HELICOPTER, ATTACK, AH-64D LONGBOW APACHE. TECHNICAL MANUAL (2002)

TM 1-1520-251-10

This text, displayed in WHITE, is provided to allow the

crew to rapidly verify procedures once immediate correc-

tive action has been taken.

WARNING

CAUTIONS

LBA−3003

Figure 2-46. In - flight Format - Active Warnings

LBA5191

and Cautions

Figure 2-47. ENG Instruments Page Emergency

Format

NOTE

2.48.4 Count-Down Timers. Count-down-timers are

displayed in WHITE on the MPD ENG Page (fig 2-48) and

Emergency procedures will be displayed on

decrement based upon the parameters listed in Table 2-4.

the ENG page emergency format only when

When a timer has decremented completely, it displays 0:0

a DTC with procedures data is installed and

until reset. In addition Engine TGT values are displayed

loaded.

on the FLT Page and the HMD based upon the parame-

ters in Table 2-5.

2.48.3

ENG Page Emergency Format.. The ENG

NOTE

page emergency format (fig 2-47) is automatically dis-

played in response to all aircraft warnings. The upper half

During multiple timer operations, the highest

of the ENG page emergency format displays the same in-

timer number (1 - 4) value will be displayed.

dications as the ground and in-flight formats (Table 2-2).

Earlier timers will continue to run. As power

The lower half of the page will display a window containing

is reduced, each preceding timer value will

the immediate action steps associated with the active

be displayed, and each timer will be reset,

warning condition.

automatically, in order.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-41

TM 1-1520-251-10

Table 2-4 Engine Page Timers (cont)

TGT ( - 701C)

Dual Engine: u

810 °C, timer

decrements from 30 minutes. (timer1).

Either engine: u

870 °C, timer decre-

ments from 10 minutes

(timer 2).

Single Engine: >810 °C, timer decre-

ments from 30 minutes (timer 1)

870 °C, timer decrements from

10 minutes (timer 2).

<878 °C timer decrements from 2

minutes and 30 seconds (timer 3).

<896 °C timer decrements from 12

seconds (timer 4).

N_G

Either engine NG u

102.2%, timer

decrements from 12 seconds.

LBA3005

N_P 1

Engine 1 N_P > 105%, timer

decrements from 12 seconds.

N_P 2

Engine 2 N_P > 105%, timer

decrements from 12 seconds.

Figure

2-48.

Engine Instrument Timers

Table

2-5. MPD FLT Page and HMD TGT

TGT ( - 701)

Dual engine: u

807 °C, timer

Table

2-4. Engine Page Timers

decrements from 30 minutes. TGT

value is displayed on HMD and FLT

page during the final 2 minutes.

PARAMETER

CONDITIONS

Single engine: u

864°C, timer

decrements from 2 minutes and 30

TORQUE

Dual Engine: u

100%, timer

seconds. TGT value is displayed on

decrements from 6 seconds.

HMD and FLT page the entire 2

Single Engine: u

110% (contingency),

minutes and 30 seconds (timer 2).

timer decrements from 2 minutes and

TGT ( - 701C)

Dual engine: u

810 °C, timer

30 seconds (timer 1).

decrements from 30 minutes. TGT

122% timer decrements from

value is displayed on HMD and FLT

6 seconds (timer 2).

page during the final 2 minutes (timer

TGT ( - 701)

Dual Engine:u

807 °C, timer

1).

decrements from 30 minutes.

870 °C, timer decrements from 10

Single Engine:u

807 °C, timer

minutes. TGT value is displayed on

decrements from 30 minutes

HMD and FLT page during the final 2

minutes (timer 2).

(timer 1).

Single Engine: >878 °C, timer decre-

>864 °C, timer decrements from

ments from 2 minutes and 30 sec-

minutes and 30 seconds (timer 2).

onds. TGT value is displayed on HMD

919 °C, timer decrements from

and FLT page the entire 2 minutes

12 seconds (timer 3).

and 30 seconds (timer 3).

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-42

Change 4

TM 1-1520-251-10

Section IV. FUEL SUPPLY SYSTEM

2.49 FUEL SUPPLY SYSTEM

into the outboard auxiliary tank forcing the fuel from the

outboard auxiliary tank to the inboard auxiliary tank and

NOTE

then to the internal fuel cells. When an internal auxiliary

fuel cell is installed, an electric pump is used to transfer

Fuel grades, specifications, and servicing

fuel from the internal auxiliary fuel cell directly into the for-

instructions are described in Servicing,

ward and aft fuel cells simultaneously. Both auxiliary fuel

Parking, and Mooring, Section XV.

subsystems are selected on/off via the MPD and man-

aged through fuel system priorities.

The fuel system (fig 2-49) provides fuel and fuel manage-

ment provisions to operate both engines and the APU.

2.53 FUEL CROSSFEED SUBSYSTEM

Fuel is stored in two crash resistant, self sealing fuel cells;

one forward and one aft of the ammunition bay. Fuel may

WARNING

be transferred from either cell to the other. Fuel levelling

can be accomplished manually or automatically. The sys-

tem is also equipped to selectively crossfeed fuel to en-

The CROSSFEED button shall be set to

gines from either the forward or aft fuel cells. The aircraft

NORM position at all times while in flight,

has provisions for carrying external fuel tanks on the wing

unless executing emergency procedures

pylon attach points. There is also a provision for an inter-

for engine 1 fuel PSI and engine 2 fuel

nal auxiliary fuel cell carried in the ammunition bay.

PSI. A malfunctioning crossfeed valve

could result in a single engine flameout.

2.50 FUEL TANKS

The crossfeed subsystem allows both engines to draw

Two internal fuel cells provide the main storage for the air-

fuel from either internal fuel cell. There are two crossfeed

craft fuel supply. The forward fuel cell holds 156 gal.; the

valves, one for each engine. The valves are three-way,

aft fuel cell holds 220 gal.

four-position valves. The crossfeed modes are crewmem-

ber selectable functions via the MPD. Crossfeed mode

2.51 EXTENDED RANGE KIT

selections are described in paragraph 2.60.4. The AFT

crossfeed is automatically activated for the engine being

Up to four auxiliary tanks can be mounted on the four wing

started by the system processor during engine start. After

pylons for extended range operations. Each auxiliary tank

start, the crossfeed is automatically reset to NORM.

holds up to 230 usable gal. The auxiliary tanks can be jet-

2.54 FUEL QUANTITY SUBSYSTEM

tisoned in the same manner as any other externally

mounted stores.

The fuel quantity subsystem measures remaining fuel and

detects low levels in the forward and aft fuel cells. The fuel

2.51A INTERNAL AUXILIARY FUEL SYSTEM (IAFS)

quantity is displayed in lb for each cell and for the total in-

ternal fuel remaining. A forward fuel low caution message

An internal auxiliary fuel cell can be installed in the am-

is displayed when the forward cell quantity drops below

munition bay for extended range operations. The internal

240 lb or when the low level switch in the fuel cell is acti-

auxiliary fuel system holds a 129, or a 98, usable gal. cell.

vated. The low level switch will be activated at a lower fuel

The fuel cell may be equipped with a fuel probe that mea-

quantity of 175 ±20 lb. An aft fuel low caution message is

sures the fuel quantity.

displayed when the aft cell quantity drops below 260 lb or

2.52 FUEL TRANSFER SUBSYSTEM

when the low level switch in the fuel cell is activated. The

low level switch will be activated at a fuel quantity of 240

The fuel transfer subsystem allows for the transfer of fuel

±20 lb. When an auxiliary tank is empty an advisory is an-

between the internal fuel cells and from both the external

nunciated.

tanks and the internal auxiliary fuel cell to the internal fuel

2.55 FUEL BOOST SUBSYSTEM

cells. The internal fuel transfer pump is a pneumatically

driven pump that allows the transfer of fuel between the

The fuel boost pump provides pressurized fuel during en-

forward and aft fuel cells, and from external fuel tanks to

gine start and is used to increase fuel pressure when

the internal forward and aft cells. The internal transfer

manually selected ON. The pneumatically driven boost

mode provides manual and automatic selections via the

pump pressurizes the fuel line out of the aft cell. Under

MPD. When external fuel tanks are mounted on both the

normal operation, fuel is drawn out of the fuel cells by the

inboard and outboard pylons, pressurized air is routed

main engine fuel pumps.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 2

2-43

TM 1-1520-251-10

BREAKAWAY VALVE

ENG

TO ENG 2

CELL

ÔÔÔÔÔ

RIGHT SIDE

BOOST PUMP

PRESSURE SWITCH

ENÔÔ

X FEED

ÔÔ

APU

APU FUEL

ENG

RIGHT CROSSFEED

FUEL BOOST

SHUTOFF

SHUTOFF VALVE

PUMP

N2 CHECK

VALVE

TO APU

XFEED

FLOW

REGULATED

AIR FROM

29 +

PRESS

PAS SYSTEM

2 PSI

AIR IN

BOOST

19+ 3 PSI

INLET

PUMP

COOLING

SHUTOFF

AIR

OUTLET

VALVE

AUX

FUEL TANK

AIR SHUTOFF

GRAVITY

BREAKAWAY

VALVE

FILLER

VALVES

PORT (CAP)

FUEL

AIR

FUEL

XFER PUMP

AIR IN

XFER

AIR VALVE

OPEN

DRAIN

ÏÏÏÏÏÏÏÏÏÔ

FUEL

BREAKAWAY VALVE

PILOT

FUEL

VALV

DRAIN

ÔÔAIR VENT

TUBE

FUEL/DEFUEL

PRESS

CHECK VALVE

OVBD

RELIEF

INLET

VENT

VALVE

OVBD

ÏPUMPT PRESS RELIEF

VENT

OUT

AIR

FUEL

OUT

XFER

(REFUEL)

VALVE

DRAIN

LOW

ÔÔ

THERMAL

LVL

BREAKAWAY

FLOW

RELIEF

SNSR

VALV

OPEN

ÔÔ

AUX

TANK

CHECK

OVBDÏÏÏÏÏ

VALVE

ÏÏÏÏÏÏÏ

DRAIN

ÓÓ

LIGHTNING

FUEL

AFT

ARRESTOR

FUEL

QTY

FUEL

CELL

ÓÓ

XMTR

OVBD

SHUTOFF

WING JOINT

4WAY

VENT

DRAIN

VALVE

CHECK

SUMP

VALVE

DRAIN

PUSH

FUEL

VALVE

ÔÔÔ

DRAIN

XFER

SHUTOFF

ÔÔ

VALVE

GRAVIT

Y FILLER

PORT

(CAP)

PULL AWAY

COUPLINGS

AUXILIARY

FUEL

TANK(S)

EMPTY

RIGHT SIDE

230 U.S. GAL

EACH

LBA02921A

Figure 2-49. Fuel System (Sheet 1 of 2)

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-44

Change 1

TM 1-1520-251-10

ENG

TO ENG 1

CELL

LEFT SIDE

ÏÏ

X FEED

ÏÏ^{BREAKAWAY VALVE}

LEFT CROSSFEED

ENG

ENG FIREWALL

SHUTOFF VALVE

CELL

XFEED

ÔÔÔ

INERT

NITROGEN

INERTING

OUT

UNIT

FUEL VENT

SHUTOFF

VALVE

OPEN

DRAIN

SINGLE

GRAVIT

AIR VENT

POINT

FUEL/

FILLER

PORT

ADAPTER

DEFUEL

PORT

BREAKAWAY

AIR VENT/

(CAP)

CHECK

(CAP)

VALVE

PRESS RELIEF

VALVE

CLOSED

VALVE

CIRCUIT

ÏÏÏÏÏ

ÏÏ

ADAPTER

PRESSURE

(CAP)

FUELING

PILOT

ÔÔÔ

VALVE

MANIFOLD

ÔÔÔ

VENT TUBE

FUEL

BREAKAWAY

PRESS

VALVES

RELIEF

PORT

ÔÔÔ

ÔÔ

SCREEN

AUX

ÔÔÔ

LIGHTNING

ÏÔÔ

TANK

CHECK

ARRESTER

VALVE

LOW

LVL

BAFFLE

SENSOR

IAFS

FORWARD

CHECK

FLOW

FUEL

MANIFOLD

VALVES

CELL

ÏÏÏ

ÓÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏ

AFT

FUEL

FWD

ÏÏ

FUEL QTY

VENT

SUMP

FUEL QTY

SHUTOFF

XMTR

VALVE

XMTR

DRAIN

OVBD

VALVE

DRAIN

CHECK/THERMAL

PUSH

RELIEF VALVE

TO DRAIN

WING JOINT

FUEL

XFER

ÔÔÔ

SHUTOFF

VALVE

ÏÏÏÏÏÏÏÏÏÏÏ

GRAVITY

ÏÏÏÏÏÏÏÏÏÏÏÏ

FILLER

PORT

VENT

(CAP)

INTERNAL AUXILIARY FUEL SYSTEM

ÏÏ

PULL AWAY

100 U.S. GAL/130 U.S. GAL

N2 Ï

COUPLINGS

FUEL IN Ï

AUXILIARY

FUEL TANK(S)

FUEL QUANTITY

FUEL OUT

LEFT SIDE

EMPTY

INDICATING TAPE

ÏÏ

220 U.S. GAL

EACH

SUMP DRAIN

ÏÏ

ÏÏÏÏÏÏÏÏÏÏÏÏ

LBA02922B

Figure 2 - 49. Fuel System (Sheet 2 of 2)

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 2

2-45

TM 1-1520-251-10

2.56 FUEL FLOW RATE SUBSYSTEM

2.60 MPD FUEL MANAGEMENT PAGE

The fuel flow rate subsystem calculates engine fuel con-

The FUEL page displays all information and controls

sumption rate in lb/hr.

required to manage the internal fuel system (fig 2-51), the

external auxiliary system, and the internal auxiliary fuel

system (para 2.61 and 2.61A). The aircraft icon is laid out

2.57 APU FUEL SUBSYSTEM

in the same relative position as the physical equipment on

the aircraft. Status windows for fuel quantity, fuel flows

The APU fuel subsystem provides fuel to the APU. It con-

and aircraft endurance are located at the lower portion of

sists of an electrically driven APU boost pump which

the display. When a crewmember selects an operation,

draws fuel from the aft fuel cell and provides pressurized

the associated fuel lines will be displayed to provide a

fuel to the APU. The APU shut-off valve controls the flow

graphic representation of fuel routing.

of fuel from the aft cell to the APU via the boost pump.

2.58 REFUEL SUBSYSTEM

The refuel subsystem allows for pressurized refueling via

a single point adapter or a closed circuit adapter. It also

allows for gravity refueling via fuel cell filler ports for the

external fuel tanks and the forward and aft internal fuel

cells. The internal auxiliary fuel cell can only be refueled

via a pressurized adapter. The external refuel panel pro-

vides direct control of the refuel, vent, level control valve

units, and internal auxiliary fuel cell.

2.59 NITROGEN INERTING UNIT (NIU)

The NIU reduces fire hazards associated with fuel cell ul-

lages (air space) by filling the ullage with oxygen depleted

air. The NIU is self contained and automatically operated

whenever pressurized air and 115 Vac is available. A

press-to-test panel (fig 2-50) is located in the aft avionics

bay. The press-to-test button will only be valid when pres-

surized air and 115 Vac is applied to the system. The

press-to-test button will indicate that the NIU is not provid-

LBA5190

ing pressurized air. The NIU utilizes pressurized air from

the pressurized air manifold and purges about 70% of the

Figure

2-51.

Fuel Page - Standard Configuration

oxygen present. This air is then regulated into the aft fuel

cell and onward into the FWD fuel cell. When transferring

Bezel button selections are as follows:

fuel from the internal auxiliary fuel cell, all air is diverted to

the internal auxiliary fuel cell.

• T1

ENG button

• T2

FLT button

NITROGEN INERT MONITOR

• T4

PERF button

7311642315

• T6

UTIL button

RESET

• L4

XFER button

• R2

BOOST ON/OFF button

• R3

CROSSFEED FWD button

BLACK OK

PRESS TO TEST

• R4

CROSSFEED NORM button

WHITE FAIL

10 SEC MIN

• R5

CROSSFEED AFT button

LBA0452

• R6

TYPE button

Figure 2-50. NIU Test Panel

• B6

CHECK button

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-46

Change 2

TM 1-1520-251-10

2.60.1

Transfer (XFER) Button. The XFER control

buttons (fig 2-52 ) allow crewmembers to balance the fuel

load between the forward and aft fuel cells. Four option

buttons, FWD, OFF, AFT, and AUTO, provide the means

to transfer fuel between the forward and aft fuel cells.

