OPERATOR’S MANUAL FOR ARMY CH-47D HELICOPTER (EIC: RCD, 2003)

TM 1-1520-240-10

WARNING

FIRE EXTINGUISHER

Personnel performing operations, procedures, and

practices which are included or implied in this tech-

Exposure to high concentrations of fire extinguish-

nical manual shall observe the following warnings.

ing agents or decomposition products should be

Disregard of these warnings and precautionary in-

avoided. The liquid should not contact the skin. It

formation can cause serious injury or death.

may cause frostbite or low temperature burns.

WARNING

ARMAMENT

STARTING ENGINES

Loaded weapons or weapons being loaded or un-

Coordinate all cockpit actions with ground observer.

loaded, shall be pointed in a direction which offers

Insure that wheels are chocked (if applicable), rotor

the least exposure to personnel or property in the

and blast areas are clear, and fire guard is posted.

event of accidental firing. Personnel shall remain

clear of the hazardous area of all loaded weapons.

WARNING

GROUND OPERATION

VERTIGO

Turn the anti-collision lights off during flight

Engines will be started and operated only by autho-

through clouds. This will eliminate light reflections

rized personnel.

from the clouds, which could cause vertigo.

WARNING

CARBON MONOXIDE

ROTOR BLADES

When smoke, suspected carbon monoxide fumes, or

Beware of moving rotor blades, particularly the

symptoms of anoxia exist, the crew should immedi-

blades of the forward rotor system.

ately ventilate the aircraft.

WARNING

HANDLING FUEL AND OIL

HIGH VOLTAGE

Turbine fuels and lubricating oils contain additives

All ground handling personnel must be informed of

that are poisonous and readily absorbed through the

high voltage hazards when making external cargo

skin. Do not allow them to remain on skin longer than

hook-ups.

necessary.

Change 1

TM 1-1520-240-10

WARNING

ELECTROMAGNETIC INTERFERENCE (EMI)

HAZARDOUS CARGO

No electrical/electronic devices of any sort, other

Items of cargo possessing dangerous physical

than those described in this manual or appropriate

properties such as explosives, acids, flammables,

airworthiness release and approved by USAATCOM,

etc. must be handled with extreme caution and in

are to be operated by crewmembers or passengers

accordance with established regulations. Ref:

during operation of this helicopter.

38-250.

WARNING

RADIOACTIVE MATERIALS

HF RADIO LIAISON FACILITY AN/ARC-220

Instrument dials on CH-47 series aircraft contain

The HF Radio Liaison Facility AN/ARC-220 in the ALE

radioactive materials. If an instrument is broken or

mode sounds (transmits short tone bursts) and re-

becomes unsealed, avoid personal contact with the

plies to ALE calls automatically without operator ac-

item. Use forceps or gloves made of rubber or poly-

tion. Anytime local flight directives forbid HF emis-

ethylene to pick up contaminated material. Place the

sions, such as ordinance loading or refeuling, or

material and the gloves in a plastic bag, seal the bag,

when personnel are working near the aircraft, ensure

and dispose of it as radioactive waste in accordance

the radio set control function switch is set to SILENT,

with AR 385-11 and TM 3-261. (Refer to TB 43-0108).

STBY, or OFF.

WARNING

NOISE LEVELS

IN ALE MODE

Sound pressure levels in this aircraft during some

The AN/ARC-220 sounds (transmit short bursts) and

operating conditions exceed the Surgeon General’s

replies to ALE calls automatically without operator

hearing conservation criteria, as defined in TB MED

action. Anytime local flight directives forbid HF

251. Hearing protection devices, such as the aviator

emissions, such as during ordance loading or refuel-

helmet or ear plugs are required to be worn by all

ing, or when personnel are working near the aircraft,

personnel in and around the aircraft during its opera-

ensure the radio set control function switch is set to

tion.

silent, STBY, or OFF.

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TM 1-1520-240-10

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Insert latest changed pages. Dispose of superseded pages in accordance with regulations.

NOTE: On a changed page, the portion of the text affected by the latest change is indicated by a vertical line, or

other change symbol, in the outer margin of the page. Changes to illustrations are indicated by miniature pointing

hands.

Dates of issue for original and changed pages are:

Original

31 January 2003

Change 2

25 July 2003

Change 1

16 April 2003

Change 3

10 May 2004

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TM 1-1520-240-10

TECHNICAL MANUAL

HEADQUARTERS

DEPARTMENT OF THE ARMY

WASHINGTON, D.C., 31 JANUARY 2003

OPERATOR’S MANUAL

FOR

ARMY MODEL CH-47D HELICOPTER

REPORTING ERRORS AND RECOMMENDING IMPROVEMENTS

You can help improve this manual. If you find any mistakes or if you know of a way to improve the

procedures, please let us know. Mail your letter or DA Form 2028 (Recommended Changes to

Publications and Blank Forms) located in the back of this manual directly to: Commander, US Army

Aviation and Missile Command, ATTN: AMSAM-MMC-MA-NP, Redstone Arsenal, AL 35898-5230. You

may also submit your recommended changes by E-Mail directly to 2028@redstone.army.mil or by fax

(256) 842-6546/DSN 788-6546. A reply will be furnished directly to you. Instruction for sending an

electronic 2028 may be found at the back of this manual immediately preceding the hard copy 2028.

DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.

Chapter/Section

Page

CHAPTER 1 INTRODUCTION .

1-1-1

SECTION I. INTRODUCTION .

1-1-1

CHAPTER 2 AIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATION .

2-1-1

SECTION I. HELICOPTER .

2-1-1

SECTION II. EMERGENCY EQUIPMENT .

2-2-1

SECTION III. ENGINES AND RELATED SYSTEMS .

2-3-1

SECTION IV. FUEL SYSTEM .

2-4-1

SECTION V. FLIGHT CONTROLS .

2-5-1

SECTION VI. HYDRAULIC SYSTEMS .

2-6-1

SECTION VII. POWER TRAIN SYSTEM .

2-7-1

SECTION VIII. ROTOR SYSTEM .

2-8-1

SECTION IX. UTILITY SYSTEMS .

2-9-1

SECTION X. HEATING, VENTILATION, COOLING, AND ENVIRONMENTAL CONTROL SYSTEMS

2-10-1

SECTION XI. ELECTRICAL POWER SUPPLY AND DISTRIBUTION SYSTEMS .

2-11-1

SECTION XII. AUXILIARY POWER UNIT .

2-12-1

SECTION XIII. LIGHTING .

2-13-1

SECTION XIV. FLIGHT INSTRUMENTS .

2-14-1

SECTION XV. SERVICING, PARKING, AND MOORING .

2-15-1

CHAPTER 3 AVIONICS .

3-1-1

SECTION I. GENERAL .

3-1-1

SECTION II. COMMUNICATIONS .

3-2-1

SECTION III. NAVIGATION EQUIPMENT .

3-3-1

SECTION IV. TRANSPONDERS .

3-4-1

TM 1-1520-240-10

Chapter/Section

Page

CHAPTER 4 MISSION EQUIPMENT .

4-1-1

SECTION I. MISSION AVIONICS .

4-1-1

SECTION II. ARMAMENT .

4-2-1

SECTION III. CARGO HANDLING SYSTEMS .

4-3-1

SECTION IV. EXTENDED RANGE FUEL SYSTEM (ERFS) AND ERFS II .

4-4-1

CHAPTER 5 OPERATING LIMITS AND RESTRICTIONS .

5-1-1

SECTION I. GENERAL .

5-1-1

SECTION II. SYSTEM LIMITS .

5-2-1

SECTION III. POWER LIMITS .

5-3-1

SECTION IV. LOADING LIMITS .

5-4-1

SECTION V. AIRSPEED LIMITS .

5-5-1

SECTION VI. MANEUVERING LIMITS .

5-6-1

SECTION VII. ENVIRONMENTAL RESTRICTIONS .

5-7-1

SECTION VIII. WATER OPERATION LIMITATIONS .

5-8-1

SECTION IX. ADDITIONAL LIMITATIONS .

5-9-1

CHAPTER 6 WEIGHT/BALANCE AND LOADING .

6-1-1

SECTION I. GENERAL .

6-1-1

SECTION II. WEIGHT AND BALANCE .

6-2-1

SECTION III. FUEL/OIL .

6-3-1

SECTION IV. PERSONNEL .

6-4-1

SECTION V. MISSION EQUIPMENT .

6-5-1

SECTION VI. CARGO LOADING .

6-6-1

SECTION VII. LOADING LIMITS .

6-7-1

CHAPTER 7 PERFORMANCE DATA .

7-1-1

SECTION I. INTRODUCTION .

7-1-1

SECTION II. EMERGENCY TORQUE AVAILABLE .

7-2-1

SECTION III. MAXIMUM TORQUE AVAILABLE .

7-3-1

SECTION IV. CONTINUOUS TORQUE AVAILABLE .

7-4-1

SECTION V. HOVER .

7-5-1

SECTION VI. TAKEOFF .

7-6-1

SECTION VII. CRUISE .

7-7-1

SECTION VIII. DRAG .

7-8-1

SECTION IX. CLIMB DESCENT .

7-9-1

SECTION X. FUEL FLOW .

7-10-1

SECTION XI. AIRSPEED CALIBRATION .

7-11-1

CHAPTER 7A PERFORMANCE DATA .

7A-1-1

SECTION I. INTRODUCTION .

7A-1-1

SECTION II. CONTINGENCY TORQUE AVAILABLE .

7A-2-1

SECTION III. MAXIMUM TORQUE AVAILABLE .

7A-3-1

SECTION IV. CONTINUOUS TORQUE AVAILABLE .

7A-4-1

SECTION V. HOVER .

7A-5-1

SECTION VI. TAKEOFF .

7A-6-1

SECTION VII. CRUISE .

7A-7-1

SECTION VIII. DRAG .

7A-8-1

SECTION IX. CLIMB DESCENT .

7A-9-1

SECTION X. FUEL FLOW .

7A-10-1

SECTION XI. AIRSPEED CALIBRATION .

7A-11-1

CHAPTER 8 NORMAL PROCEDURES .

8-1-1

TM 1-1520-240-10

CHAPTER 1

INTRODUCTION

SECTION I. INTRODUCTION

1-1-1. General.

1-1-6. Appendix B, Abbreviation.

These instructions are for use by the operator. They ap-

Appendix B is a list of the abbreviations used in this

ply to CH-47D helicopters.

manual.

1-1-7. Appendix C, Conditional Inspections.

1-1-2. WARNINGS, CAUTIONS, AND NOTES DE-

FINED.

Appendix C is a listing of conditions which require a DA

Form 2408-13-1 entry.

Warnings, cautions, and notes are used to emphasize

important and critical instructions and are used for the

1-1-8. Index.

following conditions.

The index lists in alphabetical order, every titled para-

graph, figure, and table contained in this manual.

WARNING

1-1-9. Army Aviation Safety Program.

