TECHNICAL MANUAL MAINTENANCE MANUAL FOR ARMY CH-47D HELICOPTER (1992)

TM 55-1520-240-10

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

Depending on the pilot’s seat adjustment, it may not be

two-position shoulder harness inertia reel lock lever is on

possible to reach all switches with the inertia reel locked.

the left side of each seat (22, fig.

2-4). The lever

Each pilot should check and adjust the shoulder harness

positions are LOCKED (forward) and UNLOCKED (aft).

in locked position to determine whether all switches can

The lock may be moved freely from one position to the

be reached.

other. When the lock lever is in UNLOCKED position,

the reel harness cable is released to allow freedom of

2-22.Self-Tuning Dynamic Absorbers.

movement. However. the reel will automatically lock if a

horizontal impact force of 2 to 3 g is encountered. When

The helicopter is equipped with three self-tuning dy-

the reel is locked in this manner. it stays locked until the

namic absorbers. One absorber is in the nose compart-

lock lever is moved forward to LOCKED and then

ment and the other two absorbers are under each pilot’s

returned to UNLOCKED. When the lever is at LOCKED,

seat below the cockpit floor. All three absorbers serve to

the reel is manually locked so the pilot is restrained from,

maintain a minimum vibration level through the nor- mal

bending forward. When a crash landing or ditching is

operating rotor RPM range of the helicopter. The self-

anticipated and time permits, manual locking of the

tuning feature of the dynamic absorber functions as

shoulder harness inertia reel provides added safety

follows: each dynamic absorber consists of a tuning

beyond the automatic feature of the inertia reel.

mass

Change 7 2-10.1/(2-10.2 blank)

55-1520-240-10

2-22. Self-Tuning Dynamic Absorbers.

the helicopter and the spring-mounted mass. When the

measured vibration phases differ from a built-in phase

The helicopter is equipped with three self-tuning dynamic

relationship required to assure proper tune, the electronic

absorbers. One absorber is in the nose compartment and the

circuit extends or retracts the electrical actuator to reposi-

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

tion the counterweights which, in turn,. increases or de-

cockpit floor. All three absorbers serve to maintain a

creases the resonant frequency of the spring-mounted mass

minimum vibration level through the normal operating rotor

The dynamic absorbers are constantly being adjusted (tuned)

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

to minimize helicopter vibration. A self-test box is in the

the dynamic absorber functions as follows: each dynamic

heater compartment to provide maintenance personnel with

absorber consists of a tuning mass ‘suspended by springs,

an integral testing capability for the self-tuning feature of

and electronic measuring circuit, accelerometers, counter-

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

weights, an electrical actuator and a self-test box. The

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

accelerometers sense and compare the vibration phases of

breakers on the No. 2 PDP.

NOTE:

(ON HELlCOPTERS MODIFIED BY ANVIS HUD

MWO

1-1520-240-50-56)

A64351

Figure

2-7 Canted Console (Typical)

Change 13

2-11

55-1520-240-10

SECTION II EMERGENCY EQUIPMENT

2-23. Emergency Procedures.

2-25. FIRE PULL Handles.

Refer to Chapter 9 for all emergency procedures.

WARNING

2-24. Engine Compartment Fire Extinguisher

Before flying the aircraft ensure that each

System.

FIRE PULL handle NVG filter holder can be

rotated from the closed to the open position

The engine compartment fire extinguisher system (fig. 2-14)

without causing the FIRE PULL handle to be

enables the pilot or copilot to extinguish a fire in either engine

pulled. Improper handling of the NVG filter

compartment only. It is not designed to extinguish internal

holder may cause the FIRE PULL handle to be

engine tires. The system consists of two FIRE PULL handles,

pulled unintentionally, thus fuel to the affected

an AGENT DISCH (agent discharge) switch, a FIRE DETR

engine will be shut off and the engine will shut

(fire detector) switch on the center instrument panel. and two

down. Do not use sudden or excessive force

extinguisher agent containers on the overhead structure at

when rotating the FIRE PULL handle NVG

stations 482 and 502. The containers are partially filled with

filter holder from the closed to the open posi-

or BR₃) and

tion.

pressurized with nitrogen (table

2-1 provides the range of

engine fire extinguisher pressures.) The agent in one or both

of the containers can be discharTwo control handles for

the engine

compartment. Selection of the co(fig. 2-14) are labeled FIRE PULL-FU

pulling the appropriate FIRE PULL top center section of the

center ins

ENG 1 FIRE PULL handle has been puhas a cover for the NVG filter, two

container is made by placing the necessary control switches that clos

the appropriate position. In figuvalve and ar

m the fire extinguisher

selected.

supplied for each FIRE PULL handle f

1 and No.

2 DC essential buses th

ENGINE NO. 1 and NO. 2 FUEL SHUTOFF

on the No. 1 and No. 2 PDP. Power is

of warning lights from the correspon

bus through the ENGINE NO. 1 and NO.

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

Table 2-1.

Engine Compartment Fire Extinguisher Pressures

AMBIENT TEMPERATURE

MINIMUM INDICATION

MAXIMUM INDICATION

(C)

(PSI)

271

344

275

350

292

370

320

400

355

437

396

486

449

540

15^o

518

618

593

702

691

784

785

902

Change

2-17

TM 55-1520-240-10

a fire occur in the other engine compartment. Power is

supplied from the corresponding No. 1 or No. 2 DC

essential bus through the ENGINE NO. 1 and NO.

If the FIRE PULL handle warning lights are

FIRE EXT circuit breakers on the No. 1 and No.

covered by the NVG filters during daylight

PDP.

operation, illumination of the fire warning

lights may not be apparent in the event of an

2-27. FIRE DETR Switch.

engine fire. Do not operate the aircraft with

the NVG filters covering or obscuring the fire

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

warning lights unless night vision goggles

the AGENT DISCH switch on the top center section of

are being used.

the center instrument panel (fig. 2-14). It is labeled

FIRE DETR and TEST. The toggle switch is spring-

loaded to FIRE DETR which monitors the engine fire

The NVG filter is attached to one end the FIRE PULL

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

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

checks the operation of the engine fire detection system

holder forms a tab by which the filter holder and filter

by closing relays in both controls units and the warning

may be rotated about the hinged fitting. For NVG

lights in both FIRE PULL handles illuminate. Power to

operations, the filter holder is rotated to a closed

operate the test circuit is supplied by the DC essential

position over the front of the FIRE PULL handle cover.

bus through the LIGHTING CAUTION PNL circuit

In this position, the fire warning light is NVG compat-

breaker on the No. 1 PDP.

ible. For normal operations, the filter holder is rotated

from the closed position to the fully open position. In

2-28. Hand Fire Extinguishers.

this position, the FIRE PULL handle warning lights will

be red.

Avoid prolonged exposure (5 minutes or

more) to high concentrations of fire extin-

If there is a tire in both engine compart-

guishing agent and its decomposition prod-

ments, do not pull both FIRE PULL handles

ucts because of irritation to the eyes and

simultaneously. Extinguish fire in one com-

nose. Adequate respiratory and eye relief

partment only as described below. Leave the

from excessive exposure should be sought as

FIRE PULL handle out after fire has been

soon as the primary fire emergency permits.

extinguished. Proceed in a like manner to

Use of oxygen for personnel is recommended.

extinguish fire in the other engine compart-

ment.

Three portable

6.3 pound capacity hand fire extinguish-

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

When an engine compartment fire occurs on either side,

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

the respective pair of warning lights comes on. The

extinguishers are in the cabin section. One on the

appropriate FIRE PULL handle is pulled, that engine

forward bulkhead and one in the left rear, just forward

fuel shutoff valve closes and the AGENT DISCH switch

of the ramp.

is armed.

2-29. Emergency Troop Alarm and Jump Lights.

Selection and discharge of either fire bottle is accom-

plished by placing the AGENT DISCH switch to BTL 1

Two emergency troop alarm and jump light boxes are in

or BTL 2. After depletion of the charge in the initially

the cargo compartment. The forward box is on the

selected bottle, the remaining bottle can be discharged

bulkhead above the avionics equipment shelves and the

to the same engine compartment by selecting the oppo-

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

site position on the AGENT DISCH switch. The other

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

FIRE PULL handle performs the same function for its

with a red light future on one side and a green light

respective engine compartment.

fixture on the other side. The TROOP WARN panel on

the overhead switch console is used to operate the

2-26. AGENT DISCH Switch.

emergency troop alarm and jump lights.

A three-position AGENT DISCH (discharge) switch is

The emergency troop alarm and jump lights have sev-

above the FIRE PULL handles on the center instru-

eral functions. They can be used to notify passengers

ment panel (fig. 2-14). The lever-lock momentary switch

and crew with predetermined signals in time of emer-

positions are BTL 1, neutral, and BTL 2. When BTL 1

gency. The jump lights can be used to notify flight

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

engineer during airborne delivery operations and to

into the selected engine compartment. When BTL 2 is

alert the troop commander during paratroop drop mis-

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

sions. Refer to Chapter 9 for standard use of the troop

into the selected engine compartment. Only two fire

alarm.

extinguisher agent bottles are provided. If the agent

from both bottles is used in combating a fire in one

2-30. TROOP WARN Panel. The TROOP WARN

engine compartment, agent will not be available should

(warning) panel is located on the overhead switch panel

2-19

TM 55-1520-240-10

(fig. 2-15). It has two troop jump lights labeled RED

and GREEN. Also, two switches labeled JUMP LT and

ALARM. Power to operate and control the emergency

troop alarm and jump lights is supplied 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

selected. One light is provided for each color selection

and 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

Figure 2-15. Troop Warning Panel (Typical)

overhead switch panel.

