicates that 26 VAC power is not available. All
items connected to the 26 VAC bus will be inoperative (refer to AC wiring schematic diagram in chap
ter 2 for equipment effected). Under these conditions. power must be controlled by indications of the
Nt and TGT gages. Perform the following
NOTE
Windshield defogging may be required.
1.
Power lever - IDLE.
2.
Propeller lever - HIGH RPM.
3. Flaps - APPROACH.
1.
Nt and TGT indications - Check.
4. Gear - DN.
2.
Other engine instruments - Monitor.
5.
f. Circuit Breaker Tripped. If the circuit
breaker is for a nonessential item. do not reset in
flight. If the circuit breaker is for an essential item,
the circuit breaker may be reset once. If a bus feeder
circuit breaker (on the overhead circuit breaker
panel) trips. a short is indicated. Do not reset in
flight. If a circuit breaker trips. perform as follows:
1.
BUS FEEDER breaker tripped - Do
not reset.
2.
Nonessential circuit - Do not reset.
-3.
Essential circuit - Reset once.
NOTE
Circuit breakers should not be reset more
than once until the cause of the circuit
malfunction has been determined and
corrected. Do not reset dual fed bus
feeder circuit breakers.
Airspeed - 180 KIAS maximum.
9-24. LANDING EMERGENCIES.
Structural damage may exist after landing
with brake, tire, or landing gear malfunctions. Under no circumstances shall an
attempt be made to inspect the aircraft
until jacks have been installed.
a. Landing Gear Unsafe Indication. Should
one or more of the three green landing gear indicator lights fail to indicate a safe condition. the following steps should be taken before proceeding to
extend the gear manually.
1. Gear - DN.
2.
Gear lights - Check (three green).
3. Landing gear relay and indicator circuit breaker - Check In.
g. BATTERY CHARGE Light Illuminated. If
the BATTERY CHARGE caution light illuminates
during normal cruise flight. perform the following:
NOTE
1.
Battery Volt-Ampmeter - Monitor. If
battery current continues to increase,
turn battery switch off.
If gear continues to indicate unsafe.
attempt to verify position of the landing
gear visually.
2.
Battery switch (landing gear/flap extension only) - ON.
h.
Landing Gear Emergency Extension.
9-23. EMERGENCY DESCENT.
Emergency descent is a maximum effort in
which damage to the aircraft must be considered
secondary to getting the aircraft down. The following procedure assumes the structural integrity of the
aircraft and smooth flight conditions. If structural
integrity is in doubt. limit speed as much as possible. reduce rate of descent if necessary, and avoid
high maneuvering loads. For emergency descent,
perform the following:
9-12
Change 1
Continued pumping of the handle after
GEAR DOWN position indicator lights
(3) are illuminated could damage the
drive mechanism. and prevent subsequent
gear retraction.
TM 55-1510-221-10
4. Landing gear alternate engage handle Lift and turn clockwise to the stop.
After an emergency landing gear extension has been made. do not stow the gear
ratchet handle or move any landing gear
controls or reset any switches or circuit
breakers until the cause of the malfunction has been corrected.
1.
5.
Alternate landing gear extension handle
- Pump.
6.
Gear lights - Check (three green).
c. Gear-up Landing (All Gear Up or
Unlocked). Due to decreased drag with the gear up.
the tendency will be to overshoot the approach. The
Airspeed - 130 KIAS.
2. LANDING GEAR RELAY circuit
breaker - Out.
3. Gear - DN.
Change 1
9-12.1/(9-12.2 blank)
TM 55-1510-221-10
center-of-gravity with the gear retracted is aft of the
main wheels. This condition will allow the aircraft
to be landed with the gear retracted and should
result in a minimum amount of structural damage to
the aircraft, providing the wings are kept level. It is
recommended that the fuel load be reduced and the
landing made with flaps fully extended on a hard
surface runway. Landing on soft ground or dirt is
not recommended as sod has a tendency to roll up
into chunks, damaging the underside of the aircraft’s
structure. When fuel load has been reduced, prepare
for a gear-up landing as follows:
possible and do not use brakes. Use the following
procedures:
1.
Crew emergency briefing - Complete.
2. Loose equipment - Stowed.
3. Bleed air valve switches - ENVIRO
OFF.
4. Cabin pressure dump switch - CABIN
PRESS DUMP.
5.
Cabin emergency hatch - Remove and
stow.
Loose equipment - Stowed.
6.
Seat belts and harnesses - Secured.
3.
Bleed air valve switches - ENVIRO
OFF.
7. Nonessential electrical equipment OFF.
4.
Cabin pressure dump switch - CABIN
PRESS DUMP.
NOTE
5.
Cabin emergency hatch - Remove and
stow.
Fly a normal approach to touchdown.
After landing, accomplish the following:
6.
Seat belts and harnesses - Secured.
7.
Landing gear alternate engage handle Disengaged.
1.
Crew emergency briefing - Complete.
2.
8.
Alternate landing gear extension handle
- Stowed.
9.
Gear relay circuit breaker - In.
10.
Gear - UP.
11.
Nonessential electrical equipment OFF.
12.
Flaps - As required (DOWN for landing).
NOTE
Fly a normal approach to touchdown.
After landing, accomplish the following:
13.
Power levers (runway assured) - IDLE.
14.
Condition levers - FUEL CUTOFF.
15.
Fire pull handles - Pull.
16.
Master switch - OFF.
d. Landing With Nose Gear Unsafe. If the
landing gear control switch handle warning light is
illuminated and the nose GEAR DOWN indicator
light shows an unsafe condition, the nose gear is
probably not locked down, and the gear position
should be checked visually by another aircraft, if
possible. If all attempts to lock the nose gear fail, a
landing should be made with the main gear down
and locked. Hold the nose off the runway as long as
8.
Power levers (runway assured) - IDLE.
9.
Condition levers - FUEL CUTOFF.
10.
Fire pull handle - Pull.
11.
Master switch - OFF.
e. Landing With One Main Gear Unsafe. If
one main landing gear fails to extend, retract the
other gear and make a gear-up landing. If all efforts
to retract the extended gear fail, land the aircraft on
a hard runway surface, touching down on the same
edge of the runway as the extended gear. Roll on the
down and locked gear, holding the opposite wing up
and the nose gear straight as long as possible. If the
gear has extended, but is unsafe, apply brakes lightly
on the unsafe side to assist in locking the gear. If the
gear has not extended or does not lock, allow the
wing to lower slowly to the runway. Use the following procedures:
1.