Reports necessary to comply with the Army Aviation

An operating procedure, practice, etc.,

Safety program are prescribed in AR 385-40.

which if not correctly followed, could re-

1-1-10. Destruction of Army Material to Prevent En-

sult in personal injury or loss of life.

emy Use.

For information concerning destruction of Army material

CAUTION

to prevent enemy use, refer to TM 750-244-1-5.

An operating procedure, practice, etc.,

1-1-11. Forms and Records.

which, if not strictly observed, could re-

sult in damage to or destruction of equip-

Army aviators flight record and aircraft maintenance re-

ment.

cords which are to be used by crewmembers are pre-

scribed in DA PAM 738-751 and TM 55-1500-342-23.

NOTE

An operating procedure, condition, etc.,

1-1-12. Change Symbol Explanation.

which it is essential to highlight.

Changes, except as noted below, to the text and tables,

including new material on added pages, are indicated by

1-1-3. Helicopter Description.

a vertical line. The vertical line is in the outer margin and

This manual contains the complete operating instruc-

extends close to the entire area of the material affected

tions and procedures for the CH-47D helicopters. It is

with the following exception: pages with emergency

markings, which consist of black diagonal lines around

The primary mission of the helicopter is troop and cargo

three edges, may have the vertical line or change symbol

transport. The observance of limitations, performance,

placed along the inner margins. Symbols show current

and weight and balance data provided is mandatory.

changes only. A miniature pointing hand symbol is used

Your flying experience is recognized, therefore, basic

to denote a change to an illustration. However, a vertical

flight principles are not included. It is required that THIS

line in the outer margin, rather than miniature pointing

MANUAL BE CARRIED IN THE HELICOPTER AT ALL

hands, is used when there have been extensive changes

TIMES.

made to an illustration. Change symbols are not used to

indicate changes in the following:

1-1-4. Introductory Material.

a. Introductory material.

The following paragraphs describe certain sections of

b. Indexes and tabular data where the change can-

this manual, referenced forms, manuals, and Army Re-

not be identified.

gulations. Also included is the procedure to follow to

c. Blank space resulting from the deletion of text, an

report errors or to recommend changes.

illustration, or table.

d. Correction of minor inaccuracies, such as spell-

1-1-5. Appendix A, Reference.

ing, punctuation, relocation of material, ect., unless such

Appendix A is a listing of official publications cited within

correction changes the meaning of instructive informa-

the manual applicable to and available to flight crews.

tion and procedures.

1-1-1

TM 1-1520-240-10

1-1-13. Aircraft Designation System.

conjunction with text content, paragraph titles, and il-

lustrations. Designators may be used to indicate proper

1-1-14. The designation system prescribed by AR 70-50

effectivity, unless the material applies to all models and

is used in aircraft designation as follows:

configuration within the manual. Designator symbols

Example CH-47D

precede procedural steps in Chapters 5, 8 and 9. If the

C - Mission symbol (cargo)

material applies to all series and configurations, no des-

ignator symbol will be used.

H - Basic mission and type symbol (Helicopter)

47 - Design number

DESIGNATOR

APPLICATION

SYMBOL

D - Series symbol

712

CH-47D aircraft equipped

With T55-L-712 engines.

1-1-15. Series and Effectivity Codes.

714A

CH-47D aircraft equipped

with T55-GA-714A engines

Designator symbols listed below are used to show limited

1-1-16. Use of “Shall, Should, and May”.

effectivity of airframe information material in conjunction

with text content, paragraph titles, and illustrations. Des-

Within this technical manual, the word “shall” is used to

ignators may be used to indicate proper effectivity, un-

indicate a mandatory requirement. The word “should” is

less the material applies to all models and configuration

used to indicate a nonmandatory but preferred method

within the manual. Designator symbols precede proce-

of accomplishment. The word “may” is used to indicate

dural steps Designator symbols listed below are used to

an acceptable method of accomplishment.

show limited effectivity of airframe information material in

1-1-17

1-1-2

TM 1-1520-240-10

CHAPTER 2

AIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATION

SECTION I. HELICOPTER

2-1-1. General.

2-1-4. Landing Gear Proximity Switches.

a. Two proximity switches are installed, one on each aft

The CH-47D (fig. FO-1, 2-1-1 and 2-1-2) is a twin-turbine

landing gear. Each switch is activated when its associated

engine, tandem rotor helicopter designed for transportation

shock strut is compressed during touchdown. The switches

of cargo, troops, and weapons during day, night, visual, and

improve ground handling by reducing pitch axis gain of the

instrument conditions. (Unless otherwise noted, numbers

AFCS, by canceling the longitudinal Control Position Trans-

refer to fig. FO-1.) The helicopter is powered by two

ducer (CPT), therefore longitudinal stick input, to the Differ-

T55-L-712 or T55-GA-714A shaft-turbine engines (18) on

ential Airspeed Hold (DASH) actuators, and by driving both

the aft fuselage. The engines simultaneously drive two tan-

longitudinal cyclic trim (LCT) actuators to the ground posi-

dem three-bladed counterrotating rotors (13 and 19) through

tion. In addition to the above functions, the switch on the right

engine transmissions (25), a combining transmission (16),

aft landing gear, when activated, disables the flare dispenser

drive shafting (14), and reduction transmissions (12 and 23).

to prevent accidental flare release, and enables the hold

The forward transmission is on the forward pylon above the

function of mode 4 transponder codes.

cockpit (1). The aft transmission, the combining transmis-

b. On helicopters equipped with GROUND CONTACT

sion, and drive shafting are in the aft cabin section and aft

indicating lights, activation of the proximity switches when

pylon sections (3 and 4). Drive shafting from the combining

the associated shock strut is compressed will cause the

transmission to the forward transmission is housed in a

associated GROUND CONTACT indicating light on the

tunnel along top of the fuselage. When rotors are stationary,

MAINTENANCE PANEL to illuminate.

a gas-turbine auxiliary power unit (22) drives a generator

and hydraulic pump to furnish hydraulic and electrical power.

Fuel is carried in pods on each side of the fuselage. The

CAUTION

helicopter is equipped with four non-retractable landing gear.

An entrance door (15) is at the forward right side of the cargo

Should either or both GROUND CONTACT

compartment (2). At the rear of the cargo compartment is a

indicating lights remain illuminated after

hydraulically powered loading ramp (26). The pilots seat (9)

lift-off to hover, the illuminated system(s)

and controls are at the right side of the cockpit; the copilot’s

DASH will not function properly in forward

seat (40) and controls are on the left side. See figure 2-1-3

flight. If both GROUND CONTACT indicat-

for typical cockpit and controls.

ing lights remain illuminated after lift-off,

the AUTO function of both cyclic trim sys-

tems will be inoperative and both LCT ac-

2-1-2. Gross Weight.

tuators will remain in the GND position.

c. Aft landing gear proximity switches are not actuated

in a water landing. As a result, DASH actuators will respond

The maximum gross weight of the CH-47D is 50,000

to longitudinal stick motion, producing an apparent increase

pounds. Chapters 5 and 6 provide additional weight informa-

in control sensitivity. Cyclic motion of + 3/4 inch from neutral,

tion.

if held, will drive DASH actuators hard over. If longitudi-

nal cyclic movement is required for taxing, set the AFCS

SYSTEM SEL switch to OFF.

2-1-3. Landing Gear System.

2-1-5. Steering and Swivel Lock System.

The landing gear system consists of four non-retractable

The steering and swivel lock system consists of the power

landing gears mounted on the fuselage pods. The forward

steering control box with the STEERING CONTROL panel

landing gears are a fixed-cantilever type and have twin

on the center console, utility system pressure control

wheels. The aft landing gears are of the single-wheel, full-

module, power steering actuator, power steering module,

swivel (360_) type which can be power centered and locked

swivel lock actuating cylinder, and the PWR STEER master

in trailed position. In addition, the aft right landing gear can

caution capsule. The STEERING CONTROL panel consists

be steered from the cockpit by using the steering control

of a three position SWIVEL switch and a steering control

knob on the console. Each landing gear has an individual

knob. The SWIVEL switch controls operation of power

air-oil shock strut and is equipped with tube-type tires.

steering and swivel locks.

2-1-1

TM 1-1520-240-10

The switch positions are arranged so the power steering

swivel. Setting the SWIVEL switch to UNLOCK deener-

system cannot be energized and used with swivel locks

gizes the power steering circuits in the control box and

engaged. The aft right landing gear is hydraulically steer-

the power steering actuator. It maintains the swivel locks

able and electrically controlled by the steering control

in the disengage position and both aft wheels are free to

knob.

swivel. Setting the SWIVEL switch to LOCK energizes

the swivel lock and centering cam control valve. Utility

The PWR STEER caution capsule on the master caution

system pressure is directed to the lock port of the swivel

panel indicates that power steering circuits have failed or

lock cylinder and centering cam. The aft wheels will ro-

the aft right wheel has exceeded turning limits. These

tate to neutral trail position and the swivel lock will en-

limits are set at 58_ for a left turn and 82_ for a right turn.

gage when the helicopter weight is lifted from the rear

If turning limits are exceeded, an out-of-phase switch on

wheels. AFCS heading hold is disabled at STEER and

the landing gear automatically closes the power steering

UNLOCK.

solenoid valve, lights the caution capsule, and removes

electrical power from the control box. To reenergize the

b. Steering control knob. The steering control knob

power steering system, the landing gear must be re-

has index marks around the knob to indicate degrees of

turned within operating limits and the SWIVEL switch

knob rotation LEFT and RIGHT in increments of 30_.

must be recycled.

These index marks do not represent wheel turn angle;

they are reference marks only. The knob is spring-loaded

Hydraulic power to operate the power steering actuator

to zero turn angle. Power steering is accomplished by

and the swivel locks is supplied by the utility hydraulic

rotating the knob a given amount in the desired direction.

system through the utility system pressure control mod-

When the knob is rotated, a servo valve on the power

ule and separate power steering and swivel lock module.

steering actuator regulates hydraulic pressure to extend

Electrical power to control the steering and swivel locks

or retract the actuator. A feedback variable resistor, also

system is supplied by the No. 1 DC bus through the

on the power steering actuator, stops actuator travel

BRAKE STEER circuit breaker on the No. 1 PDP.

when the selected turn radius is reached.

2-1-6. STEERING CONTROL Panel.

2-1-7. Brake System.

The four wheels of the forward landing gear, and two

The STEERING CONTROL panel (fig. 2-1-4) is on the aft

wheels of the aft landing gear, are equipped with self-ad-

end of the console. It contains the SWIVEL switch, the

justing disk brakes. Both forward and aft brakes can be

steering control knob, a fail-safe module and relay, and

applied and brake pressure maintained by depressing

a servoamplifier. The fail-safe module monitors the

the pedals. Hydraulic pressure is supplied by utility hy-

steering electrical circuits. A malfunction which could

draulic system.

cause a steering hardover will be detected by the fail-

safe module and the relay which disables the system and

2-1-8. Brake Pedals.

turns on the PWR STEER caution light.