NOTE

Selecting either FWD or AFT button will

automatically override any previous internal

transfer selection.

a. Forward (FWD) XFER Button. The FWD button

allows the manual transfer of fuel from the aft fuel cell to

the forward fuel cell.

b. Off (OFF) XFER Button. The OFF button stops

the transfer of fuel between the forward fuel cell and the

aft fuel cell.

c. Aft (AFT) XFER Button. The AFT button allows

LBA5000

the manual transfer of fuel from the forward fuel cell to the

aft fuel cell.

d. Automatic (AUTO) XFER Button. The AUTO

button (fig 2-53) allows crewmembers to select automatic

Figure

2-53.

Automatic Fuel Transfer

fuel levelling between the forward and aft fuel cell (Table

2-6).

Table 2-6.

Automatic Fuel Levelling Conditions

(with AUTO selected)

BEGIN AUTO LEVELLING TO AFT TANK

1. Engines 1 and 2 are running.

2. APU or ENG bleed air is on.

3. Aft Fuel level <814 lb

4. No Fwd Fuel Low level indication.

5. Fwd Fuel quantity >280 lb

6. If Fwd Fuel quantity >500 lb and is >100 lb more

than Aft Fuel quantity.

7. If Fwd Fuel quantity <500 lb and is >50 lb more

than Aft Fuel quantity.

HALT AUTO LEVELLING TO AFT TANK

1. Fwd Fuel low level Caution.

LBA5189

2. Fuel split <20 lb

3. APU is off and ENG bleed air is off

Figure 2-52. Fuel Transfer Options

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 1

2-47

TM 1-1520-251-10

Table 2-6. Automatic Fuel Levelling Conditions

WHITE for three seconds upon change of state, then they

(with AUTO selected) (cont)

are displayed in GREEN.

b. Normal (NORM) CROSSFEED Button. The

AUTO LEVELLING TO FWD TANK

NORM button routes fuel from the forward fuel cell to en-

1. Engines 1 or 2 are running.

gine 1 and aft fuel cell to engine 2. The aircraft icon re-

flects this state by providing a graphic representation of

2. APU or ENG bleed air is on.

the fuel cell lines. The lines are displayed in WHITE for

3. Fwd Fuel level <814 lb

three seconds upon change of state, then they are dis-

4. No Aft Fuel Low level indication.

played in GREEN.

5. Aft Fuel quantity >240 lb

c. Aft (AFT) CROSSFEED Button. The AFT button

6. If Aft Fuel quantity >500 lb and is >100 lb more

routes fuel from the aft fuel cell to engine 1 and 2. The air-

than Fwd Fuel quantity.

craft icon reflects the AFT selection by providing a graphic

representation of the aft fuel cell routing fuel to engine 1

7. If Aft Fuel quantity <500 lb and is >50 lb more

and 2. The lines are displayed in WHITE for three seconds

than Fwd Fuel quantity.

upon change of state, then they are displayed in GREEN.

2.60.5 Fuel Type Selection. The system will default to

HALT AUTO LEVELLING TO FWD TANK

fuel type JP-8. The operator will have the option to select

1. Aft Fuel low level Caution.

the fuel type other than JP-8 through the TYPE button.

2. Fuel split <20 lb

Selection of the TYPE button will display the fuel type

selections JP-8, JP-5, and JP-4. Fuel TYPE entries affect

3. Fwd tank is full

external fuel tank and non - probed internal auxiliary fuel

4. APU is off and ENG bleed air is off

cell quantity computations by changing fuel density val-

ues. Internal fuel quantity densities are sensed automati-

2.60.2 Transfer Status. When the internal transfer

cally. The fuel TYPE also affects the calculated fuel flow.

system is commanded on, the fuel transfer line will be dis-

2.60.6 Status Windows. The following status windows

played and marquee to indicate direction of transfer. The

are displayed during normal aircraft fuel configuration:

fuel transfer line is displayed in WHITE for three seconds

upon activation and then turns green. Transfer status is

a. Forward (FWD) Fuel Cell Quantity. The FWD

also indicated during automatic transfer. Fuel transfer is

fuel quantity is displayed in green at the top area of the

stopped when fuel split is less than 20 lb.

aircraft icon and provides forward fuel cell quantity. The

forward fuel cell quantity is displayed in YELLOW when

2.60.3 Boost (BOOST) ON/OFF Button. The BOOST

the cell is detected as low level. The FWD fuel cell display

button controls boost pump operation. The BOOST button

range is 0 to 1100 lb and depicts in units of 10 lb.

provides an ON dot to the right of the BOOST label when

b. Aft (AFT) Fuel Cell Quantity. The AFT fuel cell

the boost has been activated. When the boost pump has

quantity is displayed in green in the bottom area of the air-

been actioned on the system automatically sets the fuel

craft icon and provides aft fuel cell quantity. The aft fuel

CROSSFEED to AFT and the CROSSFEED AFT button.

cell quantity is displayed in YELLOW when the cell is de-

tected as low level. The AFT fuel cell display range is 0 to

2.60.4 Fuel Page Crossfeed Selections. Three se-

1500 lb and depicts in units of 10 lb.

lectable buttons, FWD, NORM and AFT, will allow selec-

c. Internal

(INT) Fuel Quantity Status Win-

tion of which cells will feed each engine. Both engines will

dow. The INT fuel quantity data field is displayed in the

feed from either the forward or aft cell when selected. This

lower left corner of the FUEL page and provides the crew-

allows the pilot an emergency means to continue flight to

member the total fuel quantity of the forward and aft fuel

a safe area after sustaining fuel system damage. Cross-

cells only. The internal fuel quantity is displayed in YEL-

feed routing can be used on the ground to control fuel feed

LOW when either the forward or aft fuel cell quantity is dis-

during hot refueling.

played in YELLOW. Internal fuel quantity is displayed

within a range of 0 to 2600 lb in increments of 10 lb. The

a. Forward (FWD) CROSSFEED Button. The

internal fuel quantity is monitored by internal sensors and

FWD button routes fuel from the forward fuel cell to engine

does not require any operator input.

1 and 2. The aircraft icon reflects the FWD selection by

providing a graphic representation of the forward fuel cell

d. Endurance (ENDR) Status Window. The ENDR

routing fuel to engine 1 and 2. The lines are displayed in

status window is displayed in the bottom right corner of

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-48

Change 1

TM 1-1520-251-10

the page and provides the amount of fuel (forward fuel cell

• L1

L AUX Fuel Transfer ON/OFF button

and aft fuel cell) remaining in time. Endurance is based on

• L5

AUX GALLONS EXT button

calculated fuel flow and will vary based upon power ap-

plication. The remaining time is displayed in WHITE when

• R1

R AUX Fuel Transfer ON/OFF button

it is less than 20 minutes.

e. Fuel Flow (CALC FLOW) Status Window. The

CALC FLOW is displayed in the bottom center of the

FUEL page and provides crewmembers with the uncor-

rected main engines calculated fuel flow in lb per hour

(LB/HR). Fuel flow is calculated based on static/free air

temperature, engine torque, pressure altitude, fuel type,

and engine type. The engine 1 & 2 fuel flow data field has

a display range of 0 to 950 LB/HR in increments of 5 LB/

HR. The total fuel flow data field has a display range of 0

to 1900 LB/HR in increments of 5 LB/HR.

f. Specific Fuel Range

(SFR) Status Win-

dow. The SFR is displayed above the CALC FLOW sta-

tus window. SFR is a calculation of ground speed in knots

divided by the total fuel uncorrected flow for the present

power setting, and is used to determine power settings for

optimum fuel economy during cruise. SFR will display only

when airspeed is ten kts or above.

2.61 FUEL PAGE EXTERNAL AUXILIARY FUEL

TANK SELECTION

LBA2565B

WARNING

Figure

2-53A.

Fuel Page - External Auxiliary Fuel

Tanks Installed

Extended Range Fuel System (ERFS)

tanks do not have any ballistic protec-

tion and are vulnerable to high-speed

2.61.1

External Auxiliary Fuel Transfer Opera-

projectiles. Projectiles passing through

tion.

External fuel transfer can be accessed on the

a fueled ERFS tank can generate a fuel

FUEL page. If tank(s) are installed on the left side of the

driven fuselage fire and or cause the

aircraft, L AUX selection is displayed. If tank(s) are

tank to detonate with the potential for

installed on the right side of the aircraft, R AUX selection

losing both thw aircrew and aircraft.

is displayed. A crewmember can select one or both auxil-

iary fuel tank transfers.

The fuel system will detect the presence and location of

2.61.2 LEFT and RIGHT AUX ON/OFF Buttons. The

external auxiliary tanks on the aircraft and display icons

AUX on/off buttons start or stop fuel transfer from left and/

representing the external tanks on the FUEL page (fig

or right tanks. The left and/or right fuel tanks must be

2-53A). A float switch in each tank indicates when that

installed on the aircraft for the AUX buttons to be dis-

tank is empty and the FUEL page displays a WHITE indi-

played. When an AUX button is activated, the respective

cator E on the center of the tank icon.

fuel line graphic will be shown and the ON dot will be dis-

Additional Bezel buttons available when auxiliary fuel

played. The lines are displayed in WHITE for three se-

tanks are installed are:

conds upon activation, then they are displayed in GREEN.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 4

2-49

TM 1-1520-251-10

NOTE

Auxiliary fuel quantity must be added

through the AUX GALLONS EXT data entry

button on the FUEL page.

a. Either left or right aux fuel transfer mode will be

commanded to off when any of the following occurs:

• A crewmember deselects aux transfer via the

MPD.

• The SP loses communication with both WPs.

• The auxiliary tanks on that side of the aircraft

indicate empty.

• AC power is lost.

b. If an internal fuel transfer is commanded, the exter-

nal auxiliary fuel transfer will be suspended until the the

forward/aft transfer is complete. The MPD R/L AUX but-

ton will display the auxiliary transfer mode as ON while in

the suspended mode. [

BLK 2

Additionally, the external

fuel transfer lines will be displayed in partial intensity

LBA5026

green while suspended.] A crewmember may select auxil-

iary transfer on and off while in the suspended mode. This

Figure 2-53B. Fuel Page - Internal Auxiliary Fuel

applies to both manual and automatic internal fuel cell

Tank Installed

transfers.

Additional Bezel buttons available when an internal auxil-

2.61A FUEL PAGE INTERNAL AUXILIARY FUEL

iary fuel cell is installed are:

TANK INSTALLED

• L2

C AUX Fuel Transfer ON/OFF button

• L6

AUX GALLONS CTR button (unprobed)

The fuel system will detect the presence and configuration

of the internal auxiliary fuel system on the aircraft and dis-

2.61A.1 Internal Auxiliary Fuel Transfer Opera-

play an icon representing the internal auxiliary fuel cell on

tion. Internal auxiliary fuel transfer can be accessed on

the FUEL page (fig 2-53B). A pressure switch indicates

the FUEL page. When an internal auxiliary fuel cell is

when the tank is empty, an advisory is displayed on the

installed, the C AUX selection is displayed.

UFD/EUFD, and the FUEL page displays a WHITE indica-

a. The internal aux fuel transfer mode will be com-

tor E on the center of the tank icon when empty.

manded to off when any of the following occurs:

• A crewmember deselects aux transfer via the

MPD.

• AC power is lost.

• The auxiliary tank is indicated as empty or failed.

b. If either an internal fuel transfer or an external aux-

iliary transfer is commanded, the internal auxiliary fuel

transfer will be suspended until the the forward/aft/exter-

nal transfer is complete. The MPD ON button will display

the auxiliary transfer mode as on while in the suspended

mode. [

BLK 2

Additionally, the external fuel transfer

lines will be displayed in partial intensity green while sus-

pended. A crewmember may select auxiliary transfer ON

and OFF while in the suspended mode.] This applies to

both manual and automatic internal fuel cell transfers.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-50

Change 2

TM 1-1520-251-10

2.61A.2 Center AUX On/Off Button. The C AUX on/

off button starts or stops the fuel transfer from the center

tank. The center fuel tank must be installed on the aircraft

for the C AUX button to be displayed. When the C AUX

button is activated, the fuel line graphic will be displayed

and the on dot will be displayed. Lines are displayed in

WHITE for three seconds upon activation, then they are

displayed in GREEN. Fuel is transferred from the internal

auxiliary fuel cell to both the forward and aft fuel cells si-

multaneously.

NOTE

Unless probed, internal auxiliary fuel quanti-

ty must be added through the AUX GAL-

LONS CTR data entry button on the FUEL

page.

2.61A.3 Status Windows. Additional status is dis-

played when auxiliary fuel tanks are installed. This data

LBA5187

field will only be displayed when the system has detected

that there are external tanks located on the aircraft.

Figure

2-54.

Fuel Page - CHECK Format

a. Total (TOT) Fuel Quantity Status. The TOT

quantity is displayed in the bottom left corner of the FUEL

page and provides the crewmember the total fuel quantity

of the forward, aft, and all auxiliary tanks. This data field

Additional bezel buttons available when CHECK is se-

will only be displayed when the system has detected that

lected:

there are external tanks located on the aircraft. The total

fuel quantity data field has a display range of 0 to 8840 lb

• R2

15 (MINUTES) button

in increments of 10 lb. Auxiliary fuel quantity must be add-

• R3

20 (MINUTES) button

ed through the AUX GALLONS EXT and/or AUX GAL-

LONS CTR data entry buttons.

• R4

30 (MINUTES) button

• R5

START/STOP button

b. Total Endurance (ENDR TOT) Status. The

ENDR TOT status window is displayed in the bottom right

2.62.2

Fuel Check Operation. Either crewmember

corner of the FUEL page and provides the total endur-

may initiate a fuel check by selecting the CHECK button,

ance based on the fuel remaining in the forward cell, aft

pressing the desired MINUTES (15, 20, or 30) button and

cell, and all auxiliary fuel tanks. Total endurance is based

then pressing the START button. A status window in the

upper left corner will display the START time, RUN time,

on calculated fuel flow and will vary based upon power ap-

and burn RATE. Upon completion of the check (fig 2-55),

plication.

a status window, in the upper right corner, will display cal-

culated BURNOUT, VFR RES, and IFR RES times in the

2.62 FUEL CHECK

currently selected time format, Zulu or local. The fuel

check may be terminated prior to the selected MINUTES

by selecting STOP. If neither crewmember is viewing the

2.62.1 Fuel CHECK Button. When selected, the fuel

CHECK format at the conclusion of the check, an advisory

CHECK button allows the calculation of (fig 2-54) burnout,

will appear on the UFD/EUFD. Selection of the CHECK

VFR reserve, and IFR reserve times, based on the fuel

format by either crewmember will remove the advisory.

quantity change, during the selected 15, 20, or 30 minute

Additional fuel checks may be initiated by selecting the

intervals.

START button.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 2

2-50.1

TM 1-1520-251-10

Section V. FLIGHT CONTROL SYSTEM

2.63 FLIGHT CONTROL SYSTEM

measure the positions of the controls and provide this in-

formation to the FMC. Hydraulic power is supplied by two

The AH-64D employs an irreversible hydromechanical

independent hydraulic pumps which are mounted on the

flight control system (fig 2-56). The hydromechanical sys-

accessory gearbox of the main transmission. The FMC

tem is mechanically activated with conventional cyclic,

provides rate damping, command augmentation, attitude

collective, and directional controls, through a series of

and altitude hold within the ± 10% (20% forward pitch) au-

push pull tubes and bellcranks which activate four air-

thority of the system. An electrically actuated horizontal

frame-mounted hydraulic servocylinders. The four hy-

stabilator is attached to the lower aft portion of the vertical

draulic servocylinders control longitudinal/lateral cyclic,

stabilizer. Movement of the stabilator is commanded by

main rotor collective, and tail rotor pitch. The servocylind-

the FMC in either a manual or automatic mode. A trim feel

ers incorporate integral Stability and Command Aug-

system is incorporated in both the cyclic and pedals pro-

mentation System (SCAS) actuators which are active

whenever the Flight Management Computer (FMC) is on.

viding a control force gradient with control displacement

Linear Variable Differential Transducers (LVDTs) are in-

from a selected trim position. A trim release switch allows

corporated into each of the flight control axes. The LVDTs

momentary disengagement of the trim feel system.

