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OPERATOR’S MANUAL FOR ARMY CH-47D HELICOPTER (EIC: RCD, 2003) - page 2

TM 1-1520-240-10

EAPS control switch located in the EAPS control panel

b. The ENG 1 and ENG 2 DOORS CLOSE-OPEN

is in the OPEN position and the doors are fully open, and

switch electrically positions the by-pass doors, open pr

will extinguish when the switch is in the CLOSE position

closed. The switches receive power from the No. 1 and

and the doors are fully closed.

No. 2 28-volt DC buses through circuit breakers marked

EAPS 1 and EAPS 2 BYPASS DOORS.

2-3-44. Differential Pressure Switch.

Each separator is equipped with a differential pressure

switch to determine the air passages of the EAPS unit are

blocked. The switch senses the pressure difference be-

tween the inside and outside of the curved panels of the

EAPS unit. When a differential pressure is detected, the

EAPS 1 FAIL or EAPS 2 FAIL caution lights illuminated.

2-3-45. EAPS Switches.

There are four, 2 position toggle switches located on the

EAPS control panel (fig. 2-3-8) situated on the overhead

switch panel.

a. The ENG 1 and ENG 2 FAN ON/OFF switches

Figure 2-3-8. EAPS Control Panel

operate the control circuit to provide AC power for the

1. EAPS fan exhaust

EAPS fans. The switches receive power from the No. 1

and No. 2 28 volt DC buses through circuit breakers

marked EAPS 1 and EAPS 2 FAN CONT.

2. EAPS rails and slides.

2-3-17

TM 1-1520-240-10

Figure 2-3-9. EAPS 1 Shown With Bypass Doors Open

2-3-18

TM 1-1520-240-10

Figure 2-3-10. EAPS Control Boxes

2-3-19/(2-3-20 blank)

TM 1-1520-240-10

SECTION IV. FUEL SYSTEM

2-4-1. Fuel Supply System.

The fuel supply system furnishes fuel to the two engines,

the heater, and the APU. Two separate systems, con-

nected by crossfeed and a pressure refueling lines are

installed. Provisions are available within the cargo

compartment for connecting Extended Range Fuel Sys-

tem (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. 2-4-1).

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

Figure 2-4-1. Fuel Control Panel

switches and the auxiliary tank low pressure indicating

A rollover vent system is installed in each tank. This

lights are on the FUEL CONTR panel, the fuel line pres-

system prevents fuel spillage from the vents should the

sure caution capsules are on the master caution panel,

helicopter roll over following a crash landing. The vent

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

The single point pressure refueling panel and nozzle

system within the tanks have a condensate drain at the

adapter are on the right side above the forward landing

aft end, however, aircraft maneuvering should never

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

force fuel into the vents. Sump drains are also installed

grades, and fuel system servicing procedures.

on the bottom forward end of each tank.

2-4-3. Controls and Indicators.

2-4-2. Fuel Tanks.

The fuel controls are the FUEL PUMP switches, XFEED

The fuel tanks are crashworthy self-sealing tanks with

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

breakaway fittings. The main fuel lines are constructed

defueling valve. Indicators include the crossfeed fuel and

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

engine fuel valve warning lights, the FUEL QUANTITY

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

indicator and caution capsules, FUEL flow indicator, AUX

activates the sealant, and close the hole.

PRESS indicating lights, and FUEL PRESS caution cap-

sules. Refer to para. 2-4-10 for a description of the pres-

Breakaway self-sealing fittings are installed where the

main fuel lines connect to the fuel tank and adjacent

sure refueling system controls and indicators.

structure. Under high impact loads, the fittings shear or

break at predetermined locations, seal themselves, re-

2-4-4. FUEL CONTR Panel.

tain the fuel, keeping fuel loss and post-crash fire hazard

The FUEL CONTR panel (fig. 2-4-1) consists of eight

to a minimum. Electrical cables having lanyard-release

two-position fuel boost pump switches, two PRESS-TO-

type connectors are installed where the cables attach to

TEST AUX PRESS indicating lights, a two-position

adjacent structure. The connectors automatically re-

XFEED switch, and a two-position REFUEL STA switch.

lease if the fuel tank breaks away from the pod.

a. FUEL PUMP Switches. Each switch controls a

Each main tank contains two fuel boost pumps, three fuel

single-speed electrically driven fuel boost pump. La-

quantity probes, a jet pump for evacuating the pressure

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

refueling system, a dual pressure refueling shutoff valve,

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

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

Each auxiliary tank contains a fuel pump with automatic

When one of these switches is at ON, power from the No.

shutoff feature, a quantity probe, a dual pressure refuel-

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

ing shutoff valve, and a fuel level control valve.

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

2-4-1

TM 1-1520-240-10

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 CONT - 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

electrically 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 10 $ 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 RH Fuel PUMP

CONT AUX FWD & AUX AFT circuit breakers on the No.

1 and No. 2 PDP.

c. XFEED Switch. The switch electrically operates

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

OPEN and CLOSE position. When the switch is at

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

Figure 2-4-2. Engine Fuel Valves, Sta. 498

valves through the XFEED CONT circuit breaker on the

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

and two PRESS-0TO-TEST FUEL VALVE WARNING

er closes the valves.

LIGHT next to FUEL VALVE # 1 ENGINE and # 2 EN-

GINE (fig. 2-4-3). They indicate the operating condition

of the individual valve and associated circuitry. Power is

2-4-5. Fuel Valves.

supplied to operate the crossfeed FUEL VALVE WARN-

ING LIGHT by the No. 1 DC bus through the XFEED

CONT circuit breaker on the No. 1 PDP. Power is sup-

There are two engine and two crossfeed fuel valves.

plied to operate the engine FUEL VALVE WARNING

LIGHT by the DC essential bus through the ENGINE NO.

a. Engine Fuel Valves. One engine fuel valve (fig.

1 and NO. 2 FUEL SHUT-OFF circuit breakers on the No.

2-4-2) is in the fuel supply line to each engine. The valve

1 and No. 2 PDP.

is electrically operated by the FIRE PULL handles and

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

The following description on when the light will illuminate

498 and labeled FUEL VALVE # 1 ENGINE and FUEL

is for the XFEED switch. The same result applies to the

VALVE # 2 ENGINE.

engine fuel valves with the FIRE PULL handles.

b. Crossfeed Fuel Valves. The crossfeed fuel

(1)

Each time the XFEED switch is moved from

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

CLOSE to OPEN or OPEN to CLOSE. After this opera-

When the valve is opened, both engine fuel feed lines are

tion, the light should extinguish, indicating the crossfeed

interconnected and fuel can be supplied from both fuel

valve is synchronized with the switch position.

tanks to feed either engine or from either tank to feed

both engines. Fuel cannot be transferred between tanks.

(2)

When a short circuit occurs, causing a signal

The valves are electrically operated by the XFEED

to be applied opposite to the valve position. However, the

switch on the FUEL CONTR panel or manually by a lever

valve will remain at the position last selected by XFEED

on the valve. They are labeled FUEL VALVE CROSS-

switch.

FEED and located at station 504.

(3)

When the crossfeed valve protection relay

c. FUEL VALVE WARNING LIGHT. There are two

fails. The crossfeed valve will remain at the last selected

PRESS-TO-TEST FUEL VALVE WARNING LIGHT next

position and the valve can be operated electrically or

to each FUEL VALVE CROSS FEED

manually, as required.