When either the pilot’s or copilot’s brake pedals are

a. SWIVEL switch. A three-position switch labeled

pressed, pressure from the master brake cylinders goes

STEER, UNLOCK, and LOCK. Setting the switch to

to a transfer valve in the brake lines. This allows indepen-

STEER applies DC power to the circuits in the power

dent braking by either pilot. From these transfer valves,

steering control box and arms the power steering actua-

pressure is directed through a parking brake valve to the

tor. Rotating the steering control knob will activate the

forward and aft wheel brakes.

power steering actuator and the aft wheel will

2-1-9. Parking Brake Handle.

A parking brake handle (4, fig. 2-1-3) is at the bottom left

corner of the pilot’s section of the instrument panel. The

brake handle is mechanically connected to the parking

brake valve. The parking brake valve is electrically con-

nected to the PARK BRAKE ON caution capsule on the

master caution panel. When the brake pedals are

pressed and the parking brake handle is pulled OUT,

pressure is trapped and maintained on forward and aft

wheel brakes. At the same time, electrical power from the

DC essential bus through the LIGHTING CAUTION PNL

circuit breaker, lights the PARK BRAKE ON caution cap-

sule.

The parking brakes must be released by applying pres-

sure to the brake pedals. This action automatically opens

the parking brake valve, retracts the parking brake han-

dle, and extinguishes the PARK BRAKE ON caution cap-

Figure 2-1-4. Steering Control Panel

sule.

2-1-5

TM 1-1520-240-10

2-1-10. Brakes and Steering Isolation Switch.

fore and aft and is locked and unlocked by a handle at the

forward end of the jettisonable door. The handle is moved

The brakes and steering isolation switch is on the HYD

forward to lock the window and aft to unlock the window.

control panel on the overhead switch panel (fig. 2-1-10

and 2-1-14). It is labeled BRK STEER, ON, and OFF. The

2-1-16. Seats.

switch isolates the brakes and steering hydraulic subsys-

tems from the rest of the utility hydraulic system in the

The pilot’s and copilot’s seats (9 and 40, fig. FO-1) are on

event of a leak in the subsystem. The normal position of

tracks to permit forward and aft, vertical and reclining posi-

the switch is ON. The switch is guarded to ON. Setting the

tion adjustments. Bungee cords in each seat exert an up-

switch to OFF, closes the power steering and brakes

ward force on the seat when it is down or tilted.

valve on the utility system pressure control module, iso-

lating the brakes and steering subsystem. With the

2-1-17. Seat Fore-and-Aft Lever.

switch at OFF, limited brake application are available due

A fore and aft control lever (14, fig. 2-1-3) for horizontal seat

to an emergency brake accumulator in the brake subsys-

adjustment is on the right side of each seat support carriage.

tem. Power to operate the isolation valve is from the No.

When the lever is pulled UP, the seat is unlocked and can

1 DC bus through the HYDRAULICS BRAKE STEER

be moved along the tracks on the cockpit floor. When the

circuit breaker on the No. 1 PDP.

lever is released, the seat is locked in position horizontally.

The total range of the horizontal movement is 4 inches in 1

2-1-11. Instrument and Control Panels.

inch increments.

NOTE

2-1-18. Seat Vertical Lever.

The NVG overhead switch panels are shown.

Vertical seat adjustment (15, fig. 2-1-3) is controlled by a

Description of control panels and operating

lever on the right side of each seat. When this lever is pulled

procedures reflect NVG configuration only.

UP, the seat is unlocked and can be moved vertically along

a track through a range of 5 inches. The range is divided into

Figures 2-1-5 and 2-1-6 show center and canted

1/2 inch increments. When the lever is released, the seat is

consoles. 712 Figures 2-1-7 through 2-1-10 show the

locked in position vertically.

copilot instrument panel, center instrument panel, pilot

instrument panel, and the NVG overhead switch panel.

714A Figures 2-1-11 through 2-1-14 show the copilot

CAUTION

instrument panel, center instrument panel, pilot instru-

ment panel, and the NVG overhead switch panel.

With the seat in the full up rotation posi-

tion (zero tilt) the seat may not be able to

2-1-12. Personnel/Cargo Doors.

be locked in the full down vertical posi-

tion. Ensure the seat is locked when ad-

Entry can be made through either the main cabin door or the

justing the vertical axis, especially when

cargo door and ramp.

the seat is in full up rotation position (zero

tilt).

2-1-13. Main Cabin Door.

The main cabin entrance (15, fig. FO-1) door is on the right

2-1-19. Seat Rotation Lever.

side of the cargo compartment. The door is divided into two

A control lever (20, fig. 2-1-3) for adjusting the seat reclining

sections: the upper section containing a jettisonable panel

position is on the left side of each seat. When this lever is

and the lower section forming the entrance step. When

pulled UP, the seat is unlocked and can be rotated through

opened, the upper section slides upward on overhead rails

a 15º tilt range divided into four equal increments. The seat,

and the lower section swings downward. When closed, the

in effect, is pivoted up and down around a horizontal axis.

two sections mate to form the complete door. Handles are

When the lever is released, the seat is locked in the selected

provided on both the outside and the inside of the door for

tilt position.

accessibility. Refer to Chapter 5 for the allowable airspeed

imposed on the helicopter while operating with the cabin

2-1-20. Armored Seats.

entrance door sections in various positions.

Both the pilot and the copilot seats are equipped with a

2-1-14. Cargo Door and Ramp.

combination of fixed and adjustable ceramic armor panels

(fig. 2-1-15). Fixed panels are installed under the back and

Chapter 6 provides a detailed description and operation of

bottom seat cushions and on the outboard side of each seat.

the cargo door and ramp.

A shoulder panel (if installed) is mounted on the outboard

side of each seat. The shoulder panel is hinged from the

2-1-15. Pilot and Copilot Sliding Windows.

seat back so it can be moved aside for ease of exit from the

The upper section of each jettisonable door (39, fig. FO-1)

helicopter. The panel is secured in its normal position by a

in the cockpit contains a sliding window. The window slides

latch and an exerciser cord.

2-1-6

TM 1-1520-240-10

2-1-21. Shoulder Harness Inertia Reel Lock Lever.

2-1-22. Self -Tuning Dynamic Absorbers.

The helicopter is equipped with three sel-tuning dynamic

A two-position shoulder harness inertia reel lock lever is on

absorbers. One absorber is in the nose compartment and

the left side of each seat (22, fig. 2-1-3) The lever positions

the other two absorbers are under each pilot’s seat below

are LOCKED (forward) and UNLOCKED (aft). The lock may

the cockpit floor. All three absorbers serve to maintain a

be moved freely from one position to the other. When the

minimum vibration level through the normal operating

lock lever is in UNLOCKED position, the reel harness cable

rotor RPM range of the helicopter. The self-tuning feature

is released to allow freedom of movement. However, the

of the the dynamic absorber functions as follows: each

reel will automatically lock if a horizontal impact force of 2 to

dynamic absorber consists of tuning mass suspended by

3 g is encountered. When the reel is locked in this manner,

springs, and electronic measuring circuit, accelerome-

it stays locked until the lock lever is moved forward to

ters, counter-weights, an electrical actuator and a self-

LOCKED and then returned to UNLOCKED. When the lever

test box. The accelerometers sense and compare the

is at LOCKED, the reel is manually locked so the pilot is

vibration phases of the helicopter and the spring-

restrained from bending forward. When a crash landing or

mounted mass. When the measured vibration phases

ditching is anticipated and time permits, manual locking of

differ from a built-in phase relationship required to assure

the shoulder harness inertia reel provides added safety be-

proper tune, the electronic circuit extends or retracts the

yond the automatic feature of the inertia reel. Depending on

electrical actuator to reposition the counterweights

the pilot’s seat adjustment, it may not be possible to reach

which, in turn, increases or decreases the resonant fre-

all switches with the inertia locked. Each pilot should check

quency of the spring-mounted mass. The dynamic ab-

and adjust the shoulder harness in locked position to deter-

sorbers are constantly being adjusted (tuned) to mini-

mine whether all switches can be reached.

mize helicopter vibration. A self-test box is in the heater

compartment to provide maintenance personnel with an

integral testing capability for self-tuning feature of the

dynamic absorbers. Power is supplied by the No. 2 AC

bus through the VIB ABSORB-LH, CTR and RH circuit

breakers in the No. 2 PDP.

Figure 2-1-6. Canted Console (Typical)

2-1-9

TM 1-1520-240-10

SECTION II. EMERGENCY EQUIPMENT

2-2-1. Emergency Procedures.

2-2-3. FIRE PULL Handles.

Refer to Chapter 9 for all emergency procedures.

WARNING

2-2-2. Engine Compartment Fire Extinguisher Sys-

tem.

Before flying the aircraft ensure that each

FIRE PULL handle NVG filter holder can be

The engine compartment fire extinguisher system (fig.

rotated from the closed to the open posi-

2-2-1) enables the pilot or copilot to extinguish a fire in

tion without causing the FIRE PULL han-

either engine compartment only. It is not designed to

dle to be pulled. Improper handling of the

extinguish internal engine fires. The system consists of

NVG filter holder may cause the FIRE

two FIRE PULL handles, an AGENT DISCH (agent dis-

PULL handle to be pulled unintentionally,

charge) switch, a FIRE DETR (fire detector) switch on the

thus fuel to the affected engine will be shut

center instrument panel, and two extinguisher agent con-

off and the engine will shut down. Do not

tainers on the overhead structure at stations 482 and

use sudden or excessive force when rotat-

502. The containers are partially filled with monobromo-

ing the FIRE PULL handle NVG filter hold-

trifluoromethane (CBrF3 or CF 3BR) and pressurized with

er from the closed to the open position.

nitrogen (table 2-2-1 provides the range of engine fire

Two control handles for the engine fire extinguisher sys-

extinguisher pressures.) The agent in one or both of the

tem (fig. 2-2-1) are labeled FIRE PULL - FUEL SHUT-

containers can be discharged into either engine

OFF on the top center section of the center instrument

compartment. Selection of the compartment is made by

panel. Each handle has a cover for the NVG filter, two

pulling the appropriate FIRE PULL handle. In figure 2-2-1

warning lights, and the necessary control switches that

the ENG 1 FIRE PULL handle has been pulled. Selection

close the engine fuel shutoff valve and arm the fire extin-

of the container is made by placing the AGENT DISCH

guisher system circuits. Power is supplied for each FIRE

switch in the appropriate position. In figure 2-2-1, BTL 1

PULL handle from the respective No. 1 and No. 2 DC

has been selected.

essential buses through the respective ENGINE NO. 1

and NO. 2 FUEL SHUTOFF circuit breakers on the No.

1 and No. 2 PDP. Power is supplied for each pair of

warning lights from the corresponding No. 1 or No. 2 AC

bus through the ENGINE NO. 1 and NO. 2 FIRE DET

circuit breakers on the No. 1 and No. 2 PDP.