2-31. First Aid Kits.

b. JUMP LT switch.

The three-position JUMP LT

Seven aeronautic first aid kits are installed in the

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

helicopter. One kit is in the passageway between the

switch is set to GREEN, the green lights on the

cockpit and cabin. The other six kits are in the cabin

emergency troop and jump light box, at both stations,

fuselage section, three on each side.

and the troop jump lights on the overhead switch panel

2-32. Emergency Entrances and Exits.

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

Refer to Chapter 9 for information on emergency

come on. OFF position turns off both sets of lights.

entrances and exits.

c. ALARM switch.

The two-position ALARM switch

2-33. Emergency Escape Axe.

is labeled OFF and ON. Moving the ALARM switch to

An emergency escape axe is provided. It is located on

ON rings the bell continuously at both stations until the

the right side of the cargo compartment slightly forward

switch is moved to OFF.

of station 200.

2-20

55-1520-240-10

SECTION Ill ENGINES AND RELATED SYSTEMS

2-34. Engines.

response to any setting of the engine controls selected by the

pilot. Engine gas producer rotor speed (Nl) and power

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

turbine speed (N2) are controlled by the fuel control unit,

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

which varies the amount of fuel delivered to the engine fuel

nacelles mounted externally on each side of the aft pylon.

nozzles. During normal operation, the fuel control unit

The engines have the capability to produce emergency

automatically controls fuel flow metering during power

power on pilot demand. See Performance Charts in Chapter

changes, thus protecting the engine from overspeed and

or Chapter 7A

overtemp. Fuel flow is automatically monitored to compen-

sate for changes in outside air temperature and compressor

2-35. General.

Each engine has a gas producer section

discharge pressure.

and a power turbine section. The gas producer supplies hot

gases to drive the power turbine. It also mechanically drives

2-39. Engine Fuel Control Units.

the engine accessory gearbox. The power turbine shaft

extends coaxially through the gas producer rotor and rotates

Each engine fuel control unit contains a single element fuel

independently of it. The gas producer section and the power

pump, a gas producer speed governor, a power turbine speed

turbine section are connected by only the hot gases which

governor, an acceleration-deceleration control, a fuel flow

pass from one section to the other.

limiter, a fuel control fuel shutoff valve, and a main

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

During engine starting, air enters the engine inlet and is

turbine (N2) lever are mounted on the fuel control unit.

compressed as it passes through seven axial stages and one

centrifugal stage of the compressor rotor. The compressed

Output power of the power turbine (a function of the speed

air passes through a diffuser. Some of the air enters the

and torque) is restricted by limiting the maximum fuel flow

combustion chamber where it is mixed with start fuel.

to the gas producer. Maximum gas producer rotor speed is

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

set by the ENG COND (engine condition) levers in the

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

cockpit. The ENG COND levers electromechanically posi-

continues to operate on metered fuel supplied to the fuel

tions the gas producer lever, which controls the fuel control

nozzles.

fuel shutoff valve and the operating level of the gas

producer. During flight, the ENG COND levers are left at

Hot expanding gases leave the combustion chamber and

FLT and the output shaft speed is regulated by the power

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

turbine speed (N2) governor.

combustion gases also drives the two-stage power turbine,

which drives the power turbine shaft to the engine trans-

The power turbine lever is electromechanically positioned

mission. The engine lubrication system has an integral oil

by the ENGINE BEEP TRIM switches, thrust control and

tank which is inside the air inlet housing and is serviced

EMERG ENG TRIM (emergency engine trim)

with approximately 12 quarts. (Refer to table 2-3.)

switches. Output shaft torques are limited by the fuel flow

limiter, which limits the maximum fuel flow. The position

2-36. Engine Inlet Screens.

of the main metering valve is determined by the gas

An engine inlet screen which minimizes foreign object

producer speed governor, power turbine speed governor, the

damage (FOD) is installed on each engine. The reduction in

acceleration-deceleration control, or the fuel flow limiter,

engine power available with screens installed is negligible.

depending on engine requirements at that time. The gover-

The engine inlet screens have bypass panels. These two

nor or the control unit demanding the least fuel flow

panels are on the aft end of each screen. Refer to Chapter 5

overrides the other in regulating the metering valve.

for information on use of bypass panels. Helicopters with

2-40. Speed Governing.

engine air particle separator (EAPS) installed, refer to TM

55-1520-240-10 EAPS SUPPLEMENT.

The power turbine speed governor senses the speed of the

power turbine and regulates the amount of fuel which is

2-37. Engine Anti-Icing.

supplied to the gas producer. This slows down or speeds up

The engine air inlet fairing and engine drive shaft fairing

the gas producer rotor so that power turbine and rotor

receive anti-icing protection from the thermal radiation

system speed remains nearly constant as loads vary.

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

At minimum rotor blade pitch, the amount of power

oil in the oil cavity of the inlet housing warms the air as it

required is at minimum. As pitch is increased, power turbine

passes into the engine inlet.

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

from the engine to maintain a constant rotor speed. The

2-38. Engine Power Control System.

power turbine speed governor senses the decrease of N2

Each engine is controlled by a separate power control

RPM and increases the flow of fuel to the gas producer.

system which includes cockpit controls and an engine fuel

Decreasing pitch causes N2 to increase. The power turbine

control unit. Each system provides automatic control of

governor senses the increase and reduces the flow of fuel to

engine gas producer rotor speed and power turbine speed in

the gas producer, thus decreasing the engine output power.

Change 13

2-21

TM 55-1520-240-10

The power turbine speed governor allows the power turbine

CAUTION

output speed to decrease (droop) approximately 10 percent

when the power loading varies from minimum to full load.

When the ENG COND lever is placed to

This is minimized by a droop eliminator linked to the thrust

GND during start sequence, the N1 actuator

control rod. The droop eliminator automatically changes the

could inadvertently go beyond the ground

power turbine lever to compensate for droop as pitch is

position. The respective ENG N1 COND

increased or decreased. Another type of droop, which is

caution capsule will illuminate. However, ig-

only transient, occurs as a result of the time required for the

nition will still occur if the start switch is

engine to respond to changing loads due to system lag.

moved to START, thus resulting in a possible

engine runaway.

2-41. ENG COND Levers

CAUTION

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

engine, are on the ENG COND panel (fig. 2-16) of the

When adjusting controls or switches on the

overhead switch panel. Each lever has three positions

overhead switch panel, make sure gloves or

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

sleeves do not catch and inadvertently move

appropriate fuel flow rates for GND, FLT, and STOP

the ENG COND levers.

(engine shutdown). Power is supplied by the DC essential

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

The ENG COND lever must be at GND before the engine

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

will start. When an ENG COND lever is advanced from

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

Each ENG COND lever is spring-loaded outboard and is

chanical actuator which establishes an appropriate fuel flow

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

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

ally control acceleration of the gas producer from STOP to

lever at GND should be 60 to 63 percent N1. When an ENG

FLT. Two engine control caution capsules are on the master

COND lever is moved to FLT, the engine is operating within

caution panel (fig. 2-51). They are labeled NO. 1 ENG N1

the N2 governing range, unless the engine is “topped out”

CONT and NO. 2 ENG N1 CONT. The capsules normally

at which time it goes back to N1 governing. The N2

illuminate when the ENG COND levers or the N1 actuators

governor then takes control to maintain selected rotor RPM

are at an intermediate position between STOP, GND, or

(RRPM) in response to the engine beep trim switches and

FLT. They extinguish when the ENG COND lever and N1

actuator positions agree. However, they remain illuminated

collective pitch changes, When an ENG COND lever is

if a component of the system (actuator, control box, or

moved to STOP, the gas producer lever closes the fuel

condition panel) has failed in other than a detent position.

control fuel shutoff valve which stops fuel flow to the gas

Power is supplied by the DC essential bus through the

producer.

LIGHTING CAUTION PNL circuit breaker on the No. 1

Each electrical system is completely separate and a failure

PDP.

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

brake holds the actuator at its last selected position if loss of

electrical power occurs.

ENG COND lever friction is

provided to reduce the possibility of overtorquing 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.

2-42. Normal Engine Beep Trim Switches.

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. 2-26). Both switches have an

RPM INCREASE, RPM DECREASE, and a neutral posi-

tion.

One switch is labeled NO. 1 & 2 which is

normally 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

910000-A6010

dual torquemeters.

Power to operate the beep trim system is supplied by

the DC and AC buses. DC power to operate a trim motor in

Figure 2-16. Engine Condition Panel

2-22

Change 13

TM 55-1520-240-10

the power turbine control box, which unbalances a control

AUTO (cover down), the normal beep trim system is

circuit, is supplied by the corresponding No. 1 or No. 2 DC

functional (115-volt AC from the AC bus is reconnected to

buses through the ENGINE NO. 1 or NO. 2 TRIM circuit

the associated engine power turbine control box). Refer to

breakers on the No. 1 or No. 2 PDP. The unbalanced control

Chapter 9 for emergency engine trim operation.

circuit causes the AC power from the No. 1 or No. 2 AC

buses through the ENGINE NO. 1 or NO. 2 TRIM &

CAUTION

TIMER circuit breakers on the No. 1 or No. 2 PDP to be

transformed and rectified to DC voltage. This DC power

Engine response is much faster when RRPM

operates the power turbine actuator on the engine fuel

is controlled with emergency engine beep

control.

trim system. It is possible to beep the rotor

speed below safe operating speed and low

NOTE

enough to disconnect the generators from the

No two engines provide matched performance

buses. The generators are disconnected at

with regard to torque, RPM, PTIT, or fuel flow.