Crew emergency briefing - Complete.
2. Loose equipment - Stowed.
3. Bleed air valve switches - ENVIRO
OFF.
4. Cabin pressure dump switch - CABIN
PRESS DUMP.
5. Cabin emergency hatch - Remove and
stow.
6.
Seat belts and harnesses - Secured.
7. Nonessential electrical equipment OFF.
9-13
TM 55-1510-221-10
8.
Touchdown - On safe main gear first.
NOTE
1.
Descend to below 25,000 feet.
2.
Cabin pressure - 4.6 PSI maximum.
3.
Do not operate more than 20 flight hours.
Fly a normal approach to touchdown.
After landing, accomplish the following:
9.
NOTE
Treat outer panel cracks which are linear
(not circular) or cracks that touch the
frame, as an inner panel crack.
Power levers (runway assured) - IDLE.
10.
Condition levers - FUEL CUTOFF.
11.
Fire pull handle - Pull.
12.
Master switch - OFF.
f. Landing With Flat Tire(s). If aware that a
main gear tire(s) is flat, a landing close to the edge
of the runway opposite the flat tire will help avoid
veering off the runway. If the nose wheel tire is flat,
use minimum braking.
9-28. CRACKED
PANEL).
CABIN WINDOW (INNER
If a cabin window inner panel crack occurs, perform the following:
1.
Oxygen - As required.
2. Cabin pressure - Depressurize.
3. Descend - As required.
9-25. LANDING WITH INOPERATIVE WING
FLAPS (UP).
The aircraft does not exhibit any unusual characteristics when landing with the wing flaps up. The
approach angle will be shallow and the touchdown
speed will be higher resulting in a longer landing
roll.
9-26. CRACKED WINDSHIELD.
a. External Crack. If an external windshield
crack is noted, no action is required in flight.
NOTE
Heating elements may be inoperative in
areas of crack.
b. Internal Crack. If an internal crack occurs,
perform the following:
1.
Descend to below 25,000 feet.
2.
Cabin Pressure - Reset pressure differential to 4 PSI or less within 10 minutes.
9-27. CRACKED CABIN WINDOW (OUTER
PANEL).
If a cabin window outer panel crack occurs, perform the following:
9-14
9-29. DITCHING.
If a decision to ditch is made, immediately alert
all crewmembers to prepare for ditching. Plan the
approach into the wind if the wind is high and the
seas are heavy. If the swells are heavy but the wind
is light, land parallel to the swells. Set up a minimum rate descent (power on or off, as the situation
dictates, airspeed - (110-120 KIAS). Do not try to
flare as in a normal landing, as it is very difficult to
judge altitude over water, particularly in a slick sea.
Leveling off too high may cause a nose low “drop
in,“ while having the tail too low on impact may
result in the aircraft pitching forward and “digging
in.“ Expect more than one impact shock and several
skips before the final hard shock. There may be
nothing but spray visible for several seconds while
the aircraft is decelerating. To prevent cartwheeling,
it is important that the wings be level when the aircraft hits the water. After the aircraft is at rest,
supervise evacuation of passengers and exit the aircraft as quickly as possible. In a planned ditching,
the life raft and first-aid kits should be secured close
to the cabin emergency hatch for easy access when
evacuating; however, do not remove the raft from its
carrying case inside the aircraft. After exiting the
aircraft, keep the raft away from any damaged surfaces which might tear or puncture the fabric. The
length of time that the aircraft will float depends on
the fuel level and the extent of aircraft damage
caused by the ditching. Refer to figure 9-3 for body
positions during ditching. Figure 9-4 shows wind
swell information. Perform the following procedures:
TM 55-1510-221-10
Do not unstrap from the seat until all
motion stops. The possibility of injury
and disorientation requires that evacuation not be attempted until the aircraft
comes to a complete stop.
1.
Radio calls/transponder - As required.
2.
Crew emergency briefing - As required.
3.
Bleed air valve switches - ENVIRO OFF.
4.
Cabin pressure dump switch - CABIN
PRESS DUMP.
5.
Cabin emergency hatch - Remove and stow.
6.
Seat belts and harnesses - Secured.
7.
Gear - UP.
8.
Flaps - DOWN.
9.
Nonessential electrical equipment - OFF.
b. Unscheduled Rudder Boost Activation. Rudder boost operation without a large variation of
power between engines indicates a failure of the system. Perform the following:
1.
Rudder boost - OFF.
NOTE
The rudder boost system may not operate
when the brake deice system is in use.
Availability of the rudder boost system
will be restored to normal when the
BRAKE DEICE switch is turned off.
IF CONDITION PERSISTS:
2. Bleed air valve switches - PNEU &
ENVIRO OFF.
3.
Rudder trim - Adjust.
c. Unscheduled Electric Elevator Trim. In the
event of unscheduled electric elevator trim, perform
the following:
10.
Approach - Normal, power on.
1.
Elevator trim switch - OFF.
11.
Emergency lights - As required.
2.
Elevator trim circuit breaker - Out.
9-30
FLIGHT CONTROLS MALFUNCTION.
Use the following procedures, as applicable, for
flight control malfunctions.
a. Autopilot/Yaw Damp Emergency Disconnection. The autopilot can be disengaged by any of
the following methods:
I.
Pressing the DISC - TRIM - AP - YD
disconnect switch (control wheels).
2.
Pressing the autopilot "AP ENGAGE"
pushbutton on the autopilot mode
selector control panel.
3.
4.
Pressing the go-around switch (left
power lever), (yaw damper will remain
on).
Pulling the AP CONTR and AFCS
DIRECT circuit breakers (overhead
control panel).
5.
Setting AVIONICS MASTER PWR
switch (overhead control panel) to the
OFF position.
6.
Setting aircraft MASTER switch (overhead control panel) to the OFF position.
9-31. BAILOUT.
When the decision has been made to abandon
the aircraft in flight, the pilot will give the warning
signal. Exit from the aircraft will be through the
main entrance door, and in the departure sequence
using the exit routes as indicated in figure 9-1. Proceed as follows if bailout becomes necessary:
1.
Notify crew to prepare to bail out.
2.