Table 2-2-1. Engine Compartment Fire Extinguisher Pressures

AMBIENT TEMPERATURE

MINIMUM INDICATION

(C)

(PSI)

-54_

271

-51_

275

-40_

292

-29_

320

-18_

355

-7_

396

449

15_

518

27_

593

38_

691

52_

785

2-2-1

TM 1-1520-240-10

compartment, agent will not be available should a fire

WARNING

occur in the other engine compartment. Power is sup-

plied from the corresponding No. 1 or No. 2 DC essential

If the FIRE PULL handle warning lights are

bus through the ENGINE NO. 1 and NO. 2 FIRE EXT

covered by the NVG filters during daylight

circuit breakers on the No. 1 and No. 2 PDP.

operation, illumination of the fire warning

lights may not be apparent in the event of

2-2-5. FIRE DETR Switch.

an engine fire. Do not operate the aircraft

A two-position FIRE DETR (detector) switch is below the

with the NVG filters covering or obscuring

AGENT DISCH switch on the top center section of the

the fire warning lights unless night vision

center instrument panel (fig. 2-2-1). It is labeled FIRE

goggles are being used.

DETR and TEST. The toggle switch is spring-loaded to

The NVG filter is attached to one end of the FIRE PULL

FIRE DETR which monitors the engine fire detection

handle by hinged fitting. The other end of the filter holder

system. When the switch is placed to TEST, it checks the

forms a tab by which the filter holder and filter may be

operation of the engine fire detection system by closing

rotated about the hinged fitting. For NVG operations, the

relays in both controls units and the warning lights in both

filter holder is rotated to a closed position over the front

FIRE PULL handles illuminate. Power to operate the test

of the FIRE PULL handle cover. In this position, the fire

circuit is supplied by the DC essential bus through the

warning light is NVG compatible. For normal operations,

LIGHTING CAUTION PNL circuit breaker on the No. 1

the filter holder is rotated from the closed position to the

PDP.

fully open position. In this position, the FIRE PULL han-

dle warning lights will be red.

2-2-6. Hand Fire Extinguishers.

CAUTION

WARNING

If there is a fire in both engine compart-

Avoid prolonged exposure (5 minutes or

ments, do not pull both FIRE PULL han-

more) to high concentrations of fire extin-

dles simultaneously. Extinguish fire in

guishing agent and its decomposition

one compartment only as described be-

products because of irritation to the eyes

low. Leave the FIRE PULL handle out after

and nose. Adequate respiratory and eye

fire has been extinguished. Proceed in a

relief from excessive exposure should be

like manner to extinguish fire in the other

sought as soon as the primary fire emer-

engine compartment.

gency permits. Use of oxygen for person-

When an engine compartment fire occurs on either side,

nel is recommended.

the respective pair of warning lights comes on. The ap-

Three portable 6.3 pound capacity hand fire extinguish-

propriate FIRE PULL handle is pulled, that engine fuel

ers are provided in the helicopter. One is in the cockpit,

shutoff valve closes and the AGENT DISCH switch is

on the floor to the right of the pilot’s seat. Two hand fire

armed.

extinguishers are in the cabin section. One on the for-

ward bulkhead and one in the left rear, just forward of the

Selection and discharge of either fire bottle is accom-

ramp.

plished by placing the AGENT DISCH switch to BTL 1 or

BTL 2. After depletion of the charge in the initially se-

2-2-7. Emergency Troop Alarm and Jump Lights.

lected bottle, the remaining bottle can be discharged to

the same engine compartment by selecting the opposite

Two emergency troop alarm and jump light boxes are in

position on the AGENT DISCH switch. The other FIRE

the cargo compartment. The forward box is on the bulk-

PULL handle performs the same function for its respec-

head and above the avionics equipment shelves and the

tive engine compartment.

aft box is on the left side of the fuselage above the ramp

at sta. 575. Each box has an electric bell in the center with

2-2-4. AGENT DISCH Switch

a red light fixture on one side and a green light fixture on

the other side. The TROOP WARN panel on the over-

A three-position AGENT DISCH (discharge) switch is

head switch console is used to operate the emergency

above the FIRE PULL handles on the center instrument

troop alarm and jump lights.

panel (fig. 2-2-1). The lever-lock momentary switch posi-

tions are BTL 1, neutral, and BTL 2. When BTL 1 is

The emergency troop alarm and jump lights have several

selected, the agent is discharged from the No. 1 bottle

functions. They can be used to notify passengers and

into the selected engine compartment. When BTL 2 is

crew with predetermined signals in time of emergency.

selected, the agent is discharged from the No. 2 bottle

The jump lights can be used to notify flight engineer

into the selected engine compartment. Only two fire ex-

during airborne delivery operations and to alert the troop

tinguisher agent bottles are provided. If the agent from

commander during paratroop drop missions. Refer to

both bottles is used in combating a fire in one engine

Chapter 9 for standard use of the troop alarm.

2-2-3

TM 1-1520-240-10

2-2-8. TROOP WARN Panel

The TROOP WARN (warning) panel is located on the

overhead switch panel (fig. 2-2-2). It has two troop jump

lights labeled RED and GREEN. Also, two switches la-

beled JUMP LT and ALARM. Power to operate and con-

trol the emergency troop alarm and jump lights is sup-

plied by the DC essential bus through the TROOP

ALARM BELL and TROOP ALARM JUMP LT circuit

breakers on the No. 2 PDP.

a. Troop jump lights. The troop jump lights provides

the pilots a visual indication of the troop jump light selec-

ted. One light is provided for each color selection and

Figure 2-2-2. Troop Warning Panel (Typical)

comes on when the respective light is selected. The

brightness of the lights is controlled by the PLT INST

rotary control switch on the PLT LTG panel of the over-

2-2-9. First Aid Kits.

head switch panel.

Seven aeronautic first aid kits are installed in the helicop-

ter. One kit is in the passageway between the cockpit and

b. JUMP LT switch. The three-position JUMP LT

cabin. The other six kits are in the cabin fuselage section,

switch is labeled GREEN, OFF, and RED. When the

three on each side.

switch is set to GREEN, the green lights on the emergen-

cy troop and jump light box, at both stations, and the

2-2-10. Emergency Entrances and Exits.

troop jump lights on the overhead switch panel come on.

When the switch is set to RED, the red lights come on.

Refer to Chapter 9 for information on emergency en-

OFF position turns off both sets of lights.

trances and exits.

2-2-11. Emergency Escape Axe.

c. ALARM switch. The two-position ALARM switch

is labeled OFF and ON. Moving the ALARM switch to ON

An emergency escape axe is provided. It is located on

rings the bell continuously at both stations until the switch

the right side of the cargo compartment slightly forward

is moved to OFF.

of station 200.

2-2-4

TM 1-1520-240-10

SECTION III. ENGINES AND RELATED SYSTEMS

2-3-1. Engines.

fuel control unit. Each system provides automatic control

of engine gas producer rotor speed and power turbine

The CH-47D is powered by either two T55-L-712 or two

speed in response to any setting of the engine controls

T55-GA-714A engines. The engines are housed in sepa-

selected by the pilot. Engine gas producer rotor speed

rate nacelles mounted externally on each side of the aft

(N1) and power turbine speed (N2) are controlled by the

pylon. The engines have the capability to produce emer-

fuel control unit, which varies the amount of fuel delivered

gency power on pilot demand. See Performance Charts

to the engine fuel nozzles. During normal operation, the

in Chapter 7 712 or Chapter 7A 714A .

fuel control unit automatically controls fuel flow metering

during power changes, thus protecting the engine from

overspeed and overtemp. Fuel flow is automatically

2-3-2. General

monitored to compensate for changes in outside air tem-

Each engine has a gas producer section and a power

perature and compressor discharge pressure.

turbine section. The gas producer supplies hot gases to

2-3-6. Engine Fuel Control Units. 712

drive the power turbine. It also mechanically drives the

engine accessory gearbox. The power turbine shaft ex-

Each engine fuel control unit contains a single element

tends coaxially through the gas producer rotor and ro-

fuel pump, a gas producer speed governor, a power tur-

tates independently of it. The gas producer section and

bine speed governor, an acceleration-deceleration con-

the power turbine section are connected by only the hot

trol, a fuel flow limiter, a fuel control shutoff valve, and a

gases which pass from one section to the other.

main metering valve. A gas producer (N1) lever and a

power turbine (N2) lever are mounted on the fuel control

During engine starting, air enters the engine inlet and is

unit.

compressed as it passes through seven axial stages and

one centrifugal stage of the compressor rotor. The com-

Output power of the power turbine (a function of the

pressed air passes through a diffuser. Some of the air

speed and torque) is restricted by limiting the maximum

enters the combustion chamber where it is mixed with

fuel flow to the gas producer. Maximum gas producer

start fuel.

rotor speed is set by the ENG COND (engine condition)

The mixture ignited by four igniter plugs. Some of the air

levers in the cockpit. The ENG COND levers electrome-

is directed to the fuel nozzles. After the engine is started,

chanically positions the gas producer lever, which con-

it continues to operate on metered fuel supplied to the

trols the fuel control fuel shutoff valve and operating level

fuel nozzles.

of the gas producer. During flight, the ENG COND levers

Hot expanding gases leave the combustion chamber and

are left at FLT and the output shaft speed is regulated by

drive a two-stage gas producer turbine. Energy from the

the power turbine speed (N2) governor.

combustion gases also drives the two-stage power tur-

The power turbine lever is electromechanically posi-

bine, which drives the power turbine shaft to the engine

tioned by the ENGINE BEEP TRIM switches, thrust con-

transmission. The engine lubrication system has an inte-

trol, and EMERG ENG TRIM (emergency engine trim)

gral oil tank which is inside the air inlet housing and is

712 switches. Output shaft torques are limited by the

serviced with approximately 12 quarts. (Refer to table

fuel flow limiter, which limits the maximum fuel flow. The

2-15-1.)

position of the main metering valve is determined by the

2-3-3. Engine Inlet Screens.

gas producer speed governor, power turbine speed gov-

ernor, the acceleration-deceleration control, or the fuel

An engine inlet screen which minimizes foreign object

flow limiter, depending on engine requirements at that

damage (FOD) is installed on each engine. The reduc-

tion in engine power available with screens installed is

time. The governor or the control unit demanding the

negligible. The engine inlet screens have bypass panels.

least fuel flow overrides the other in regulating the meter-

These two panels are on the aft end of each screen.

ing valve.

Refer to Chapter 5 for information on use of bypass pan-

2-3-7. Speed Governing.

els. Helicopters with engine air particle separator (EAPS)

installed, refer to TM 1-1520-240-10 EAPS SUPPLE-

The power turbine speed governor senses the speed of

MENT.

the power turbine and regulates the amount of fuel which

is supplied to the gas producer. This slows down or

2-3-4. Engine Anti-Icing.

speeds up the gas producer rotor so that power turbine

The engine air inlet fairing and engine drive shaft fairing

and rotor system speed remains nearly constant as loads

receive anti-icing protection from the thermal radiation

vary.

produced by the oil tank in the engine inlet housing. The

hot oil in the oil cavity of the inlet housing warms the air

At minimum rotor blade pitch, the amount of power re-

as it passes into the engine inlet.

quired is at minimum. As pitch is increased, power tur-

bine speed (N2) starts to decrease since more power is

2-3-5. Engine Power Control System. 712

required from the engine to maintain a constant rotor

speed. The power turbine speed governor senses the

Each engine is controlled by a separate power control

decrease of N2 RPM and increases the flow of fuel to the

system which includes cockpit controls and an engine

gas producer. Decreasing pitch causes N2 to increase.