UPPER FLIGHT CONTROLS

LONGITUDINAL SERVOCYLINDER AND BUCS SHEAR PIN

COLLECTIVE SERVOCYLINDER AND BUCS SHEAR PIN

LONGITUDINAL

LATERAL SERVOCYLINDER AND

MECHANICAL CONTROL LINKAGE

BUCS SHEAR PIN

PILOT CYCLIC CONTROL STICK

(LVDT, AND ARDD ASSEMBLIES

UNDER COVER)

PILOT DIRECTIONAL CONTROL PEDALS

(LVDT, AND ARDD ASSEMBLIES UNDER

COVER)

CPG CYCLIC CONTROL STICK

(LVDT, AND ARDD,

ASSEMBLIES UNDER COVER)

LATERAL

MECHANICAL CONTROL LINKAGE

PILOT COLLECTIVE CONTROL STICK

(LVDT, ARDD ASSEMBLIES AND 1G

SPRING ASSEMBLIES UNDER COVER).

(PILOT ONLY) LATERAL AND LONGITUDINAL

FEEL SPRING ASSEMBLIES AND MAGNETIC BRAKES

CPG COLLECTIVE CONTROL STICK

(LVDT, ARDD ASSEMBLIES AND 1G

CPG DIRECTIONAL CONTROL PEDALS

SPRING ASSEMBLIES

(LVDT, AND ARDD ASSEMBLIES

UNDER COVER)

LBA0475

UNDER COVER)

Figure 2-56. Flight Controls System

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-51

TM 1-1520-251-10

2.64 CYCLIC CONTROL SUBSYSTEM

Conventional cyclic control sticks (fig 2-57) are attached

to individual support assemblies which house the primary

longitudinal and lateral control stops and two LVDT’s. The

cyclic stick grips incorporate the force trim release, hold

mode switch, and the FMC release button. The CPG cy-

clic stick incorporates a fold down mechanism to minimize

interference when the CPG is not flying the aircraft. The

CPG cyclic stick is fully effective in either the up or down

position.

BUCS SELECT

TRIGGER

STABILATOR

(CPG ONLY)

CONTROL SWITCH

FORCE

TRIM/HOLD

MODE

LBA2082

SWITCH

Figure 2-58. Collective Grip Flight Control

Switches

2.66 DIRECTIONAL CONTROL SUBSYSTEM

The directional control system consists of pedals in each

crew station which activate the tail rotor servocylinder lo-

cated in the vertical stabilizer. Attached to each directional

pedal assembly is a control stop and one LVDT. Pedal ad-

justment is accomplished through a lock/unlock lever lo-

FMC

RELEASE

cated between and aft of the pedals. Wheel brakes are in-

corporated into the upper portion of the control pedals.

LBA2081

2.67 FORCE TRIM SUBSYSTEM

WARNING

Figure 2-57. Cyclic Grip Flight Control Switches

Even with the force trim ON, the flight

controls must be monitored at all times

2.65 COLLECTIVE CONTROL SUBSYSTEM

while the rotors are turning. When force

trim is selected OFF from the A/C UTIL

page, no force gradient will be available

The collective pitch control system consists of identical

to position the cyclic. Hands off cyclic

collective grips (fig 2-58) in each cockpit. Movement of the

capability will NOT be available.

collective is transmitted to the collective servocylinder for

control of main rotor blade pitch angle and to the load de-

A force trim system is incorporated in the longitudinal, lat-

mand spindle of each engines hydromechanical unit for

eral and directional control axes. The force trim system is

load anticipation. Located at the base of each collective

designed to keep the controls in the position that the pilot

control is the primary control stop, an LVDT, and a 1G bal-

or CPG selects. The force trim system is located beneath

ance spring. The 1G spring counterbalances the weight of

the pilots floor. The Force Trim Release (FTR) switch al-

the collective control sticks. Both collective control sticks

lows the crewmember to momentarily interrupt the sys-

employ a friction adjustment twist grip. The collective con-

tem. When interrupted, the force trim brakes allow move-

trol grips also provide the stabilator control switch.

ment of the controls without any resistance. When

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-52

TM 1-1520-251-10

engaged, the magnetic brakes lock and the spring assem-

c. The “?” symbol is displayed in red just above the

blies resist control movement. DC electrical power is re-

symbol when the stabilator has been detected as failed

quired for operation of the system and a complete failure

and the position of the stabilator is unknown.

would disable the system and allow the cyclic and pedals

to move freely without resistance from the trim feel

springs.

2.68 HORIZONTAL STABILATOR

The stabilator provides pitch trim angle control and im-

proves over-the-nose visibility at low airspeeds. The stabi-

lator has both an automatic and a manual mode. The au-

tomatic mode is engaged following power-up of the

aircraft and is controlled by the FMC. Two modes are

available within the automatic control system. The auto -

mode provides automatic scheduling in accordance with

collective position, airspeed, and pitch rate. The Nap Of

the Earth/Approach (NOE/A) mode commands the stabi-

lator to 25° trailing edge down, up to a speed of 80 KTAS.

At speeds greater than this, the stabilator schedule re-

verts to the auto - mode. A manual mode is selectable at

airspeeds less than 80 KTAS, or is engaged when the au-

tomatic system fails. Manual control or stabilator reset is

affected through the stabilator control switch on the collec-

tive flight grip. Depressing the stabilator control switch will

LBA2083A

reset the stabilator to the AUTO mode. Stabilator position-

ing is accomplished by two tandem DC motor actuators.

Stabilator position information (in degrees) is presented

Figure

2-59.

FLT Page - Stabilator Symbol

on the SYS page, and relative position information on the

the FLT page and on the FLT SET page. The stabilator

trailing edge incorporates Gurney flaps for increased

2.68.2 Engine SYS Page. Indications in the Stabilator

aerodynamic stabilization.

Status Window on the engine SYS page (fig 2-60) are as

follows:

2.68.1 FLT and FLT SET Page. Stabilator indications

on the FLT and FLT SET (fig 2-59) pages are as follows:

a. The stabilator trailing edge angle is displayed at a

range from - 10° up to 35° down. A “?” symbol is displayed

when the stabilator has been detected as failed and the

a. The stabilator symbol is displayed in WHITE with-

position of the stabilator is unknown.

out airspeed information when the stabilator is operating

in the manual mode. When the stabilator has been de-

tected as failed, the stabilator symbol is displayed in YEL-

b. The stabilator orientation is presented as “UP” or

LOW along with the limiting TAS for the current stabilator

“DN” (down). This indication is not presented when the

position. The stabilator symbol and TAS limit are dis-

stabilator has been detected as failed.

played in RED when the current TAS is greater than the

TAS limit.

c. The nominal TAS limit is displayed to indicate the

maximum operating airspeed limit for the current trailing

b. The nominal true airspeed (TAS) limit is displayed

edge angle. The limit is displayed in WHITE when the sta-

just below the symbol to indicate the maximum operating

bilator is operating in the manual mode; the limit is dis-

airspeed limit when the stabilator has been detected as

played in YELLOW when the stabilator has been detected

failed. When the position of the stabilator is unknown, the

as failed. When the position of the stabilator is unknown,

nominal airspeed limit will default to the TAS equivalent of

the nominal airspeed limit will default to the TAS equiva-

90 KIAS. When FMC data is not valid, the nominal IAS

lent of 90 KIAS. When FMC data is not valid, the nominal

limit is displayed.

IAS limit is displayed.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-53

TM 1-1520-251-10

the radar or barometric altitude hold mode. Attitude and

Altitude hold switches are momentary ON/OFF. Selecting

the switch once turns the mode ON; selecting the switch a

second time turns the mode OFF. The collective control

flight grip incorporates a switch for control of the horizontal

stabilator. The three position switch allows the stabilator

to be moved throughout its complete range and to be re-

set to the automatic mode. When the stabilator is in the

NOE/APPR mode, pressing the stabilator RESET button

will disengage the NOE/APPR mode.

2.70 MPD FMC CONTROLS

The A/C UTIL page (fig 2-61) displays controls for the

FMC. The controls have two possible initialization condi-

tions; ground and in - flight. If startup occurs in - flight, the

FMC controls will be set to the state they were in before

the power interrupt. If startup occurs on the ground, all

modes and controls are on except the NOE/Approach

mode. The Stability and Command Augmentation System

LBA2540A

(SCAS) will automatically be engaged following EGI align-

ment.

Figure 2-60. SYS Page - Stabilator Status

2.69

CONTROL SWITCHES

Control switches are integrated into the cyclic and collec-

tive to allow hands-on cyclic and collective manipulation of

the systems. The cyclic grip (fig 2-57) incorporates an

Flight Management Computer (FMC) release button and

the Force Trim/Hold Mode switch. The FMC release but-

ton immediately disengages the SCAS if pressed. The

Force Trim/Hold Mode switch is a five position center

maintained switch. In the center position, SCAS, force trim

and the hold mode, if selected, are active. Pressing the

switch up (forward) to the Release (R) position interrupts

force trim and the attitude hold mode references if the hold

modes are selected. Releasing the force trim switch re-

engages the force trim at the new location and resets the

attitude hold modes to the new conditions. Pressing the

switch down to the Disengage (D) position disengages

any selected hold modes. Pressing the switch to the left to

the attitude (AT) position engages/disengages the posi-

LBA2083

tion, velocity, or attitude hold mode. Pressing the switch to

the right to the altitude (AL) position engages/disengages

Figure

2-61.

UTIL Page

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-54

TM 1-1520-251-10

2.71 STABILITY AND COMMAND AUGMENTATION

The hold modes are divided into altitude and attitude hold

SUBSYSTEM (SCAS)

(fig 2-62). The modes are activated through the five - posi-

tion Force Trim / Hold Mode switch. Upon activation, the

NOTE

pilot is provided with a visual indication that the mode is

The FMC is susceptible to electrical switch-

engaged (see Section XIV). Upon disengagement, the

ing transients at voltage levels below that

flight cues are removed and a tone sounds. The attitude

found on the aircraft. Susceptibility is char-

hold mode can be engaged anytime the SCAS channels

acterized by uncommanded disengagement

and Force Trim are on. However, the mode is only active

of SCAS functions in some or all axes.

in-flight. The pilot can “fly through” the hold modes in any

or all axes. When the FMC senses control motion above

The SCAS has three functions; Stabilization System,

the breakout in any axis or the force trim is momentarily

Command System, and the Hold modes. The Stabilization

released, that axis reference is no longer held. When the

System provides rate damping in all axes. Additionally, the

controls are returned to trim, the FMC captures and holds

Stabilization System automatically provides turn coor-

the new reference.

dination at speeds greater than 40 kts, and dampens any

atmospheric upsets to the airframe in order to stabilize the

aircraft. The Command System provides a uniform aircraft

response for a given control input at all airspeeds. The

Hold modes are designed to provide limited hands-off

flight and decrease pilot workload. The FMC monitors all

inputs and will disengage on an axis-by-axis basis if a fail-

ure occurs.

2.72 HOLD MODES

WARNING

• Hold modes operate through the

SCAS system, and as such provide

only limited capability. The system is

not capable of maintaining the se-

lected flight condition in all flight

conditions. The pilot shall continually

monitor the aircraft’s flight condition

to ensure safe operation.

• Fly the aircraft to a stabilized trimmed

state before engaging attitude, hover,

or altitude hold. Alternatively, fly the

aircraft with the trim gradient (FTR

switch released) until the aircraft is

stabilized at the desired trimmed

state. Failure to follow this procedure

may result in undesirable and/or un-

suspected aircraft responses.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-55

TM 1-1520-251-10

2.72.1 Use of Force Trim during Hold Mode Opera-

NOTE

tions.

Yaw breakout ranges (Table 2-7) increase in

order to provide good pedal response at a

hover. They also de - sensitize the pedals in

WARNING

order to prevent disengaging the heading

hold mode during cruise flight.

Even with the force trim ON, the flight

controls must be monitored at all times

while the rotors are turning. When force

Table 2-7. Flight Control Breakout Values

trim is selected OFF from the A/C UTIL

page, no force gradient will be available

Pitch and Roll cyclic displacement w

0.25 in.

to position the cyclic. Hands off cyclic

Collective displacement w

0.50 in.

capability will NOT be available.

Pedal displacement depends on the Attitude Hold

submode as follows:

The force trim release provides two functions to the hold

modes: 1) it provides a means to communicate to the flight

Position hold w

0.10 in.

control system that the pilot wants to establish a new ref-

Velocity hold w

0.20 in.

erence condition to hold, and 2) it provides a mechanism

Attitude hold w

0.30 in.

to allow the SAS sleeves to re - center, providing the great-

est possible margin for maintaining any of the desired hold

2.72.2 Attitude Hold. This mode is divided into four

conditions.

submodes of operation: position hold, velocity hold, atti-

tude hold, and heading hold. Activation of a submode is

a. Force Trim Operational Theory. When the SAS

dependent on inertial ground speed.

is attempting to hold a position or attitude it develops a

bias towards one side. As the wind or other disturbances

a. Position hold submode. Position hold may be

continue to upset the aircraft this can lead to a case where

engaged below 5 kts ground speed provided the force trim

more response is required than SAS can provide (the SAS

is operational. The FMC uses velocity inputs from the EGI

has only ± 10% of the total actuator authority in roll, yaw,

to approximate and maintain a position. Heading hold is

and collective actuators, and 20% and - 10% in the pitch

provided and, if Altitude Hold is engaged, a three dimen-

actuator). When all of the SAS authority is used up, the

sional hover hold is available. If the aircraft drifts out of a

SAS is saturated. When this occurs a flight control tone

48 foot boundary (one rotor diameter) and the hold mode

will sound and a message will appear on the UFD/EUFD,

is engaged, the Flight Controls tone will sound and a UFD/

meaning that the SAS is saturated in one or more axes. At

EUFD advisory message will be displayed.

this point, the SAS will be totally ineffective in the axis that

is saturated.

b. Velocity hold submode. Velocity hold may be

engaged at ground speeds between 5 kts and 40 kts (30

b. Force Trimming Procedure. When the force

kts when decelerating in attitude hold mode). The force

trim release is actioned it causes the SAS command to

trim system must be operational and ON. The FMC uses

decay or “washout” at a one second rate. Actioning or

input from the EGI to maintain a constant velocity. Head-

“bumping” the force trim release for only a fraction of a se-

ing Hold is active and Altitude Hold may be engaged.

cond resets the mechanical portion of the force trim to a

zero force state, but does not provide enough time to re-

NOTE

center the SAS. The best procedure to follow after receipt

of a SAS saturated message, or entering a hold mode, is

This is not a terrain following mode. The ra-

to fly the aircraft to the desired state (hover, velocity, or at-

dar altimeter provides only distance from

titude) then press and hold the FTR button for 3 seconds.

the ground directly below the aircraft and

This allows the SAS time to re-center providing the most

does not provide any approaching terrain

margin for any external disturbances to the aircraft.

variation information.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 2

2-57

TM 1-1520-251-10

c. Attitude Hold Submode. The FMC uses EGI

Heading Hold Submode.

rates and attitudes to maintain the selected aircraft atti-

tude. Attitude hold will either maintain the attitude and

NOTE

heading that existed when the mode was selected or the

attitude/heading that existed when the force trim was re-

There is no pilot symbology or tone indica-

leased. Heading hold is provided in this mode until the pi-

tions of heading hold engagement or disen-

lot initiates a turn. Upon initiating a turn, the turn coordina-

gagement.

tion mode will engage while maneuvering. When the turn

1. Heading Hold submode is engaged when Atti-

is completed, heading hold will automatically re-engage.

tude Hold mode is ON when:

Attitude Hold Engagement. Attitude Hold may

a. Pilot displacement of the pedals is t the

be engaged when ground speed u

40 kts; pitch

breakout values listed in table 2-7.

attitude t

± 30°; roll attitude is t

± 60°; and

pitch and roll rates are t

5°/second. Attitude

b. Yaw rate is t

3°/second.

hold will re - reference to a new pitch or roll atti-

c. Pilot displacement of roll cyclic input is t

tude when the force trim release button is no

0.25 in. from trim position.

longer being asserted and the aircraft is within

the attitude limits defined previously. If the roll t

d. Roll attitude is t

± 3° from level.

3°, the roll reference will be automatically set to

zero (wings level).

2. Heading Hold submode is engaged when Atti-

tude Hold mode is OFF (SAS only) when:

Turn Coordination. Yaw turn coordination oper-

ates in conjunction with heading hold at speeds

a. Ground speed is t

40 kts.

40 kts ground speed in Attitude Hold sub-

b. Pilot displacement of the pedals is t

0.1 in.

mode. Heading hold will revert to turn coordina-

from trim position.

tion when:

c. Yaw rate is t

3°/second.

a. Pilot displaces pedals u

0.3 in. from trim

position.

d. One second has passed after heading hold

was disengaged.

b. Pilot applies roll cyclic input u

0.25 in. from

trim position.