Distress message - Transmit.
3.
Voice security - ZEROIZE.
4.
Transponder - 7700.
5.
Mode 4 - Zeroize.
6.
Flaps - DOWN.
7.
Airspeed - 100 KIAS.
8.
Trim - As required.
9.
Autopilot - Engage.
10.
Cabin pressure switch - DUMP.
11.
Parachute - Attach to harness.
12.
Cabin door - Open.
13.
Abandon the aircraft.
9-15
TM 55-1510-221-10
Table 9-1. Ditching
PLANNED DITCHING
PILOT
A. ALERT OCCUPANTS
B. ORDER TO PREPARE SURVIVAL GEAR FOR
AERIAL DROP
C. TRANSMIT DISTRESS MESSAGE
D. LIFE VEST - CHECK (DO NOT INFLATE)
E. DISCHARGE MARKER
F. LAND AND DITCH AIRCRAFT
G. ABANDON AIRCRAFT
IMMEDIATE DITCHING
PILOT
A. WARN OCCUPANTS
B. TRANSMIT DISTRESS MESSAGE
C. LIFE VEST - CHECK (DO NOT INFLATE)
D. APPROACH - NORMAL
E. NOTIFY OCCUPANTS TO BRACE FOR
DITCHING
F. LAND AND DITCH AIRCRAFT
G. ABANDON AIRCRAFT AFTER COPILOT
THROUGH CABIN EMERGENCY HATCH
COPILOT
COPILOT
A. REMOVE CABIN EMERGENCY HATCH
B. LIFE VEST - CHECK (DO NOT INFLATE)
C. ABANDON AIRCRAFT (TAKE LIFE RAFT AND
FIRST AID KIT)
A. REMOVE CABIN EMERGENCY HATCH
B. LIFE VEST - CHECK (DO NOT INFLATE)
C. ABANDON AIRCRAFT (TAKE LIFE RAFT AND
FIRST AID KIT)
PASSENGERS
PASSENGERS
A. SEAT BELTS - FASTEN
B. LIFE VEST - CHECK ( DO NOT INFLATE
C. ON PILOTS SIGNAL - BRACE FOR DITCHING
D. ABANDON AIRCRAFT THROUGH CABIN
DOOR (TAKE LIFE RAFT AND FIRST AID KIT
9-16
A.
B.
C.
D.
SEAT BELTS - FASTEN
LIFE VEST - CHECK (DO NOT INFLATE)
ON PILOTS SIGNAL -BRACE FOR DITCHING
ABANDON AIRCRAFT THROUGH CABIN
DOOR (TAKE LIFE RAFT AND FIRST AID KIT)
TM 55-1510-221-10
Figure 9-3. Emergency Body Positions
9-17
TM 55-1510-221-10
Figure 9-4. Wind Swell Ditch Heading Evaluation
9-18
TM 55-1510-221-10
APPENDIX A
REFERENCES
Reference information for the subject material contained in this manual can be found in the following
publications:
AR 70-50
Designating and Naming Defense Equipment, Rockets, and Guided
Missles
AR 95-1
Army Aviation - General Provisions and Flight Regulations
AR 95-16
Weight and Balance - Army Aircraft
AR 380-40
Safeguarding COMSEC Information
AR 385-40
Accident Reporting and Records
AR 700-26
Aircraft Designation System
DA PAM 738-751
Functional User’s Manual for the Army Maintenance Management System Aviation (TAMS-A)
FAR Part 91
General Operating and Flight Rules
FM 1-5
Instrument Flying and Navigation for Army Aviators
FM 1-30
Meterology for Army Aviators
TB AVN 23-13
Anti-icing, Deicing and Defrosting Procedures for Parked Aircraft
TB MED 501
Noise and Conservation of Hearing
TB 55-9150-200-24
Engine and Transmission Oils, Fuels, and additives for Army Aircraft
TM 9-1095-206-13&P
Operator’s Aviation Unit Maintenance and Aviation Intermediate
Maintenance Manual (Including Repair Parts and Special Tools List)
to Dispenser, General Purpose Aircraft: M-130
(C) TM 11-5825-252-15
Operator, Organizational, DS, GS, and Deport Maintenance Manual:
RC-12H Aircraft Mission Equipment, (V)
TM 11-5841-291-12
Operator and Organizational Maintenance Manual, Radar Warning
System, AN/APR-44(V)1
TM 11-5841-283-20
Organizational Maintenance Manual for Detection Set, Radar Signal
AN/APR-39(V)1.
TM 11-6140-203-14-2
Operator’s Organizational, Direct Support, General Support and Depot
Maintenance Manual Including Repair Parts and Special Tools List:
Aircraft Nickel-Cadmium Batteries
TM 11-6940-214-12
Operator and Organization Maintenance Manual, Simulator, Radar Signal, SM-756/APR-44(V)
TM 55-410
Aircraft Maintenance, Servicing and Ground Handling Under Extreme
Environmental Conditions
TM 55-1500-204-25/1
General Aircraft Maintenance Manual Maintenance Manual: Army
Model RC-12H Aircraft
TM 55-1500-314-25
Handling, Storage, and Disposal of Army Aircraft Components Containing Radioactive Materials
TM 55-1500-342-23
Army Aviation Maintenance Manual: Weight and Balance
TM 55-1510-200-PM
Phased Maintenance Checklist
TM 55-1510-219-23
Aviation Unit and Aviation Intermediate
A-1
TM 55-1510-221-10
TM 750-244-1-5
A-2
Procedures for the Destruction of Aircraft and Associated Equipment
to Prevent Enemy Use
TM 55-1510-221-10
APPENDIX B
ABBREVIATIONS AND TERMS
For the purpose of this manual, the following abbreviations and terms apply. See appropriate technical
manuals for additional terms and abbreviations.
Airspeed Terminology.
CAS
Calibrated airspeed is indicated airspeed corrected for position and
instrument error.
FT/MIN
Feet per minute.
GS
Ground speed, though not an airspeed, is directly calculable from true
airspeed if the true wind speed and direction are known.
IAS
Indicated airspeed is the speed as shown on the airspeed indicator and
assumes no error.
KT
Knots.
TAS
True airspeed is calibrated airspeed corrected for temperature, pressure, and compressability 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.
V fe
Maximum flap extended speed is the highest speed permissible with
wing flaps in a prescribed extended position.