2-3-1

TM 1-1520-240-10

The power turbine governor senses the increases and

CAUTION

reduces the flow of fuel to the gas producer, thus de-

creasing the engine output power.

When the ENG COND lever is placed to

The power turbine speed governor allows the power tur-

GND during start sequence, the N1 actua-

bine output speed to decrease (droop) approximately 10

tor could inadvertently go beyond the

percent when the power loading varies from minimum to

ground position. The respective ENG N1

full load. This is minimized by a droop eliminator linked

COND caution capsule will illuminate.

to the thrust control rod. The droop eliminator automati-

However, ignition will still occur if the start

cally changes the power turbine lever to compensate for

switch is moved to START, thus resulting

droop as pitch is increased or decreased. Another type

in a possible engine runaway.

of droop, which is only transient, occurs as a result of the

time required for the engine to respond to changing loads

due to system lag.

CAUTION

2-3-8. ENG COND Levers 712

When adjusting controls or switches on

Two ENG COND (engine condition) levers, one for each

the overhead switch panel, make sure

engine are on the ENG COND panel (fig. 2-3-1) of the

gloves or sleeves do not catch and inad-

overhead switch panel. Each lever has three positions

vertently move the ENG COND levers.

labeled STOP, GND, and FLT. They are used to select

appropriate fuel flow rates for GND, FLT, and STOP (en-

The ENG COND lever must be at GND before the engine

gine shutdown). Power is supplied by the DC essential

will start. When an ENG COND lever is advanced from

buses through the ENGINE NO. 1 and NO. 2 COND

STOP to GND, power is then supplied to the electrome-

CONT circuit breakers on the No. 1 and No. 2 PDP.

chanical actuator which establishes an appropriate fuel

flow rate at ground idle. The speed of the gas producer

Each ENG COND lever is spring-loaded outboard and is

with the lever at GND should be 60 to 63 percent N1.

inhibited by lock gates. They allow the pilot to proportion-

When an ENG COND lever is moved to FLT, the engine

ally control acceleration of the gas producer from STOP

is operating within the N2 governing range, unless the

to FLT. Two engine control caution capsules are on the

engine is “topped out” at which time it goes back to N1

master caution panel (fig 2-14-5). They are labeled NO.

governing. The N2 governor then takes control to main-

1 ENG N1 CONT and NO. 2 ENG N1 CONT. The cap-

tain selected rotor RPM (RRPM) in response to the en-

sules normally illuminate when the ENG COND levers or

gine beep trim switches and collective pitch changes,

the N1 actuators are at an intermediate position between

When an ENG COND lever is moved to STOP, the gas

STOP, GND, or FLT. They extinguish when the ENG

producer lever closes the fuel control fuel shutoff valve

COND lever and N1 actuator positions agree. However,

which stops fuel flow to the gas producer.

they remain illuminated if a component of the system

(actuator, control box, or condition panel) has failed in

Each electrical system is completely separate and a fail-

other than a detent position. Power is supplied by the DC

ure in one system will not affect the other. A built-in me-

essential bus through the LIGHTING CAUTION PNL cir-

chanical brake holds the actuator at its last selected posi-

cuit breaker on the No. 1 PDP.

tion if loss of electrical power occurs. ENG COND lever

friction is provided to reduce the possibility of overtorqu-

ing the engine transmissions by resisting movement of

the ENG COND levers. The ENG COND lever friction

brake cannot be adjusted by the pilot and a force of 4 to

5 pounds is needed to move them.

2-3-9. Normal Engine Beep Trim Switches.

712 On 712 engine installations engine beep trim

switches are active at all times during normal operation.

Two momentary switches are on the auxiliary switch

bracket of each THRUST CONT lever and are labeled

ENGINE BEEP TRIM (fig. 2-5-1). Both switches have an

RPM INCREASE, RPM DECREASE, and a neutral posi-

tion. 712 One switch is labled NO. 1 & 2 which is normal-

ly used to select desired RRPM. The second switch is

labeled NO. 1 which will only affect the No. 1 engine and

is used to match engine loads which are indicated by the

Figure 2-3-1. Engine Condition Panel 712

dual torquemeters.

2-3-2

TM 1-1520-240-10

712 Power to operate the beep trim system is supplied

switches and two momentary emergency engine trim

by the DC and AC buses. DC power to operate a trim

switches.

motor in the power turbine control box, which unbalances

a. Normal Engine Trim System Disable

a control circuit, is supplied by the corresponding No. 1

Switches. The guarded switches permit the pilot to dis-

or No. 2 DC buses through the ENGINE NO. 1 or NO. 2

able either or both normal beep trim systems. This pre-

TRIM circuit breakers on the No. 1 or No. 2 PDP. The

vents unwanted signals from the normal beep trim sys-

unbalanced control circuit causes the AC power from the

tem to interfere with the operation of the emergency

No. 1 or No. 2 AC buses through the ENGINE NO. 1 or

engine trim system. Each switch is labeled AUTO and

NO. 2 TRIM & TIMER circuit breakers on the No. 1 or No.

MANUAL. When either switch is at MANUAL, the respec-

2 PDP to be transformed and rectified to DC voltage. This

tive normal beep trim system is disabled (115-volt AC

from AC bus to the engine power turbine control box is

DC power operates the power turbine actuator on the

interrupted). When the switch is at AUTO (cover down),

engine fuel control.

the normal beep trim system is functional (115-volt AC

NOTE

from the AC bus is reconnected to the associated engine

power turbine control box). Refer to Chapter 9 for emer-

No two engines provide matched perfor-

gency engine trim operation.

mance with regard to torque, RPM, PTIT, or

fuel flow. With torque matched all other pa-

rameters may not be matched.

CAUTION

712 Holding the No. 1 & 2 switch forward (RPM IN-

CREASE) will increase the RRPM. Holding the switch aft

Engine response is much faster when

(RPM DECREASE) will decrease the RRPM. When the

RRPM is controlled with emergency en-

switch is released, it returns to the center or neutral posi-

gine beep trim system. It is possible to

tion. The switch electrically controls both power turbines

beep the rotor speed below safe operating

by movement of the N2 actuator through each engine

speed and low enough to disconnect the

power turbine control box.

generators from the buses. The genera-

The procedure for matching engine load requires that

tors are disconnected at 85% to 82%

NO. 1 & 2 engine beep switch be used in conjunction with

RRPM after a 3 to 7 second time delay.

NO. 1 engine beep switch. When NO. 1 engine beep

switch is moved forward (RPM INCREASE), the torque

b. Emergency Engine Trim Switches. Each mo-

of No. 1 engine increases. At the same time RRPM in-

mentary switch is used to change the power turbine

creases, even though No. 2 engine torque decreases

speed of its respective engine if the power turbine control

slightly. Moving NO. 1 & 2 engine beep trim switch aft

box (normal beep trim system) malfunction.

(RPM DECREASE) causes both engine torques to de-

crease and reduce RRPM. If torques are still not

When the normal trim system fails, the droop eliminator

matched, this procedure is continued until torques are

also fails to function. Both switches have an INC, DECR,

matched and desired RRPM is attained. The opposite

and spring-loaded center position. When one of the

action occurs when NO. 1 engine beep switch is moved

switches is held at INC, power from the essential DC bus

aft.

goes directly to the respective power turbine actuator

The engine beep trim switches should not be used during

and increases the lever setting and the power turbine

power changes initiated by thrust lever movement be-

speed. When the switch is held at DECR, the lever set-

cause RRPM droop should only be momentary. The en-

ting is decreased, and the power turbine speed is de-

gine beep trim system adjusts engine RPM only if the

creased.

respective ENG COND lever is at FLT. At STOP or GND,

it is possible to move the power turbine lever by moving

The emergency engine trim switches are to be used

the engine beep trim switches to RPM DECREASE or

when the normal beep trim system is disabled. If one of

RPM INCREASE, but in either case, engine RPM will not

the switches is used while the respective power turbine

be affected because the engine is not operating in the N2

control box is functioning normally, the power turbine

governing range.

actuator setting will temporarily change but will return to

its original setting when the switch is released. Power to

2-3-10. EMERG ENG TRIM Panel 712

operate the emergency engine beep trim switches and

The EMERG ENG TRIM (emergency engine) panel is

actuators is supplied by the essential DC bus through the

located on the center console (fig, 2-3-2). The panel con-

NO. 1 and NO. 2 EMERG ENG TRIM circuit breakers on

sists of two guarded normal engine trim system disabled

the No. 1 and No. 2 PDP.

2-3-3

TM 1-1520-240-10

stops are stowed on the right side of the center console

aft of the pedals.

2-3-12. EMERGENCY POWER Panel. 712

The EMERGENCY POWER panel is located on the over-

head switch panel (fig. 2-3-3). It consists of an emergen-

cy power indicator and a digital timer for each engine.

They are labeled NO. 1 and NO. 2 ENGINE. The timer

counts the minutes that emergency power is in use.

2-3-13. Oil Supply System.

The oil supply system is an integral part of the engine.

The oil tank is part of the air inlet housing and the filler

neck is on the top of the housing. An oil level indicator is

on the left side of the engine inlet housing. Refer to table

2-15-1 for the tank capacity. If the oil level decreases to

Figure 2-3-2. Emergency Engine Trim Panel 712

about 2 quarts usable, the corresponding ENG OIL LOW

caution capsule will illuminate.

2-3-11. Emergency Power System. 712

CAUTION

To prevent damage, monitor the torque

and the PTIT indicators when operating

with emergency power. Failure to observe

these indicators could result in serious

damage to the drive train and engine.

An emergency power system is included with T55-L-712

engines. With the emergency power system, increased

Figure 2-3-3. Emergency Power Panel 712

power is available on pilot demand and is actuated by

raising the thrust control into the emergency power

range. Refer to Chapter 5 for limitations on its use.

2-3-14. Engine Start System. 712

When fuel flow increases to the point where PTIT is 890_

The engine start system includes the hydraulic starters

to 910_C, the EMERG PWR lights will illuminate on the

on each engine, the engine start valves and the solenoid-

copilot and pilot instrument console (17, fig. 2-1-7 and 18,

operated pilot valves on the utility system pressure con-

fig. 2-1-9). If temperature is maintained in this range for

trol modules, the START switch, and the start fuel sole-

more than 5 seconds, the associated indicator will apply

noids and ignition exciters on the engines.