Altitude Hold Submode. The Altitude Hold sub-

c. Pilot actions force trim release.

mode may be engaged at any airspeed. Vertical velocity

must be less than 200 fpm at a hover, or 400 fpm at cruise

d. Aircraft rolls u

± 7° from level.

for the mode to engage. The Altitude Hold mode will auto-

Turn coordination with the Attitude Hold mode

matically disengage when:

OFF operates continuously at speeds u

40 kts

ground speed. Turn coordination will revert to

1. Pilot displaces the collective more than 0.50 in.

YAW SAS when:

from the reference position.

a. Pilot displaces pedals u

0.3 in. from trim

2. Rotor speed is u

104%.

position.

3. Rotor speed is t

97%.

b. Pilot actions force trim release.

4. Either engine torque exceeds 100%.

c. Sideslip is re-referenced only when the

force trim release is actioned.

5. Either engine TGT exceeds 867 °C.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-58

Change 4

TM 1-1520-251-10

2.73.1

Description. The Back-Up Control System

CAUTION

(BUCS) is a single - channel, non - redundant, fly- by- wire

control system that can electronically operate all four control

In the event the radar altitude hold func-

axes. The BUCS is modeled to duplicate the mechanical

tion fails due to radar altimeter or sub-

flight controls without the Stability Augmentation System

system failure and the BAR HOLD and

(SAS) engaged. This provides identical control authority and

RADHLD FAIL advisories are not dis-

handling qualities as SAS-OFF flight in any given flight con-

played, the Altitude Hold mode should

trol axis. A shear pin is located on each of the flight control

be disengaged to prevent an inadvertent

actuators and an Automatic Roller Detent Device (ARDD) is

loss of altitude. The Altitude Hold mode

located on each control, in each crewstation, to allow the con-

may be subsequently re-engaged as

trols to decouple in the event of a control jam. This system

long as the proper UFD/EUFD advisories

allows the controls to decouple regardless of where the jam

are present.

takes place. The FMC uses Linear Variable Displacement

f. Radar Altitude to Barometric Altitude Hold

Transducers (LVDT) control position and actuator position to

Changeover. When the altitude hold mode is engaged

calculate the equivalent mechanical control command when

and the aircraft is within 0 - 40 kts groundspeed and be-

in BUCS. Rotary Variable Displacement Transducers (RVDT)

tween 0 - 1428 ft AGL, the aircraft will be in the Radar Alti-

are used to sense a decoupling of the ARDDs, and LVDTs

tude hold mode; anything outside these parameters and

are used to sense a mistrack and send control movement in-

the aircraft will be in Barometric hold mode.

formation to the FMC. The FMC electrically controls the

2.73 BACK - UP CONTROL SYSTEM (BUCS)

BUCS servo valve on the primary side of the affected flight

control actuator. The primary hydraulic system and the FMC

CAUTION

must be operational for BUCS to operate.

• Engagement of the BUCS system may

occur without Automatic Roller Detent

WARNING

Decoupler (ARDD) breakout during

crew force fights if a CPG BUCS select

switch failure has also occurred.

Servoactuator LVDT probe migration

• When BUCS is engaged, do not re-

may aggravate the force fight condi-

lease the flight controls until the flight

tion. If a BUCS engagement occurs af-

has been completed and the main ro-

ter a control force fight, attempt to re-

tor has come to a complete stop.

cover the aircraft without severing the

Force trim may or may not be available

mechanical controls and follow estab-

in the BUCS ON axis.

lished procedures for BUCS ON flight.

• If during a force fight with the CPG

• BUCS can become engaged when a

BUCS select switch activated, and a

mistrack between the controls and the

breakout of the ARDD occurred, the

actuator is sensed. This commonly

BUCS would not provide a three sec-

occurs when external power is pro-

ond easy on and would subject the air-

vided to the aircraft without hydraulic

craft to a large control transient.

power. Applying hydraulic power and

• Breakout of the ARDD at the base of

completing an FMC IBIT may clear the

the collective control will eliminate the

BUCS engagements.

normal mass of the control system

and may cause the collective to move

slightly in response to rotor system

vibration. This slight movement will

2.73.2 Operation. When BUCS is engaged, an FMC

be detected by the LVDT and can pro-

disengaged caution and a SAS DISENGAGED (A/C UTIL

duce an unwanted heave (collective)

page) for the axis in BUCS will be displayed. Several

application. Increasing collective fric-

types of engagements are possible, and are covered, by

tion can eliminate this characteristic.

crewstation, in the following paragraphs.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 4

2-59

TM 1-1520-251-10

(2) Severances.

WARNING

(a) Severance aft of the pilot’s control. A

A BUCS jam engagement in the aft cyclic

severance can occur due to malfunction or battle damage

longitudinal control direction or in the

between the pilots’ controls and the actuator. In this case

left or right pedal tail rotor control direc-

the BUCS will automatically engage when the pilot

tion may result in limited control author-

achieves the proper mistrack with respect to RAM LVDT

ity (from the point of the jam) because

(17.5 %, except in longitudinal aft where the mistrack must

the actuator is unable to break the actua-

be 27.5%). The BUCS flight control logic will wash in full

tor shear pin. The worst case jam condi-

authority over a period of approximately one (1) second

tion in the longitudinal aft cyclic (jam oc-

allowing the pilot to transition to BUCS controlled flight.

curs at the actuator) will result in a 12%

SAS will be OFF in the engaged axis. Force trim will be

aft control authority from the point of the

ON in all axes. Hold modes will be available in the axes

jam. Forward cyclic can break the shear

that SAS is available. The UFD/EUFD will display the

pin and gain full control. The yaw axis

BUCS caution messages.

may be limited to +/ - 25% control author-

ity from the point of jam. If flight control

(b) Severance between crewstations. If a

authority is insufficient for a hover, at-

severance occurs between the crewstations the pilot will

tempt a run - on landing.

retain full mechanical control. The pilot should keep flying

on the mechanical flight controls. When a mistrack is de-

tected, a master caution light and flight control tone will be

CAUTION

presented and, on the UFD/EUFD, a BUCS FAIL X (X = P,

R, Y, or C) and FMC DISENGAGED caution messages

will be displayed. The CPG controls will not follow the pilot

After a BUCS engagement and transition

control inputs in the axis that is severed. SAS will be OFF

to BUCS controlled flight, a flight con-

in the severed axis. Force trim feel will not be available to

trols controllability check (small control

the CPG in in the severed axis. Hold modes will be avail-

inputs in each axis to check for correct

able in the axes that SAS is available.

response) should be conducted to es-

tablish if any other control axis has been

b. Control System Engagement Logic (Copilot

affected.

crewstation - CPG flying the aircraft).

Control System Engagement Logic (Pilot

(1) Jams.

crewstation - pilot flying the aircraft).

The CPG will have to decouple the ARDD. BUCS will au-

(1) Jams. A jam can occur anywhere in the flight

tomatically engage in the affected axis. Breakout values

control system through malfunction or battle damage.

will be approximately 15% higher than the pilots’ values.

Should the pilot be unable to move the controls, in one or

The CPG can expect the control to displace to full throw

more axes, he should aggressively decouple into the ap-

when the ARDD releases. The FMC will wash in full au-

propriate axis. The pilot can expect the control to displace

thority over a period of approximately three (3) seconds

to full throw when the ARDD releases. The FMC will wash

allowing the CPG to center the controls and transition to

in full authority over a period of approximately three (3) se-

BUCS controlled flight. SAS and force trim will be OFF in

conds allowing the pilot to center the controls and transi-

the engaged axis. Hold modes will be available in the axes

tion to BUCS controlled flight. SAS will be OFF in the en-

that SAS is available. The UFD/EUFD will display the

gaged axis. Force trim will be ON in all axes. Hold modes

BUCS caution messages.

will be available in the axes that SAS is available. The

UFD/EUFD will display the BUCS caution messages.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-60

Change 4

TM 1-1520-251-10

(2) Severances.

(b) Severance between the crewsta-

tions. If sufficient mistrack is detected between the pilot

(a) Severance Aft of the Pilot’s control. A

control position and CPG control position, the master cau-

severance can occur due to malfunction or battle damage

tion light and flight control tone will be presented in both

between the pilots’ controls and the actuator. Because the

crewstations, and the UFD/EUFD will display BUCS FAIL

two crewstations are still mechanically linked together,

X (X = P, R, Y, or C) and FMC disengaged caution mes-

this type of severance will operate the same as in the pilot

sages. In this instance, the pilot still has full mechanical

station. BUCS will automatically engage when both the pi-

control. The CPG should transfer control of the aircraft to

lot and CPG LVDTs achieve the proper mistrack with re-

the pilot. If it is necessary for the CPG to the fly the aircraft,

spect to the RAM LVDT. The FMC will wash in full author-

the BUCS trigger should be engaged and BUCS control

ity over a period of approximately one (1) second allowing

will be established after sufficient mistrack between CPG

either crewmember to transition to BUCS controlled flight.

control position and the RAM position is reached. The

SAS will be OFF in the engaged axis. Force trim will be

BUCS FAIL message will be replaced by the appropriate

ON in all axes. Hold modes will be available in the axes

BUCS ON caution message. The FMC will wash in full au-

that SAS is available. The UFD/EUFD will display the

thority over a period of approximately one second, allow-

BUCS caution messages.

ing the CPG to transition to BUCS controlled flight. SAS

and force trim will be off in the severed axis. Hold modes

CAUTION

will be available in the axis that SAS is available.

The effect of the CPG assuming control

c. BUCS - Flight Transfer of Controls. Transfer of

with the BUCS trigger select is to trans-

controls should only be accomplished if the flying crew-

fer control of the aircraft from a flight

member is incapacitated or in the case of the CPG flying

control that still retains some integrity to

the aircraft and a severance occurs between the crewsta-

that of a non-redundant electronic

tions. Table 2 - 8 lists the procedures for transfer of control

means of flight control. This shall only

between the crewmember in BUCS and the opposite

be activated if the pilot is incapable of

crewmember.

flying the aircraft.

NOTE

The CPG must create the proper mistrack. If

the CPG repositions the flight control (af-

fected axis) to a position that is is less than

the appropriate mistrack value and engages

the BUCS trigger, engagement will not oc-

cur. The CPG may engage the trigger and

hold it prior to achieving the mistrack. BUCS

engagement will occur when the proper mis-

track is reached.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 4

2-60.1/(2-60.2 blank)

TM 1-1520-251-10

Section VI. HYDRAULIC AND INTEGRATED PRESSURIZED AIR

SYSTEMS (IPAS)

2.74 HYDRAULIC SYSTEMS

2.77.3 Manifold Pressure and Return Filters. Filters

on both manifold pressure and return sides have mechan-

ical dirty filter indicators for visual inspection. These indi-

The hydraulic systems consist of two independent sys-

cators operate on differential pressure. Only the return fil-

tems: the primary system and the utility system. They are

ter has bypass valve provision. In addition to the

similar but not identical, and have separate as well as

mechanical/visual indicators both filters contain electrical

shared functions.

switches that provide signals to generate a caution mes-

sage on the UFD/EUFD.

2.75 PRIMARY HYDRAULIC SYSTEM

The primary hydraulic system (fig 2-63) provides hydraulic

2.77.4 Fluid Level Indicator. A fluid level indicator in

power to the primary side of all four flight control servocy-

the manifold housing allows visual inspection of the reser-

linders. Only the primary sides of these servo actuators

voir fluid level.

have electrohydraulic valves that allow the Flight Manage-

ment Computer (FMC) to affect the flight controls. Conse-

quently, failure of the primary hydraulic system will result

2.77.5 Primary System Pressure Sensing Switch. A

in the loss of FMC. The primary hydraulic equipment in-

pressure switch senses primary system pressure and in-

cludes a hydraulic pump, manifold, and servo actuators.

forms the pilot and CPG of a low fluid pressure condition

by generating a caution message on the UFD/EUFD.

2.76 PRIMARY HYDRAULIC PUMP

The primary hydraulic pump is mounted on the accessory

2.77.6 Pressure Transducer. A pressure transducer

drive case of the main transmission (left side). The pump

measures hydraulic pressure on the pressure side of the

is a constant pressure, variable displacement design,

manifold and transmits this value to the MPD.

driven by the transmission accessory gear train.

2.77 PRIMARY HYDRAULIC MANIFOLD

2.77.7 Pressure Transducer. A pressure transducer

measures hydraulic pressure on the pressure side of the

The primary manifold is installed on the left forward quad-

manifold and transmits this value to the MPD.

rant of the transmission deck. Its function is to store, filter,

and regulate the flow of hydraulic fluid as well as provide

analog pressure, dirty filter, and low level indications. The

2.78 UTILITY HYDRAULIC SYSTEM

manifold reservoir is pressurized on the return side by

IPAS air acting on the manifold reservoir piston. This pre-

vents pump inlet cavitation. Servicing crews introduce

The utility hydraulic system (fig 2-64) provides hydraulic

fluid to the reservoir through the Ground Support Equip-

power to the utility side of all four flight control servocylind-

ment (GSE) connections or the hand pump. The primary

ers. This system also provides hydraulic power to the rotor

hydraulic system fluid capacity is 6 pt. The reservoir

brake, area weapon turret drive, ammunition handling

stores approximately 1 pt.

system, APU start motor, tail wheel unlock actuator, exter-

nal stores elevation actuators and emergency hydraulic

2.77.1 Air Bleed Valve. The air bleed valve is used to

system. Equipment includes a hydraulic pump, manifold,

deplete the pressurized air from the manifold reservoir for

and servocylinders. The pump is mounted on the acces-

system repair or service.

sory drive case of the main transmission (right side). The

significant difference in the primary and utility hydraulic is

2.77.2 Reservoir Low Level Indicating Switch. A res-

the manifold. Additional components in the system are the

ervoir low level indicating switch is activated by the man-

accumulator, rotor brake, and the utility hydraulic return

ifold reservoir piston. The UFD/EUFD displays the mini-

accumulator that dampens hydraulic pressure surges

mum operating level caution message.

caused by sudden actuation of the gun turret.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-62

Change 2

TM 1-1520-251-10

LATERAL

COLLECTIVE

LONGITUDINAL

SERVOCYLINDER

DIRECTIONAL

SERVOCYLINDER

ÏÏ

ROTOR

ÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏ

BRAKE

PRESS FILTER

TAIL

WHEEL

RTN FILTER

UTILITY

LEGEND

LOCK

MANIFOLD

TO PAS

ÏÏÏÏÏÏ

PRESSURE

SYSTEM

ROTOR

BRAKE

CONTROL

ÏÏÏÏÏÏ

VALVE

GSE

ACC

AUX

ÏÏ

ÏÏÏÏÏÏRETURN

PRESS

UTILITY

START

GSE

PUMP

ÏÏ

RTN

NITROGEN

CHECK

ÏÏ

ÏÏÏÏ

VALVE

ÏÏÏÏÏÏÏÏÏÏÏÏÏ

Ï Ï

SHUTOFF

ÏÏÏÏ

HAND PUMP

Ï Ï

VALVE

APU

ÏÏÏÏÏÏÏÏÏÏÏÏÏ

START

MOTOR

ÏÏÏ

ÏÏ

CHECK VALVE

ÏÏÏÏ

ÏÏ

UTILITY

ÏÏ

ACCUMULATOR

UTILITYÏÏGAGESURE

NITROGEN

GAS

RESERVOIR

ÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏ

ÏÏ

ÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏ

ÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏ

RETURN

ÏÏÏÏÏÏ

ACCUMULATOR

ÏÏ

55 PSI

ÏÏ

ÏÏÏÏÏ

CARRIER

ÏÏÏ

DRIVE

ÏÏ

ÏÏÏÏÏ

MOTOR

ASSEMBLY

LEFT

ÏÏ

INBOARD

ÏÏÏÏÏÏÏÏÏ

TURRET

ÏÏ

RIGHT

ELEVATION

RIGHT

OUTBOARD

INBOARD

ÏÏ

Ï SERVOCYLINDER

ÏÏ

ÏÏÏ

ÏÏÏÏÏ

TURRET

Ï Ï

AZIMUTH

SERVOCYLINDER

ÏÏÏ

ÏÏÏÏÏ

LEFT

OUTBOARD

LBA−2542A

Figure 2-64. Utility Hydraulic System

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-64

Change 3

TM 1-1520-251-10

2.79 UTILITY HYDRAULIC MANIFOLD

utility hydraulic system. A pressure transducer measures

hydraulic pressure on the pressure side of the manifold

The utility manifold is installed on the aft main fuselage

and transmits this value for display on the MPD.

deck on the right side. It stores, filters, supplies, and regu-

lates the flow of utility hydraulic fluid. Demands on the util-

2.79.6 Rotor Brake Solenoid Valves. The solenoid

ity system are much greater than those on the primary

valves are controlled by the RTR BRK switch on the pilot

system, and the utility manifold is therefore larger. The

POWER lever quadrant (fig 2-42) adjacent to the POWER

utility manifold incorporates the utility accumulator hy-

levers. When this switch is positioned to BRK, utility sys-

draulic pressure transducer and rotor brake solenoids not

tem pressure is applied to stop the rotor brake disc on the

duplicated on the primary manifold.

main transmission. When positioned to LOCK, the brake

OFF solenoid valve traps pressure between the manifold

2.79.1 Low Level and Auxiliary Isolation Valves. The

and the utility system accumulator.

low level and auxiliary isolation valves permit hydraulic

fluid to flow to external stores, ammo carrier drive, and

2.79.7 Reservoir Low Level Indicating Switch. A res-

area weapon turret. If reservoir fluid level decreases sig-

ervoir low level indicating switch is activated by the man-

nificantly, the reservoir piston, driven by IPAS air, closes

ifold reservoir piston. A minimum operating level caution

the low level valve. The auxiliary isolation valve, which

message is presented on the UFD/EUFD sent from the

normally requires two sources of pressure to permit fluid

low level indicating switch.

flow, then closes and denies hydraulic power to the area

weapon turret, external stores actuator, and ammo carrier

2.79.8 Manifold Pressure and Return Filters. Filters

drive.

on both manifold pressure and return sides have mechan-

ical dirty filter indicators for visual inspection. These indi-

cators operate on differential pressure. Only the return fil-

2.79.2 Shutoff Valve. A shutoff valve in the pressure

ter has bypass valve provision. In addition to the

line to the directional servo and tail wheel unlock actuator

mechanical/visual indicators both filters contain electrical

is actuated by the low level switch in the utility system res-

switches that provide signals to generate a caution mes-

ervoir. The utility side of the directional servo actuator and

sage on the UFD/EUFD

the tail wheel unlock actuator become inoperative if a low

utility system fluid level is sensed.