2-4-2

TM 1-1520-240-10

2-4-6. Manual Defueling Valve.

A manual defueling valve is in the aft cargo compartment

next to FUEL VALVE # 2 ENGINE. The valve should only

be used by maintenance personnel to defuel the helicop-

ter or adjust fuel load.

2-4-7. Fuel Quantity Indicator and Selector.

An indicator calibrated to measure fuel quantity in

pounds and seven position selector switch (fig. 2-4-4) is

on the center instrument panel. Power is supplied to the

indicator through the FUEL QUANTITY selector switch

by the No. 1 AC bus through the FUEL QTY circuit break-

er on the No. 1 PDP.

a. FUEL QUANTITY Indicator. The indicator pro-

vides two types of display. One display is in the 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 selector switch

is selected. Then, the pointer will indicate fuel remaining

in that tank. The fuel quantity indicator is electrically con-

nected to 10 capacitance-type measuring units in the

tanks.

b. Fuel Quantity Selector Switch. The fuel quantity

selector switch has seven positions labeled TOTAL, L

Figure 2-4-4. Fuel Quantity Indicator and Selector

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

Switch

position other than TOTAL causes the indicator pointer

caution panel (fig. 2-14-5) of the center instrument con-

to display the fuel remaining in that tank. The digital read-

sole. Each light is electrically connected to a thermistor

out is not affected during individual tank readings.

sensor on a measuring unit in the respective main tank.

These lights are labeled L FUEL LOW an R FUEL LOW.

2-4-8. Fuel System Cautions.

When the is 20 percent of fuel remaining in the main tank,

Four caution capsules are dedicated to the fuel system.

the caution capsule for that main tank illuminates (20

percent of fuel is equal to 320 to 420 pounds.) Power for

a. L and R FUEL LOW. Two fuel quantity caution

these capsules is supplied by the DC essential bus

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

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 ar on the

master caution panel. Each caution capsule is electrical-

ly connected to a fuel pressure switch between the main

tank and the engine fuel valves. When one of these cap-

sules illuminates, it indicates that fuel pressure in the

respective fuel line is below 10 $ 1 psi. Fuel pressure is

measured after the fuel boost pumps and not at the en-

gine driven pump. When fuel pressure caution illumi-

nants, 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 CAUTION PNL circuit break-

er on No. 1 PDP.

2-4-9. FUEL Flow Indicators.

A dual fuel-flow indicator (fig. 2-4-5), on the center instru-

ment panel, indicates fuel flow to each engine in pounds

Figure 2-4-3. Fuel Valve Warning Light, Sta. 500

per hours. The indicator dial is graduated from 0 to 3,000

2-4-3

TM 1-1520-240-10

pounds per hour in 100 pound increments. The signal to

c. Pressure Refueling Manifold. The pressure re-

drive the indicator is derived from a fuel

fueling manifold connects to all tanks to the pressure

refueling receptacle. It does not include projectile resist-

ant features because the fuel is evacuated before flight

by the jet pumps.

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.

2-4-11. 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. 2-4-6).

2-4-12. PWR Control Switch.

The PWR (power) control 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

Figure 2-4-5. Fuel Flow Indicator.

ON light will illuminate, the fuel quantity indicator will

register the quantity of fuel in the tanks, and the REFUEL

flow transmitter in the fuel line of each engine at the quick

VALVE POSN lights will illuminate momentarily. When

disconnect shelf. Power to operate the No. 1 indicator is

placed to OFF, electrical power is removed.

from the No. 1 AC bus through the ENGINE NO. 1 FUEL

FLOW circuit breaker on the No. 1 DP. Power to operate

2-4-13. REFUEL STA Switch.

the No. 2 indicator is from the No. 2 AC bus through

ENGINE NO. 2 FUEL FLOW circuit breaker on the No.

The REFUEL STA switch is on the cockpit FUEL CONTR

2 PDP.

panel (fig. 2-4-1) when placed to ON, applies electrical

power from the DC switched battery bus to the PWR ON

2-4-10. Pressure Refueling System.

switch on the refueling station panel. Setting the switch

to OFF after pressure refueling, closes the refuel valves

The pressure refueling system permits rapid refueling of

and discontinues electrical power to the refueling panel.

all fuel tanks simultaneously or selective refueling of any

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

tank or combination of tanks. Maximum fueling rate is

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

300 gallons per minute at 55 psi. The system control

panel and refueling nozzle receptacle are on the right

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

side of the helicopter above the forward right landing

refuel station quantity indicator is inoperative, and there

gear (fig. 2-4-6).

is no precheck capability.

In addition to the control panel and refueling receptacle,

2-4-14. Fuel Quantity Indicator and Selector

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

Switch.

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

tank, and pressure refueling manifold.

The pressure refueling station fuel quantity indicator and

selector switch (fig. 2-4-4) are identical to those in the

a. Dual Fuel Level Control Valves and Dual Fuel

cockpit. The indicator at the refueling station indicates

Shutoff Valves. The dual fuel level control valves control

fuel quantity only when the REFUEL STA switch on the

the operation of the fuel fuel shutoff valves. When fuel in

cockpit FUEL CONTR panel (fig. 2-4-1) is at ON and the

a tank rises to the full level during pressure refueling, the

PWR switch on the refueling station panel (fig. 2-4-6) is

floats in the control valve close and apply a signal to the

at PWR ON. Electrical power to drive the indicator is AC

shutoff valve, closing it. the floats can also be closed

electrically to stop fuel flow into a tank at some intermedi-

from a solid-state inverter in the cabin at sta 220. The

ate level. The floats are controlled by the FUEL CELL

inverter, in turn, is powered by the DC switched battery

SHUTOFF VALVE TEST switches on the refueling con-

bus through the FUEL REFUEL circuit breaker on the No.

trol panel.

1 PDP.

b. Jet Pumps. The jet pump installed in each main

2-4-15. FUEL CELL SHUTOFF VALVE TEST

tank evacuates the refueling manifold and discharges

Switches.

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

activated when the forward boost pump in each main

Seven three-position FUEL CELL SHUTOFF VALVE

rank is first turned ON following pressure refueling.

TEST switches are on the refueling control panel (fig.

2-4-4

TM 1-1520-240-10

Figure 2-4-6. Pressure Refueling Station

2-4-6

TM 1-1520-240-10

SECTION V. FLIGHT CONTROLS

2-5-1. Flight Control System.

pedals by imputting equal but opposite lateral cyclic pitch

to the blades. Lateral control is obtained by application

The helicopter is controlled by changing the pitch of the

of equal lateral cyclic pitch to the blades with the cyclic

blades either collectively or cyclically. Pitch changes are

control stick. The helicopter is controlled longitudinally

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

with the cyclic stick through application of differential

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

collective pitch.

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

terconnected with the copilot’s controls.

In addition, the helicopter has an advanced flight control

system (AFCS). AFCS provides the following features:

Flight control movements are transmitted through a sys-

tem of bellcranks, push-pull tubes, and actuators to a

a. Rate damping in all axes and sideslip stability.

mixing unit just aft of the cockpit, next to the forward

b. Pitch and roll attitude hold and heading hold.

transmission. The control movements are mixed to give

the correct lateral cyclic and collective pitch motions to

c. Airspeed hold.

the rotors through dual hydraulic actuators. These dual

boost actuators are under each swashplate. Each set of

d. Improved control response in pitch, roll, and

dual boost actuators is normally powered by both flight

yaw.

control hydraulic systems.

e. Barometric and radar altitude hold.