28-volt DC from the ENGINE NO. 1 and / or NO. 2 START

& TEMP circuit breaker to the EMERGENCY POWER

When the start switch is moved to MTR, the respective

panel. With 28-volt DC applied to the panel, the applica-

engine STARTER ON indicator light illuminates and the

ble emergency power timer will start, and the indicator

start valve opens (fig. 2-3-4). The start valve applies

will display a black-and-white flag. When thrust is re-

utility system pressure from the APU to the engine start-

duced below the emergency power level, the emergency

er: rotating the engine starter and compressor. At 15

power light will extinguish and the timer will stop. Howev-

percent N1, the ENG COND lever is moved to GND. The

er, the emergency power indicator will continue to display

start switch is immediately moved to START, energizing

the black-and white flag. The flag can be reset on the

the ignition exciter. Start fuel is sprayed into the combus-

ground only.

tor and combustion begins. Before PTIT reaches 200_C,

the START switch is manually released to MTR. At MTR,

WARNING

the start fuel valve is closed and the ignition exciter is

deenergized.

Before flight, be sure the two topping

The engine then accelerates to ground idle speed. At 50

stops are in their stowed position on the

percent N1, the START switch is manually moved to the

right side of the console. If stops are not

locked OFF position. At OFF, the pilot valve closes, clos-

stowed, be sure the stops are not installed

ing the start valve and deenergizing the STARTER ON

on the fuel controls before you start the

indicator light. A relay in each engine start circuit is ener-

engine. Failure to check may result in in-

gized when either START switch is at MTR or START.

ability to achieve emergency power in an

The relay, when energized, disables the start circuit of

emergency.

the opposite engine, thus preventing simultaneous dual

Topping stops are stowed on each helicopter. The stops

engine starts. Power is supplied by the No. 1 and No. 2

are installed on the N1 control of each engine for mainte-

DC essential buses through the ENGINE NO. 1 and NO.

nance engine topping checks. The stops provide an es-

2 START & TEMP AND IGN CIRCUIT BREAKERS ON

tablished fuel flow when topping. When not in use, the

THE No. 1 AND No. 2 PDP.

2-3-4

TM 1-1520-240-10

box section of each engine. 712 The outer scale of the

tachometer is calibrated from 0 to 100 in increments of

two. The smaller, vernier scale is calibrated from 0 to 10,

in increments of one. 714A The tachometer is calibrated

from 0 to 110.

2-3-19. Torquemeter.

One torquemeter is on the copilot instrument panel and

the other on the pilot instrument panel (1, fig. 2-1-7 and

17, fig. 2-1-9). Each torquemeter has two pointers, one

for each engine, labeled 1 and 2. Each torquemeter has

Figure 2-3-4. ENGINE START PANEL 712

a range of 0 to 150 percent. The system consists of a

2-3-15. START Panel. 712

power output shaft, torquemeter head assembly, power

supply unit,

714A ratio detector power supply unit

The START panel is located on the overhead switch pan-

(RDPS), and a torquemeter junction box. Power to oper-

el (fig. 2-3-4). It consists of the ENG 1 and ENG 2 START-

ate the torquemeter is provided by No. 1 and No. 2 AC

ER ON indicator lights and two start switches.

buses through the ENGINE NO. 1 and NO. 2 TORQUE

a. Start Switches. The switches are labeled OFF,

circuit breakers on the No. 1 and No. 2 PDP. Power for

MTR, and START. They are locked in OFF, detented in

the power supply unit 714A and RDPS is provided by

MTR and spring-loaded from START to MTR. At MTR,

the No. 1 and No. 2 DC buses through the DC ENGINE

the engine is rotated by the starter, but ignition and start

NO. 1 and NO. 2 TORQUE circuit breakers on the No. 1

fuel circuits are deenergized. At START, the engine is

and No. 2 PDP.

rotated with start fuel and the ignition circuits are energi-

2-3-20. Power Turbine Inlet Temperature Indica-

zed. MTR is selected during starting, in case of engine

tors.

fire or to clear the combustion chamber.

Two power turbine inlet temperature (PTIT) indicators,

b. STARTER ON Indicator Lights. The STARTER

one for each engine, are on the center instrument panel

ON indicator lights will illuminate when the associated

(7, fig. 2-1-8, 7, fig. 2-1-12). Each indicator is calibrated

START switch is moved to MTR or START. The light

from 0_ to 1,200_C. The temperatures registered on the

alerts the pilots when the START switch is inadvertently

PTIT indicator are transmitted by chromel-alumel ther-

left at MTR. Power is supplied by the No. 1 and No. 2 DC

mocouples. the thermocouples sense gas temperature

essential buses through the ENGINE NO. 1 and NO. 2

at the power turbine inlet and transmit an average gas

START & TEMP circuit breakers on the No. 1 and No. 2

temperature reading to the PTIT indicator in the cockpit.

PDP.

712 When power turbine inlet temperature increases to

2-3-16. Ignition Lock Switch.

the emergency power range, the EMERG PWR indicator

light will illuminate and DC power is supplied to the

An ignition system lock switch (11, fig. 2-1-3) is installed

EMERGENCY POWER panel. 714A When power turb-

on the right side of the console forward of the thrust lever.

ine inlet temperature increases to the contingency power

The key-operated switch prevents unauthorized use of

range, the ENG CONT PWR master caution advisory

the helicopter. When the switch is off, the circuits of the

panel capsule will illuminate.

ignition exciters and the start fuel solenoids of both en-

2-3-21. Engine Oil Pressure Indicator.

gines are open. Therefore, the engines cannot be star-

ted. Be sure both START switches are OFF before turn-

An engine oil pressure indicator on the center instrument

ing the ignition lock switch ON or OFF.

panel is provided for each engine (17, fig. 2-1-8 and

2-1-12). Each indicator relates pressure sensed at No. 2

2-3-17. Engine Instruments and Cautions.

bearing by an oil pressure transmitter mounted near the

engine. Each engine oil pressure indicator displays a

The engine instruments are the gas producer tachome-

pressure range from 0 to 200 psi. Power to operate the

ter, the dual torquemeter, power turbine inlet tempera-

engine oil pressure circuit is supplied by the AC instru-

ture (PTIT), fuel flow, oil pressure and oil temperature

ment buses through the ENGINE NO. 1 and NO. 2 OIL

indicators. The caution capsules are the NO. 1 and NO.

PRESS circuit breakers on the No. 1 and No. 2 PDP.

2 ENGINE OIL LOW and the NO. 1 and NO. 2 ENG CHIP

DET.

2-3-22. Engine Oil Temperature Indicator.

Two engine oil temperature indicators are on the center

2-3-18. Gas Producer Tachometer.

the instrument panel (18, fig. 2-1-8 and 2-1-12). Each

Two gas producer tachometers (N1), one for each en-

engine oil temperature indicator is calibrated from -70_

gine, are on the center instrument panel (6, fig. 2-1-8 and

to + 150_C. A temperature probe within the lubrication

6, 2-1-12), above the PTIT indicators. Each tachometer

lines of the engine, before the fuel-oil cooler, is the point

displays gas producer turbine speed in percent of N1.

at which the temperature is sensed. Power to operate the

Each tachometer operates from power supplied by a gas

resistance-type oil temperature circuit is supplied by the

producer tachometer generator on the accessory gear

No. 1 and No. 2 DC buses through the ENGINE NO. 1

2-3-5

TM 1-1520-240-10

and NO. 2 OIL TEMP circuit breakers on the No. 1 and

detector is bridged by ferrous metal particles, which may

No. 2 PDP.

indicate impending engine or engine transmission fail-

ure, the corresponding NO. 1 or 2 ENG CHIP DET cau-

2-3-23. Engine Caution Capsules.

712 The follow-

tion capsule will illuminate. Also, the associated ENGINE

ing items are in reference to Fig. 2-14-5:

CHIP DETECTOR or ENGINE TRANSMISSION CHIP

DETECTOR magnetic indicator on the MAINTENANCE

a. NO. 1 (2) ENGINE OIL LOW. This is illuminated

PANEL (fig. 2-9-2) will latch. Refer to Chapter 9 for emer-

when approximately 2 quarts of usable oil is remaining in

gency procedures.

the engine oil tank.

2-3-26. Engine Chip Detector Fuzz Burn-Off.

b. NO. 1 (2) ENG CHIP DET. This is illuminated if

a detector is bridged by ferrous metal particles which

Helicopters equipped with the chip detector fuzz burn -off

may indicate impending engine or engine transmission

system in the engine are identified by a module labeled

failure.

PWR MDL CHIP BURN-OFF located below the MAIN-

TENANCE PANEL. The chip detector fuzz burn-off sys-

c. NO. 1 (2) ENG N1 CONT. This is illuminated

tem employs an automatically operated fuzz burn-off

when the ECL is not in the STOP, GROUND or FLIGHT

electrical circuit with the ability to eliminate nuisance chip

detent or when the ECL position does not agree with the

lights caused by minute ferrous metallic fuzz or ferrous

N1 actuator position.

metallic particles on the engine accessory gear box

(AGB) chip detectors. The response time of the fuzz

2-3-24. Engine CAUTION/ADVISORY Capsules.

burn-off circuit is more rapid than that of the helicopter

warning system; thus a successful fuzz burn-off will be

714A The following items are referenced in Fig. 2-14-6:

accomplished before any caution capsule on the master

a ENG 1 (2) FAIL. Active when the engine failure

caution panel illuminates. Should the particle or particles

logic in the DECU detects a failed engine condition. The

not burn-off, the NO. 1 or NO. 2 ENG CHIP DET caution

engine failure logic is active when N1 is greater than 60%

capsule will illuminate. Also, the corresponding ENGINE

and the ECL position is greater than 50_(within 10_ of

CHIP DETECTOR or ENGINE TRANSMISSION CHIP

DETECTOR magnetic indicator on the MAINTENANCE

FLT position). The engine failure logic in each DECU is

PANEL will latch. Power for the PWR MDL CHIP BURN-

used to recognize any of the following:

OFF is supplied by the No. 1 DC bus through the HY-

(1)

Power turbine shaft failure. N2 is greater

DRAULICS MAINT PNL circuit breaker on the No. 1 PDP.

than RRPM by more than 3 percent.

2-3-27. Engine Interstage Air Bleed.

(2)

N1 underspeed.

NOTE

(3)

Engine flameout.

Bleed band oscillations at low torque settings

(4)

Over temperature start abort (Primary mode

(approximately 30% torque per engine), indi-

only).

cated by fluctuating RRPM and torque, can

occur and are not cause for engine rejection.

(5)

During normal shutdown as the N1 goes be-

To aid compressor rotor acceleration and prevent com-

low 48 percent the ENG 1 (2) FAIL caution is illuminated

for 12 seconds, this is a BIT self system check.

pressor stall, an interstage air bleed system is provided

on each engine. A series of vent holes through the com-

b. FADEC 1 (2). Active if Primary FADEC System

pressor housing at the sixth stage vane area allows pres-

hard fails.

surized air to bleed from the compressor area. This en-

ables the compressor rotor to quickly attain preselected

c. REV 1 (2). Active if Reversionary FADEC system

RPM. The pneumatic interstage air bleed actuator con-

hard fails.

trols operation of the air bleed by tightening or loosening

d. ENG 1 (2) OIL LVL. Active when approximately 2

a metal band over the vent holes. Should the bleed band

quarts of usable oil is remaining in the engine oil tank.

malfunction and remain open, there would be a notice-

able loss in power. 712 The interstage air bleed system

e. ENG 1 (2) CHIP DETR. Active if a detector is

operates automatically when the ENG COND levers or

bridged by ferrous metal particles which may indicate

the engine beep trim switches are used to govern RPM.

impending engine or engine transmission failure.