2.79.9 Fluid Level Indicator. A fluid level indicator in

the manifold housing allows visual inspection of the reser-

2.79.3 Accumulator Isolation Valve. The accumulator

voir fluid level.

isolation valve normally isolates accumulator pressure

from the rest of the utility system but allows system flow

2.79.10 Utility System Pressure Sensing Switch. A

from the pump to pass through a portion of the valve to the

pressure switch senses utility system pressure and in-

utility side of the tandem servocylinders.

forms the pilot and CPG of a low fluid pressure condition

by generating the caution message on the UFD/EUFD.

2.79.4 Override Solenoid. An override solenoid, de-

energized to the closed position, permits crew manage-

2.80 UTILITY HYDRAULIC ACCUMULATOR

ment of accumulator reserve pressure. Upon activation of

the EMERG HYD pushbutton, the override solenoid valve

NOTE

energizes open and accumulator fluid passes to the accu-

mulator isolation valve via emergency routing. In this

The accumulator should be checked on pre-

case, another portion of the accumulator isolation valve

flight and thru-flight inspections for a mini-

permits accumulator fluid to flow to the utility side of the

mum of 2600 psi prior to APU start.

servocylinders.

The accumulator stores hydraulic fluid at 3000 psi. The

2.79.5 Pressure Transducer. An accumulator hydrau-

accumulator provides damping for fluid pressure

lic pressure transducer in the manifold provides the pilot

changes, hydraulic power for rotor brake application, APU

with a continuous indication of accumulator pressure on

starting, and emergency flight control operation. The start

the accumulator hydraulic pressure indicator. During nor-

valve opens when the APU ON button is pressed to ON

mal operation, the indicated pressure is the same as the

and closes automatically at 60% APU speed.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 3

2-65

TM 1-1520-251-10

2.81 UTILITY HYDRAULIC RETURN ACCUMULATOR

The utility hydraulic return accumulator stores hydraulic

fluid at 55 psi. The accumulator dampens hydraulic pres-

sure surges caused by sudden actuation of the gun turret.

2.82 HYDRAULIC SYSTEM HAND PUMP

A hand pump is installed, next to the primary system GSE

panel, on the right side of the aircraft. The pump provides

one method of charging fluid pressure in the utility accu-

mulator as well as access for the ground crew to fill the

primary and utility reservoirs. The control lever may be

moved to any of three positions. This opens one of three

check valves to the accumulator or to either reservoir.

2.83 HYDRAULIC SYSTEM CONTROLS AND

DISPLAYS

LBA2540A

Operation of the hydraulic system is automatic except in

emergency situations, tail wheel locking/unlocking, and

rotor brake activation. Hydraulic system indications are

displayed on the ENG page ground format (fig. 2-65) and

Figure

2-66.

SYS Page

SYS page (fig 2-66).

2.83.1 ENG Page Ground Format.. See Section III for

descriptions of ENG page ground indications.

2.83.2

SYS Page. The following hydraulic pressures

(HYD PSI); primary (PRI), utility (UTIL) and accumulator

(ACC) are indicated on the ENG SYS page:

NOTE

The following descriptions of SYS page hy-

draulic system indications give ranges of

those displays. Refer to Chapter 5 for sys-

tem limits and restrictions.

PRI/UTIL/ACC PSI

0 to 6000

displayed in increments of 10 psi.

3410 - 6000

YELLOW ≥5 seconds RED w/box

3310 - 3400

YELLOW ≥5 minutes RED w/box

1260 - 6000

GREEN

0-1250

RED w/box

2.83.3 Emergency Hydraulics. An override solenoid

LBA3000

valve, normally de-energized closed, permits crew man-

agement of accumulator reserve pressure. Upon activa-

tion of the EMER HYD ON button on the EMERGENCY

Figure

2-65.

ENG Page Ground Format

panel (fig 2-67), the button will become

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-66

TM 1-1520-251-10

illuminated, the valve energizes open and accumulator

from being locked when the power levers are in any posi-

fluid passes via emergency routing to the utility side of the

tion except IDLEor OFFWhen the switch is set atBRK so-

servo actuators.

lenoid valves in the utility hydraulic manifold operate and

applies 337 psig to actuate the the brake. When rotors are

stopped, the switch may be set to LOCKwhich causes the

solenoid valves in the manifold to de-energize and all

available utility hydraulic system or accumulator pressure

to be applied to the brake. With the switch atOFF, the only

hydraulic pressure to the brake is 30 psig from the pres-

surized air system which, when operating, pressurizes the

return side of the utility hydraulic system. If helicopter

power is lost, the rotor brake, if previously set at LOCK re-

mains locked as long as accumulator pressure is avail-

LBA0021A

able.

Figure 2-67. EMERG HYD Pushbutton

2.83.5 Tail Wheel Lock. Utility hydraulic pressure is

used by the tail wheel lock actuator to unlock the tail

wheel. The tail wheel can be locked or unlocked from ei-

CAUTION

ther crew station via theTAIL WHEEL panel or collective

grip switch (refer to Section I). The tail wheel can also be

Do not place theRTR BRK switch in LOCK

locked or unlocked by ground crew using a handle pro-

position with rotors turning.

vided on the locking device.

NOTE

2.84 INTEGRATED PRESSURIZED AIR SYSTEM

When engaging rotor lock, pause in the

(IPAS)

BRK position until the RTR BRK advisory

message is displayed prior to placing the

Pneumatic power for the IPAS (fig 2-68) is generated by

switch in the LOCK position. The POWER

dual engine bleed air, APU or AGPU air. The IPAS pres-

levers will not advance past the ground idle

detent with the rotor brake switch in the

surizes, regulates, and distributes air to the following:

LOCK position.

• Air turbine starter

2.83.4 Rotor Brake. The rotor brake is a disc brake

• Fuel boost and transfer pumps

mounted at the aft end of the main transmission. The rotor

• External fuel tanks

brake reduces turnaround time for aircraft loading and

servicing and prevents windmilling of the rotor system

• Hydraulic reservoirs

during gusty wind conditions. TheRTR BRK switch on the

pilotPOWERlever quadrant has three modes:OFF, BRK,

• Engine inlet anti-ice

and LOCK. The BRK mode reduces time required to stop

• Ice detect probe aspirator

the main rotor system after engine shutdown. TheLOCK-

mode is used to prevent windmilling in strong winds and

• Nitrogen inerting unit

for locked rotor dual engine starts. When both engines are

• Engine firewall/cooling

at idle and the switch set atLOCK (full 3000 psig utility hy-

draulic system pressure), the brake prevents the drive

• Utility receptacle

train and power turbine from being driven by the gas tur-

bine. A system of three interlocks prevents the rotor brake

• Environmental Control System (ECS)

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-67

TM 1-1520-251-10

FIREWALL

ENG INLET

AIR TURBINE

ANTI−ICE

STARTER (ATS)

ATS PRESSURE

SOV

REGULATOR SOV

HIGH PRESSURE

CHECK VALVE

TO ENG COOLING

ENGINE 2

DOOR ACTUATOR

(RH)

RESTRICTOR

VENTURI

ENG BLEED

FLOW SHARE

PRESSURE

REGULATOR

SENSE LINE

ACCESSORY

SOV

ENG BLEED

GEARBOX

AIR CHECK

INTERFACE

VALVE

APU

FIREWALL

HYDRAULIC

APU

APU LOAD

RESERVOIR

CREWSTATION

CHECK

CONTROL

CHECK

CAUTION INDICATION

ATS PRSOV POSITION

VALVE

SYSTEM

VALVE

ENGINE Ng

PRESSURE

ATS PRSOV CLOSE

PROCESSOR

OVERTEMP SIGNAL

HIGH PRESSURE

ENGINE

CHECK VALVE

BLEED AIR

FIREWALL

PRECOOLER

AND FAN

BLEED AIR

OVERTEMP

ENG BLEED

SWITCH

ENG BLEED

AIR CHECK

PRESSURE

VALVE

REGULATOR

FLOW SHARE

SOV

SENSE LINE

RESTRICTOR

VENTURI

PRECOOLER

BYPASS

PRECOOLER BYPASS

TO ENG COOLING

ENGINE 1

CONTROL

TEMPERATURE

DOOR ACTUATOR

VALVE

THERMOSTAT

(LH)

FIREWALL

ATS PRESSURE

ENG INLET

AIR TURBINE

REGULATOR

ANTI−ICE

STARTER (ATS)

SOV

NOTES:

UTILITY

1. SYMBOL

DENOTES

RECEPTACLE

28 Vdc SOLENOID

GROUND CART

ACTUATED VALVE.

CONNECTION

LBA02911

Figure 2-68. Integrated Pressurized Air System (Sheet 1 of 2)

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-68

TM 1-1520-251-10

2.84.1 Dual Engine Bleed Air. The IPAS subsystem

primary pneumatic power source is dual engine bleed air.

Bleed air is provided by both engines during normal op-

eration.

2.84.2

Single Engine Bleed Air. The secondary

source of pneumatic power for the IPAS is single main en-

gine bleed air.

2.84.3 Power Up. During aircraft power up, APU bleed

air is the pressurized air source and is used to start main

engines. External air from an AGPU or another aircraft

may be used to start engines via the external air recep-

tacle. Each engine has a low and high pressure bleed air

port. The high pressure port is used exclusively to pres-

surize the hydraulic reservoirs and the low pressure port

supports the remaining functions. Low pressure flow and

pressure is controlled by the engine bleed pressure regu-

lator and shutoff valve. High pressure flow and pressure is

controlled by a restrictor and a regulator.

2.84.4 IPAS Contro. Control of IPAS is an integrated

LBA - 2083

function provided by system processors, display proces-

Figure 2-69. A/C UTIL Page (Pilot)

sors, and the Electrical Power Management System

(EPMS). The A/C UTIL page (fig 2-69) provides BLEED

2.84.5 PAS Cautions. Crewmembers are provided the

AIR 1 and 2ON/OFF pushbuttons.

following IPAS advisory messages via the Warning/Cau-

tion/Advisory system:

a. Engine Bleed Air Fail. This advisory message

occurs when engine 1 and/or engine 2 primary shutoff

valve is in the commanded position and the system pro-

cessor commands the engine 1 and/or engine 2 bleed air

shutoff valve open or closed and it does not move.

b. Engine Bleed Air Overtemperture. When the

system processor receives a bleed air overtemperature

signal, it will send a caution message to the UFD/EUFD.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-70

Change 2

TM 1-1520-251-10

Section VII. DRIVE TRAIN SYSTEM

2.85 INTRODUCTION

• Tail rotor gearbox

The drive train (fig 2-70) transmits engine power to the ro-

tors and to accessories mounted on the transmission. The

2.86 MAIN ROTOR DRIVE SYSTEM

drive train includes the following:

• Two engine nose gearboxes

2.86.1

Engine Nose Gearboxes. One engine nose

gearbox is mounted on the front of each engine. They re-

• Two input shafts

duce drive shaft speed and change the angle of the drive.

• APU drive shaft and couplings

Both nose gearboxes have self contained pressurized oil

systems with provisions to ensure limited operation if a to-

• Couplings and input clutch to the main transmis-

tal loss of pressurized lubrication occurs. The input drive

sion

shafts have flexible couplings that require no lubrication.

• Main transmission

Sensors and detectors monitor the nose gearboxes and

provide information to crewmembers about oil tempera-

• Main rotor drive shaft

ture, oil pressure, and the presence of metal chips. High

• Tail rotor drive shafts

oil temperature, low oil pressure, and presence of metallic

chips in the gearbox cause cautions to be annunciated in

Intermediate gearbox

the crew stations.

TAIL ROTOR

STATIC MAST

TAIL ROTOR

GEARBOX

FLEXIBLE COUPLING

(TYPICAL)

TAIL ROTOR AND

INTERMEDIATE BEARBOX

COOLING FAN

MAIN ROTOR

DRIVE SHAFT

TAIL ROTOR

DRIVE SHAFT

HANGER BEARING

STATIC MAST

ANTIFLAIL

INTERMEDIATE

SLEEVE

GEARBOX

ANTI−FLAIL

SLEEVE

COOLING

FANS (2)

DRIVE SHAFT

DAMPER

DRIVE SHAFT

DAMPER

APU DRIVE SHAFT

MAIN

TRANSMISSION

NOSE

GEARBOX (2)

INPUT DRIVE SHAFTS (2)

LBA0115

Figure 2-70. Drive Train

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-71

TM 1-1520-251-10

2.86.2 Main Transmission. The main transmission is

mounted below the main rotor static mast base which al-

lows its removal without removing the upper controls,

mast or hub. The main transmission combines the two en-

gine nose gearbox inputs, reduces output RPM, and pro-

vides drive to the main rotor, tail rotor, accessories, and

rotor brake disc. Sensors in the transmission monitor

conditions and provide crewstation caution messages.

a. Main Rotor Drive Shaft. The main rotor drive

shaft is designed to carry torque loads only. The rotor hub

is on a static mast which carries vertical or bending loads.

The drive shaft rotates inside the static mast.

b. Reduction Gearing. The main transmission has

a three stage primary reduction gearing with two engine

inputs, main rotor, and tail rotor power output. An over run-

ning clutch provides APU drive to the accessory section of

the transmission when the rotor is stopped.

LBA2610A

c. Accessory Gearbox. The accessory gearbox is

driven by the APU drive shaft or engine drive shafts and

Figure 2-71. SYS Page

provides shaft power to the main generators and the hy-

draulic pumps while the rotor is stationary or rotating. The

oil pressure in the gearbox is monitored by the accessory

NGB1 OR 2 OIL PRESSURE (PSI)

oil pressure switch. Reduction gearing and drive shaft

coupling is provided.

0 to 100 psi Resolution 1 psi.

30 - 100 psi Normal operation (GREEN)

d. Main Transmission Lubrication. The main

30 psi Minimum (<30 RED w/box)

transmission has two independent oil systems. Each sys-

tem has its own sump, pump, filter, and heat exchanger.

Oil level sight gages are located in the transmission hous-

NGB1 OR 2 OIL TEMPERATURE (°C)

ing at each oil sump. These systems are not totally inde-

- 32 to 149

Resolution 1_

pendent in the usual sense because during normal opera-

134

Maximum (>134 RED w/box)

tion, the oil mixes. If oil loss occurs in either sump or in

either heat exchanger, the diverter (float) valve will seal off

0-134

Normal operation (<134 GREEN)

that sump to prevent a total loss of oil.