The helicopter is vertically controlled with the THRUST

f.

Automatic coupled turns.

CONT lever through application of equal pitch to all bla-

des. Directional control is obtained with the directional

g. Longitudinal cyclic trim scheduling.

2

THRUST CONTROL LEVER

THRUST CONTROL LEVER

(ON HELICOPTER WITH

(ON HELICOPTER WITH

712 ENGINE

714A ENGINE

Figure 2-5-1. Thrust Control Lever

2-5-1

TM 1-1520-240-10

2-5-6. Directional Pedals.

The directional pedals (7 and 24, fig. 2-1-3) are used for

directional control of the helicopter during flight and while

taxiing with the forward gear off the ground.

When the right pedal is displaced forward, the forward

rotor disk tilts to the right and the aft rotor disk tilts to the

left. The opposite action occurs when the left pedal is

displaced forward. An ILCA is installed to assist the pilot

in moving the pedals.

The pedals are adjusted individually fore and aft by

pressing a lever mounted on the pedal support and mov-

ing the pedal to a new position before repositioning the

lever. Insure that both pedals are adjusted equally (left

and right pedals in same respective hole position) and

pedal adjustment lockpins are engaged. A balance

spring is installed to reduce control sensitivity.

2-5-7. Advanced Flight Control System. (AFCS)

a. The Advanced Flight Control System (AFCS) sta-

bilizes the helicopter about all axes and enhances control

response. It automatically maintains desired airspeed,

altitude, bank angle, and heading. An automatic turn fea-

ture, coupled to the pilot or copilot HSI (horizontal situa-

tion indicator) is also included in the AFCS.

b. Built In Test Equipment (BITE) is installed in each

AFCS computer. This equipment is intended for ground

troubleshooting purposes only. An interlock circuit

through the engine condition control box prevents BITE

use anytime either ECL is out of STOP.

Figure 2-5-2. Cyclic Stick Grip

c. Power is supplied to the HDG ENGAGED, BARO

ALT and RAD ALT ENGAGED lights from the DC essen-

tial bus through the CAUTION PNL circuit breaker on the

No. 1 PDP. The No. 1 AFCS receives AC and DC buses

2-5-5. AFCS Trim Switch.

respectively through the AFCS NO. 1 circuit breakers on

the No. 1 PDP. The No. 2 AFCS receives AC and DC

power from the No. 2 AC and DC buses respectively

NOTE

through the AFCS NO. 2 circuit breakers on the No. 2

PDP.

If the longitudinal CCDA fails, it can be recog-

nized by loss of pitch trim or failure of the

d. The AFCS consists of the following compo-

centering devise to release. A centering

nents:

spring in the pitch axis allows these forces to

be over-come.

(1)

A cockpit control panel.

(2)

Two AFCS computers in the avionics

The AFCS trim switch (fig. 2-5-2) is used to make small

compartment.

changes in the pitch (airspeed) and roll attitude while the

AFCS is operating. The switch is spring-loaded to center

(3)

Three ILCA’s in the flight control closet.

off position. Moving the switch forward or aft from center

(4)

A differential airspeed hold (DASH) actuator

off position commands an increase (forward) or de-

in the flight control closet.

crease (aft) in airspeed by driving a trim motor in the

longitudinal CCDA.

(5)

Two longitudinal cyclic trim (LCT) actuators

are installed, one in the forward upper controls, the other

Moving the switch left or right commands the roll ILCA to

in the aft upper controls.

bank the helicopter in the selected direction without mov-

ing the stick. Power is supplied to drive the pitch trim

(6)

Roll and yaw magnetic brakes, a longitudi-

motor from No. 1 AC bus through CLTV DRIVER ACTR

nal CCDA, and a thrust CCDA are all located in the flight

circuit breaker on the No. 1 PDP.

controls closet.

2-5-3

TM 1-1520-240-10

(1)

HDG Switch. The HDG (heading) switch is

FWD and AFT switches are used to extend or retract the

used in conjunction with the CMD SEL switch either HSI

appropriate cyclic trim actuator.

MODE SELECT panel and the heading bug on either HSI

to select coupled turns. The switch can be used only

(1)

AUTO and MANUAL switch. A two-position

when airspeed is above 40 knots. When the switch is

switch which is normally placed in AUTO.

pressed and either CMD SEL switch is pressed, the heli-

copter will automatically turn to and capture the heading

(a) AUTO Mode. In this mode, No. 1 AFCS

bug on the selected HSI. In addition, the ENGAGED

controls the forward actuator and the No. 2 AFCS con-

legend will illuminate. The switch is disengaged by press-

trols the aft actuator.

ing it again.

Heading intercept will be at standard rate of 3_ per sec-

(b) MANUAL Mode. In this mode, the ac-

ond up to a bank angle limit of 20_ at 133 knots. The

tuator can be controlled with separate FWD and AFT

helicopter must be trimmed before engaging the mode

actuator control switches, using the airspeed indicator

and cyclic stick control inputs should be avoided except

and CYCLIC TRIM indicators.

for longitudinal AFCS trim inputs to adjust airspeed.

(2)

FWD and AFT switched. Three- position

switched that can be placed in the EXT (extend) or RET

(retract) position. These switches are spring-loaded to

the center off position. If the cyclic trim actuators fail to

extend or retract as indicated on the CYCLIC TRIM indi-

cators, MANUAL mode can be selected.

2-5-16. Cyclic Trim Indicators.

WARNING

If the longitudinal cyclic trim actuators fail

at the full retract position or are manually

Figure 2-5-3. Advanced Flight Control System

selected to the full retract position, do not

Panel

exceed the airspeed limitations shown in

fig. 5-7-1.

Heading select is disengaged if either CENTERING DE-

VICE RELEASE switch is pressed.

The FWD and AFT CYC (cyclic) TRIM indicators (fig.

(2)

BARO ALT and RAD ALT Switches. The

2-5-4) are on the center instrument panel. The indicators

BARO ALT and RAD ALT are used to select altitude hold

are labeled 60 RET, GND, 150 EXT. The indicators dis-

mode. An interlock prevents both switches from being

play position of the forward and aft LCT actuators relative

engaged at the same time. When pressed, the EN-

to airspeed. During ground operations, the pointer will be

GAGED legend will illuminate. RAD ALT hold is used

at GND to indicate activation of the landing gear proximi-

ty switches.

below 1,500 feet AGL. BARO ALT hold is used in forward

flight to maintain a constant cruise altitude or may be

used in HOGE.

2-5-17. AFCS OFF Caution.

b. SYSTEM SEL Switch. The SYSTEM SEL switch

is a five position rotary switch labeled OFF, 1 BOTH, 2

Two AFCS OFF caution capsules are on the master cau-

OFF. Normally, the switch is at BOTH. In this position,

tion/advisory panel

( 712 ,fig.

2-14-5,

714A , fig.

both AFCS are operating at one-half gain. If one system

2-14-6). They are labeled 712 NO. 1 AFCS OFF and

should fail, the good system is selected and that system

NO. 2 AFCS OFF, 714A AFCS 1 and AFCS 2. These

operates at 3/4 gain. At OFF, both systems are inopera-

cautions will illuminate when the associated AFCS is

tive except for CYCLIC TRIM.

manually shutoff or has failed or the associated DASH is

c. CYCLIC TRIM Switches. The AUTO and MANU-

in a low rate condition. Refer to Chapter 8 AFCS Off Flight

AL switch selects the mode of cyclic trim operation. The

Characteristics.