V le
Maximum landing gear extended speed is the maximum speed at
which an aircraft can be safely flown with the landing gear extended.
V 1o
Maximum landing gear operating speed is the maximum speed at
which the landing gear can be safely extended or retracted.
V lof
V mea
Lift off speed (takeoff airspeed).
V mo
Maximum operating limit speed.
V ne
Never exceed speed.
vr
Rotation speed.
vs
Power off stalling speed or the minimum steady flight speed at which
the aircraft is controllable.
V
s o
V sse
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).
Stalling speed or the minimum steady flight speed in the landing configuration.
The safe one-engine inoperative speed selected to provide a reasonable
margin against the occurence of an unintentional stall when making
intentional engine cuts.
B-1
TM 55-1510-221-10
Best angle of climb speed.
Best single-engine angle of climb speed.
Best rate of climb speed.
The best single engine rate of climb speed.
Meteorological Terminology.
Altimeter Setting
°C
°F
FAT
Indicated Pressure Altitude
ISA
Pressure Altitude
SL
Wind
Beta Range
Barometric pressure corrected to sea level.
Degrees Celsius.
Degrees Fahrenheit.
Free Air Temperature is the free air static temperature obtained either
from the temperature indicator (IFAT), adjusted for compressibility
effects, or from ground meteorlogical sources.
The number actually read from an altimeter when, the barometric scale
(Kollsman window) has been set to 29.92 inches of mercury (1013 millibars).
International Standard Atmosphere in which:
a. The air is a dry perfect gas.
b. The temperature at sea level is 59 degrees Fahrenheit, 15
degrees Celsius.
c. The pressure at sea level is 29.92 inches Hg.
d. The temperature gradient from sea level to the altitude at which
the temperature is -69.7 degrees Fahrenheit is -0.003566 Fahrenheit per foot and zero above that altitude.
Indicated pressure altitude corrected for altimeter error.
Sea level.
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).
The region of the power lever control which is aft of the idle stop and
forward of reversing range where blade pitch angle can be changed
without a change of gas generator RPM.
Cruise Climb
Is the maximum power approved for normal climb. This power is
torque or temperature (ITT) limited.
High Idle
Obtained by placing the condition lever in the HIGH IDLE position.
HP
Low Idle
Maximum Cruise Power
Maximum Power
Horsepower.
Obtained by placing the condition lever in the LO IDLE position.
Is the highest power rating for cruise and is not time limited.
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).
Obtained by lifting the power levers and moving them aft of the beta
range.
Reverse Thrust
B-2
TM 55-1510-221-10
RPM
Revolutions Per Minute.
Takeoff Power
The maximum power available from an engine for takeoff, limited to
periods of five minutes duration.
Control and Instrument Terminology.
Condition Lever (Fuel Shut-off
Lever)
Interstage Turbine Temperature
(ITT)
Nt Tachometer (Gas Generator
RPM)
Power Lever (Gas Generator N1
RPM)
Propeller Control Lever (N2
RPM)
Propeller Governor
Torquemeter
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.
Eight probes wired in parallel indicate the temperature between the
compressor and power turbines.
The tachometer registers the RPM of the gas generator with 100% representing a gas generator speed of 37,500 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.
This lever requests the control to maintain RPM at a selected value
and, in the maximum decrease RPM position, feathers the propeller.
This Governor will maintain the selected propeller speed requested by
the propeller control lever.
The torquemeter system determines the shaft output torque. Torque
values are obtained by tapping into two outlets on the reduction gear
case and recording the differential pressure from the outlets.
Graph and Tabular Terminology.
AGL
Best Angle of Climb
Best Rate of Climb
Clean Configuration
Demonstrated Crosswind
Gradient
Landing Weight
Maximum Zero Fuel Weight
MEA
Obstacle Clearance Climb
Speed
Ramp Weight
Route Segment
Above ground level.
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.
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.
Gear and flaps up regardless of mission antenna installation.
The maximum 90° crosswind component for which adequte control of
the aircraft during takeoff and landing was actually demonstrated during certification tests.
The ratio of the change in height to the horizontal distance, usually
expressed in percent.
The weight of the aircraft at landing touchdown.
Any weight above the value given must be loaded as fuel.
Minimum Enroute Altitude.
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.
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
grond roll to liftoff.
A part of a route. Each end of that part is identified by:
a. A geographic location; or
B-3
TM 55-1510-221-10
Service Ceiling
Takeoff Weight
b. A point at which a definite radio fix can be established.
The altitude at which the minimum rate of climb of 100 feet per minute can be attained for existing aircraft weight.
The weight of the aircraft at liftoff from the runway.
Weight and Balance Terminology.
Arm
Approved Loading Envelope
Basic Empty Weight
Center-of-Gravity
CG Limits
Datum
Engine Oil
Empty Weight
Landing Weight
Maximum Weight
Moment
Standard
Station
Takeoff Weight
Unusable Fuel
Usable Fuel
Useful Load
The distance from the center of gravity of an object to a line about
which moments are to be computed.
Those combinations of aircraft weight and center of gravity which
define the limits beyond which loading is not approved.
The aircraft weight with unusable fuel, full oil, and full operating fluids.
A point at which the weight of an object may be considered concentrated for weight and balance purposes.
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.
A vertical plane perpendicular to the aircraft longitudinal axis from
which fore and aft (usually aft) measurements are made for weight and
balance purposes.
That portion of the engine oil which can be drained from the engine.
The aircraft weight with fixed ballast, unusable fuel, engine oil, engine
coolant, hydraulic fluid, and in other respects as required by applicable
regulatory standards.
The weight of the aircraft at landing touchdown.
The largest weight allowed by design, structural, performance or other
limitations.
A measure of the rotational tendency of a weight, about a specified
line, mathematically equal to the product of the weight and the arm.
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.
The longitudinal distance from some point to the zero datum or zero
fuselage station.
The weight of the aircraft at liftoff.
The fuel remaining after consumption of usable fuel.
That portion of the total fuel which is available for consumption as
determined in accordance with applicable regulatory standards.
The difference between the aircraft ramp weight and basic empty
weight.
Miscellaneous Abbreviations.