85% to 82% RRPM after a 3 to 7 second time

With torque matched all other parameters may

delay.

not be matched.

b. Emergency Engine Trim Switches.

Each momentary

Holding the NO. 1 & 2 switch forward (RPM

switch is used to change the power turbine speed of its

INCREASE) will increase the RRPM. Holding the switch

respective engine if the power turbine control box (normal

aft (RPM DECREASE) will decrease the RRPM. When the

beep trim system) malfunctions.

switch is released, it returns to the center or neutral position.

The switch electrically controls both power turbines by

When the normal trim system fails, the droop eliminator

movement of the N2 actuator through each engine power

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

turbine control box.

and a spring-loaded center position. When one of the

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

The procedure for matching engine load requires that NO. 1

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

goes directly to the respective power turbine actuator and

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

increases the lever setting and the power turbine speed.

moved forward (RPM INCREASE), the torque of No. 1

When the switch is held at DECR, the lever setting is

engine increases. At the same time RRPM increases, even

decreased, and the power turbine speed is decreased.

though No. 2 engine torque decreases slightly. Moving NO.

The emergency engine trim switches are to be used when

1 & 2 engine beep trim switch aft (RPM DECREASE)

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

causes both engine torques to decrease and reduce RRPM.

switches is used while the respective power turbine control

If torques are still not matched, this procedure is continued

box is functioning normally, the power turbine actuator

until torques are matched and desired RRPM is attained.

The opposite action occurs when NO. 1 engine beep switch

setting will temporarily change but will return to its original

is moved aft.

setting when the switch is released. Power to operate the

emergency engine beep trim switches and actuators is

The engine beep trim switches should not be used during

supplied by the essential DC bus through the NO. 1 and NO.

power changes initiated by thrust lever movement because

2 EMERG ENG TRIM circuit breakers on the No. 1 and

RRPM droop should only be momentary. The engine beep

No. 2 PDP.

tim system adjusts engine RPM only if the respective ENG

COND lever is at FLT. At, STOP or GND, it is possible to

move the power turbine lever by moving the engine beep

trim switches to RPM DECREASE or RPM INCREASE,

but in either case, engine RPM will not be affected because

the engine is not operating in the N2 governing range.

2-43. EMERG ENG TRIM Panel.

The EMERG ENG TRIM (emergency engine) panel is

located on the center console (fig. 2-17). The panel consists

of two guarded normal engine trim system disable switches

and two momentary emergency engine trim switches.

a. Normal Engine Trim System Disable Switches.

The

guarded switches permit the pilot to disable either or both

normal beep trim systems. This prevents unwanted signals

from the normal beep trim system to interfere with the

A61206

operation of the emergency engine trim system. Each switch

is labeled AUTO and MANUAL. When either switch is at

MANUAL, the respective normal beep trim system is

Figure 2-17.

Emergency Engine Trim Panel

disabled (115-volt AC from the AC bus to the engine power

turbine control box is interrupted). When the switch is at

Change 13

2-23

TM 55-1520-240-10

2-44. Emergency Power System.

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 engines.

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

Figure 2-18. Emergency Power Panel

engines. With the emergency power system, increased

power is available on pilot demand and is actuated by

tank capacity. If the oil level decreases to about 2 quarts

raising the thrust control into the emergency power range.

usable, the corresponding ENG OIL LOW caution capsule

Refer to Chapter 5 for limitations on its use.

will illuminate.

2-47. Engine Start System.

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

910°C, the EMERG PWR lights will illuminate on the

The engine start system includes the hydraulic starters on

copilot and pilot instrument console (17, fig. 2-8 and 18, fig

each engine, the engine start valves and the solenoid-

2-10) If temperature is maintained in this range for more

operated pilot valves on the utility system pressure control

than 5 seconds, the associated indicator will apply 28-volt

modules, the START switch, and the start fuel solenoids and

DC from the ENGINE NO. 1 and /or NO. 2 START &

ignition exciters on the engines.

TEMP circuit breaker to the EMERGENCY POWER panel.

With 28-volt DC applied to the panel, the applicable

When the start switch is moved to MTR, the respective

emergency power timer will start, and the indicator will

engine STARTER ON indicator light illuminates and the

display a black-and-white flag. When thrust is reduced

start valve opens (fig. 2-19). The start valve applies utility

below the emergency power level, the emergency power

system pressure from the APU to the engine starter; rotating

light will extinguish and the timer will stop. However, the

the engine starter and compressor. At 15 percent N1, the

emergency power indicator will continue to display the

ENG COND lever is moved to GND. The start switch is

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

immediately moved to START, energizing the ignition

only.

exciter. Start fuel is sprayed into the combustor and com-

bustion begins. Before PTIT reaches 200°C, the START

WARNING

switch is manually released to MTR. At MTR, the start fuel

valve is closed and the ignition exciter is deenergized.

Before flight, be sure the two topping stops

are in their stowed position on the right side

The engine then accelerates to ground idle speed. At 50

of the console. If the stops are not stowed, be

percent N1, the START switch is manually moved to the

sure the stops are not installed on the fuel

locked OFF position. At OFF, the pilot valve closes, closing

controls before you start the engine. Failure

the start valve and deenergizing the STARTER ON indicator

to check may result in inability to achieve

light. A relay in each engine start circuit is energized when

emergency power in an emergency.

either START switch is at MTR or START. The relay, when

energized, disables the start circuit of the opposite engine,

Topping stops are stowed on each helicopter. The stops are

thus preventing simultaneous dual engine starts. Power is

installed on the N1 control of each engine for maintenance

supplied by the No. 1 and No.

2 DC essential buses through

engine topping checks. The stops provide an established

the ENGINE NO. 1 and NO. 2 START & TEMP and IGN

fuel flow when topping. When not in use, the stops are

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

stowed on the right side of the center console aft of the

pedals.

2-45. EMERGENCY POWER Panel.

The EMERGENCY POWER panel is located on the over-

head switch panel (fig. 2-18). It consists of an emergency

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.

2-46. 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 2-3 for the

Figure 2-19.

Engine Start Panel

2-24

Change 13

TM 55-1520-240-10

2-48. START Panel.

712 The START panel is located

power supply unit

714A and RDPS is provided by the No.

on the overhead switch panel (fig. 2-19). It consists of the

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

ENG 1 and ENG 2 STARTER ON indicator lights and two

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

start switches.

PDP.

a. Start Switches.

The switches are labeled OFF, MTR,

2-53. Power Turbine Inlet Temperature Indicators.

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

Two power turbine inlet temperature (PTIT) indicators, one

spring-loaded from START to MTR. At MTR, the engine is

for each engine, are on the center instrument panel (7, fig.

rotated by the starter, but ignition and start fuel circuits are

2-9). Each indicator is calibrated from

0° to 1,200°C. The

deenergized. At START, the engine is rotated with start fuel

temperatures registered on the PTIT indicator are transmit-

and the ignition circuits are energized . MTR is selected

ted by chromel-alumel thermocouples. The thermocouples

during starting, in case of engine fire or to clear the

sense gas temperature at the power turbine inlet and

combustion chamber.

transmit an average gas temperature reading to the PTIT

b. STARTER ON Indicator Lights.

The STARTER ON

indicator in the cockpit.

712 When power turbine inlet

indicator lights will illuminate when the associated START

temperature increases to the emergency power range, the

switch is moved to MTR or START. The light alerts the

EMERG PWR indicator light will illuminate and DC ower

pilots when the START switch is inadvertently left at MTR.

is supplied to the EMERGENCY POWER panel.

714A

Power is supplied by the No. 1 and No. 2 DC essential buses

When power turbine inlet temperature increases to the

through the ENGINE NO. 1 and NO. 2 START & TEMP

contingency power range, the ENG CONT PWR master

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

caution advisory panel capsule will illuminate.

2-49. Ignition Lock Switch.

2-54. Engine Oil Pressure Indicator.

An engine oil

pressure indicator on the center instrument panel is provided

An ignition system lock switch (11, fig. 2-4) is installed on

for each engine (17, fig.

2-9). Each indicator relates pressure

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

sensed at No. 2 bearing by an oil pressure transmitter

key-operated switch prevents unauthorized use of the heli-

mounted near the engine. Each engine oil pressure indicator

copter. When the switch is off, the circuits of the ignition

displays a pressure range from

0 to 200 psi. Power to

exciters and the start fuel solenoids of both engines are

operate the engine oil pressure circuit is supplied by the AC

open. Therefore, the engines cannot be started. Be sure both

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

START switches are OFF before turning the ignition lock

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

switch ON or OFF.