714A

The interstage air bleed system operates auto-

ENG CONT PWR. Active when power turbine

matically through the FADEC system.

inlet temperature is in the contingency power range.

2-3-28. Engine Drain Valves.

2-3-25. Engine Chip Detectors.

Pressure-operated engine drain valves are in the bottom

The engine accessory section oil sump and engine trans-

of each engine combustion housing. The valves auto-

mission chip detectors are electrically connected to the

matically drain unburned fuel from the combustion cham-

corresponding NO. 1 or NO. 2 ENG CHIP DET caution

ber following an aborted start or whenever the engine is

capsule on the master caution panel (fig. 2-14-6). If a

shut down. One valve is at the forward end of the com-

2-3-6

TM 1-1520-240-10

bustion camber and the other is at the aft end to ensure

q. The Digital Electronic Control Unit (DECU) in-

complete drainage.

cludes a primary mode and a reversionary section for

backup (fig. 2-3-7).

2-3-29. FADEC Description.

The Hydromechanical Metering Assembly

714A Each engine is controlled by its own Full Authority

(HMA), includes Hydromechanical Fuel Metering Unit

Digital Electronic control System (FADEC) which pro-

(HMU) and fuel pump unit for all fuel metering to support

vides the following features:

both primary and reversionary fuel metering, a self-con-

tained alternator for powering the FADEC electronics, a

a. Automatic start scheduling.

primary and revisionary compressor bleed air control,

and redundant speed sensing.

1 and 2 engine load sharing.

c. Power turbine speed governing.

s. ENG COND panel (fig. 2-3-5).

d. Transient load anticipation (using rotor speed

FADEC control panel (fig. 2-3-6).

and collective pitch rates).

u. RPM INC/DEC (Beep) switch THRUST CONT

e. Transient torque smoothing (using N2 rates).

Lever.

Contingency power capability to meet aircraft de-

(1)

On 714A engine installations, engine beep

mand.

switches are only active when in reversionary mode.

g. Acceleration and deceleration control.

(2)

Each switch is labeled NO. 1 or NO. 2 which

is used to adjust RRPM when in reversionary mode.

h. Engine temperature limiting throughout the oper-

ating range.

(3)

Operation of the beep switches on the

714A in the reversionary mode are the same as for the

Surge avoidance.

712 except that each switch operates respective en-

Compressor bleed band scheduling..

gine independently. If only one engine is in reversionary

mode, the RRPM will not change, as it is governed by the

k. Fuel flow limiting.

engine in primary mode.

Engine fail detection.

2-3-30. Reversionary System. 714A

m. Power assurance test.

NOTE

n. Engine history/fault recording.

Aircrew should be alert to the possibility of

o. Engine-to-engine communication (via data

abrupt NR and engine power changes when

bus).

operating the FADEC in single or dual engine

REV mode (s).

p. Automatic switchover to reversionary backup in

the event of a FADEC primary system failure.

The reversionary (backup mode) automatically takes

control of the engine if the primary mode fails or if se-

The FADEC provides automatic engine start, simulta-

lected by the operator via the FADEC panel, REV switch.

neously sequencing ignition, start fuel, and stabilized

operation at idle. A data link between 1 and 2 engine

When an engine is operating in reversionary mode,

FADEC systems transmits signals to achieve load sha-

FADEC provides engine and rotor control through N1

ring. It also provides control of N1 speed and NR (N2)

speed governor, beep control, and thrust pitch

output shaft speed to maintain the rotor system at a near

compensator.

constant RRPM throughout all flight power demand con-

When both engines are in reversionary mode, RRPM will

ditions. FADEC provides smooth acceleration and over-

temperature protection when ECLs (both together) are

require more pilot attention since proportional rotor

moved from GROUND to FLIGHT. Overtemperature

speed governor will not hold speed as accurately as the

protection is provided (through the DECU temperature

primary systems. With large collective changes, the rotor

limiting function) by control system thermocouple inter-

speed can change up to $3 percent from a nominal

face at the power turbine inlet. The control system

setting.

compares PTIT temperature signals with reference limits

The reversionary system provides the following control

to calculate and provide appropriate N1 acceleration.

functions:

During starts, an absolute 816_C limit is set and if ex-

ceeded an engine out indication and shutdown will occur.

a. Automatic start sequencing including over tem-

If compressor performance deteriorates for any reason,

perature protection, but not start abort.

surge detection automatically allows recovery from com-

pressor instability while protecting the engine from dam-

b. Pilot controlled start fuel enrichment/derichment,

age due to overtemperature.

if required, through ECL modulation.

The FADEC system consists of:

c. Ground idle set 55 "5 percent with ECL at GND.

2-3-7

TM 1-1520-240-10

d. RRPM droop compensation based on thrust le-

2-3-31. ENG COND Panel. 714A

ver position.

CAUTION

e. Beep capability becomes active for load match to

When adjusting controls or switches on

other engine.

the overhead switch panel, make sure

gloves or sleeves do not catch and inad-

Full contingency power capability.

vertently move the ENG COND levers.

The ENG COND panel is located in the overhead switch

g. Over temperature protection throughout opera-

panel (fig. 2-3-5).

tion.

NO.1 and NO. 2 ENG COND Levers. The ENG COND

Levers (ECL) provide the pilot with proportional accelera-

h. Engine shutdown in response to ECL being

tion and deceleration authority. The ECLS are spring-

placed at STOP

loaded outboard creating a gated motion when ad-

vanced from the STOP to GND and to FLT positions.

ENG COND lever friction is provided to reduce the possi-

Tracking of primary mode during normal primary

bility of over-torque transmissions by resisting rapid

mode operation allowing a smooth switchover when se-

movement of the levers.

lected.

2-3-32. FADEC Panel. 714A

If in reversionary mode for any reason (training) and

The FADEC panel is located on the overhead switch

there is a reversionary failure, the FADEC will not auto-

panel (fig. 2-3-6). It comprises of the following:

matically switch back to the primary mode. The pilot must

a. NR % Switch. The NR% switch controls a rheo-

manually select PRI mode.

stat which allows the operator to select any RRPM be-

tween 97% and 103%. There are detents at 97%. 100%

and 103%. With the ECL(s) in FLT, NR will be maintained

at the selected speed. 100% is the normal position.

CAUTION

b. PRI/REV Switches. The primary mode is the nor-

mal mode of operation. The REV (reversionary) mode is

selected as a backup mode or is automatically selected

If both the primary and reversionary sys-

if the primary system has a hard fault failure. A hard fault

tem fail, the engine remains at the fuel flow

failure is defined as one in which normal primary system

being used at the time of the failure. When

performance might be jeopardized. Other failures are

a failed fixed fuel flow condition exists, the

classified as soft failures when the system is fault tolerant

ECL and the beep trim switch for the af-

and can continue fully operational with the fault signal

fected engine is inoperative, therefore

present.

there is no proportional control through

ECL except under some conditions STOP.

c. BU/PWR Switch. The Back-Up Power switch

Engine shutdown may be accomplished

when ON connects the aircraft battery relay and essen-

by moving the ECL from its present posi-

tial relay.

tion to STOP. Under these conditions, the

ENG 1 (2) FAIL and/or FADEC 1 (2) cau-

tions are illuminated.

NOTE

If FADEC and/or REV and ENG FAIL lights

are illuminated (which signifies a fuel flow

fixed condition), toggling the PRI/REV switch

reboots the microprocessor in the DECU. If it

is a spurious fault, the lights extinguish except

the REV light. If the hard fault is real, the REV

and ENG FAIL lights illuminate again but not

the FADEC light.

When taking off with one engine in reversionary mode

the procedure is, before lift-off, the engine still in PRI

mode is used to set the correct rotor speed via the

FADEC NR% switch. The operator then uses the beep

switch of the engine in reversionary mode to match en-

gine torque.

Figure 2-3-5. Engine Condition Panel 714A

2-3-8

TM 1-1520-240-10

This provides the FADEC system with back up electrical

2-3-34. BIT. 714A

power in the event of a HMU integral alternator failure

The DECU contains a two-digit BIT display. When active,

thus preventing loss of the PRI mode. Placing the B/U

the display indicates the operating status of the FADEC

PWR switch to OFF will reduce operating time on the

system and power assurance test results. A complete list

FADEC circuitry. The B/U PWR switch should always be

of the FADEC BIT fault codes are located at Table 2-3-1.

ON during engine operation.

The fault monitoring carried out by the DECU consists of:

d. OSPD 1, 2 Switch. The Over Speed test switch is

a. Power up tests.

a three position switch used to test the FADEC over-

b. Fault tests designed to discover dormant faults.

speed system. In the event of a NR overspeed of 114.8

c. A set of repeated monitoring tests to detect faults

percent, FADEC reduces fuel flow to a ground idle condi-

occurring during normal operation.

tion. The system remains activated until the overspeed

Fault information for the previous or current engine cycle

condition no longer exists, and will re-activate as soon as

can be seen on the DECU BIT display. The last engine

an overspeed re-occurs. The system contains provisions

cycle is reset on the first occurrence of start mode and not

to inhibit overspeed trip command if the other engine has

on engine shutdown. During engine shutdown (when

experienced a overspeed trip condition. To prevent inad-

ECL is at STOP or N1 is less than 10 percent) faults are

vertent operation during flight, this test is locked out if NR

not stored. Fault indications are stored in the DECU and

is greater than 81.3 percent. When performing an over-

are retained throughout the life of the control unit. How-

speed test with the engine running and the RRPM

ever, fault information prior to the previous cycle can only

79.0$1%, pressing the test switch to 1 or 2, lowers the

be accessed with specialized test equipment. During en-

overspeed trip threshold to 79.0$1% NR. At this time the

gine start the DECU BIT displays 88 for satisfactory test

system senses an overspeed and reduces the fuel flow.

or if the test fails, a fault code. Faults are classified as

e. LOAD SHARE, PTIT/TRQ Switch. The primary

either “HARD” or “SOFT”. In primary mode a hard fault

FADEC system provides pilot selectable engine torque

will cause the FADEC to transfer to Reversionary , while

PTIT matching to govern the engines. Torque matching

in Reversionary a hard fault will cause the FADEC to “fail

is normally the preferred option. The selected parameter

fixed” to a constant power condition. If a hard fault occurs

is constantly compared between the two engines until the

in Primary after a hard fault exists in reversionary then

RRPM stabilizes at datum figure. The PTIT option may

the primary will fail fixed. In the event of a soft fault the

be used when one engine is running hot. N1 matching is

FADEC will remain in the mode it was in prior to the fault

engaged automatically if the selected matching mode

but there may be some degradation or redundancy. All

fails.

soft faults are less severe than a hard fault since the

FADEC will not switch modes due to a soft fault.