Deg
DN
FT
FT LB
B-4
Degrees
Down
Foot or feet
Foot-pounds
TM 55-1510-221-10
GAL
Gallons
HR
kHz
LB
MAX
MHz
MIN
NAUT
Hours
NM
PSI
R/C
Kilohertz
Pounds
Maximum
Megahertz
Minimum
Nautical
Natucial miles
Pounds per square inch
Rate of climb
B-5/(B-6 blank)
TM 55-1510-221-10
INDEX
Paragraph, Figure,
Table Number
Subject
Paragraph, Figure,
Table Number
Subject
A
A
Abort Start ...................................................... 8-32
AC Power Supply ............................................ 2-75
Accelerate-Go Distance Over 50-Foot
Obstacle (Flaps 0%) .................................... 7-10
Accelerate-Stop (Flaps 0%) .............................. 7-7
Accelerometer .................................................. 2-85
Action Codes and Recommended
Actions .......................................................... T3-5
Additional Data (Normal
Procedures) .................................................... 8-8
After Emergency Action .................................... 9-4
After Landing .................................................. 8-51
After Takeoff .................................................... 8-42
Air Conditioning System ................................ 2-70
Air Induction Systems - General .................... 2-18
Aircraft Compartment and
Stations ................................................ 6-3, F6-1
Aircraft Dimensions ........................................ F6-1
Aircraft Designation System .......................... 1-11
Aircraft Systems ................................................ 9-1
Airspeed Indicators .......................................... 2-81
Airspeed Limitations ........................................ 5-19
Altitude Limitations .......................................... 5-28
Altitude Select Controller ................................ 3-23
Ammunition ........................................................ 4-5
Antenna Deicing System ................................ 2-52
Anti-Icing, Deicing and Defrosting
Treatment .................................................. 2-100
Appendix A, References .................................... 1-4
Appendix B, Abbreviations and Terms ............ 1-4
Application of External Power ...................... 2-101
Approved Fuels .............................................. T2-10
Approved Military Fuels, Oil, Fluids
and Unit Capacities ...................................... T2-9
Army Aviation Safety Program ........................ 1-7
Arrival Briefing ................................................ 8-75
Audio Control Panels .............................. 3-6, F3-1
Autoignition System ........................................ 2-30
Automatic Flight Control System .................... 3-29
Automatic Direction Finder
(DF-203) ............................................ 3-27, F3-18
Autopilot Controller ...................................... F3-23
Autopilot System Limits .................................. T3-1
Autopilot Limitations .......................................... 5-9
Autopilot Approaches ...................................... 8-61
Avionics Equipment Configuration .................... 3-2
B
Bailout ..............................................................
Balance Definitions ............................................
Bank and Pitch Limits ....................................
Before Exterior Check ..................................
Before Landing ................................................
Before Leaving Aircraft ..................................
Before Starting Engines ..................................
Before Takeoff ................................................
Before Taxiing ..................................................
Brake Deice Limitations ..................................
Brake Deice System ........................................
9-31
6-8
5-27
8-12
8-47
8-53
8-29
8-39
8-36
5-11
2-56
C
Cabin and Cargo Doors ..................................
Cabin Door Caution Light Illuminated ............
Cabin Pressure Limits ....................................
Caution/Advisory Annunciator Panel
Legend ..........................................................
Center of Gravity Limitations ................ 5-17,
Center Section, Right Side ............................
Center Section Left Side ................................
Chart C - Basic Weight and
Balance Record .................................... 6-9,
Charts and Forms ..............................................
Checklist ............................................................
Checks (Checklist Symbols) ............................
Chemical Toilet ................................................
Chip Detector Warning Light
Illuminated ......................................................
Cigarette Lighters and Ash Trays ..................
Class (Aircraft) ..................................................
Climb ................................................................
Cockpit ............................................................
F2-9
9-14
5-33
T2-7
6-13
8-21
8-18
F6-2
6-5
8-9
8-11
2-66
9-9
2-65
6-2
8-43
F2-8
Index-1
TM 55-1510-221-10
Paragraph, Figure,
Table Number
Subject
Paragraph, Figure,
Table Number
Subject
C
D
Cold Weather Operations ................................ 8-69
Comments Pertinent to the Use of
Performance Graphs .................................... 7-16
Condition Levers .............................................. 2-23
Conditions (At Stapleton
International) ..................................................
7-2
Control Pedestal .............................................. F2-7
Control Wheels .................................... 2-37, F2-17
Copilot’s Encoding Altimeter .......................... 2-82
Copilot’s Horizontal Situation
Indicator ......................................................
F3-11
Copilot’s Encoding Altimeter .............. 3-34, F3-29
Copilot’s Gyro Horizon Indicator ........ 3-20, F3-13
Cracked Cabin Window (Inner Panel) ............ 9-27
Cracked Cabin Window (Outer Panel) .......... 9-26
Cracked Cabin Window/Windshield .............. 5-34
Cracked Windshield ........................................ 9-26
Crew Briefings .......................................... 8-6, 8-73
Crossfeed Fuel Flow ...................................... F2-15
Crosswind Limitation ...................................... 5-31
Cruise ..............................................................
8-44
Cylinder Capacity vs Pressure
and Temperature ........................................ F2-20
Ditching .................................................. 9-29, F9-2
8-64
Diving ................................................................
Draining Moisture from Fuel System .............. 2-92
Duct Overtemp Caution Annunciator
Light Illuminated .......................................... 9-10
DC Electrical System .................................... F2-22
DC Power Supply ............................................ 2-74
D
Definition of Landing Terms ............................
Defrosting System ..........................................
Density Variation of Aviation Fuel ..................
Departure Briefing ..........................................
Descent ............................................................
Descent-Arrival ................................................
Description (Electrical System) ......................
Description (Navigation) ..................................
Description (Communications) ..........................
Description (Manual) ..........................................
Description (Propellers) ..................................
Description (Flight Controls) ..........................
Description (Emergency Equipment) ..............
Desert Operation and Hot Weather
Operation ......................................................
Destruction of Army Materiel to
Prevent Enemy Use ........................................
Dimensions (Aircraft) ........................................