2-5-5

TM 1-1520-240-10

2-5-18. Command Select Switch.

The CMD SEL switch is on the pilot and copilot HSI

MODE SELECT panel (Chapter 3). The switches are

used to select the HSI which will provide the refenced

heading when the HDG switch is engaged. Only one

CMD SEL switch may be selected at a time. If the other

CMD SEL switch is selected during heading select op-

erations, the HDG switch on the AFCS panel will disen-

gage and heading select will be disabled until the HDG

switch is again pressed. When selected, the SEL legend

on the switch illuminates.

Figure 2-5-4. Forward and Aft Cyclic Trim

Indicators

2-5-6

TM 1-1520-240-10

SECTION VI. HYDRAULIC SYSTEMS

2-6-1. Hydraulic Power Supply System.

open, and No. 2 system is pressurized. Simultaneously,

No. 1 solenoid valve closes and No. 1 system is turned

The hydraulic power supply system consists of three

off.

separate systems. They are the No. 1 flight control sy-

stem, No. 2 flight control system, and a utility system.

2-6-4. Utility Hydraulic System.

Each system includes a variable delivery pump and res-

ervoir cooler. In addition, each flight control system has

The utility hydraulic system supplies hydraulic power to

a power control module, and the utility system has a

the wheel brakes, power steering actuator, swivel locks,

pressure control module. Each flight control system is

centering cams, ramp actuating cylinders, hydraulic car-

connected to the utility system by a Power Transfer Unit

go door motor, actuator for the center cargo hook, cargo/

(PTU). All systems are serviced by a common fill module

rescue winch control valve, two engine starters. PTU;s

and are pressurized to prevent pump cavitation.

and APU start circuit. When the APU is running, the utility

hydraulic system is pressurized by an APU driven pump.

2-6-2. Flight Control Systems.

When the APU is not running and the rotors are turning,

The No. 1 and No. 2 flight control systems are identical.

the utility hydraulic system is pressurized by an aft trans-

they ar parallel in operation, hydraulically separated, and

mission driven pump.

electrically integrated. The flight control system operate

The utility hydraulic system incorporates a pressure con-

at approximately 3,000 psi, which is reduced to 1,500 psi

trol module which isolates utility subsystems from each

for ILCA operation. They power four upper dual boost

other. When a failure occurs in one utility hydraulic sub-

actuators (3,000 psi) and four ILCAs (1,500 psi). Each

flight control system powers one piston of each actuator.

system, the remaining subsystems continue to operate

normally if the BRK STEER and RAMP PWR switches in

No. 1 flight control system is pressurized by a pump on

the cockpit are set to OFF.

the forward transmission. No. 2 system is pressurized by

a pump on the aft transmission. The power control mod-

The APU starting subsystem of the utility hydraulic sys-

ules consist of pressure-line and return-line filters. No.1

tem includes three accumulators which accelerate the

system power control module is in the forward pylon. No.

APU to start, maintain reservoir pressure throughout the

2 system power control module is in the aft pylon. The

start cycle, and control operation of the APU motor pump.

accumulators dampen low frequency pressure surges

The APU starting subsystem also includes a two-stage

and provide stored hydraulic power for peak loads.

hand pump for charging the APU start accumulators. The

APU is normally recharged by the APU motor-pump after

The PTU in each system allows ground checkout of the

the APU is started. An additional accumulator in the

flight control systems with the rotors stopped. Each PTU

brake system provides for limited brake operation in the

consists of a pump driven by a hydraulic motor which is

event of utility hydraulic system failure. the steering sys-

pressurized by the utility hydraulic system. The PTU’s

are controlled by the PWR XFER 1 and 2 switches on the

tem also has an accumulator to keep the swivel locks

HYD panel in the overhead switch panel.

engaged when the BRK STEER switch is OFF.

Normal operating pressure range for the utility hydraulic

2-6-3. FLT CONTR Switch.

system is 2500 to 3500 psi. During APU operation pres-

The FLT CONTR (flight control) switch is located on the

sure is increased to approximately 3350 psi for engine

HYD panel in the overhead switch panel (fig. 2-6-1). It is

starting. (See table 2-15-1 for flight control and utility

a three-position center locked switch labeled 2 ON,

hydraulic system capacities and fig. 2-15-3 for accumula-

BOTH, and 1 ON. this switch can be used to turn off one

tor precharge pressures.)

of the flight control systems, provided the other one is

operating. Turning off one of the flight control hydraulic

systems disables the corresponding AFCS and causes

the remaining AFCS to make full corrections. In addition,

the respective AFCS OFF and HYD FLT CONTR caution

capsules will illuminate. The FLT CONTR switch shall be

set to BOTH during all flight conditions.

At BOTH , both solenoid valves are deenergized open

and both flight control systems are pressurized. When

the FLT CONTR switch is set to 1 ON, the two-way sole-

noid valve on No. 2 power control module is energized

closed. This causes No. 2 pressure-operated valve to

close, depressurizing No. 2 system. When the FLT

CONTR switch is moved to 2 ON, the two-way solenoid

valve on No. 2 power control module is deenergized

Figure 2-6-1. Hydraulics Control Panel

2-6-1

TM 1-1520-240-10

No. 2 engine transmission oil reservoir on the combining

may not immediately indicate a transmission problem.

(mix) transmission through the right pump assembly, fil-

Loss of oil or low oil pressure may not be accompanied

ter, cooler, jet protection screen, and to the jets which

by high oil temperature indication.

spray the oil onto the various gears and bearings. One of

the scavenge elements of the right pump assembly re-

turns the oil from the No. 2 engine transmission sump

through a debris indicating screen and back to the No. 2

engine transmission oil reservoir. Engine transmissions

do not have auxiliary lubrication systems.

All No. 1 and No. 2 engine transmission lubrication sys-

tem components are on the combining (mix) transmis-

sion except the jet protection screen and jets. Separate

oil jets are utilized for each transmission lubrication oil

system. The individual oil cooler for the combining (mix)

and both engine transmissions are mounted on the com-

bining (mix) transmission and utilize a common transmis-

sion driven fan for cooling air.

2-7-6. Transmission Main Oil Pressure Indicator.

A transmission main oil pressure indicator is located on

the center instrument panel (fig. 2-7-1). It indicates either

the lowest main oil pressure in any one of the transmis-

sions or only the oil pressure in the transmission selected

by the pilot. The indicator is electrically connected to

each transmission. In addition, each transmission and

the aft rotor shaft bearing has a separate low pressure

switch. These switches are connected to the XMSN OIL

PRESS caution capsule on the master caution and the

TRANSMISSION MAIN OIL PRESS indicating lights on

the MAINTENANCE PANEL (fig. 2-10-1). Power to oper-

ate the indicator is supplied by the No. 1 AC bus through

the XMSN OIL PRESS circuit breaker on the No. 1 PDP.

2-7-7. Transmission Main Oil Pressure Selector

Switch.