2-55. Engine Oil Temperature Indicator. T

en-

2-50. Engine Instruments and Cautions.

gine oil temperature indicators are on the center the instru-

ment panel (18, fig. 2-9). Each engine oil temperature

The engine instruments are the gas producer tachometer, the

indicator is calibrated from

-70° to +150°C. A temperature

dual torquemeter, power turbine inlet temperature (PTIT),

probe within the lubrication lines of the engine, before the

fuel flow, oil pressure arid oil temperature indicators. The

fuel-oil cooler, is the point at which the temperature is

caution capsules are the NO. 1 and NO. 2 ENGINE OIL

sensed. Power to operate the resistance-type oil temperature

LOW and the NO. 1 and NO. 2 ENG CHIP DET.

circuit is supplied by No. 1 and No. 2 DC buses through the

ENGINE NO. 1 and NO. 2 OIL TEMP circuit breakers on

2-51. Gas Producer Tachometer.

Two gas producer

the No. 1 and No. 2 PDP.

tachometers (Nl), one for each engine, are on the center

instrument panel, above the PTIT indicators. Each tachom-

2.56. Engine Caution Capsules.

712 The following

eter displays gas producer turbine speed in percent of N1.

items are in reference to Fig. 2-51:

Each tachometer operates from power supplied by a gas

producer tachometer generator on the accessory gear box

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

section of each engine.

712 The outer scale of the tachom-

when approximately 2 quarts of usable oil is remaining in

eter is calibrated from

0 to 100 in increments of two. The

the engine oil tank.

smaller, vernier scale is calibrated from

0 to 10, in incre-

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

ments of one.

714A The tachometer is calibrated from

0 to

detector is bridged by metal particles which may indicate

110.

impending engine or engine transmission failure.

2-52. Torquemeter.

One torquemeter is on the copilot

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

instrument panel and the other on the pilot instrument panel

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

(1, fig. 2-8 and 17, fig. 2-10). Each torquemeter has two

when the ECL position does not agree with the N1 actuator

pointers, one for each engine, labeled 1 and 2. Each

position.

torquemeter has a range of

0 to 150 percent. The system

consists of a power output shaft, torquemeter head assem-

2-56.1. Engine

CAUTION/ADVISORY Capsules.

bly, power supply unit,

714A ratio detector power supply

714A The following items are in reference to Fig. 2-51.1:

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

operate the torquemeters is provided by No. 1 and No. 2 AC

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

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

in the DECU detects a failed engine condition. The engine

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

failure logic is active when N1 is greater than 60% and the

Change 13

2-25

TM 55-1520-240-10

ECL position is greater than 50°. The engine failure logic in

2-59. Engine Interstage Air Bleed.

each DECU is used to recognize any of following:

NOTE

(1) Power turbine shaft failure. N2 is greater than

Bleed band oscillations at low torque settings

RRPM by more than 3 percent.

(approximately

30%

torque per engine), indi-

(2) N1 underspeed. N1 speed drops below 48 per-

cated by fluctuating engine RPM and torque,

cent.

can occur and are not cause for engine rejection.

(3) Engine flameout.

To aid compressor rotor acceleration and prevent compres-

(4) Over temperature start abort (Primary mode

sor stall, an interstage air bleed system is provided on each

only).

engine. A series of vent holes through the compressor

housing at the sixth stage vane area allows pressurized air to

(5) During normal shutdown as the N1 goes below

bleed from the compressor area. This enables the compres-

48 percent the ENG 1 (2) FAIL caution is illuminated for 12

sor rotor to quickly attain a preselected RPM. The pneu-

seconds, this is a BIT self system check.

matic interstage air bleed actuator controls operation of the

b. FADEC 1 (2). Active if Primary FADEC system hard

air bleed by tightening or loosening a metal band over the

fails.

vent holes. Should the bleed band malfunction and remain

open, there would be a noticeable loss in power.

712 The

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

interstage air bleed system operates automatically when the

hard fails.

ENG COND levers or the engine beep trim switches are

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

used to govern RPM.

714A

The interstage air bleed

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

system operates automatically through the FADEC system.

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

2-60. Engine Drain Valves.

by metal particles which may indicate impending engine or

engine transmission failure.

Pressure-operated engine drain valves are in the bottom of

each engine combustion housing. The valves automatically

f. ENG CONT PWR. Active when power turbine inlet

dram unburned fuel from the combustion chamber follow-

temperature is in the contingency power range.

ing an aborted start or whenever the engine is shut down.

2-57. Engine Chip Detectors.

The engine accessory

One valve is at the forward end of the combustion chamber

section oil sump and engine transmission chip detector is

and the other is at the aft end to ensure complete drainage.

electrically connected to the corresponding NO. 1 or NO. 2

ENG CHIP DET caution capsule on the master caution

2-60.1. FADEC Description.

714A

Each engine is

panel (fig 2-51). If a detector is bridged by metal particles

controlled by its own Full Authority Digital Electronic

which may indicate impending engine or engine transmis-

Control system (FADEC) which provides the following

sion failure, the corresponding NO. 1 or 2 ENG CHIP DET

features:

caution capsule will illuminate. Also, the associated EN-

a. Automatic start scheduling.

GINE CHIP DETECTOR or TRANSMISSION CHIP DE-

TECTOR magnetic indicator on the MAINTENANCE

b. 1 and 2 engine load sharing.

PANEL (fig. 2-34) will latch. Refer to Chapter 9 for

c. Power turbine speed governing.

emergency procedures.

d. Transient load anticipation (using rotor speed and

2-58. Engine Chip Detector Fuzz Burn-Off.

collective pitch rates).

Helicopters equipped with the chip detector fuzz bum-off

e. Transient torque smoothing (using N2 rates)

system in the engine are identified by a module labeled

f. Contingency power capability to meet aircraft de-

PWR MDL CHIP BURN-OFF located below the MAIN-

mands.

TENANCE PANEL. The chip detector fuzz bum-off system

employs an automatically operated fuzz burn-off electrical

g. Acceleration and deceleration control.

circuit with the ability to eliminate nuisance chip lights

h. Engine temperature limiting throughout the operating

caused by minute ferrous metallic fuzz or ferrous metallic

range.

particles on the engine accessory gear box (AGB) chip

detectors. The response time of the fuzz bum-off circuit is

i. Surge avoidance.

more rapid than that of the helicopter warning system; thus

j. Compressor bleed band scheduling.

a successful fuzz bum-off will be accomplished before any

k. Fuel flow limiting.

caution capsule on the master caution panel illuminates.

Should the particle or particles not bum-off, the NO. 1 or

l. Engine fail detection.

NO. 2 .ENG CHIP DET caution capsule will illuminate.

m. Power assurance test.

Also, the corresponding ENGINE CHIP DETECTOR or

TRANSMISSION CHIP DETECTOR magnetic indicator

n. Engine history/fault recording.

on the MAINTENANCE PANEL will latch. Power for the

o. Engine-to-engine communication (via data bus).

PWR MDL CHIP BURN-OFF is supplied by the No. 1 DC

bus through the HYDRAULICS MAINT PNL circuit breaker

p. Automatic switchover to reversionary backup in the

on the No. 1 PDP.

event of a FADEC primary system failure.

2-26

Change

TM 55-1520-240-10

The FADEC provides automatic engine start, simulta-

When both engines are in reversionary mode, RRPM will

neously sequencing ignition, start fuel, and stabilized op-

require more pilot attention since proportional rotor speed

eration at idle. A data link between 1 and 2 engine FADEC

governor will not hold speed as accurately as the primary

systems transmits signlas to achieve load sharing. It also

systems. With large collective changes, the rotor speed can

provides control of N1 speed and NR (N2) output shaft

change up to ±

3 percent from a nominal setting.

speed to maintain the rotor system at a near constant RRPM

throughout all flight power demand conditions. FADEC

If a fault is detected and to ensure positive engagement of

provides smooth acceleration and overtemperature protec-

the reversionary mode, set the FADEC control REV/PRI

tion when ECLs (both together) are moved from GROUND

switch for the affected engine to REV

to FLIGHT. Overtemperature protection is provided

The reversionary system provides the following control

(through the DECU temperature limiting function) by con-

functions:

trol system thermocouple interface at the power turbine

inlet. The control system compares PTIT temperature sig-

a. Automatic start sequencing including over temperature

nals with reference limits to calculate and provide appro

protection, but not start abort.

priate N1 acceleration. During starts, an absolute

816°C

limit is set and if exceeded an engine out indication and

b. Pilot controlled start fuel enrichment/derichment, if

shutdown will occur. If compressor performance deterio-

required, through ECL modulation.

rates for any reason, surge detection automatically allows

c. Ground idle set at

50 to 59 percent

with ECL at GND.

recovery from compressor instability while protecting the

engine from damage due to overtemperature..

d. RRPM droop compensation based on thrust lever

The FADEC system consists of:

position.

q. The Digital Electronic Control Unit (DECU) includes

e. Beep capability becomes active for load match to other

a primary mode and a reversionary section for backup (fig

engine.

2-19.3).

f. Full contingency power capability.

r. The Hydromechanical Metering Assembly (HMA),

includes Hydromechanical Fuel Metering Unit (HMU) and

g. Over temperature protection throughout operation.

fuel pump unit for all fuel metering to support both primary

and reversionary fuel metering, a self-contained alternator

h. Engine shutdown in response to ECL being placed at

for powering the FADEC electronics, primary and rever-

STOP.

sionary compressor bleed air control, and redundant speed

i. Tracking of the primary mode during normal primary

sensing.

mode operation allowing a smooth switchover when se-

s. ENG COND panel (fig. 2-19.1).

lected.

t. FADEC control panel (fig. 2-19.2).

If in reversionary mode for any reason (training) and there

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

is a reversionary failure, the FADEC will not automatically

Lever.

switch back to the primary mode. The pilot must manually

(1) On 714A engine installations, engine beep

select PRI mode.

switches are only active when in reversionary mode.