2.87.2 Transmission Oil Status Indications. The fol-

2.87 DRIVE TRAIN CONTROLS AND DISPLAYS

lowing main transmission oil status indications are dis-

played on the SYS page:

Engine nose gearbox and main transmission systems in-

dications are displayed on the SYS page (fig 2-71).

XMSN 1 OR 2 OIL PRESSURE (PSI)

0 to 100 psi Resolution 1 psi

NOTE

30 - 100 psi Normal operation (GREEN)

The following descriptions of SYS page en-

30 psi Minimum (<30 RED w/box)

gine(s) NGB and main transmission indica-

tions give ranges of those displays. Refer to

Chapter 5 for system limits and restrictions.

XMSN 1 OR 2 OIL TEMPERATURE (°C)

- 32 to 149

Resolution 1_

2.87.1 Engine NGB Oil Status Indications. The fol-

134

Maximum (>134 RED w/box)

lowing engine oil status indications are displayed on the

- 32 - 134

Normal operation (<134 GREEN)

SYS page:

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-72

TM 1-1520-251-10

2.88 TAIL ROTOR DRIVE SYSTEM

2.88.2 Intermediate Gearbox. The intermediate gear-

box reduces RPM and changes the angle of drive to the

tail rotor. The intermediate gearbox is a grease lubricated

The tail rotor drive system consists of the tail rotor drive

sealed unit. Four thermistors monitor temperature and an

shafting, couplings, hanger bearing, dampers, anti-flail

accelerometer measures vibration to provide crewmem-

assemblies, and intermediate and tail rotor gearboxes.

bers with UFD/EUFD caution messages.

2.88.1 Tail Rotor Drive Shaft. There are four tail rotor

drive shaft sections. Three tail rotor drive shaft sections

CAUTION

lead from the transmission to the intermediate gearbox.

Two are of equal length. The fourth section is installed on

Prolonged OGE hover (20-30 minutes)

the vertical stabilizer between the intermediate and tail ro-

with outside air temperature above 755 F

tor gearboxes. Hanger bearings support the longer shafts.

(245 C) may cause the tailrotor gearbox

The two equal length shafts incorporate friction dampers

to overheat.

and anti-flail assemblies. Flexible couplings, attached to

the shaft ends, are capable of accommodating shaft mis-

2.88.3

Tail Rotor Gearbox. The tail rotor gearbox,

alignments throughout the power range.

mounted on the vertical stabilizer, reduces the output rpm

and changes the angle of drive. The tail rotor output shaft

passes through the gearbox static mast. All tail rotor loads

CAUTION

are transmitted to the static mast. The output shaft trans-

mits only torque to the tail rotor. Lubrication of this gear-

Prolonged OGE hover (20-30 minutes)

box is identical to that of the intermediate gearbox. As with

with outside air temperature above 755 F

the intermediate gearbox, four thermistors monitor tem-

(245 C) may cause the intermediate gear-

perature and an accelerometer measures vibration to pro-

box to overheat.

vide crewmembers with UFD/EUFD caution messages.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 2

2-73

TM 1-1520-251-10

Section VIII. ROTORS

2.89 ROTOR SYSTEM

move horizontally. The pin goes through the “V” portion of

each strap within the pitch housing.

The rotor system (fig 2-72) consists of a four bladed, fully

articulated main rotor and a four bladed tail rotor assem-

2.90.3 Damper Assembly. Two damper assemblies con-

bly with two teetering rotor hubs.

trol lead-lag movement of each main rotor blade. Each

damper attaches outboard to a link lug and inboard to a

2.90 MAIN ROTOR ASSEMBLY

trunnion at the pitch housing. The dampers contain elas-

tomeric elements that distort to allow the blade to lead or

The main rotor has four removable blades. The rotor head

lag.

allows the four blades to flap, feather, lead, or lag inde-

pendently. The head consists of a hub assembly, pitch

2.90.4 Main Rotor Blades. The outboard tip is swept

housings, rotor dampers, and lead-lag links. The main ro-

aft 20° and tapers to a thinner section. Tip weights are in-

tor is controlled by the cyclic and collective control sticks

stalled in the blades. Each blade is secured to its lead-lag

through a swashplate mounted about the static mast. This

link by two blade attachment pins. These pins can be re-

arrangement allows the static mast, rather than the main

moved without the use of tools. Sets of five doublers are

rotor drive shaft, to assume all flight loads. The hub is

located on the upper and lower surfaces of the blade at

splined to the main rotor drive shaft by a drive plate adapt-

the blade root.

er bolted to the hub. The hub is secured to the static mast

by a large locknut secured with multiple bolts. The hub

2.91 TAIL ROTOR ASSEMBLY

houses two sets of grease lubricated, sealed roller bear-

ings that transfer hub loads to the static mast. Mechanical

droop stops limit blade droop. When blade droop occurs,

The tail rotor system is of semi-rigid, teetering design. Two

a striker plate on the pitch housing contacts a roller. The

pairs of blades, each pair fastened to its own delta hinged

roller presses a plunger against a droop stop ring on the

hub, provide anti-torque action and directional control. A

lower portion of the hub.

titanium fork houses four elastomeric teetering bearings

and drives the rotating swashplate through an attached

2.90.1 Pitch Housing. The pitch housing permits blade

scissors assembly. The tail rotor assembly is splined to,

pitch changes in response to flight control movements

and driven by, the tail rotor gearbox drive shaft which

transmitted through the swashplate. This is made possi-

passes through a static mast. Blade pitch changes are

ble within the four pitch housings by “V” shaped stainless

made when directional control inputs cause the non-rotat-

steel strap assemblies that twist and flap to permit blade

ing swashplate to act upon the rotating swashplate. One

feathering, flapping, and carry the centrifugal force load.

pitch link for each blade, attached to the rotating swash-

Cyclic and collective stick inputs are transmitted to the

plate and pitch horn, causes blade movement about two

pitch housing horns by pitch links attached to the swash-

pitch change bearings in the blade root. Centrifugal forces

plate. Feather bearings are installed inboard on the pitch

are carried by strap assemblies attached outboard to the

housing to allow vertical and horizontal loads to be trans-

blade root and inboard at the hub center. An elastomeric

ferred from pitch housing to the hub. Centrifugal loads are

bearing assembly positions the hub and strap pack in the

transmitted by each strap assembly to the hub.

tail rotor fork. Each blade has one stainless steel spar and

two aluminum spars. Doublers, adhesive, and rivets at-

2.90.2 Lead Lag Links. Lead-lag links are connected

tach the blade to the blade root. Brackets on the root fitting

to the outboard end of each pitch housing and are secured

hold chord-wise balance weights. Span-wise balance

in place by a pin and two bearings allowing the links to

weights are installed in blade tip caps.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-74

TM 1-1520-251-10

Section IX. UTILITY SYSTEMS

2.92 ANTI-ICE SYSTEM

a. AUTO Mode. When in the AUTO mode, ANTI-

ICE system buttons are not selectable when an icing con-

ditions is detected. The buttons are selectable in this

The anti-ice protection systems provides crewmembers

mode when no icing condition is detected. The SP uses

with automatic and manual means to prevent ice accu-

the ice status signal from the ice detector signal processor

mulation (anti-ice). Anti-ice subsystems are installed in

and FAT status from the FMC to determine when to acti-

the airspeed sensors, engine inlets, engine nose gear-

vate the anti-ice systems. Once activated, anti-ice sys-

boxes, canopy, sensor shrouds and windows of TADS/

tems are not automatically set to off; the systems must be

PNVS.

manually selected to off.

2.92.1 Ice Detect Probe. The Ice Detect Probe aspira-

tor will become active when the Free Air Temperature

b. MANUAL Mode. When in MANUAL mode, the

(FAT) decreases to 5° C or less. When the ice detect

ANTI-ICE system buttons are selectable. When entering

probe senses ice, it sends a discrete ice detect signal and

the MANUAL mode, any anti-ice system that is currently

an analog icing rate signal to the SP. In the AUTO mode,

ON (activated in the AUTO mode), will remain in an ON

the SP commands all anti-icing functions to an ON state

state.

when the icing rate signal indicates an icing condition. The

probe will be deactivated when the FAT increases to 7° C

c. FAT Status Window. The current FAT value is

or higher. The SP will not command the anti - icing func-

displayed in the FAT status window. The range value of

tions to OFF. In MANUAL control, the anti-ice system op-

FAT is from -50° C to +50° C.

eration is independent of icing conditions.

2.92.2 Anti-Ice System Controls. The A/C UTIL page

d. ICE Status Window. The current ICE status is

(fig 2-73) ANTI-ICE SYSTEM button toggles the system

displayed in the ICE status window. The state of the ICE

between MANUAL and AUTO modes.

status is displayed in conditions of TRACE (GREEN),

LIGHT (WHITE), MODER (YELLOW), or SEVERE

(RED).

2.92.3 Airspeed Sensors Anti-Ice. The airspeed sen-

sors anti-ice system prevents formation of ice that could

cause false indications from the pitot tubes and Air Data

Sensors. When PITOT is selected, power is applied to

heat the pitot tubes and ADS, which prevents ice forma-

tion on the tubes and sensors.

2.92.4 Engine INLET Anti-Ice System. The engine

anti-ice system includes the engine, engine inlet fairings,

and nose gearbox fairings. Engine fifth stage bleed air is

used to heat the swirl vanes, nose splitter, and engine inlet

guide vanes on each engine. The nose gearbox fairing is

electrically heated to prevent ice from forming on the sen-

sors.

2.92.5 CANOPY Anti-Ice. Canopy windshield anti-ice

is incorporated into the pilot and CPG middle forward

LBA0651A

looking windshields. Heating elements and sensors are

embedded into the windshield laminates. When CANOPY

is selected, heat produced by the elements prevents ice

Figure

2-73.

A/C UTIL Page (Pilot)

formation on the windshields.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-76

TM 1-1520-251-10

2.93.1 Windshield wipers. Two electrically driven wip-

WARNING

ers are mounted on the canopy frame to remove moisture

from the two windshields. The wipers have two speeds

Do not touch TADS/PNVS shroud win-

and a park position and are controlled by the WIND-

dows. Electrical shock can result, and

SHIELD panel WIPER rotary switch (fig 2-74) in each

heaters in these fairings can cause seri-

crew station.

ous burns. If shock or burns occur, seek

medical aid.

2.93.2 Canopy Defog. Pressurized hot air is mixed

with crew station conditioned air and directed against the

2.92.6 TADS/PNVS SENSOR Anti-Ice.

canopy side panels to defog them. Canopy defog is con-

trolled by the WINDSHIELD panel DEFOG pushbutton

a. Inflight. SENSOR anti-ice prevents ice formation

switch (fig 2-74) in each crew station.

on turret shrouds, boresight and sensor modular win-

dows. Anti-icing is accomplished via thermostatically con-

trolled heating elements in the shrouds and electronically

regulated power through the conductive window coatings.

ÑÑÑ

b. GND Position. The SENSOR anti-ice is inhibited

from operation when the helicopter is on the ground. This

ÑÑÑ

inhibit may be overridden by using the sensor anti-ice

ÑÑÑ

ground (GND) override on/off button (fig 2-73).

2.93 RAIN REMOVAL

LBA0024

The rain removal system consists of two Windshield wip-

ers and canopy defog.

Figure 2-74. Windshield Panel

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-77

TM 1-1520-251-10

Section X. ENVIRONMENTAL CONTROL SYSTEM

2.94 ENVIRONMENTAL CONTROL SYSTEM (ECS)

• Operating the aircraft in a loose grass en-

vironment could cause the ECS condens-

er inlet to become clogged. This could re-

The ECS (fig 2-75) provides crew station ventilation, heat-

sult in a single or double condenser

ing, and Vapor Cycle Cooling System (VCCS) air condi-

overtemp condition.

tioning. The VCCS is the primary source of cool air for the

Extended Forward Avionics Bays (EFAB) and the TADS/

PNVS. No heating is required in the EFABs.

2.94.1 ECS Normal Operation. Environmental cooling

is provided by two independently operated VCCS sys-

tems. Crewstation heating uses bleed air from the main

NOTE

engines or APU via the IPAS. System control is performed

by the Digital Control Unit (DCU). One VCCS system will

• During engine starts in the heating mode,

service the aft right and left hand EFAB and the pilot crew

conditioned air will cease to be provided

station. The other system will service the forward right and

until starter drop-out. Non-conditioned

left hand EFAB, CPG crew station and the TADS/PNVS.

airflow will continue.

The aft avionics bays are supplied ambient air circulation

• When a canopy door is unlatched, cool-

via a fan. Temperature sensors, located in each crew sta-

ing air will not be provided to that

tion and EFAB, provide temperature status to the ap-

crewstation (heating mode is unaffected

propriate VCCS system allowing independent crew sta-

by canopy position).

tion and EFAB cooling.

RECEIVER/FILTER/DRYER (2 PL)

AFT EVAPORATOR WITH

THERMAL EXPANSION

VALVE/FAN (2 PL)

EVAPORATOR WITH

THERMAL EXPANSION

VALVE/BLOWER (2 PL)

CONTROLLER (2 PL)

BLOWER (4 PL)

CONDENSER (2 PL)

OIL SEPARATOR (2 PL)

COMPRESSOR (2 PL)

FWD EVAPORATOR WITH

THERMAL EXPANSION

VALVE/FAN (2 PL)

LBA2391

Figure 2-75. Environmental Control System (ECS)

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-78

TM 1-1520-251-10

2.94.2 ECS Controls and Displays. Controls and dis-

b. TEMP Set Button. After crewmember selects

plays required to manage the ECS are provided on the

TEMP button, the desired crew station temperature is en-

A/C UTIL page (fig 2-76) and the SYS page. ECS controls

tered in 1° increments, 50° F to 90° F, via the Keyboard

consist of an ECS ON/OFF button, a data entry TEMP

Unit (KU). The crewmember then selects ENTER on the

(temperature) set button and a current crew station TEMP

KU and the new temperature setting will be displayed

status window located on the A/C UTIL page. Individual

above the TEMP button. Upon aircraft power up, TEMP

crew compartment air temperature readouts are dis-

will default to the last value set prior to shutdown. This

played on the SYS page. ECS variables involving the

data entry is independent in each crewstation.

equipment, EFABS, and crewstations can be viewed on

the ECS page.

c. Temperature Status Windows. The current

crew station ambient temperature is displayed in the

TEMP status window on the A/C UTIL page (fig 2-76).

The range value for the PLT/CPG CKPT TEMP status

window is from - 65° F to +160° F in 1° increments. Each

crewstation and EFAB compartment (left/right/forward

and aft) temperature is displayed in the ECS TEMP° F sta-

tus window on the SYS page (fig 2-77). An EFAB compart-

ment temperature is displayed in YELLOW when it is

greater than 105°. The range value for the ECS TEMP° F

window is from - 65° F to +160° F and displayed in 1° in-

crements.

LBA2547

Figure 2-76. A/C UTIL Page (Pilot)

a. ECS ON/OFF Button. The ECS button toggles

LBA2392A

the ECS between ON and OFF. Upon aircraft power up,

the ECS defaults ON.

Figure 2-77. SYS Page

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-79

TM 1-1520-251-10

2.95 CREWSTATION BACKUP COOLING

CAUTION

During most ECS failure modes, the EFAB and crewsta-

tion blowers will continue to circulate air to each compart-

To obtain optimum performance from the

ment, and the Air Particle Separator (APS) will provide

heating and cooling system, the gaspers

ambient air to the crewstations. There is one interconnect

must be full open. Adjusting the temper-

valve located in the air distribution ducting between the

ature level rather than closing gaspers

two crewstations. In the event of a cockpit cooling failure,

will increase the reliability of the blower.

the Digital Control Unit (DCU) will send a signal to the in-

Closing two or more gaspers will cause

terconnect valve to fully open the valve, thus allowing air-

the blower to operate in a stall region, in-

flow between the two cockpits. The crewstation blower to

creasing noise levels. Operating with all

the affected crewstation will shut down and allow the re-

gaspers closed may cause the blower to

maining crewstation blower to supply both cockpits with

overheat and shutdown.

conditioned air. It is possible with an Air Particle Separator

and ECS failure that the blowers may be disabled and no

d. Crewstation Gasper Adjustment. Each

fresh air is provided.

crewstation air distribution system features two torso

gaspers, two head gaspers and two leg vents. The torso

2.95.1 ECS Failure. The system begins to monitor for

and head gaspers may be manually adjusted open or

ECS failure when generator power is applied. The ECS

closed for airflow regulation and may be positioned up to

failure caution occurs when the system processor detects

60° off centerline in any direction for crew comfort. The

various ECS failures. Failures of the forward and aft ECS

torso gaspers are located on the front instrument panel

systems are usually temporary and the blowers continue

facing the crewmember, while the head gaspers are lo-

to operate. Failure of the ECS control is usually perma-

cated over each shoulder of the crewmember. The leg

nent and the blowers may be disabled. The UFD/EUFD

vents are located above both legs of each crewmember

and MPD will display the related caution message.

and are not adjustable.