ENG START Switch. It is a three position switch,

The activation of the BIT display is dependent upon the

spring loaded to the center position, labeled 1 and 2. It is

position of the ECL as follows:

used to commence the start sequence on the respective

engine.

a. With the ECL at STOP, the fault information for

the last engine cycle and current faults are displayed.

2-3-33. DECU Unit. 714A

b. When the ECL is positioned at GROUND only

current faults are displayed.

The two airframe mounted DECU’s, one for each engine,

contain the primary and reversionary mode electronics.

c. When the ECL is positioned at FLIGHT the dis-

The DECUs are located on the left (sta 390) and right (sta

play will be turned off except as required for Power Assur-

410) side of the aft cabin.

ance Test (PAT).

2-3-35. Starting in Primary Mode. 714A

CAUTION

The (P3) compressor pressure signal line

going to the DECU contains a manually

operated moisture drain valve. This valve

shall not be drained while the engine is

running.

NOTE

Engine may not start if REV fail caution is

illuminated.

In primary mode, engine start is initiated with the ECL in

the GND position. Select and hold the respective ENG

Figure 2-3-6. FADEC Panel 714A

START switch and allow the engine to accelerate to 10

2-3-9

TM 1-1520-240-10

primary mode except that the over temperature start

abort facility is not provided.

Alternate Reversionary Starting. For most conditions, a

start is successfully completed with the ECL held at the

GND position. However, if the engine fails to start due to

either a rich or lean hung start condition, the pilot may use

the ECL to increase or decrease the start flow as required

to complete a successful start.

a. Reversionary Rich Hung Start. A rich hung start

is characterized by N1 holding at about 40 percent and

PTIT climbing above 600_C. If a rich hung start is experi-

enced.

(1)

Set the affected engine ECL to STOP.

(2)

Allow PTIT to decay to 260_C or below (mo-

tor engine as required).

(3)

Check N1 0 percent.

(4)

Advance ECL half the distance between

STOP and GND (15_).

(5)

Motor engine using ENG START switch until

10% N1 and then release switch.

(6)

After engine ignition (PTIT rising), slowly ad-

Figure 2-3-7. DECU Panel 714A

vance ECL to GND. Check that N1 is stabilized at ground

idle.

percent N1 then release the ENG START switch. The

automatic start sequence has been energized and FA-

b. Reversionary Lean Hung Start. A lean hung start

DEC will complete the start. When an engine start is

is characterized by N1 hanging approximately 30 percent

energized, FADEC turns on the engine start solenoid to

and PTIT remaining below 500_C. If a lean hung start is

introduce fuel flow and energize the engine igniters. Suc-

experienced.

cessful engine ignition is immediately indicated to FA-

(1)

Slowly advance the hung engine ECL to

DEC by an increase of PTIT or compressor speed (N1).

achieve acceleration (maximum of one-third travel from

Engine temperature is monitored throughout the se-

GND to FLT).

quence and will result in a fuel flow reduction if the tem-

perature exceeds 650_C with a full cutback to minimum

(2)

Retard the ECL to GROUND as ground idle

flow limit at 760_C. The starter motor and igniter are

speed is approached. Check that N1 is stabilized at

automatically turned off when N1 speed exceeds 48 per-

ground idle.

cent. Ground idle is 50 to 59 percent and is corrected for

temperature. The engine is allowed to take 45 seconds

2-3-38. Starting Cycle, Aborting and Motoring.

to stabilize at ground idle.

714A

A starting cycle can be aborted at any time by moving the

2-3-36. Engine Start Abort. 714A

ECL to STOP.

PTIT above 816_C will cause immediate fuel shutoff to

2-3-39. Power Assurance Test Switch. 714A

below minimum fuel flow. The ECL must be retarded to

STOP to achieve total fuel shutoff. A start abort results in

The primary FADEC will perform a BIT whenever the

the ENG 1 (2) FAIL warning to be illuminated until the

PWR ASSURANCE TEST switch has been place to the

abort is reset by moving the ECL to STOP.

desired engine position. The switch is located below the

maintenance panel at station 524. The switch is label

2-3-37. Starting in Reversionary Mode. 714A

1/OFF/2 and spring loaded to OFF. The results of the test

are displayed in the DECU BIT window.

The initial start sequence in reversionary mode is the

same as in primary mode except, when N1 reaches 8

2-3-40. Engine Wash System. 714A

percent, the control system turns on the engine start fuel

solenoid to provide an initial altitude biased fuel flow and

The helicopter is equipped with an engine wash system

activates the igniters and latches the starter motor. The

for each engine. Air and water connections are externally

pilot can modulate fuel flow to the engine with ECL to

mounted inboard of each engine work platform. A series

start the engine at a desired acceleration rate. Tempera-

of spray nozzles are installed at the engine inlet and air

ture and temperature rate limiters are the same as in

lines are routed to the bleed band actuator.

2-3-10

TM 1-1520-240-10

Table 2-3-1. T55-GA-714A DECU BIT Fault Code List/Matrix

The DECU fault code matrix will be consulted when anything other than 88 is displayed on the DECU Hex display during

the DECU Pre-Start, Start, or Shutdown BIT checks. The matrix lists DECU fault codes, type of fault, aircraft operational

effect, pilot actions require, and overall mission impact.

Abort mission is defined as: perform a normal engine shutdown and make appropriate entries on the DA Form

2408-13-1.

Continue mission is defined as: the crew may complete the present day’s mission but maintenance troubleshooting

actions are required prior to the next mission.

DECU Fault

Code(s)

Fault

Operational Effect

Pilot Action

Mission Impact

DECU internal hard fault.

FADEC light

N/A

Abort Mission

Aircraft won’t start in Primary

DECU internal hard fault.

FADEC light

N/A

Abort Mission

Aircraft won’t start in Primary

DECU internal soft fault.

N/A

Make an entry on

Continue Mission

the DA Form

2408-13-1 even if

fault code clears.

DECU internal soft fault.

N/A

Make an entry on

Continue Mission

the DA Form

2408-13-1 even if

fault code clears.

DECU internal soft fault.

N/A

Continue Mission

DECU internal soft fault.

N/A

Continue Mission

DECU internal soft fault.

N/A

Make an entry on

Continue Mission

the DA Form

2408-13-1 even if

fault code clears.

DECU internal soft fault.

Without FADEC light

N/A

Continue Mission

DECU hard fault.

With FADEC light

N/A

Abort Mission

DECU internal soft fault.

N/A

Make an entry on

Continue Mission

the DA Form

2408-13-1 even if

fault code clears.

DECU internal soft fault.

Without FADEC light

N/A

Continue Mission

DECU hard fault.

With FADEC light

N/A

Abort Mission

DECU internal hard fault.

With FADEC light

N/A

Abort Mission

Unable to start in Primary

DECU internal hard fault.

FADEC light

N/A

Abort Mission

DECU TQ soft fault.

N/A

Select PTIT

Continue Mission

Check TQ AC/

DC CBs

Make an entry on

the DA Form

2408-13-1 even if

fault code clears.

2-3-11

TM 1-1520-240-10

Table 2-3-1. T55-GA-714A DECU BIT Fault Code List/Matrix (Continued)

DECU Fault

Code(s)

Fault

Operational Effect

Pilot Action

Mission Impact

DECU NR set soft fault.

N/A

Continue Mission

DECU thrust hard fault.

REV light

N/A

Abort Mission

Unable to start in Primary and

Reversionary

DECU NR soft fault.

N/A

Continue Mission

DECU ECL soft fault.

N/A

Abort Mission

DECU emergency bus

N/A

Check REV

If CB in, Continue

soft fault.

CONT CB

Mission

If CB out and

can’t reset, Abort

Mission

DECU airframe bus soft

N/A

Check PRI

If CB in, Continue

fault.

CONT CB

Mission

Back-up power

If CB out and

can’t reset, Abort

Mission

DECU internal N1B hard

With REV light,

N/A

Abort Mission

fault.

Unable to start in Primary or

Reversionary

DECU N1B soft fault.

Without REV light

N/A

Continue Mission

DECU N2B soft fault.

Overspeed system compro-

Make an entry on

Continue Mission

mised

the DA Form

2408-13-1 even if

fault code clears.

DECU PTIT soft fault.

N/A

Perform REV

Continue Mission

(DECU Pre-

mode selected

start BIT

start. Switch to

check)

PRI when ground

idle is achieved.

Make an entry on

the DA Form

2408-13-1 even if

fault code clears.

DECU PTIT soft fault.

N/A

Continue Mission

(All other

DECU BIT

checks)

DECU thrust soft fault.

Without REV light

N/A

Continue Mission

DECU thrust hard fault.

With REV light,

N/A

Abort Mission

Unable to start in Primary or

Reversionary

2-3-12

TM 1-1520-240-10

Table 2-3-1. T55-GA-714A DECU BIT Fault Code List/Matrix (Continued)

DECU Fault

Code(s)

Fault

Operational Effect

Pilot Action

Mission Impact

DECU ECL soft fault.

N/A

Abort Mission

DECU PLA hard/soft fault.

With or without REV light

N/A

Abort Mission

Unable to start in Primary or

Reversionary

DECU cold junction com-

N/A

Make an entry on

Continue Mission

pensation (CJCV) soft

the DA Form

fault.

2408-13-1 even if

fault code clears.

DECU internal soft fault.

N/A

Continue Mission

DECU internal soft fault.

N/A

Continue Mission

DECU 400 Hz hard/soft

With or without REV light

N/A

Abort Mission

fault.

DECU communication soft

N/A

Continue Mission

fault.

DECU communication soft

N/A

Continue Mission

fault.

DECU communication soft

N/A

Continue Mission

fault.

DECU communication soft

Unable to TQ share

Switch to PTIT

Continue Mission

fault.

sharing

DECU communication soft

N/A

Continue Mission

fault.

DECU communication soft

N/A

Continue Mission

fault.

DECU communication soft

N/A

Continue Mission

fault.

More than

DECU communication link

N/A

Abort Mission

two codes,

redundancy failure.

C1 through

DECU communication link

N/A

Abort Mission

failure.

DECU N1 from the other

Unable to N1 load share

N/A

Continue Mission

engine fails.

DECU load share hard

FADEC light

N/A

Abort Mission

fault.

Primary mode failed

DECU overspeed system

Overspeed system compro-

N/A

Continue Mission

soft fault.

mised

DECU P3 sensor soft

Loss of surge protection

N/A

Abort Mission

fault.

DECU P1 sensor hard

With FADEC light

N/A

Abort Mission

fault.

DECU P1 sensor soft

Without FADEC light

N/A

Continue Mission

fault.

DECU internal soft fault.

N/A

Continue Mission

2-3-13

OPERATOR’S MANUAL FOR ARMY CH-47D HELICOPTER (EIC: RCD, 2003) - page 1