A transmission oil pressure selector switch is located on

the center instrument panel (fig. 2-7-1). The switch posi-

tions are labeled TEST, SCAN, FWD, AFT, MIX, LEFT,

Figure 2-7-1. Transmission Main Oil Pressure

and RT. When the switch is set to TEST, the pointer on

Indicator and Selector Switch

the transmission pressure indicator will drop to zero or

below. When the switch is set to SCAN, the lowest main

oil pressure among all the transmission will be indicated.

Each temperature probe incorporates a high oil tempera-

The remaining positions are used to select a particular

ture switch which is independent of the temperature indi-

transmission oil pressure indication. When selecting a

cator and is triggered at 140_C, lighting the XMSN OIL

particular switch position, be sure the switch is in detent.

HOT caution capsule on the master caution panel and to

If the switch is not in detent, the pressure gage will indi-

the TRANSMISSION OVERTEMP magnetic indicator on

cate zero.

the MAINTENANCE PANEL. Power to operate the indi-

2-7-8. Transmission Main Oil Temperature Indica-

cator is supplied by the No. 1 AC bus through the XMSN

tor.

OIL TEMP circuit breaker on the No. 1 PDP.

A transmission oil temperature indicator is located on the

2-7-9. Transmission Main Oil Temperature Selector

center instrument panel (fig. 2-7-2). It reads from -70_ to

Switch

+150_C. It indicates the highest oil temperature among

all the transmissions or only the oil temperature of the

A transmission oil temperature selector switch is on the

selected transmission. A temperature probe is located in

center panel below the transmission oil temperature indi-

the forward and aft transmission sumps and in each

cator (fig. 2-7-2). The switch positions are labeled TEST,

compartment of a three-compartment oil tank for the

SCAN, FWD, AFT, MIX, LEFT, and RT. When the switch

combining (mix) transmission and in each engine trans-

is set to TEST, the pointer on the transmission oil temper-

mission. The temperature probes in the three tank

ature indicator deflects full scale toward low temperature.

compartments measure oil temperature in the tank and

.When the switch is set to SCAN, the highest oil tempera-

2-7-2

TM 1-1520-240-10

ture among all transmissions is indicated. The remaining

a. XMSN OIL HOT Caution. It illuminates when the

positions are used for selecting a particular transmission

main oil temperature in the sump of the forward, aft and

fied by the oil temperature selector switch oil temperature

reservoir of the combining (mix) or either engine trans-

indication. When selecting a particular switch position,

mission exceeds 140_C. The hot transmission is identi-

be sure the switch is in detent. If the switch is not detent,

fied by the oil temperature selector switch and indicator

the oil temperature indicator will indicate -70_C.

and the TRANSMISSION OVERTEMP magnetic indica-

tors on the MAINTENANCE PANEL.

b. XMSN OIL PRESS Caution. It illuminates when

main oil pressure drops below 20 psi in any transmission

or aft rotor shaft pressure drop below 10 psi. The low-

pressure system is identified by the transmission oil pres-

sure selector switch and indictors on the center instru-

ment panel and the TRANSMISSION MAIN OIL PRESS

indicating lights on the MAINTENANCE PANEL. If the

XMSN OIL PRESS caution capsule illuminates and the

affected transmission cannot be determined using the

selector switch, the condition may be caused by loss of

aft rotor shaft oil pressure. Low oil pressure at the aft rotor

shaft is indicated by the illumination of the TRANSMIS-

SION AFT SHAFT MAIN OIL PRESS indicating light on

the MAINTENANCE PANEL.

c. XMSN AUX OIL PRESS Caution. It is activated

by individual aux oil switches and illuminates when auxil-

iary oil pressure drops below 20 psi in the fwd or aft

transmission and 10 psi in the combining (mix) transmis-

sion. The transmission with the low pressure is identified

by a lit TRANSMISSION AUX OIL PRESS indicating light

on the MAINTENANCE PANEL.

d.

712 NO. 1 or NO. 2 ENG XMSN HOT, 714A

ENG 1 XMSN HOT or ENG 2 XMSN HOT Caution. They

illuminate if oil temperature in either engine transmission

exceeds about 190_C. The capsules are activated by a

thermoswitch in each engine transmission. The thermo-

switch monitors oil temperature in the transmission, not

in the reservoir. It is part of a chip detector and tempera-

ture assembly in each engine transmission.

2-7-11. Transmission Chip Detectors.

Figure 2-7-2. Transmission Main Oil Temperature

Chip detectors are installed in all transmission and aft

Switch and Indicator

rotor shaft thrust bearing lubrication systems. All trans-

mission chip detectors, except those in the engine trans-

mission, are connected to the XMSN CHIP DET caution

capsule on the master caution panel. Engine transmis-

2-7-10. Transmission Oil Cautions.

sion chip detectors are connected to the corresponding

712 NO. 1 or NO. 2 ENG CHIP DET, 714A ENG 1 or

Five transmission oil caution capsules are on the master

ENG 2 CHIP DET caution capsules.

caution panel. The capsules are labeled XMSN OIL HOT,

XMSN OIL PRESS, XMSN AUX OIL PRESS, and

All transmissions and aft rotor shaft chip detectors are

712 NO. 1 AND NO. 2 ENG XMSN HOT, 714A ENG 1

also connected to the TRANSMISSION CHIP DETEC-

or ENG 2 XMSN HOT. These cautions, in conjunction

TOR magnetic indicators on the MAINTENANCE PAN-

with the transmission oil pressure and temperature indi-

EL. When a chip detector is bridged by ferrous particles,

cators on the center instrument panel and the TRANS-

the XMSN CHIP DET, or the 712 NO. 1 and/or NO. 2

MISSION OVERTEMP magnetic indicators, MAIN OIL

ENG CHIP DET, 714A ENG 1 and/or ENG 2 CHIP DET

PRESS, and AUX OIL PRESS indicating lights on the

caution capsule illuminates. At the same time, the corre-

MAINTENANCE PANEL, alert the crew to impending

sponding TRANSMISSION CHIP DETECTOR indicator

transmission lubrication problems. The cautions operate

on the MAINTENANCE PANEL will trip and change from

independently of the pressure and temperature indica-

an all-black indication to a black-and-white indication,

tors on the center instrument panel.

identifying the transmission.

2-7-3

TM 1-1520-240-10

SECTION IX. UTILITY SYSTEMS

2-9-1. Anti Icing Systems.

CAUTION

Anti icing is provided for the pitot tubes, AFCS yaw ports,

and pilot and copilot windshields. The center windshield is

If windshield bubbling or delamination oc-

not anti-iced, it is only defog.

curs around the sensor element, immedi-

ately place switch to OFF for that wind-

shield.

2-9-2. ANTI ICE Panel.

When any switch is moved to ON, current flows to the

The ANTI ICE panel is located on the overhead switch

associated temperature controller and then to the wind-

panel (fig. 2-9-1). It has three two-position W/S (wind-

shield. As the temperature of the windshield rises to a

shield) switches labeled CPLT, CTR, and PLT. The

preset value (about 44_), as sensed by the sensor ele-

switches positions are OFF and ON. In addition, a two-

ment, the electrical current to the windshield is inter-

position PITOT heat switch is in this panel. The switch

rupted by the temperature control relay. Once the wind-

positions OFF and ON.

shield has cooled sufficiently, electrical current is

reapplied. This causes a cycling effect which maintains

Power for the pilot and center windshields is from the No.

windshield temperature within operating limits.