2.96 VENTILATING SYSTEM

In the event of an ECS failure, the EFABS and crew sta-

tion blowers will continue to circulate air to each compart-

ment, and the Air Particle Separator (APS) will provide

ambient air to the crew stations. In the event of an ECS

and APS failure the blowers may be disabled and no fresh

air is provided.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-80

Change 2

TM 1-1520-251-10

Section XI. ELECTRICAL SYSTEM

2.97 INTRODUCTION

ing Modules (HPSM) and two Electrical Load Centers

(ELC). On each HPSM is mounted a Circuit Breaker Panel

The electrical system (fig 2-78) produces and distributes

(CBP). The EPMS and the CBPs provide distribution for

all of the electrical power required for operation of the heli-

AC, DC, battery power, and ground power. EPMS opera-

copter. The electrical system provides: AC power genera-

tion is fully automated. Because of redundancy, there are

tion, DC power generation, and battery power.

no normal indications for the EPMS subsystem). Failures

within the EPMS are displayed on the MPD Data Manage-

2.98 ELECTRICAL POWER MANAGEMENT SYSTEM

ment System (DMS) page (refer to Section XVI). Busses

(EPMS)

1, 3 and 5 are tied together in each of the respective pow-

er distribution systems as are busses 2, 4 and 6

The EPMS (fig 2-79) consists of two High Power Switch-

(Figs 2-80,2-81 ,2-82 ,2-83).

CIRCUIT BREAKER

PANEL q1 AND

TRANSFORMER

PANEL q2 AND

HIGH POWER

RECTIFIER q2

HIGH POWER

SWITCHING

MODULE q1

MODULE q2

GENERATOR 2

GENERATOR 1

BATTERY

CHARGER

BATTERY

ELECTRICAL

LOAD CENTER q2

GENERATOR

CONTROL UNITS

ELECTRICAL

LOAD CENTER q1

TRANSFORMER

LIGHTING

RECTIFIER q1

CONTROLLER

LBA2396

Figure 2-78. Electrical System Components

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-81

TM 1-1520-251-10

115/200 VAC 3PHASE 400 HERTZ

AVIATION GROUND POWER UNIT

115/200 VAC 3PHASE 400 HERTZ

NO. 1 AC

NO. 2 AC

GENERATOR

EXTERNAL

POWER

GCU 1

MONITOR

GCU 2

SQUAT

SWITCH

OVERSPEED TEST PANEL

NO. 1 & NO. 2

GENERATOR

RESET

HIGH POWER

CHK OVSP TEST

GEN RST

HIGH POWER

ÍÍ

ENG 1

SWITCHING

OFF

GEN 1

GEN 2

SWITCHING

CKT A

ENG 2

CKT B

MODULE NO. 1

MODULE NO. 2

ÍÍ

FAIL

NO. 1 REGULATED

NO. 2 REGULATED

TRANSFORMER

AC CONTACTOR

TRANSFORMER

RECTIFIER UNIT

DC BUS TIE

FAIL

115/200 VAC

FIBROUS

115/200 VAC

3PHASE

DC CONTACTOR

NICKEL

DC CONTACTOR

3PHASE

TO 28 VDC

CADMIUM

TO 28 VDC

BATTERY

FAIL

DC BUS NO. 1

DC BUS NO. 2

CAUTIONS

WARNINGS

ADVISORIES

PILOT’S POWER

UPFRONT DISPLAY

QUADRANT

VIEW

RTS

ENG1 OUT CAUTIONS ADVISORIES

ÄÄBAT BUS NO. 1Ä

ÄBAT BUS NO. 2ÄÄ

MASTER IGNITION

H8G65

139.50

149.500249.500

RTS

P3R56 C3 GJ8L52 C5 H

240.07545.07

45.550

LAST

KEY SWITCH

F1F2

B5Z23 C2

49.50XPNORM 5000 B

10:56:01 L

69.525

ÄÄÄÄÄ

ÍRTR B

OFFBRK

LOCK

GEN1 FAIL

LOCK

OUT

BAT CONTACTOR

GEN2 FAIL

MSTR IGN

FLY

OFF

INC

RECT1 FAIL

BATT

RECT2 FAIL

EXT

FAIL

PWR

BATTERY

APU

CHARGER

IDLE

DECR

MULTIPURPOSE DISPLAY

ENG START

STARTOFF

OFF

BATTERY CHARGER

DAY

BRT

NIGHT

CON

ÍNO.IGN ORIDE. 2 Í

SYMBRT

AVIONICS DATA BUS

FCR

VID

WPN

COM

TSD

A/C

BATTERY FAIL

BATTERY CHARGER FAIL

GENERATOR 1 FAIL

GENERATOR 2 FAIL

LEGEND:

GENERATOR CONTROL UNIT 1 FAIL

GENERATOR CONTROL UNIT 2 FAIL

TRANSFORMER/RECT 1 FAIL

TRANSFORMER/RECT 2 FAIL

HPSM 1 _ _ _ _ FAIL

DC POWER

HPSM 2 _ _ _ _ FAIL

ELECTRIC LOAD

ELC 1 _ _ _ _ FAIL

CENTER NO. 1

ELECTRIC LOAD

BATT POWER

ELC 2 _ _ _ _ FAIL

CENTER NO. 2

LBA2561

Figure 2-80. Electrical Power Distribution

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-83

TM 1-1520-251-10

• ECS COMP 1

• ECS COMP 2

• FCR

• ELC2 AC

• ELC1 AC

• AIR DATA SENSOR HTR

• P2 1760 AC

• P1 1760 AC

• CPG FLT CONT (REF B)

• AIR DATA SENSOR HTR

• WP/SP1&2/STAB SYN (REF C)

• P2 1760 AC

• ICE PROBE HEATER

• CPG FLT CONT (REF B)

•

AWS AC

• P3 1760 AC

• P4 1760 AC

• GUN MOTOR

• AFT COND BLOWER B

• ANTICOLL/ FORM LTS

• DISPL PROC q2

• ENG1 OVSP

• IHADSS DEU

• LASER EU

• NITROGEN INERT

• PNVS EU

• ENG2 OVSP

• DISPL PROC q1

• AFT COND BLOWER A

• NOSE GRBX HEATER 1

• ORT

• PLT LH MPD

• CPG LH MPD

• IHADSS SEU

• VIDEO RCDR

• CPG RH MPD

• GND SVC RECPT AC

• FWD COND BLWR A

• AMMO MOTOR

• FWD LH/RH FAB FANS

• AFTLH/RH FAB FANS

• CPG CKPT FAN

• FANENG BLEED AIR

• AIR PARTICLE SEP FAN

• NOSE GRBX HEATER 2

• FWD COND BLWR B

• PLT RH MPD

• RH/LH FANSAFT BAYS

• BATTERY CHARGER

• TADS PWR SUPPLY

• PLT CKPT FAN

LBA2562

• CANOPY HEATER

Figure 2-81.

[

BLK 1

AC Power Distribution]

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-84

Change 2

TM 1-1520-251-10

•

AFT LH FAB FAN

•

AFT COND BLOWER A

•

AIR PARTICLE SEP

•

AFT COND BLOWER B

•

ANTI COLL/FORM LTS

•

AFT RH FAB FAN

•

CPG CKPT FAN

•

AMMO MOTOR

•

CPG RH MPD

•

BATTERY CHARGER

•

DISPL PROC #1

•

CANOPY HEATER

•

ECS COMP 1

•

CPG LH MPD

•

ELC1 AC

•

DISP PROC #2

•

ENG 1 OVSPD

•

ECS COMP 2

•

FCR

•

ELC2 AC

•

FWD COND BLWR A

•

ENG2 OVSPD

•

FWD COND BLWR B

•

FANENG BLEED AIR

•

FWD LH FAB FAN

•

FWD RH FAB FAN

•

GUN MOTOR

•

GND SVC RECPT AC

•

HF ARC220 FAN

•

IHADSS DEU

•

IHADSS SEU

•

NITROGEN INERT

•

LASER EU

•

NOSE GRBX HEATER 2

•

NOSE GRBX HEATER 1

•

PLT CKPT FAN

•

PLT LH MPD

•

PLT RH MPD

•

PNVS EU

•

RH/LH FANS AFT BAYS

•

TADS PWR SUPPLY

• AIR DATA SENSOR

• AWS AC

• IAFS FUEL PUMP

• CPG FLT CONT (REF B)

• ICE PROB HEATER

• P1 1760 AC

• P2 1760 AC

• P3 1760 AC

• P4 1760 AC

• WP&SP 1&2/STAB SYN

• PLT FLT CONT (REF A)

LBA5167

Figure 2-81A.

[

BLK 2

AC Power Distribution ]

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 2

2-84.1

TM 1-1520-251-10

DC BUS NO. 1

DC BUS NO. 2

•

IR JAM SOURCE/ELX/FAN

•

ELC2 DC

•

PSP/LPRF

•

RFI

•

ELC1 DC

DC BUS NO. 3

DC BUS NO. 4

•

ENG1 AI/START BL V/ NGB HTR 1

•

ENG2 AI/START BL VLV/ NGB HTR 2

•

LW ROCKET ARM

•

IHADSS DEU PWR

•

LWT 1760 DC2

•

RW PITOT HTR

•

RDR WRN RCVR

•

RWT 1760 DC2

•

TADS STANDBY

•

RW ROCKET ARM

•

CANOPY HEAT/DEFOG V

•

ICE SIG PROC/ ASP SOV

•

CHAFF ARM

•

RWT 1760 DC1

•

LWT 1760 DC1

•

IHADSS SEU PWR

•

FUEL VALVE

•

RDR ALTM PWR

•

P2 1760 DC1

•

P3 1760 DC1

•

P2 1760 DC2

•

P3 1760 DC2

•

P1 1760 DC1

•

P4 1760 DC1

•

P1 1760 DC2

•

P4 1760 DC2

•

ELC1 DC CONT

•

ELC2 DC CONT

•

LW PITOT HTR

•

GUN ARM

•

PNVS STANBY

•

DPLR SENSOR

•

LASER WRN RCVR

DC BUS NO. 5

DC BUS NO. 6

•

EGI 1A

•

PRI LT CPG

•

STAB ACTR RH

•

NAV LTS/ANTI COL

•

DCU 1

•

PRI LT PLT

•

P3 ACTR PWR

•

P3 PIU PWR

•

P2 PIU PWR

•

GRBX XDCR

•

WSHLD WPR CPG

•

EGI 2B

•

ENG CHOP

•

RADAR JAM

•

BUCS/FMC CMPTR

•

KEYBD CPG

•

KYBD PLT

•

EGI ANT SPLITTER

•

PNVS EU

•

WPNS PROC q2

•

STAB ACTR LH

•

CPG SIGHT SW

•

AREA WPNS SYS DC

•

GND SVC RECPT DC

•

P2 ACTR PWR

•

PLT SIGHT SW

•

STAB CNTRL

•

FUZZ BURNER

•

SELECT JTSN

•

EGI 2A

•

P4 ACTR PWR

•

SIDE LOADER

•

FAN AFT BAY

•

IR JAM CONTRL

•

P4 PIU PWR

•

HADS

•

PLT CBR FLTR BLWR

•

TADS/PNVS

•

BATT HTR

•

EGI 1B

•

WPNS PROC q1

•

TESS TECU

•

DCU 2

•

ADF

•

P1 ACTR PWR

•

TESS TLIA

•

CPG CBR FLTR BLO

•

WSHLD WPR PLT

•

P1 PIU PWR

•

TESS LRFD

LBA2563

Figure

2-82.

[

BLK 1

DC Power Distribution ]

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-84.2

Change 2

TM 1-1520-251-10

DC BUS NO. 1

DC BUS NO. 2

•

DCU 1

• PLT CBR FLTR BLWR

• ADF

•

NAV LTS/ANTI COL

•

EGI 1A

• PLT SIGHT SW

• AREA WPNS SYS DC

•

P3 ACTR PWR

•

EGI 2A

• PRI LT CPG

• BATT HTR

•

P3 PIU PWR

•

EGI ANT SPLITTER

• PRI LT PLT

• BUCS/FMC CMPTR

•

P4 ACTR PWR

•

ELC1 DC

• PSP/LPRF

• CPG CBR FLTR BLO

•

PNVS EU

•

ENG CHOP

• RADAR JAM

• CPG SIGHT SW

•

PR PIU PWR

•

FAN AFT BAY

• RFI

• DCU 2

•

SIDE LOADER

•

GRBX XDCR

• SELECT JTSN

• EGI 1B

•

STAB ACTR RH

•

IAFS FUEL QTY

• STAB ACTR LH

• EGI 2B

•

STAB CNTRL

•

IR JAM CONTROL

• TADS/PNVS

• ELC2 DC

•

VIDEO RCDR

•

KYBD PLT

• TESS LRFD

• FUZZ BURNER

•

WPNS PROC #2

•

P1 ACTR PWR

• TESS TECU

• GND SVC RECPT DC

•

WSHLD WPR CPG

•

P1 PIU PWR

• TESS TLIA

• HADS

•

WSHLD WPR PLT

•

P2 ACTR PWR

• WPNS PROC #1

• KEYBD CPG

•

P2 PIU PWR

DC BUS NO. 3

DC BUS NO. 4

•

CANOPY HEAT

•

DPLR SENSOR

•

CHAFF ARM

•

ELC2 DC CONT

•

ELC1 DC CONT

•

ENG2 AI/START BL

•

ENG1 A1/START BL

•

GUN ARM

•

FUEL VALVE

•

ICE SIG PROC/ASP SOV

•

IAFS CTRL

•

IHADSS DEU PWR

•

LASER WRN RCVR

•

IHADSS SEU PWR

•

LW PITOT HTR

•

P3 1760 DC1

•

LW ROCKET ARM

•

P3 1760 DC2

•

LWT 1760 DC1

•

P4 1760 DC1

•

LWT 1760 DC2

•

P4 1760 DC2

•

P1 1760 DC1

•

RDR ALTM PWR

•

P1 1760 DC2

•

RESERVED FOR ROTOR

•

P2 1760 DC1

•

RW PITOT HTR

•

P2 1760 DC2

•

RW ROCKET ARM

•

PNVS STANDBY

•

RWT 1760 DC1

•

RDR WRN RCVR

•

RWT 1760 DC2

•

TADS STANDBY

LBA5168

Figure 2-82A.

[

BLK 2

DC Power Distribution ]

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 2

2-84.3

TM 1-1520-251-10

ÄÄÄBATT BUS NO. 1ÄÄÄÄ

ÄÄÄBATT BUS NO. 2ÄÄÄ

ÄÄÄÄÄÄÄÄÄÄÄ

ÄÄÄÄÄÄÄÄÄÄ

ÄÄÄBATT BUS NO. 3ÄÄÄÄ

ÄÄÄBATT BUS NO. 4ÄÄÄ

• ELC1 BAT CONT

• ELC2 BAT CONT

• ENG1 ATS

• FUEL CRSFD2

• FUEL CRSFD1

• APU FUEL BOOST AND SOV

• ZEROIZE PNL

• FUEL BOOST SOV

• ENG1 OIL PRESS

• ENG2 ATS

• ENG1 BL LOUVER ON/OFF

• XFER AIR VALVE

• APU ECU PWR/START

ÄÄÄBATT BUS NO. 5ÄÄÄÄ

• ENG2 BL LOUVER ON/OFF

Ä•ÄFIRE DET ENG q1ÄÄÄÄÄ

• ENG2 OIL PRESS/TRANS XDCR

• LD MAINT/PRCS SEL

• VIDEO CONT PNL/BPIA

• IFM AMP

ÄÄÄÄÄÄÄÄÄÄÄ

• SQUAT SWITCH

ÄÄÄÄÄÄÄÄÄÄÄ

• FM2 ARC201

•

LDG SRCH LIGHT

• IDM

•

CIU

• TAIL WHL LOCK

•

UHF ARC 164

• RTR/TADS BRAKES

•

XPNDR/IFF

• EMER JTSN

•

UTIL LTS MAINT

• SEC/CPG ADV LTS

•

FIRE EXT PLT

• GEN 2 BACKUP

•

DATA XFR UNIT

• FIRE EXT CPG

•

SYS PROC q2

• FM1 ARC201

•

HYD PRESS XDCRS

• SYS PROC q1

•

INT GRBX TEMP ALM

• UPFRONT DSPL PLT

•

UTIL RLY PNL

• CPG PWR LEVER

•

VHF ARC 186

•

KY58

ÄÄÄÄÄÄÄÄÄÄÄ

•

REFUEL PNL

ÄÄÄBATT HOT BUSÄÄÄÄ

•

EMG S/O VALVE

• COLD RFL

•

PLT ADV LT

• MIK SW

•

UPFRONT DSPL CPG

• HPSM1 CONT

•

FIRE DET ENG q2

• HPSM2 CONT

•

FIRE DET APU

•

STBY INSTR

•

MDR

NOTE: ALL EQUIPMENT LISTED ABOVE ON BATTERY BUSSES 1

THROUGH 6 WILL BE POWERED BY BATTERY POWER IN

•

GEN 1 BACKUP

THE EVENT OF AN ELECTRICAL FAILURE.