Index-2
9-3
2-50
F6-4
8-74
8-45
8-46
2-73
3-16
3-4
1-3
2-42
2-36
2-12
8-70
1-8
2-3
E
Electrical System Emergencies ...................... 9-22
Emergency Entrance .......................................... 9-6
Emergency Exits and Equipment ............ 9-5, F9-1
Emergency Body Positions ............................ F9-3
Emergency Descent ........................................ 9-23
Emergency Locator Transmitter (ELT)
3-15, F3-8
Emergency Lighting ........................................ 2-78
Empennage, Area 5 ........................................ 8-26
Engine Bleed Air System Failure .................... 9-12
Engine Chip Detection System ...................... 2-28
Engine Clearing ................................................ 8-33
Engine Compartment Cooling ........................ 2-17
Engine Fire Extinguisher System .......... 2-26, T2-1
Engine Fire Detection System ........................ 2-25
Engine Fuel Control System .......................... 2-21
Engine Ice Protection Systems ...................... 2-20
Engine Ignition System .................................... 2-29
Engine Instruments .......................................... 2-32
Engine Limitations .......................................... 5-13
Engine Malfunction ............................................ 9-7
Engine Runup .................................................. 8-38
Engine Shutdown ............................................ 8-52
Engine Starter-Generators .............................. 2-31
Engines ............................................................
2-16
Entrance and Exit Provisions .......................... 2-10
Environmental Controls .................................. 2-72
Environmental Systems ................................ F2-21
Exceeding Operational Limits .......................... 5-3
Exhaust and Propeller Danger Area .............. F2-5
Exhaust Danger Area ........................................ 2-6
Explanation of Change Symbols .................... 1-10
Extent of Coverage (Weight, Balance
and Loading) .................................................. 6-1
Exterior Inspection .......................................... F8-1
TM 55-1510-221-10
Subject
Paragraph, Figure,
Table Number
Paragraph, Figure,
Table Number
Subject
E
G
Exterior Check.
8-13
Exterior Lighting. . . . . . . . . . . . . . . . . . . . . . . . . 2-76.F2-27
General (Aircraft) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2
General (Instrument Flight) . . . . . . . . . . . . . . . . . . . . . .8-54
General (Introduction) . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-1
General (Operating Limits) . . . . . . . . . . . . . . . . . . . . . . .5-2
General (Servicing, Parking and Mooring) . . . . . . . . . .2-89
General Exterior Arrangement . . . . . . . . . . . . . . . . . . . F2-1
General Interior Arrangement . . . . . . . . . . . . . . . . . . . .F2-2
Generator Limits . . . . . . . . . . . . . . . . . . . . . . . . .5-16. T5-2
Go-Around. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .8-50
.
Gravity Feed Fuel Flow . . . . . . . . . . . . . . . . . . . . . . . F2-16
Ground Handling . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-103
Ground Turning Radius . . . . . . . . . . . . . . . . . . . . .2-4,F2-4
Gyromagnetic Compass Systems . . . . . . . . . . . . . . . . . 3-22
F
Feathering Provisions . . . . . . . . . . . . . . . . . . . . . . . . . . 2-43
Ferry Chair. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-64
Ferry Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-35
Filling Fuel Tanks. . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-91
Fire.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9-20
First Aid Kits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-13
First Engine Start (Battery Start) . . . . . . . . . . . . . . . . . .8-30
First Engine Start (GPU Start) . . . . . . . . . . . . . . . . . . . 8-34
Flight Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .8-66
Flight Controls Lock . . . . . . . . . . . . . . . . . . . . . 2-39. F2-18
Flight Controls Malfunction . . . . . . . . . . . . . . . . . . . . .9-30
Flight Director. . . . . . . . . . . . . . . . . . . . . . . . . F3-21,F3-22
Flight Envelope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . F5-2
Flight Plan. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-5
Flight Planning. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .7-12
Flight Under IMC (Instrument
Meteorological Conditions. . . . . . . . . . . . . . . . . . . . 5-30
Foreign Object Damage Control . . . . . . . . . . . . . . . . . .2-19
Forms and Records . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-9
Friction Lock Knobs . . . . . . . . . . . . . . . . . . . . . . . . . . .2-24
Fuel and Oil Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-12
Fuel Handling Precautions. . . . . . . . . . . . . . . . . . . . . . .2-90
F u e l L o a d . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-11
Fuel Management Panel . . . . . . . . . . . . . . . . . . . . . . F2-14
Fuel Quantity Data. . . . . . . . . . . . . . . . . . . . . . . . . . . . T2-2
Fuel Sample . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .8-14
Fuel Sump Drain Locations . . . . . . . . . . . . . . . . . . . . . T2-3
Fuel Supply System. . . . . . . . . . . . . . . . . . . . . . . . . . . .2-33
Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .9-21
Fuel System Anti-Icing. . . . . . . . . . . . . . . . . . . . . . . . .2-57
Fuel System Limits. . . . . . . . . . . . . . . . . . . . . . . . . . . .5-10
Fuel System Management . . . . . . . . . . . . . . . . . . . . . . .2-34
Fuel System Schematic . . . . . . . . . . . . . . . . . . . . . . . . F2-13
Fuel Types . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-93
Fuselage. Left Side, Area 6 . . . . . . . . . . . . . . . . . . . . . .8-27
Fuselage, Right Side, Area 4 . . . . . . . . . . . . . . . . . . . . .8-25
Fuselage. Underside. . . . . . . . . . . . . . . . . . . . . . . . . . . .8-19
H
HF Communication Set (KHF-950). . . . . . . . . . . . . . .
HFControl Panel (718 U-5). . . . . . . . . . . . . . . . . . . .
Hand-Operated Fire Extinguisher . . . . . . . . . . . . . . . .
Heating System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Horizontal Situation Indicators . . . . . . . . . . . . . . . . . .