2 AC bus through the WSHLD ANTI ICE HEAT PILOT

and CTR circuit breakers. Power for the copilot wind-

Operating temperature is on in less than 1 minute after

shield is from the No. 1 AC bus through the WSHLD

the switch is placed to on. When the switch is placed to

COPLT HEAT circuit breaker on the No. 1 PDP. Anti-ice

OFF, the anti-icing system is deenergized.

control for the pilot and center windshield is from the

28-volt No. 2 DC bus through the WSHLD ANTI ICE

b. PITOT Heat Switch. Heating elements prevent

CONT CTR and PILOT circuit breakers on the No. 2 PDP.

ice accumulation in the pitot tubes and the yaw ports.

Anti-ice control for the copilot windshield is from the

When the PITOT switch is placed to ON, power to the

28-volt No. 1 DC bus through the WSHLD COPLT CONT

heater elements in the pitot tubes and yaw ports is ap-

circuit breaker on the No. 1 PDP. Power to operate the

plied. When the switch is placed to OFF, the heating

heater elements in the pitot tubes and yaw ports is sup-

elements are deenergized.

plied by the No. 2 AC bus through the PITOT HEAT and

YAW PORT HEAT circuit breakers on the No. 2 PDP.

2-9-3. MAINTENANCE PANEL.

a. W/S Switches. The pilot and copilot windshields

The MAINTENANCE PANEL is on the right side of the

are anti-iced and defogged electrically. The center wind-

cabin above the ramp (fig. 2-9-2). The panel is provided

shield is defogged but not anti-iced. The laminated wind-

to assist in the identification of system malfunction or

shield panels are heated electrically by current which

condition that may require servicing or other mainte-

passes through a transparent conductive coating em-

nance. The panel is divided into four sections. They are

bedded between the layers.

labeled TRANSMISSION, HYDRAULICS, ENGINE, and

GROUND CONTACT.

2-9-4. TRANSMISSION Section.

This section monitors the FWD, COMB, AFT, AFT

SHAFT, LEFT, and RIGHT transmissions. It consists of

six CHIP DETECTOR magnetic indicators, six DEBRIS

SCREEN a magnetic indicators, six MAIN OIL PRESS

indicating PRESS-TO-TEST lights, three AUX OIL

PRESS indicating PRESS-TO-TEST lights, and five

OVERTEMP magnetic indicators. Power to operate the

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

HYDRAULIC MAINT PNL circuit breaker on the No. 1

PDP.

a. CHIP DETECTOR Magnetic Indicators. When

the corresponding CHIP DETECTOR is bridged by fer-

rous particles, the associated chip detector indicator

changes from all-black to black-and-white. In addition,

the XMSN CHIP DET or ENG CHIP DET caution capsule

Figure 2-9-1. Anti Ice Panel

illuminates on the master caution panel.

2-9-1

TM 1-1520-240-10

LEFT

RIGHT

Figure 2-9-2. Maintenance Panel

b. DEBRIS SCREEN Magnetic Indicators. There is

d. AUX OIL PRESS indicating Lights. If auxiliary oil

one indicator each for the FWD transmission, AFT trans-

pressure drops below 20 psi in the FWD or AFT trans-

mission, and the Left (No. 1) and Right (No. 2) engine

mission or 10 psi in the COMB transmission, the corre-

transmissions. There are two indicators for the COMB

sponding indicating light will illuminate. In addition, the

transmission. one indicator for the left sump and one indi-

XMSN AUX OIL PRESS caution will illuminate on the

cator for the right sump.

master caution panel.

e. OVERTEMP Magnetic Indicators. Each OVER

NOTE

TEMP magnetic indicator is electrically connected to a

Their is no cockpit indication of a latched DE-

temperature probe in the reservoir of each transmission.

BRIS SCREEN magnetic indicator. If a DE-

If oil temperature in the transmission reservoir exceeds

BRIS SCREEN magnetic indicator latches,

140_C, a switch closes. When the switch closes, the

the flight engineer shall advise the pilot imme-

XMSN OIL HOT caution illuminates on the master cau-

diately.

tion panel and trips the corresponding OVERTEMP mag-

netic indicator on the MAINTENANCE PANEL, thus iden-

The indicators are electrically connected to screens in

tifying the hot transmission.

the sumps of each transmission. If the screen mesh is

bridged with conductive particles, the indicating circuit

closes and trips the corresponding DEBRIS SCREEN

2-9-5. HYDRAULICS Section.

magnetic indicator on the MAINTENANCE PANEL.

This section monitors the FLT CONT NO 1, FLT CONT

NO 2, and UTILITY hydraulic systems. It consists of three

c. MAIN OIL PRESS indicating Lights. If main oil

PRESSURE indicators, three fluid TEMPERATURE indi-

pressure drops below 20 psi in any transmission or 10

cators, two RESERVOIR LEVEL indicators, six FILTER

psi in the aft shaft bearing, the corresponding indicating

CHANGE indicating PRESS-TO-TEST lights, and four

light will illuminate. in addition, the XMSN OIL PRESS

PUMP FAULT indicating PRESS-TO-TEST lights. Power

caution will illuminate on the master caution panel.

to operate the indicators is supplied by the No. 2 DC bus

2-9-2

TM 1-1520-240-10

SECTION X. HEATING, VENTILATION, COOLING, AND ENVIRONMENTAL

CONTROL SYSTEMS

2-10-1. Heating and Ventilating System.

A 200,000 btu/hr capacity internal combustion heating

system is provided. it consists of a heater unit, a fuel

control unit, an ignition assembly, a blower, control re-

lays, and air pressure and temperature control circuits.

Ducting carries heated air or ventilating air to the cockpit

and the cabin. The heater consumes approximately 15

pounds of fuel per hour from the right main fuel tank.

The heater and blower are mounted vertically on the right

side of the helicopter, immediately aft of the forward cab-

in section bulkhead. Air for the system is provided by the

blower which draws air from an inlet on the forward upper

side of the fuselage. If sufficient air is not available for

proper heater operation, an automatic differential pres-

sure switch in the heater circuit will stop the heater.

Both ventilating and combustion air enters the heater

inlet. The heating air passes over the heated metal walls

of the combustion chamber and is directed to a network

of ducting. The air entering the combustion chamber is

combined with atomized fuel and, after combustion that

heats the metal walls, the exhaust is discharged through

an outlet on the forward upper side of the fuselage. Pow-

er to operate the blower is supplied by the No. 2 AC bus

through the CABIN HEATER BLOWER circuit breaker on

the No. 2 PDP. Power to the rest of the system is supplied

by the No. 2 DC bus through the CABIN HEATER CONT

circuit breaker on the No. 2 PDP.

Figure 2-10-1. HTG Panel

2-10-2. HTG Panel.