CAUTION

(2) Each switch is labeled NO. 1 or 2 which is used

to adjust RRPM when in reversionary mode.

If both the primary and reversionary system

(3) Operation of the beep switches on the 714A in

fail, the engine remains at the fuel flow being

the reversionary mode are the same as for the 712 except

used at the time of the failure. When a failed

that each switch operates respective engine independently.

fixed fuel flow condition exists, the ECL and

If only one engine is in reversionary mode, the RRPM will

the beep trim switch for the affected engine is

not change, as it is governed by the engine in primary mode.

inoperative, therefore there is no propor-

tional control through the ECL except under

2-60.2. Reversionary System.

714A

some conditions STOP. Engine shutdown may

be accomplished by moving the ECL from its

NOTE

present position to STOP Under these con-

Aircrew should be alert to the possibility of

ditions, the ENG 1 (2) FAIL and FADEC 1 (2)

abrupt NR and engine power changes when

cautions are illuminated.

operating the FADEC in single or dual engine

REV mode (s).

When taking off with one engine in reversionary mode the

The reversionary (backup mode) automatically takes control

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

of the engine if the primary mode fails or if selected by the

used to set the correct rotor speed via the FADEC NR%

operator via the FADEC panel, REV switch.

switch. The operator then uses the beep switch of the engine

When an engine is operating in reversionary mode, FADEC

in reversionary mode to match engine torque.

provides engine and rotor control through an N1 speed

governor, beep control, and a thrust pitch compensator.

Change

2-26.1

TM 55-1520-240-10

This provides the FADEC system with back up electrical

2-60.3. ENG COND Panel. 714A

power in the event of a HMU integral alternator failure thus

CAUTION

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

switch to OFF will reduce operating time on the FADEC

When adjusting controls or switches on the

circuitry. The B/U PWR switch should always be ON during

overhead switch panel, make sure gloves or

engine operation.

sleeves do not catch and inadvertently move

d. OSPD 1, 2 Switch.

The Over Speed test switch is a

the ENG COND levers

three position switch used to test the FADEC overspeed

The ENG COND panel is located in the overhead switch

system. In the event of a NR over-speed of 114.8. percent,

panel (fig. 2-19.1).

FADEC reduces fuel flow to a ground idle condition. The

NO. 1 and NO. 2 Levers. The ENG COND Levers (ECL)

system remains activated until the over-speed condition no

provide the pilot with proportional acceleration and decel-

longer exists, and will re-activate as soon as an overspeed

eration authority. The levers are spring-loaded outboard

re-occurs. The system contains provisions to inhibit over-

creating a gated motion when advanced from the STOP to 1

speed trip command if the other engine has experienced a

/ 2 GND and to FLT positions. ENG COND lever friction is

overspeed trip condition. To prevent inadvertent operation

provided to reduce the possibility of over-torquing trans-

during flight, this test is locked out if NR is greater than

missions by resisting rapid movement of the levers.

81.3 percent. When performing an overspeed test with the

engine running and the RRPM 79.0 ±1% and the function is

2-60.4. FADEC Panel.

714A

The FADEC panel is

locked out above 81.3 percent, pressing the test switch to 1

located in the overhead switch panel (fig. 2-19.2). It

or 2, lowers the overspeed trip threshold to 79 ±1% NR. At

comprises of the following:

this time the system senses an overspeed and reduces the

fuel flow.

a. NR% Switch.

The NR% switch controls a rheostat

which allows the operator to select any RRPM between

e. LOAD SHARE, PTIT/TRQ Switch.

The primary FA-

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

DEC system provides pilot selectable engine torque or PTIT

With the ECL(s) in FLT, NR will be maintained at the

matching to govern the engines. Torque matching is nor-

selected speed. 100% is the normal position.

mally the preferred option. The selected parameter is

constantly compared between the two engines until the

b. PRI/REV Switches.

The primary mode is the normal

RRPM stabilizes at a datum figure. The PTIT option may be

mode of operation. The REV (reversionary) mode is select-

used when one engine is running hot. N1 matching is

able as a backup mode or is automatically selected if the

engaged automatically if the selected matching mode fails.

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

is defined as one in which normal primary system perfor-

f. ENG START Switch.

It is a three position switch,

mance might be jeopardized. Other failures are classified as

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

soft failures when the system is fault tolerant and can

used to commence the start sequence on the respective

continue fully operational with the fault signal present.

engine.

c. B/U PWR Switch.

The Back-Up Power switch when

ON connects the aircraft battery relay and essential relay.

2-60.5. DECU Unit.

714A The two airframe mounted

DECUs, one for each engine, contain the primary and

A60879

Figure 2-19.1. Engine Condition Panel 714A

Figure 2-19.2.

FADEC Panel 714A

2-26.2

Change 13

TM 55-1520-240-10

reversionary mode electronics. The DECUs are located on

the left and right side of the aft cabin at station 400 (fig.

2-19.3).

2-60.6. BIT.

714A

The DECU contains a two-digit BIT

display. When active, the display indicates the operating

status of the FADEC system and power assurance test

results. A complete list of the FADEC BIT fault codes are

located at Table 2-1.1. The fault monitoring carried out by

the DECU consists of:

a. Power up tests.

b. Fault tests designed to discover dormant faults.

c. A set of repeated monitoring tests to detect faults

occurring during normal operation.

Fault information for the previous or current engine cycle

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

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

engine shutdown. During engine shutdown (when a ECL is

at STOP or N1 is less than 10 percent) faults are not stored.

Fault indications are stored in the DECU and are retained

throughout the life of the control unit. However, fault

information prior to the previous cycle can only be accessed

with specialized test equipment. During engine start the

DECU BIT displays 88 for a satisfactory test or if the test

A73381

fails, a fault code. Faults are classified as either “HARD“ or

"SOFT". In primary mode a Hard fault will cause the

FADEC to transfer to Reversionary, while in Reversionary a

Hard fault will cause the FADEC to "fail fixed" to a constant

Figure

2-19.3. DECU Panel

714A

power condition. If a hard fault occurs in Primary after a

percent N1 then release the ENG START switch. The

hard fault exists in reversionary then the primary will fail

automatic start sequence has been energized and FADEC

fixed. In the event of a soft fault the FADEC will remain in

the mode it was in prior to the fault but there may be some

will complete the start. When an engine start is energized,

degradation or redundancy. All soft faults are less severe

FADEC turns on the engine start solenoid to introduce fuel

than a Hard fault since the FADEC will not switch modes

flow and energize the engine igniters. Successful engine

due to a soft fault.

ignition is immediately indicated to FADEC by an increase

of PTIT or compressor speed (N1). Engine temperature is

The activation of the BIT display is dependent upon the

monitored throughout the sequence and will result in a fuel

position of the ECL as follows:

flow reduction if the temperature exceeds

650°C with a full

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

cutback to minimum flow limit at

760°C. The starter motor

last engine cycle and current faults are displayed.

and igniter are automatically turned off when N1 speed

exceeds

48 percent. Ground idle is

50 to 59 percent and is

b. When the ECL is positioned at GROUND IDLE only

corrected for temperature. The engine is allowed to take 45

current faults are displayed.

seconds to stabilize at ground idle.

c. When the ECL is positioned at FLIGHT the display

will be turned off except as required for power assurance

2-60.8. Engine Start Abort.

714A

PTIT above

816°C

test

will cause immediate fuel shutoff to below minimum fuel

flow. The ECL must be retarded to STOP to achieve total

fuel shutoff. A start abort results in the ENG 1 (2) FAIL

warning to be illuminated until the abort is reset by moving

the ECL to STOP

2-60.7. Starting in Primary Mode.

714A

CAUTION

2-60.9. Starting in Reversionary Mode.

714A

The

initial start sequence in reversionary mode is the same as in

The (P3) compressor pressure signal line

primary mode except, when N1 reaches

8 percent, the

going to the DECU contains a manually

control system turns on the engine start fuel solenoid to

operated moisture drain valve. This valve

provide an initial altitude biased fuel flow and activates the

should not be drained while the engine is

igniters and latches the starter motor. The pilot can modulate

running.

fuel flow to the engine with ECL to start the engine at a

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

desired acceleration rate. Temperature and temperature rate

the GND position. Select and hold the respective ENG

limiters are the same as in primary mode except that the

START switch and allow the engine to accelerate to 10

over temperature start abort facility is not provided.

Change 13

2-26.3

TM 55-1520-240-10

Alternate Reversionary Starting. For most conditions, a start

(2) Retard the ECL to GROUND as ground idle

is successfully completed with the ECL held at the GND

speed is approached. Check that N1 is stabilized at ground

position. However, if the engine fails to start due to either a

idle,

rich or lean hung start condition, the pilot may use the ECL

to increase or decrease the start flow as required to complete

2-60.10. Starting Cycle, Aborting . and Motoring.

a successful start. If the engine fails to start using the normal

714A A starting cycle can be aborted at any time by

starting procedure, proceed as follows:

moving the ECL to STOP. If a motoring is required:

a. Reversionary Rich Hung Start. A rich hung start is

a. Set the ECL to STOP.

characterized by N1 holding at about 40 percent and PTIT

climbing above

600°C. If a rich hung start is experienced:

b. ENG START switch to 1 (2).

(1) Set the affected engine ECL to STOP.

c. Hold ENG START switch until PTIT decreases below

(2) Allow PTIT to decay to 260°C or below (motor

26O°C.

engine as required).