.3-14
.F3-7
.2-14
.2-69
.3-18
I
Ice and Rain (Typical) . . . . . . . . . . . . . . . . . . . . . . . . . .8-72
Icing (Severe). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .8-72A
Icing Limitations (Typical) . . . . . . . . . . . . . . . . . . . . 5-30A
Icing Limitations (Severe) . . . . . . . . . . . . . . . . . . . . . 5-30B
Ice Vane Failure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9-11
Immediate Action Emergency Checks . . . . . . . . . . . . . . 9-2
I n d e x . . . . . . . . . . . . . . . . . -. . . . . . . . . . . . . . . . . . . . . . 1-6
l
Inertial Navigation System . . . . . . . . . . 3-30. F3-24, F3-25
Inflating Tires. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-97
Installation of Protective Covers . . . . . . . . . . . . . . . . .2-105
Instrument Approaches. . . . . . . . . . . . . . . . . . . . . . . . .8-60
Instrument Climb. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .8-57
Instrument Cruise. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-58
Instrument Descent. . . . . . . . . . . . . . . . . . . . . . . . . . . .8-59
Instrument Flight Procedures . . . . . . . . . . . . . . . . . . . . 8-55
Instrument Landing System Limits . . . . . . . . . . . . . . . .5-36
Instrument Marking Color Codes . . . . . . . . . . . . . 5-6. F5-1
Instrument Markings, . . . . . . . . . . . . . . . . . . . . . . . . . . .5-5
Instrument Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . F2-29
Instrument Takeoff. . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-56
Intentional Engine Out Speed . . . . . . . . . . . . . . . . . . . ,5-37
Interior Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .8-28
Change 5
Index-3
TM 55-1510-221-10
Paragraph, Figure,
Table Number
Subject
Paragraph, Figure,
Table Number
Subject
I
M
Interior Lighting .............................................. 2-77
Introduction (Aircraft & Systems
Description and Operation) .......................... 2-1
Introduction (Avionics) ...................................... 3-1
Introduction to Performance ............................ 7-1
Introduction (Adverse Environmental
Conditions) .................................................... 8-68
Maximum Design Sink Rate ............................ 5-35
Maximum Glide ................................................ F9-2
Maximum Takeoff Weight Permitted
By Enroute Climb .......................................... 7-5
Maximum Weights ............................................ 2-5
Microphones, Switches and Jacks .................. 3-5
Minimum Crew Requirements .......................... 5-4
Minimum Oil Temperature Required
For Flight ............................................ 5-39, F5-3
Minimum Single-Engine Control
Airspeed (V,,) .............................................. 5-24
Miscellaneous Instruments .............................. 2-88
Mission Avionics Coverage .............................. 4-1
Mission Control Panel Annunciator
Legend .......................................................... T2-8
Mission Control Panel ............................ 4-2, F4-1
Mission Equipment ........................................ F2-25
Mission Planning ................................................ 8-1
Mooring .............................................. 2-106, F2-33
L
Landing ............................................................
Landing on Unprepared Runway ....................
Landing with Inoperative
Wing Flaps (UP) ..........................................
Landing Emergencies ......................................
Landing Gear Extension Speed ......................
Landing Gear System ........................................
Landing Gear Retraction Speed ....................
Landing Information ........................................
Left Engine and Propeller ..............................
Left Main Landing Gear ..................................
Left Wing, Area 1 ............................................
Level Flight Characteristics ............................
Line Up ............................................................
Load Planning ..................................................
Loading Procedure ..........................................
Loss of Pressurization
(Above 10,000 Feet) ....................................
Low Oil Pressure ..............................................
8-49
5-38
9-25
9-24
5-21
2-7
5-22
7-15
8-17
8-16
8-15
8-67
8-40
6-14
6-15
9-13
9-8
M-130 Flare and Chaff Dispensing
System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-3, F4-2, F4-3
Malfunction Indications and
Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . T3-4
Malfunction Code Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . T3-3
Maneuvering Flight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-65
Maneuvers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-26
Marker Beacon Receiver . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-26
Marker Beacon Audio Control Panel . . . . 3-7, F3-2
Maximum Allowable Airspeed . . . . . . . . . . . . . . . . . . . . . . . . 5-20
Maximum Design Maneuvering Speed . . . . . . . . . . 5-25
Index-4
N
Nose Section, Area 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-20
NAV 1 - NAV 2 Control Panel .. . . . . . . .. . . .. . . ... . .. F3-17
O
Obstacle Clearance Approach and
Minimum Run Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-48
Occupants Useful Load ) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . T6-1
Oil Supply System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-27
Operating Procedures and Maneuvers . . . . . . . . . . . . 8-7
Operating Limits and Restrictions . . . . . . . . . . . . . . . . . . . . 8-2
Outside Air Temperature (OAT) Gage . . . . . . . . . . . . 2-86
Overhead Control Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . F2-12
Overhead Circuit Breaker Panel . . . . . . . . . . . . . . . . . . F2-26
Over-temperature and Overspeed
Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-14, T5-1
Oxygen Cylinder Capacity Example
Problem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-62
Oxygen Duration Example Problem . . . . . . . . . . . . . . 2-61
Oxygen Duration In Minutes
Foot System ...............................................
T2-5
TM 55-1510-221-10
Paragraph, Figure,
Table Number
Subject
Paragraph, Figure,
Table Number
Subject
O
Oxygen Flow Planning Rates
vs Altitude .................................................... T2-4
Oxygen Requirements .................................... 5-32
Oxygen System ................................................ 2-60
Oxygen System Schematic .......................... F2-19
Oxygen System Servicing Pressure .............. F2-31
P
2-104
Parking ..........................................................
Parking Brake .................................................... 2-9
Parking, Covers, Ground Handing,
and Towing ................................................ F2-32
8-4
Performance ......................................................
Performance Example ...................................... 7-4
F3-10
Pilot’s ..............................................................
Pilot’s Altimeter Indicator ........ 2-83, 3-33, F3-28
Pilot’s Attitude Director Indicator ...... 3-19, F3-12
Pitot and Static System ...................... 2-79, F2-28
Pitot and Stall Warning Heat System ............ 2-54
Pitot Heat Limitations ...................................... 5-12
1-13
Placard Items ..................................................
Power Definitions for Engine Operations ...... 5-15
Power Levers .................................................. 2-22
Power Source .................................................... 3-3
Pressure Altitude .............................................. 7-3
Pressurization System .................................... 2-59
Principal Dimensions ...................................... F2-3
Propeller Test Switches .................................. 2-45
Propeller Governors ........................................ 2-44
Propeller Tachometers .................................... 2-49
Propeller Reversing ........................................ 2-48
Propeller Levers .............................................. 2-47
Propeller Synchrophaser ................................ 2-46
Propeller Electrothermal Anti-Ice
2-53
System ..........................................................
Propeller Limitations .......................................... 5-7
Propeller Failure (Over 2080 RPM) ................ 9-19
Purpose (Operating Limits and
5-1
Restrictions) ....................................................
........................
6-4
Purpose (Weight and Balance)
..................
.
.
.......................