The circuit to the thermostat heater winding is also inter-

The HTG (heating) panel (fig. 2-10-1) is located on the

rupted, allowing the winding to cool and the mercury

overhead switch panel (fig. 2-1-10). It consists of a rheo-

column to contract, thus reenergizing the temperature

stat-type CABIN TEMP SEL rotary switch, a three-posi-

controller relay. this creates a cycling effect, the rate of

tion heater function switch, and a spring-loaded push-

which can be varied by increasing or decreasing the

buttton HTR START switch.

resistance between the temperature selector and he

thermostat heating winding. Resistance is varied by turn-

a. CABIN TEMP SEL Rotary Switch. The CABIN

ing the CABIN TEMP SEL rotary switch. this increase or

TEMP SEL rotary switch is labeled COOL and WARM.

decrease in resistance directly varies the time the heater

This switch operates in conjunction with the temperature

is allowed to operate before being automatically cycled.

controller relay in the heater circuit and with a cabin ther-

b. Heater Function Switch. The heater function

mostat. One set contacts on the temperature controller

switch is labeled BLWR ONLY , OFF, and HTR ON. The

relay closes to complete a circuit to the fuel control sole-

switch selects the desired feature of the heating and

noid valve. This allows fuel to be delivered to the heater.

ventilating system. When the switch is set to BLWR

ONLY, the blower forces unheated air into both the cock-

The second set of contacts on the temperature controller

pit and cabin. Further movement of the heater controls is

relay closes to complete the circuit to the heater windings

not required. Selecting HTR ON energizes the various

in the cabin thermostat. The heater windings heat a col-

units of the heater once the HTR START switch is pres-

umn of mercury in the thermostat, causing it to rise.

sed. The heating and ventilating system is shut down

When the mercury column reaches a 34_C contact, the

when the switch is set to OFF.

temperature control relay is shunted, causing its contacts

to open and interrupt the circuit to the fuel control sole-

c. HTR START Switch. When HTR ON is selected

noid valve. This stops heater operation by shutting off the

on the heater function switch and the HTR START switch

fuel supply to the heater.

is pressed, the heater control circuits are energized. The

2-10-1

TM 1-1520-240-10

SECTION XI. ELECTRICAL POWER SUPPLY AND DISTRIBUTION SYSTEMS

2-11-1. Electrical Power Supply System.

2-11-2. AC System.

Alternating current (AC) is the primary source of power

The AC system supplies

115/200-volt three-phase

to operate the electrical and electronic equipment. Three

400-Hz power from No. 1 AC generator to No. 1 three-

AC generators, two driven by the aft transmission and

phase AC bus and from No. 2 AC generator to the No. 2

one driven by the APU, produce 115/200-volt 3-phase

three-phase AC bus (fig. FO-3 and FO-6). The AC equip-

400-Hz power. The system develops 28-volt DC through

ment is powered by these buses. Some of the equipment

two transformers rectifiers (RECT) one each in the for-

is operated by 115-volt single-phased AC and some

ward section of the left and right fuselage pods. DC is

equipment by 26-volt AC power supplied through the

also supplied by a 24-volt nickel-cadmium battery.

transformers.

The AC system is protected from overvoltage, undervol-

Both 115/200-volt 3-phase AC and 28-volt DC can be

tage, and underfrequency conditions by generator con-

supplied by operating the APU or by connecting an AC

trol units. The generators will be disconnected from the

external power source to the external power receptacles

AC buses any time the RRPM drops below 82 to 85

(fig. 2-11-1). If the APU is running or AC external power

percent for more than 3 to 7 seconds. The AC power

is connected, DC power is supplied by the helicopter

distribution system has four power sources, a contactor,

transformer rectifiers (RECT). If only DC external power

control circuit, an AC power transfer circuit, and two AC

is supplied, AC power is not available on the helicopter.

buses.

Circuits are protected by circuit breakers (fig. FO-2 and

FO-5). The electrical load is divided between the two AC

The No. 1 and No. 2 generator power sources are two

generators (fig. FO-3 and FO-6). Should one generator

main generators driven directly by the aft transmission.

fail, the other will automatically take over the entire load.

The APU generators is driven directly by the APU. The

When APU is running, its single generator powers the

external power source is an AC power supply connected

entire load.

to the helicopter.

No. 1 and No. 2 generators feed their respective buses.

If No. 1 and No. 2 generator fails (or are shut down), the

failed generator is isolated from its bus and the operating

generator feeds both buses. When No. 1 or No. 2 or both

generators are operating, APU generator and external

power are blocked from the AC buses.

When the APU generator is operating and the main gen-

erators are shut down (or rotors turning below about

84%) or switched off, the APU generator feeds both bu-

ses. When the APU generator is operating, external pow-

er is blocked from the AC buses. When external power

is applied to the helicopter (GEN APU, GEN 1, and GEN

2 are OFF), the external power source feeds both buses.

The Generator Control Unit (GCU) also provides genera-

tor feeder fault protection. If a fault occurs between the

feeder and the airframe, the GCU will disable the genera-

tors. This prevents structural damage to the airframe

when a ground fault occurs.

The Permanent Magnet Generator (PMG) section within

the generator is used to power the main contactors (re-

lays) in the distribution system. A pickoff coil within the

PMG provides an RPM signal for the rotor tachometer

indicators. This tachometer signal is available whenever

the rotor are turning.

2-11-3. Generator Control Switches.

The generator control switches are located on the ELEC

panel of the overhead switch panel (fig. 2-11-2). The

three switches are labeled GEN 1, GEN 2, and GEN

APU. The switch positions are TEST, OFF RESET, and

Figure 2-11-1. External Power Receptacles

ON.

2-11-1

TM 1-1520-240-10

When the switches are ON, the respective main relay

essential bus through the LIGHTING CAUTION PNL cir-

operates, which energizes and connects the generator to

cuit breaker on the No. 1 PDP.

the buses. At OFF RESET, the generator is deenergized

and disconnected from the bus. This position is also used

2-11-6. DC System.

to reset a generator. The TEST position is provided to

allow the generator to be energized but disconnected

The direct current (DC) power supply system supplies

from the bus to determine whether the AC produced is of

28-volt DC from the No. 1 transformer rectifier (RECT) to

proper frequency and voltage, except the APU genera-

No. 1 DC bus and the No. 2 RECT supplies power to the

tor.

No. 2 DC bus (fig. FO-4 and FO-7).RECT convert 200

VAC power to 28-volt DC power for use in the DC distribu-

tion system.

Cooling air for the RECTS is obtained from within the

cabin.The air inlets are located at sta. 176 on the left and

right side of the cabin behind the troop seats.If the inlets

are blocked, the RECTS will overheat.

A bus-tie relay is between No. 1 and No. 2 DC buses.If

either RECT fails, the respective RECT failure relay op-

erates and the bus-tie relay closes automatically to con-

nect the unpowered bus to operating RECT.In addition to

No. 1 and No. 2 DC buses, the DC system includes an

essential bus, a switched battery bus, and a hot battery

Figure 2-11-2. Electrical Power Panel

bus.

2-11-4. GEN OFF Cautions.

During normal operation, the essential bus and the

switched battery bus are energized by No. 1 DC bus.If

Two generator caution capsules labeled

712 NO. 1

both DC buses fail or if NO. 1 DC bus fails and does not

GEN OFF and NO. 2 GEN OFF, 714A GEN 1 AND GEN

bus-tie to No. 2 DC bus, the essential bus, the switched

2 are on the master caution/advisory panel ( 712 fig.

battery bus, and the hot battery bus will be energized by

2-14-5 714A fig. 2-14-6).These caution capsules illumi-

the battery as long as the BATT switch is ON.These

nate whenever the generators are in operative.The cap-

buses provide power to emergency, ground mainte-

sules are controlled by the main generator contactors

nance, and communications components.The hot bat-

when the generator control switches are in either ON or

tery bus and switched battery bus are energized as long

OFF RESET.In TEST, the capsules are controlled by the

as the battery is connected.The hydraulic reservoir level

generator control switch and will extinguish if generator

indicators and the emergency APU control circuits and

output has the proper frequency and voltage.Power to

cabin and maintenance lights are on these buses.

operate the generator capsules is supplied by the DC

essential bus through the LIGHTING CAUTION PNL cir-

The 24-volt nickle cadmium battery is located in the left

cuit breakers on the No. 1 PDP.

forward electrical compartment.The battery capacity is

11 ampere-hours.A battery charger is connected to the

2-11-5. EXT PWR Caution.

battery.The battery charger receives power from No. 1

AC bus, rectifies the AC and applies the DC to the battery

to maintain a charge on the battery.