2-60.11. Power Assurance Test Switch.

714A

The

(3) Check N1 0 percent.

primary FADEC will perform a BIT whenever the PWR

(4) Advance ECL half the distance between STOP

ASSURANCE TEST switch has been placed to the desired

and GND (15°).

engine position. The switch is located below the mainte-

(5) Motor engine using ENG START switch until

nance panel at station 524. The switch is labeled 1/OFF/2

10% N1 and then release switch.

and spring loaded to OFF’. The results of the test are

displayed in the DECU BIT window.

(6) After engine ignition (PTIT rising), slowly ad-

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

2-60.12. Engine Wash System.

714A

The heli-

idle.

copter is equipped with an engine wash system for each

b. Reversionary Lean Hung Start. A lean hung start is

engine. Air and water connections are externally mounted

characterized by N1 hanging at approximately 30 percent

inboard of each engine work platform. A series of spray

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

nozzles are installed at the engine inlet and air lines are

experienced:

routed to the bleed band actuator.

(1) Slowly advance the hung engine ECL to achieve

acceleration (maximum of one-third travel from GND to

FLT).

2-26.4

Change 13

TM 55-1520-240-10

Table

2-1.1. FADEC Bit Fault Code List

714A

FAULT

CODE

FAULT DESCRIPTION

CODE

FAULT DESCRIPTION

Microprocessor hard fault

Communication line soft fault on NR (0) sig-

Non-volatile random access memory (RAM)

nal

check sum hard fault

Communication link soft fault

Non-volatile RAM engine history data soft

Communication line soft fault on N1B (0)

fault

signal

Non-volatile RAM fault data soft fault

Loss of load share signals hard fault

Non-volatile RAM accumulated fault data soft

Overspeed drive soft fault

fault

P3 transducer soft fault

Non-volatile RAM write test soft fault

P1 transducer soft fault

Non-volatile RAM storage incomplete

28V “OR” diodes soft fault

Non-volatile RAM history data inconsistent

+ 1OV reference hard fault

Minor cycle not competed hard or soft fault

±15V hard fault

Engine monitoring system cycle not com-

+12V reversionary or ±12V overspeed soft

pleted soft fault

fault

Analog-to-digital conversion not completed

+5V hard fault

hard fault or soft fault

Primary and reversionary cold junction com-

RAM failure hard fault

pensation for temperature signal soft fault

Opcode error hard fault

+24V regulator soft fault

Q sensor soft fault

+5V reversionary soft fault

N2 set potentiometer soft fault

Reversionary system soft fault or idle check

Primary and reversionary collective pitch

soft fault

angle linear variable displacement transformer

T4.5 calibration soft fault

(LVDT) soft fault

Over-speed check soft fault

NR sensor soft fault

Primary and reversionary 400Hz resolver ref-

Primary and reversionary ENG COND lever

erence soft fault

resolver soft fault

Watchdog timer test hard fault

Airframe emergency 28V DC supply soft fault

E1*

Primary and reversionary T4.5 sensor soft

Airframe +28V DC supply soft fault

fault

Primary or reversionary N1B sensor soft fault

E2*

T1 sensor soft fault

Primary or reversionary N2B sensor soft fault

E3*

N2A sensor soft fault

Primary or reversionary T4.5 sensor soft fault

E4*

Primary and reversionary N2B sensor hard

Primary or reversionary collective pitch angle

fault

LVDT soft fault

E5*

N2A/N2B difference soft fault

Primary or reversionary ENG COND lever

N1A sensor soft fault

resolver soft fault

Primary and reversionary N1B sensor hard

Primary or reversionary power level angle

fault

potentiometer soft fault

N1A/N1B difference soft fault

Primary or reversionary cold junction com-

MV potentiometer hard or soft fault

pensation for temperature signal soft fault

Fuel flow stepcount difference hard fault

Reversionary +28V soft fault

Primary and reversionary power level angle

Reversionary T4.5 calibration soft fault

potentiometer soft fault

Primary or reversionary 400Hz resolver refer-

Bleed valve solenoid hard fault

ence soft fault

Primary/Reversionary solenoid hard fault

Communication line soft fault on T4.5 (0)

Alternator voltage soft fault

signal

FA*

Start fuel solenoid soft fault

Communication line hard or soft fault on P1

Reversionary step count soft fault

(0) signal

Communication line hard or soft fault on T1

*Denotes engine related faults.

(0) signal

Communication line soft fault on Q (0) signal

Communication line soft fault on N2 SET (0)

signal

Communication line soft fault on collective

pitch angle (0) signal

Change 13

2-26.5/(2-26.6 blank)

TM 55-1520-240-10

SECTION IV

FUEL SYSTEM

2-61. Fuel Supply System.

quantity probes, a jet pump for evacuating the pressure

refueling system, a dual pressure refueling shutoff valve,

The fuel supply system furnishes fuel to the two engines,

a dual fuel level control valve, and a gravity filler port.

the heater, and the APU. Two separate systems,

Each auxiliary tank contains a fuel pump with automatic

connected by crossfeed and a pressure refueling lines

shutoff feature, a quantity probe, a dual pressure

are installed. Provisions are available within the cargo

refueling shutoff valve, and a fuel level control valve.

compartment for connecting Extended Range Fuel

System (ERFS) and ERFS II to the two fuel systems.

Each fuel system consists of three fuel tanks contained

in a pod on each side of the fuselage. The tanks are

identified as forward auxiliary, main, and aft auxiliary

tanks. During normal operation, with all boost pumps

operating, fuel is pumped from the auxiliary tanks into the

main tanks, then from the main tanks to the engine. A

simplified fuel flow diagram is engraved on the FUEL

CONTR (control) panel on the overhead switch panel

(fig. 2-20).

When the fuel is consumed in an auxiliary tank, the fuel

pump is automatically shut off and a check valve closes

to prevent fuel from being pumped back into that tank.

Should a fuel pump fail in an auxiliary tank, the fuel in that

tank is not usable. However, should both boost pumps

fail in a main tank, fuel will be drawn from the main tank

as long as the helicopter is below 6,000 feet pressure

altitude (PA).

Fuel is delivered to the APU from the left main tank and

to the heater from the right main tank. Fuel system

switches and the auxiliary tank low pressure indicating

lights are on the FUEL CONTR panel, the fuel line

pressure caution capsules are on the master caution

panel, and the fuel flow meter is on the center instrument

Figure 2-20. Fuel Control Panel

panel. The single point pressure refueling panel and

A rollover vent system is installed in each tank. This

nozzle adapter are on the right side above the forward

system prevents fuel spillage from the vents should the

landing gear. Refer to Section XV for fuel tank capacities,

helicopter roll over following a crash landing. The vent

fuel grades, and fuel system servicing procedures.

system within the tanks have a condensate drain at the

aft end, however, aircraft maneuvering should never

2-62. Fuel Tanks.

force fuel into the vents. Sump drains are also installed

The fuel tanks are crashworthy self-sealing tanks with

on the bottom forward end of each tank.

breakaway fittings. The main fuel lines are constructed

of self-sealing material. Penetration of the tank wall or a

2-63. Controls and Indicators.

fuel line by a projectile exposes the sealant to the fuel,

The fuel controls are the FUEL PUMP switches, XFEED

activates the sealant, and close the hole.

fuel valve switch, the engine fuel valve, and the manual

Breakaway self-sealing fittings are installed where the

defueling valve. Indicators include the crossfeed fuel and

main fuel lines connect to the fuel tank and adjacent

engine fuel valve warning lights, the FUEL QUANTITY

structure. Under high impact loads, the fittings shear or

indicator and caution capsules, FUEL flow indicator,

break at predetermined locations, seal themselves,

AUX PRESS indicating lights, and FUEL PRESS caution

retain the fuel, keeping fuel loss and post-crash fire

capsules. Refer to para. 2-70 for a description of the

hazard to a minimum. Electrical cables having

pressure refueling system controls and indicators.

lanyard-release type connectors are installed where the

2-64. FUEL CONTR Panel. The FUEL CONTR panel

cables attach to adjacent structure. The connectors

(fig. 2-20) consists of eight two-position fuel boost pump

automatically release if the fuel tank breaks away from

switches, two PRESS-TO-TEST AUX PRESS indicating

the pod.

lights, a two-position XFEED switch, and a two-position

Each main tank contains two fuel boost pumps, three fuel

REFUEL STA switch.

Change 14

2-27

TM 55-1520-240-10

a. FUEL PUMP Switches.

Each switch controls a

single-speed electrically driven fuel boost pump. La-

beled next to each switch is the name of the pump which

it operates. Each switch has an ON and OFF position.

When one of these switches is at ON, power from the

No. 1 or No. 2 DC bus closes the respective pump relay

connecting power from the No. 1 or No. 2 AC bus to

energize the pump. When switch is at OFF, the relay

circuits open and power from the No. 1 and No. 2 AC

bus is de-energized thus shutting off the pump. Power is

supplied for these relay circuits by the No. 1 and No. 2

DC bus through the LH and RH FUEL PUMP CON-

T - AUX AFT, MAIN AFT, MAIN FWD, and AUX

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

Power is supplied to the pumps circuits by the No. 1 and

No.

2 AC bus through the LH and RH FUEL

PUMPS -

MAIN FWD, MAIN AFT, AUX FWD,

and AUX AFT circuit breakers on the No. 1 and No. 2

PDP.

b. AUX PRESS Indicating Lights.