F2-11
PT6A-41 Engine
Radar Signal Detecting Set
(AN/APR-39(V)1) ........................ 4-6, F4-4, F4-5
Radar Warning Receiver
(AN/APR-44() (V3) ................................ 4-7, F4-6
Radio Altimeter Indicator .................... 3-24, F3-16
Radio Magnetic Indicators (RMI) .......... 3-17, F3-9
Recommended Fluid Dilution Chart ............ T2-12
Relief Tube ...................................................... 2-68
Required Equipment Listing .................. 5-40, F5-4
Reserve Fuel .................................................... 7-13
Responsibility (Wgt and Balance) .................... 6-6
Right Engine and Propeller ............................ 8-22
Right Main Landing Gear ................................ 8-23
Right Wing, Area 3 .......................................... 8-24
Rudder System ................................................ 2-38
S
2-11, F2-10
Seats ..................................................
Second Engine Start (GPU Start) .................. 8-35
Second Engine Start (Battery Start) .............. 8-31
Securing Loads ................................................ 6-16
Servicing Oil System ...................................... 2-95
F2-30
Servicing ........................................................
Servicing Oxygen System ............................ 2-102
Servicing the Air Conditioning
2-99
System ..........................................................
Servicing the Chemical Toilet ........................ 2-98
Servicing Hydraulic Brake System
2-96
Reservoir ......................................................
Single Phase AC Electrical System .............. F2-23
Single-Engine Go-Around ................................ 9-18
Single-Engine Landing Check ........................ 9-17
Single-Engine Before Landing ........................ 9-16
Single-Engine Descent/Arrival ........................ 9-15
8-63
Spins ................................................................
Stall Warning System ...................................... 2-55
8-62, F8-2
Stalls ......................................................
Standard Alternate and Emergency
T2-11
Fuels ............................................................
Standby Magnetic Compass .......................... 2-87
Starter Limitations ............................................ 5-8
F2-6
Subpanels ........................................................
2-67
Sun Visors ........................................................
Index-5
TM 55-1510-221-10
Paragraph, Figure,
Table Number
Subject
Paragraph, Figure,
Table Number
Subject
S
V
Surface Deicing System .................................. 2-51
Survival Kits ....................................................
2-15
System Daily Preflight/Re-Arm Test ................ 4-4
Various Values for UTM Grid
Coefficients . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . T3-2
Vertical Velocity Indicators . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-84
Voice Order Wire (AN/ARC-164) . . . . . . . . . . . . . . . . . . . . . . 3-9
Voice Security System TSEC/KY-58
(Provisions Only) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-13
Voice Security System TSEC/KY-28
(Provisions Only) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-12, F3-6
VHF AM-FM Command Set
(AN/ARC- 186) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-11, F3-5
VHF-AM Communications (VHF-2OB) . . 3-10, F3-4
VOR/LOC Navigation System . . . . . . . . . . . . . . . . . . . . . . . . 3-25
T
Takeoff ..............................................................
8-41
Takeoff Climb Gradient - One
Engine Inoperative ...................................... 7-11
Takeoff Distance (Flaps 0%) ............................ 7-8
Takeoff Flight Path Example .................. 7-9, F7-1
Takeoff Weight to Achieve Positive ................ 7-6
Taxiing ..............................................................
8-37
Temperature Limits .......................................... 5-29
Three Phase AC Electrical System .............. F2-24
Transponder Set (APX-100) ................ 3-32, F3-27
Trim Tabs ........................................................
2-40
Turbulence and Thunderstorm Operation ...... 8-71
Turn and Slip Indicators .......... 2-80, 3-21, F3-14
TACAN Systems ...................... 3-28, F3-19, F3-20
U
Unpressurized Ventilation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Use of Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Use of Fuels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Use of Words Shall, Will, Should,
and May . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
UHF Command Set (AN/ARC-164) . . . . . . . . . . 3-8,
2-71
8-10
2-94
1-12
F3-3
W
Warning Annunciator Panel Legend . . . . . . . . . . . . . . T2-6
Warnings, Cautions, and Notes . . . . . . . . . . . . . . . . . . . . . . . . 1-2
Weather Radar Set (AN/APN-215) . . . . 3-31, F3-26
Weight and Balance Clearance
Form F, DD Form 365-4 . . . . . . . . . . . . . . . . . . . . . . 6-9, F6-3
Weight Definitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-7
Weight Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-18
Weight, Balance, and Loading . . . . . . . . . . . . . . . . . . . . . . . . . . 8-3
Wind Swell Ditch Heading
Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . F9-4
Windows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10
Windshield Wipers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-63
Windshield Electrothermal
Anti-Ice System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-58
Wing Flap Extension Speeds . . . . . . . . . . . . . . . . . . . . . . . . . . 5-23
Wing Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-41
Zero Fuel Weight Limitation . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-14
Index-6
TM 55-1510-221-10
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
WILLIAM J. MEEHAN II
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, -10 & CL Maintenance requirements for RC-12H Airplane, Reconnaissance.
* U.S. GOVERNMENT PRINTING OFFICE : 1994 0 - 300-769 (12205)
*U.S. GOVERNMENT PRINTING OFFICE
1994-300-769-12205
These are the instructions for sending an electronic 2028
The following format must be used if submitting an electronic 2028. The subject line must be exactly the
same and all fields must be included; however only the following fields are mandatory: 1, 3, 4, 5, 6, 7, 8,
9, 10, 13, 15, 16, 17, and 27.
From:
To:
‘Whomever” <[email protected] .mil>
[email protected] .mil
Subject: DA Form 2028
1. From: Joe Smith
2. Unit: home
3. Address: 4300 Park
4. City: Hometown
5. St: MO
6. Zip: 77777
7. Date Sent: 19-OCT-93
8. Pub no: 55-2840-229-23
9. Pub Title: TM
10. Publication Date: 04-JUL-85
11. Change Number: 7
12. Submitter Rank: MSG
13. Submitter FName: Joe
14. Submitter MName: T
15. Submitter LName: Smith
16. Submitter Phone: 123- 123- 1234
17. Problem: 1
18. Page: 2
19. Paragraph: 3
20. Line: 4
21. NSN: 5
22. Reference: 6
23. Figure: 7
24. Table: 8
25. Item: 9
26. Total: 123
27. Text:
This is the text for the problem below line 27.
PIN: 065798-000
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