CAUTION

Sensors in the battery charger detect battery or battery

charger overtemperature, short or open circuits or cell

When external power is used, a visual

imbalance.If any of these conditions occur, the battery

check shall be made by the crew to ensure

charger will stop functioning and activates the BATT SYS

that the external power unit has been dis-

MAL caution capsule on the master caution panel.

connected from the helicopter before taxi-

ing.

External DC power is supplied to the DC buses of the

An external power caution capsule labeled EXT PWR is

helicopter by connecting the external DC power source

on the master caution panel (fig. 2-14-5 and fig.

to the DC external power receptacle (fig. 2-11-1).Applica-

2-14-6).This capsule illuminates and remains illuminated

tion of external power operates the DC external power

whenever external power is connected.The light is con-

relay which connects the power source to No. 1 DC

trolled by the AC external power contactor and the DC

bus.No. 2 bus is energized when the bus tie relay opera-

power relay.The capsule extinguishes when the genera-

tes.If the polarity of the external power is reversed, a

tors are supplying current to the buses.Power to operate

blocking diode in the circuit prevents the external power

the external power caution capsule is supplied by the DC

relay from closing.

2-11-2

TM 1-1520-240-10

SECTION XII. AUXILIARY POWER UNIT

2-12-1. General.

2-12-4. APU ON Caution.

The gas turbine auxiliary power unit T62-T-2B (APU) (fig.

The APU ON caution capsule is on the mater caution

2-12-2) is mounted in the aft cabin above the ramp. The

panel (fig. 2-14-5 and 2-14-6). Normally, the APU is in-

basic components of the APU are the gas turbine engine,

tended for ground operation only. It is not intended for

hydraulic motor-pump, fuel control, accessory drive, and

operation during flight. If the caution remains illuminated

AC generator. An APU Electronic Sequencing Unit

following take-off, it alerts the pilot to shut down the APU.

(ESU) which monitors APU operation is on the left side

When the caution is illuminated it indicates the APU is up

of the cabin above the ramp. the ESU is also labeled APU

to speed and the exhaust gas temperature is normal. It

CONTROL BOX.

does not necessarily indicate that APU hydraulic pump

or generator output is normal. If rotors are not turning,

The motor-pump on the APU pressurizes the utility and

check the UTIL HYD SYS and RECT OFF cautions to

hydraulic system for main engine starting and ground

evaluate output of the APU hydraulic pump and genera-

checks. The APU also drives an AC generator which

tor. The APU ON caution is controlled by the ESU.

supplies power to the No. 1 and No. 2 electrical systems.

Refer to Section VI for further information on the hydrau-

2-12-5. Emergency APU Fluid Shut Off Vavle.

lic systems. The APU oil supply is intergal and contained

The EMERGENCY APU FLUID SHUT OFF VALVE is in

within the sump of the accessory drive assembly. The

the fuel supply line to the APU (fig. 2-12-1). It is located

APU receives fuel from the left main fuel tank through a

inside the aft cabin above and to the left of the ramp

booster pump, a manual fuel shutoff valve, and a sole-

interphone station. The valve can also be reached from

noid valve.

the outside through an access door labeled ACCESS

APU EMER FLUID SHUT OFF. The knob on the valve

2-12-2. Electronic Sequencing Unit.

has an OPEN and CLOSE position. Placing the knob to

CLOSE shuts off fuel to the APU.

The ESU is mounted on the left side of the cabin above

the ramp. The unit monitors APU starting and operation.

in addition, it monitors APU speed and exhaust gas tem-

perature. the unit continuously compares these parame-

ters with limits programmed into ESU circuits. If a limit is

exceeded, the ESU will automatically shut down the

APU.

NOTE

The BITE indicators indicate engine condition

only. They will not indicate a defective hydrau-

lic motor-pump or generator.

Four magnetic built-in-test-equipment (BITE) indicators

are on the ESU. These indicators are either black or

white. A label on the ESU explains the various BITE

indications and their meaning.

2-12-3. APU Switch.

The APU switch is on the ELEC panel of the overhead

switch panel (fig. 2-11-2). It is a three-position switch

labeled OFF, RUN, and START. The switch is spring

loaded from START to RUN. Normally, power to operate

the APU is supplied by the DC essential bus through the

APU CONT NORM circuit breaker on the No. 1 PDP.

Emergency power to operate the APU is from the battery

bus through the APU CONT EMERG circuit breaker on

Figure 2-12-1. Emergency APU Fluid Shut OFF

the No. 1 PDP.

Valve

2-12-1

TM 1-1520-240-10

Figure 2-12-2. Auxiliary Power Unit

2-12-2

TM 1-1520-240-10

SECTION XIII. LIGHTING

2-13-1. Position Lights.

NOTE

The crew chief must inform the pilot when the

Three position lights (1, 4, and 6, fig. 2-13-1) are installed

AFT POS LIGHT switch has been changed to

on the helicopter. On the right side of the fuselage is a

the OFF position.

green light (1); on the left, red (6); and on the vertical

panel of the aft pylon, white (4). Power to operate the

b. AFT POS LIGHT Switch. The AFT POS LIGHT

position lights is supplied by the No. 2 DC bus through the

switch is located in the cabin at sta. 534 near the MAIN-

LIGHTING POS circuit breaker on the No. 2 PDP.

TENANCE PANEL (fig. 2-13-4). The guarded two-posi-

tion switch is labeled OFF and ON. It allows the aft posi-

tion light to be turned off during aided (NVG) operations

2-13-2. Position Light Switches.

and on during unaided night operations.

a. POSN Light Switch. The POSN (position)

2-13-3. Formation Lights.

switches located on the EXT LTG (exterior lighting) panel

on the left side of the overhead switch panel (fig. 2-1-10).

There are five electroluminescent panels for normal

The three position switch is labeled DIM, OFF, and BRT.

night formation operations (2 and 5, fig. 2-13-1) and eight

It adjusts the intensity of the position lights. When the

NVG compatible formation lights for NVG formation op-

switch is OFF, the position light system is deenergized.

erations (9, fig. 2-13-1)

Figure 2-13-1. Exterior Lights

2-13-1

TM 1-1520-240-10

Figure 2-13-2. Cockpit Lighting and Control

2-13-4

TM 1-1520-240-10

on. The circuit monitors electrical failures and landings in

excess of 3 to 4g’s. The light from the charge indicator

lamps can be seen emitting through two pin holes at the

base of the main light reflector. When the switch is set

from ARM to TEST, the main light comes on, powered by

the batteries. When the switch is set to DISARM, the

indications are the same as for the ARM position, except

the circuit does not monitor electrical or failures hard

landing.

Figure 2-13-3. Emergency Exit Light

2-13-7

TM 1-1520-240-10

Figure 2-13-4. Cabin Lighting and Controls

2-13-8

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