Each light is elec-

trically connected to the forward and aft auxiliary tank

pressure switches. When this indicating light illumi-

nates, it indicates that the fuel pressure in-either the

forward or aft auxiliary fuel line is below

± 1 psi. The

auxiliary tank fuel boost pump switches must be at ON

to provide electrical power to the indicating light. The

light intensity can be adjusted by turning the light

housing. Power is supplied to operate the indicating

light by the No. 1 and No. 2 DC bus through the LH or

Figure 2-21. Engine Fuel Valves, Sta. 498

RH FUEL PUMP CONT AUX FWD & AUX AFT

and two PRESS-TO-TEST FUEL VALVE WARNING

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

LIGHT next to FUEL VALVE # 1 ENGINE and # 2

c. XFEED Switch.

The switch electrically operates

ENGINE (fig. 2-22). They indicate the operating con-

two fuel valves in the crossfeed line. The switch has an

dition of the individual valve and

^-associated circuitry.

OPEN and CLOSE position. When the switch is at

Power is supplied to operate the crossfeed FUEL

OPEN, power from the No. 1 DC bus opens the fuel

VALVE WARNING LIGHT by the No. 1 DC bus

valves through the XFEED CONT circuit breaker on

through the XFEED CONT circuit breaker on the No.

No. 1 PDP. When the switch is at CLOSE, electrical

1 PDP. Power is supplied to operate the engine FUEL

power closes the valves.

VALVE WARNING LIGHT by the DC essential bus

through the ENGINE NO. 1 and NO. 2 FUEL SHUT-

2-65. Fuel Valves. There are two engine and two

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

crossfeed fuel valves.

The following description on when the light will illumi-

a. Engine Fuel Valves.

One engine fuel valve (fig.

2-21) is in the fuel supply line to each engine. The valve

nate is for the XFEED switch. The same result applies

is electrically operated by the FIRE PULL handles and

to the engine fuel valves with the FIRE PULL handles.

manually by a lever on the valve. They are located at sta.

(1) Each time the XFEED switch is moved from

498 and labeled FUEL VALVE # 1 ENGINE and

CLOSE to OPEN or OPEN to CLOSE. After this

FUEL VALVE # 2 ENGINE.

operation, the light should extinguish, indicating the

b. Crossfeed Fuel Valves.

The crossfeed fuel valves

crossfeed valve is synchronized with the switch position.

connects the No. 1 and No. 2 engine fuel lines. When

(2) When a short circuit occurs, causing a signal

the valve is opened, both engine fuel feed lines are

to be applied opposite to the valve position. However,

interconnected and fuel can be supplied from both fuel

the valve will remain at the position last selected by the

tanks to feed either engine or from either tank to feed

XFEED switch.

both engines. Fuel cannot be transferred between tanks.

The valves are electrically operated by the XFEED

(3) When the crossfeed valve protection relay

switch on the FUEL CONTR panel or manually by a

fails. The crossfeed valve will remain at the last selected

position and the valve can be operated electrically or

lever on the valve. They are labeled FUEL VALVE

CROSS FEED and located at station 504.

manually, as required.

c. FUEL VALVE WARNING LIGHT.

There are

2-66. Manual Defueling Valve. A manual defueling

two PRESS-TO-TEST FUEL VALVE WARNING

valve is in the aft cargo compartment next to FUEL

LIGHT next to each FUEL VALVE CROSS FEED

VALVE # 2 ENGINE. The valve should only be used

2-28

TM 55-1520--240-10

by maintenance personnel to defuel the helicopter or

adjust fuel load.

2-67. Fuel Quantity indicator and Selector

Switch. An indicator calibrated to measure fuel quan-

tity in pounds and a seven position selector switch (fig.

2-23) is on the center instrument panel. Power is

supplied to the indicator through the FUEL QUAN-

TITY selector switch by the No. 1 AC bus through the

FUEL QTY circuit breaker on the No. 1 PDP.

a. FUEL QUANTITY Indicator.

The indicator pro-

vides two types of display. One display is in digital form

and the other is a pointer, The digital readout contin-

uously indicates the total amount of fuel remaining in

all the fuel tanks. The pointer remains hidden until one

of the tank positions on the FUEL QUANTITY selec-

tor switch is selected. Then, the pointer will indicate

fuel remaining in that tank. The fuel quantity indicator

is electrically connected to 10 capacitance-type measur-

ing units in the tanks.

b. Fuel Quantify Selector Switch.

The fuel quantity

selector switch has seven positions labeled TOTAL, L

(left) and R (right) FWD, MAIN, and AFT. Selecting

any position other than TOTAL causes the indicator

pointer to display the fuel remaining in that tank. The

digital readout is not affected during individual tank

readings.

2-68. Fuel System Cautions. Four caution capsules

Figure 2-23. Fuel Quantity Indicator and Selector

are dedicated to the fuel system.

Switch

a. L and R FUEL LOW.

Two fuel quantity caution

caution panel (fig. 2-51) of the center instrument con-

capsules, one for each main tank, are on the master

sole. Each light is electrically connected to a thermistor

sensor on a measuring unit in the respective main tank.

These lights are labeled L FUEL LOW and R FUEL

LOW. When there is

20 percent of fuel remaining in the

main tank, the caution capsule for that main tank

illuminates (20 percent of fuel is equal to

320 to 420

pounds.) Power for these capsules is supplied by the DC

essential bus through the LIGHTING CAUTION PNL

circuit breaker on the No. 1 PDP.

b. L and R FUEL PRESS.

Two caution capsules

labeled L FUEL PRESS and R FUEL PRESS are on

the master caution panel. Each caution capsule is

electrically connected to a fuel pressure switch between

the main tank and the engine fuel valves. When one of

these capsules illuminates, it indicates that fuel pressure

in the respective fuel line is below

± 1 psi. Fuel

pressure is measured after the fuel boost pumps and not

at the engine driven pump. When a fuel pressure

caution illuminates, it does not represent a possible

engine flameout, unless flight is being conducted

above

6,000 feet PA. Power for these capsules is supplied by

the DC essential bus through the LIGHTING CAU-

TION PNL circuit breaker on No. 1 PDP.

2-69. FUEL Flow Indicators.

A dual fuel-flow indicator (fig. 2-24), on the center

instrument panel , indicates fuel flow to each engine in

pounds per hours. The indicator dial is graduated from

0 to 3000 pounds per hour in

100 pound increments.

Figure 2-22. Fuel Valve Warning Light, Sta. 500

The signal to drive the indicator is derived from a fuel

2-29

TM 55-1520-240-10

Electrical power is applied to the system only when the

REFUEL STA switch on the cockpit FUEL CONTR

panel is placed to ON. Power to operate the pressure

refueling system is supplied by the DC switched battery

bus through the REFUEL circuit breaker on the No. 1

PDP.

2-71. Controls and Indicators.

Except for the REFUEL STA switch on the cockpit

FUEL CONTR panel, all pressure refueling system

controls and indicators are on the pressure refueling

station panel (fig. 2-25).

2-72. PWR Control Switch. The PWR (power) con-

trol switch is labeled ON and OFF. When placed to ON,

electrical power is applied to the pressure refueling

system and to the refueling station quantity indicator

provided the REFUEL STA switch on the cockpit

FUEL CONTR panel is at ON. Also, the PWR ON light

Figure 2-24. Fuel Flow Indicator

will illuminate, the fuel quantity indicator will register

flow transmitter in the fuel line of each engine at the

the quantity of fuel in the tanks, and the REFUEL

quick disconnect shelf. Power to operate the No.

VALVE POSN lights will illuminate momentarily. When

indicator is from the No. 1 AC bus through the EN-

placed to OFF, electrical power is removed,

GINE NO. 1 FUEL FLOW circuit breaker on the No.

2-73. REFUEL STA Switch. The REFUEL STA

1 PDP. Power to operate the No. 2 indicator is from the

switch is on the cockpit FUEL CONTR panel (fig. 2-20)

No. 2 AC bus through the ENGINE NO. 2 FUEL

when placed to ON, applies electrical power from the

FLOW circuit breaker on the No. 2 PDP.

DC switched battery bus to the PWR ON switch on the

2-70. Pressure Refueling System.

refueling station panel. Setting the switch to OFF after

pressure refueling, closes the refuel valves and discon-

The pressure refueling system permits rapid refueling of

tinues electrical power to the refueling panel. When

all fuel tanks simultaneously or selective refueling of

pressure refueling, be sure the switch is at ON at all

any tank or combination of tanks. Maximum fueling rate

times. If the switch is at OFF, the aft auxiliary tanks will

is 300 gallons per minute at

55 psi. The system control

not fill, the remaining four tanks will fill to maximum,

panel and refueling nozzle receptacle are on the right

the refuel station quantity indicator is inoperative, and

side of the helicopter above the forward right landing

there is no precheck capability.

gear (fig. 2-25).

2-74. Fuel Quantity Indicator and Selector

In addition to the control panel and refueling recepta-

Switch. The pressure refueling station fuel quantity

cle, the system consists of a dual fuel level control valve,

indicator and selector switch (fig. 2-25) are identical to

a dual fuel shutoff valve in each tank, a jet pump in each

those in the cockpit. The indicator at the refueling

main tank, and pressure refueling manifold.

station indicates fuel quantity only when the REFUEL

a. Dual Fuel Level Control Valves and Dual Fuel

STA switch on the cockpit FUEL CONTR panel (fig.

Shutoff Valves. The dual fuel level control valves con-

2-20) is at ON and the PWR switch on the refueling

trol the operation of the dual fuel shutoff valves. When

station panel (fig. 2-25) is at PWR ON. Electrical power

fuel in a tank rises to the full level during pressure

to drive the indicator is AC from a solid-state inverter in

refueling, the floats in the control valve close and apply

the cabin at sta 220. The inverter, in turn, is powered by

a signal to the shutoff valve, closing it. The floats can

the DC switched battery bus through the FUEL RE-

also be closed electrically to stop fuel flow into a tank at

FUEL circuit breaker on the No. 1 PDP.

some intermediate level. The floats are controlled by

the FUEL CELL SHUTOFF VALVE TEST switches

2-75. FUEL CELL SHUTOFF VALVE TEST

on the refueling control panel.

Switches. Seven three-position FUEL CELL SHUT-

OFF VALVE TEST switches are on the refueling

b. Jet Pumps.

The jet pump installed in each main

control panel (fig. 2-25). The switches are used to test

tank evacuates the refueling manifold and discharges

the automatic shutoff features in each tank and to stop

the displaced fuel into the main tank. The jet pump is

refueling when the desired fuel level is attained in each

activated when the forward boost pump in each main

or all tanks. Six of the seven switches are connected to

tank is first turned ON following pressure refueling.

fuel level control valves in a specific tank. The seventh

c. Pressure Refueling Manifold.

The pressure refuel-

switch, labeled ALL TEST, is electrically connected to

ing manifold connects all tanks to the pressure refueling

the fuel level control valves in all six tanks. Setting any

receptacle. It does not include projectile resistant fea-

of the six switches to PRI OFF or SEC OFF raises the

tures because the fuel is evacuated before flight by the

corresponding primary or secondary float in the fuel

jet pumps.

level control valve. This action simulates a high fuel

2-30

TM 55-1520-240-10

level and causes the fuel shutoff valve in that tank to

opened and allow fuel flow from the refueling system

close. Setting the ALL TEST switch to PRI OFF or SEC

into the aft tank,

OFF raises the corresponding float in all six tanks and

The valves are controlled by the refueling station PWR

shuts off the fuel flow into all tanks simultaneously.

ON switch. When the switch is ON, the valves are

2-76. REFUEL VALVE POSN Indicating Lights. The

opened and the indicating lights will illuminate momen-

two amber PRESS-TO-TEST REFUEL VALVE POSN

tarily indicating valve transitioned from close to open.

Conversely, when the switch is OFF, the valves are

(position) lights (fig. 2-25) indicate the status of the two

closed and the indicating lights will illuminate momen-

refueling valves in the pressure refueling system. The

tarily indicating valve transitioned from open to close. A

valves are normally closed and prevent fuel feedback

continuously illuminated light, with the switch at OFF,

into the aft auxiliary tanks when the aft tank pumps are

indicates the associated valve is opened and the fuel in

operating. While pressure refueling, the valves are

that tank will not be available.

2-32

TM 55-1520-240-10

SECTION V

FLIGHT CONTROLS

2-77. Flight Control System.

by imparting equal but opposite lateral cyclic pitch to the

blades. Lateral control is obtained by application of equal

The helicopter is controlled by changing the pitch of the

lateral cyclic pitch to the blades with the cyclic control stick.

blades either collectively or cyclically. Pitch changes are

The helicopter is controlled longitudinally with the cyclic

made by the pilot’s movement of the flight controls which

stick through application of differential collective pitch.

include a THRUST CONT (control) lever, a cyclic control

stick, and directional pedals. The pilot’s controls are inter-

In addition, the helicopter has an advanced flight control

connected with the copilot’s controls.

system (AFCS). AFCS provides the following features:

Plight control movements are transmitted through a system

a. Rate damping in all axes and sideslip stability.

of bellcranks, push-pull tubes, and actuators to a mixing unit

just aft of the cockpit, next to the forward transmission. The

b. Pitch and roll attitude hold and heading hold.

control movements are mixed to give the correct lateral

c. Airspeed hold.

cyclic and collective pitch motions to the rotors through

d. Improved control response in pitch, roll, and yaw.

dual hydraulic actuators. These dual boost actuators are

under each swashplate. Each set of dual boost actuators is

e. Barometric and radar altitude hold.

normally powered by both flight control hydraulic systems.

f. Automatic coupled turns.

The helicopter is vertically controlled with the THRUST

CONT lever through application of equal pitch to all blades.

g. Longitudinal cyclic trim scheduling.

Directional control is obtained with the directional pedals

THRUST CONTROL LEVER

(TYPICAL)

(ON HELICOPTERS MODIFIED BY

(ON HELICOPTERS MODIFIED

MWO 1-1520-240-50-56, HUD)

WITH 714A FADEC)

A73382

Figure 2-26.

Thrust Control Lever

Change 13

2-33

TM 55-1520-240-10

2-78. THRUST CONT Lever.

A detent capsule establishes a ground operation detent to

reduce droop stop pounding. A viscous damper in the thrust

Either THRUST CONT lever

712 (fig. 2-26) or 714A

control system improves control feel. Mounted on each

[fig. 2-26) is used to apply equal pitch simultaneously to

THRUST CONT lever is an auxiliary switch bracket con-

both rotors, thus controlling ascent and descent of the

taining a SEARCH LIGHT control switch, a SLT-FIL

helicopter. Raising the THRUST CONT lever increases

(search light filament) switch, two ENGINE BEEP TRIM

pitch. Lowering the THRUST CONT lever decreases pitch.

switches, and a HUD control switch.

An integrated lower control actuator (ILCA) is installed

2-79. Cyclic Stick

between the THRUST CONT lever and the mixing unit.

Each cyclic stick (fig. 2-27) is used for lateral and longitu-

This actuator assists the pilot in moving the THRUST

dinal control of the helicopter. Moving the cyclic stick to the

CONT lever. A cockpit control driver actuator (CCDA) is

right tilts both rotor disks equally to the right and causes the

also installed in the thrust control system. This actuator

helicopter to roll to the right in flight. Moving the cyclic

responds to signals from the AFCS and increases or de-

stick to the left causes the opposite movement. When

creases collective pitch on the blades to maintain a constant

moving the cyclic stick forward, the pitch of the fwd rotor

altitude. In addition, a balance spring is installed that

blades is decreased collectively while the pitch of the aft

counteracts the downward imbalance of the THRUST

rotor blades is increased collectively, thus causing a nose-

CONT level.

down helicopter attitude in flight. Moving the cyclic stick

aft causes the opposite movement resulting in a nose-up

NOTE

attitude.

If the THRUST CONT lever CCDA fails, the

Two ILCA’s, one for lateral control and one for longitudinal

THRUST CONT lever will slip when a force

control, are installed to assist the pilot in moving the cyclic

between 7 and 23 pounds is applied.

stick. In addition to these actuators, viscous dampers are

A BRAKE TRIGGER switch under each THRUST CONT

installed. One damper is for longitudinal control and one for

grip controls the magnetic brake of the CCDA in the flight

lateral control to improve control feel.

control closet. Pressing the switch applies electrical power

Located on the pilot and copilot cyclic stick grips are a

to release the magnetic brake in the THRUST CONT lever

CENTERING DEVICE RELEASE switch, an AFCS trim

CCDA. The THRUST CONT lever can then be freely

switch, a CARGO HOOK RELEASE switch, interphone-

moved.

transmitter TRIGGER switch, and a FLARE DISP (dis-

When barometric or radar altitude

hold has been selected,

penser) control switch.

pressing the trigger will disengage

altitude hold. When the

switch is released, power is applied through the simplex

2-80. CENTERING DEVICE RELEASE Switch.

The

clutch to the THRUST CONT lever CCDA and the AFCS

CENTERING DEVICE RELEASE switch (fig. 2-27) is

will hold the altitude. Power is supplied to operate the

used to simultaneously release the force feel trim magnetic

THRUST CONT lever magnetic brake from the DC essen-

brakes for the lateral, longitudinal, and directional flight

tial bus through the THRUST BRAKE circuit breaker on the

controls. In addition, it disengages bank angle hold, heading

No. 1 PDP

hold, and heading select functions when AFCS is operating.

Power is supplied to operate the magnetic brakes from the

The

712 THRUST CONT lever is also electrically linked

DC switched battery bus through the CONT CENTER

to the power turbine actuator through the droop eliminator

circuit breaker on the No. 1 PDP.

system. An upward movement of the THRUST CONT lever

A centering spring and a magnetic brake for each control

electrically increases the power turbine governor

speed

provide a sense of force feel to hold the control in a trim

setting to compensate

for inherent engine droop and main-

position. However, the pilot can override the force manually

tain engine speed as rotor loads are increased. A downward

while maneuvering the helicopter. When the switch is

movement of the THRUST CONT lever electrically de-

pressed, electrical power is applied to release the magnetic

creases the power turbine governor speed setting.

brakes. Each centering spring assumes a new trim position

where the control forces are nulled. Releasing the switch

The

714A

714A system includes both thrust lever position

removes electrical power and applies the magnetic brakes.

compensation and thrust lever rate compensation.

The centering springs are retained in their new positions.

2-34

Change

TECHNICAL MANUAL MAINTENANCE MANUAL FOR ARMY CH-47D HELICOPTER (1992) - page 2