•

FLT CNTRL TRIM

LBA2564

Figure 2-83.

[

BLK 1

Battery Power Distribution ]

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-84.4

Change 2

TM 1-1520-251-10

ÄÄÄÄÄÄBATT BUS NO. 1ÄÄÄÄÄ

ÄÄÄÄÄBATT BUS NO. 2ÄÄÄÄÄ

Ä•ÄÄÄÄÄÄÄ•ÄÄÄÄÄÄ

•ÄCIUÄÄÄÄÄÄ•ÄMDRÄÄÄÄ

CPG PWR LEVER

IDM

• EMER JTSN

• IFM AMP

• DATA XFR UNIT

• PLT ADV LTS

• FIRE DET ENG #1

• LD MAINT PRCS SEL

• EMG S/O VALVE

• REFUEL PNL

• FIRE EXT CPG

• RTR/TADS BRAKES

• FIRE DET APU

• STDBY INSTR

•

FM1 ARC201

• SEC/CPG ADV LTS

• FIRE DET ENG #2

• SYS PROC #2

• FM2 ARC201

• SQUAT SWITCH

• FIRE EXT PLT

• UHF ARC164

• GEN 2 BACKUP

• SYS PROC #1

• FLT CNTRL TRIM

• UPFRONT DSPL CPG

• HF ARC220

• TAIL WHL LOCK

• GEN 1 BACKUP

• UTIL LTS MAINT

• HF KY100

• UPFRONT DISPL PLT

• HYD XDCR PRI/EMER

• UTIL RLY PNL

• HYD XDCR UTIL

• INT GRBX TEMP ALM

• VHF ARC186

• KY58

• XPNDR/IFF

• LDG SRCH LIGHT

ÄÄÄÄÄÄÄÄÄÄ

ÄÄÄÄÄÄÄÄÄÄÄ

ÄÄÄBATT BUS NO. 3ÄÄÄ

ÄÄÄBATT BUS NO. 4ÄÄÄÄ

• ELC1 BAT CONT

• APU ECU PWR/START

• ENG1 ATS

• APU FUEL BOOST AND

• ENG1 BL/LOUVER

• ELC2 BAT CONT

• ENG1 OIL PRESS

• ENG2 OIL

• FDMU PROVISION

• ENG2 ATS

• FUEL CRSFD1

• ENG2 BL/LOUVER

• ZEROIZE PNL

• FUEL BOOST SOV

• FUEL CRSFD2

• VIDEO CONT PNL/BPIA

• XFER AIR VALVE

ÄÄÄÄÄÄÄÄ

• COLD RFL PANEL

• HPSM1 CONT

• HPSM2 CONT

• MIK SW & MDR

LBA5169

Figure 2-83A.

[

BLK 2

Battery Power Distribution ]

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 2

2-84.5/(2-84.6 blank)

TM 1-1520-251-10

2.98.1 [

BLK 1

Circuit Breaker Panel (CBP) 1. CBP

switched secondary loads on the No. 1 busses. CBP 1 is

1 receives 115/200 Vac, 28 Vdc, and 24/28 Vdc input from

mounted on the forward face of HPSM 1. Figure 2-84

HPSM 1. It contains the CBs necessary to supply the un-

shows which CBs are located on CBP 1. ]

SEC/CPG

EMERG

RTR/TADS

TAIL WHL

IDM

FM2

SQUAT

IFM

LO MAINT

FIRE DET

ADV LTS

JISN

BRAKES

LOCK

ARC201

SWITCH

AMP

PRCS SEL

ENG q1

CPG PWR

UPFRONT

SYS PROC

FM1

FIRE EXT

GEN 2

LEVER

DSPL PLT

ARC201

CPG

BACKUP

SELECT

P2 ACTR

STAB

KYBD

ENG

P2 PIU

OCU 1

EGI 1A

JTSN

PWR

ACTR LH

PLT

CHOP

PWR

RADAR

GRBX

PRI LT

P1 PIU

P1 ACTR

WPN PROC

PLT CBR

FAN AFT

JAM

XDCR

PLT

CPG

PWR

FLTR BLWR

BAY

TADS/

EGI 2A

PLT SIGHT

CPG

EGI ANT

TESS

PNVS

SIGHT SW SPLITTER

TECU

TLIA

LRFD

ENG1

ANTI COLL

GUN

IR JAM

OVSP

FORM LTS

MOTOR

CONTROL

NOSE GRBX

DISPL PROC

PNVS

LASER

HEATER 1

FWD COND

CPG RH

IHADS

PLT LH

BLWR A

MPD

SEU

MPD

FWD COND

AIR PARTICLE

CPG CKPT

FWD LH/RH

BLWR B

SEP FAN

FAN

FAB FANS

TADS PWR

RH/LH FANS

SUPPLY

AFT BAYS

LBA23981A

CIRCUIT BREAKER PANEL 1

Figure 2-84.

[

BLK 1

Circuit Breaker Panel 1 ]

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 2

2-85

TM 1-1520-251-10

2.98.1A [

BLK 2

Circuit Breaker Panel

(CBP)

ply the unswitched secondary loads on the No. 1 busses.

1. CBP 1 receives 115/200 Vac, 28 Vdc, and 24/28 Vdc

CBP 1 is mounted on the forward face of HPSM 1. Figure

input from HPSM 1. It contains the CBs necessary to sup-

2-84A shows which CBs are located on CBP 1. ]

SEC/CPG

EMERG

RTR/TADS

TAIL WHL

IDM

FM2

SQUAT

IFM

LO MAINT

FIRE DET

ADV LTS

JISN

BRAKES

LOCK

ARC201

SWITCH

AMP

PRCS SEL

ENG q1

CPG PWR

UPFRONT

SYS PROC

FM1

FIRE EXT

GEN 2

HYD XDCR

LEVER

DSPL PLT

ARC201

CPG

BACKUP

UTIL

KY100

SELECT

P2 ACTR

STAB

KYBD

ENG

P2 PIU

OCU 1

EGI 1A

JTSN

PWR

ACTR LH

PLT

CHOP

PWR

ARC220

RADAR

GRBX

PRI LT

P1 PIU

P1 ACTR

WPN PROC

PLT CBR

FAN AFT

JAM

XDCR

PLT

CPG

PWR

FLTR BLWR

BAY

TADS/

EGI 2A

PLT SIGHT

EGI ANT

TESS

PNVS

SPLITTER

TECU

TLIA

LRFD

ENG1

ANTI COLL

GUN

IR JAM

OVSP

FORM LTS

MOTOR

CONTROL

NOSE GRBX

DISPL PROC

PNVS

LASER

HEATER 1

FWD COND

CPG RH

IHADS

PLT LH

BLWR A

MPD

SEU

MPD

FWD COND

AIR PARTICLE

CPG CKPT

FWD LH

BLWR B

SEP FAN

FAN

FAB FAN

TADS PWR

RH/LH FANS

FWD RH

HF ARC220

SUPPLY

AFT BAYS

FAB FAN

FAN

LBA5170

CIRCUIT BREAKER PANEL 1

Figure 2-84A.

[

BLK 2

Circuit Breaker Panel 1 ]

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-86

Change 2

TM 1-1520-251-10

2.98.2 [

BLK 1

Circuit Breaker Panel (CBP) 2. CBP

switched secondary loads on the No. 2 busses. CBP 2 is

2 receives 115/200 Vac, 28 Vdc, and 24/28 Vdc input from

mounted on the forward face of HPSM 2. Figure 2-85

HPSM 2. It contains the CBs necessary to supply the un-

shows which CBs are located on CBP 2.]

INT GRBX HYD PRESS

SYS PROC

DATA XFR

FIRE

UTIL LTS

XPNDR/

VHF

CIU

LDG SRCH

IFF

ARC186

TEMP ALM XDCRS

UNIT

EXT PLT

MAINT

LIGHT

STDBY

FIRE DET

UPFRONT

PLT ADV

EMG S/O

REFUEL

KYSB

UHF

UTIL

INSTR

ARC164

APU

DSPL CPG

LTS

VALVE

PNL

RLY PNL

FLT CNTRL GEN 1

MDR

STAB

AREA WPNS

PNVS

BUCS/FMC

WSHLD

P3 PAC

STAB

TRIM

SYS DC

BACK UP

CNTRL

CMPTR

WPR CPG

PWR

ACTR RH

WPN PROC

KEYBD

EGI ZB

P3 PIU

NAV LTS/

CPG CBR

DCU 2

BATT

P4 PIU

P4 ACTR

CPG

HTR

PWR

ANTICOL

FLTR BLO

WSHLD

ADF

EGI 1B

HADS

SIDE

FUZZ

GND SVC

WPR PLT

LOADER

BURNER

RECPT DC

ENG2

NITROGEN

IHADSS

DISPL PROC

AFT COND

OVSP

INERT

DEU

BLOWER B

VIDEO

CPG LH

ORT

AFT COND

RCDR

MPD

BLOWER A

FANENG

FWD LH/RH

AMMO

GND SVC

BLEED AIR

FAB FANS

MOTOR

RECPT AC

PLT CKPT

BATTERY

PLT RH

NOSE GRBX

FAN

CHARGER

MPD

HEATER 2

CANOPY

HEATER

CIRCUIT BREAKER PANEL 2

LBA23982A

Figure 2-85.

[

BLK 1

Circuit Breaker Panel 2 ]

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 2

2-87

TM 1-1520-251-10

2.98.2A [

BLK 2

Circuit Breaker Panel

(CBP)

ply the unswitched secondary loads on the No. 2 busses.

2. CBP 2 receives 115/200 Vac, 28 Vdc, and 24/28 Vdc

CBP 2 is mounted on the forward face of HPSM 2. Figure

input from HPSM 2. It contains the CBs necessary to sup-

2-85A shows which CBs are located on CBP 2.]

INT GRBX HYD PRESS

SYS PROC

DATA XFR

FIRE

UTIL LTS

XPNDR/

VHF

CIU

LDG SRCH

IFF

ARC186

TEMP ALM PRI/EMER

UNIT

EXT PLT

MAINT

LIGHT

STDBY

FIRE DET

UPFRONT

PLT ADV

EMG S/O

REFUEL

KYSB

UHF

UTIL

INSTR

ARC164

APU

DSPL CPG

LTS

VALVE

PNL

RLY PNL

FLT CNTRL GEN 1

MDR

AREA WPNS

PNVS

BUCS/FMC

WSHLD

P3 PAC

STAB

TRIM

SYS DC

BACK UP

CMPTR

WPR CPG

PWR

ACTR RH

WPN PROC

KEYBD

EGI ZB

P3 PIU

NAV LTS/

CPG CBR

DCU 2

BATT

P4 PIU

P4 ACTR

CPG

HTR

PWR

ANTICOL

FLTR BLO

CPG

STAB

WSHLD

ADF

EGI 1B

HADS

SIDE

FUZZ

VIDEO

GND SVC

SIGHT SW

CNTRL

WPR PLT

LOADER

BURNER

RCDR

RECPT DC

ENG2

NITROGEN

IHADSS

DISPL PROC

AFT COND

OVSP

INERT

DEU

BLOWER B

CPG LH

AFT COND

MPD

BLOWER A

FANENG

AFT RH

AMMO

GND SVC

BLEED AIR

FAB FAN

MOTOR

RECPT AC

PLT CKPT

BATTERY

PLT RH

NOSE GRBX

FAN

CHARGER

MPD

HEATER 2

CANOPY

AFT LH

HEATER

FAB FAN

CIRCUIT BREAKER PANEL 2

LBA5171

Figure 2-85A.

[

BLK 2

Circuit Breaker Panel 2 ]

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-88

Change 2

TM 1-1520-251-10

2.98.3

Electrical Load Center 1. ELC 1 receives

ary loads on the No. 1 busses. It also receives commands

115/200 Vac, 28 Vdc, and 24/28 Vdc input from HPSM 1.

from the system processors for positioning the remote

It contains the Circuit Breakers (CB) and remote con-

controlled switches. ELC 1 is located in the left forward

trolled switches necessary to supply the switched second-

EFAB. Figure 2-86 shows the CBs located on ELC 1.

IAFS

1760

ÔÔ

1760

FUEL

PUMP

TADS

LASER

CHAFF

RDR

SPARE

WP*SP 1&2

AIR DATA

STANDBY

WRN

ARM

WRN

/STAB SYN

SENSOR

RCVR

(REFC)

HEATER

LWT

ELCI

ENG1

ENG 1 BL

ROCKET

1760

BAT

OIL

LOUVER

ARM

DC2

DC1

DC2

DC1

DC2

CONT

PRESS

ON/OFF

PNVS

IAFS

FUEL

CANOPY

ENG1 A1

ELCI

FDMU

SPARE

ENG 1

ZEROIZE

FUEL

STANDBY

VALVES

PITOT

HEAT\

START BL

(PROV)

ATS

PANEL

CRSFD1

CTRL

VLV/NOSE

HTR

DEFOG

CONT

VALVE

GBX HTR1

ELECTRICAL LOAD CENTER NO. 1

LBA2395A

Figure 2-86. ELECTRICAL LOAD CENTER NO 1 Circuit Breakers

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 2

2-89

TM 1-1520-251-10

2.98.4

Electrical Load Center 2. ELC 2 receives

ary loads on the No. 2 busses. It also receives commands

115/200 Vac, 28 Vdc, and 24/28 Vdc input from HPSM 2.

from the system processors for positioning the remote

It contains the Circuit Breakers (CB) and remote con-

controlled switches. ELC 2 is located in the right forward

trolled switches necessary to supply the switched second-

EFAB. Figure 2-87 shows the CBs located on ELC 2.

SPARE

1760

SPARE

IHADSS

RDR

DPLR

GUN

AWS

PLT FLT

ICE

DEU PWR

SEU

ALTM

SENSOR

ARM

CONT

PROBE

PWR

(REF B)

(REF A)

HEATER

RWT

ELC2

ENG2 OIL

ENG2 BL

VIDEO

SPARE

ROCKET

1760

BAT

PRESS

LOUVER

CONT

ARM

DC2

DC1

DC2

DC1

DC2

CONT

ON/OFF

PNL/PB1A

SPARE

ELC2

ICE SIG

ENG2 A1/

APU ECU

XFER

ENG2

FUEL

APU FUEL

FUEL

PROC/

PITOT

START BL

PWR/

AIR

ATS

BOOST

CRSFD2

CONT

ASP SOV

HTR

VLV/NOSE

START

VALVE

AND SOV

GBX HTR2

ELECTRICAL LOAD CENTER NO. 2

LBA2397

Figure 2-87. ELECTRICAL LOAD CENTER NO 2 Circuit Breakers

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-90

TM 1-1520-251-10

2.99 BATTERY

the internal temperature of the battery drops below 10° C

the battery heater is activated. The internal heater is func-

tional between 10° C and - 40° C, therefore, an external

The battery subsystem consists of one 24 Vdc, 15 ampere

heater blanket is not required. The battery heater is pow-

hour, sealed cell, Fibrous Nickel Cadmium (FNC) battery

ered by the aircraft 28 Vdc bus via CBP 2.

and associated battery charger. Both the battery and bat-

tery charger are located in the right center EFAB. The bat-

2.101 DC ELECTRICAL SYSTEM

tery provides power during ground refueling, APU start-

ing, and engine starting. The battery also provides backup

power in the event of a loss of all other DC power (fig

2-83). In that event, the battery will supply the normal in

flight battery loads for a minimum of 12 minutes provided

Two regulated 28 Vdc, 350 amp TRUs are the primary DC

the battery is at least 80% charged. During normal opera-

power source, each capable of supplying managed power

tion the battery is isolated from the battery busses and is

to all DC loads. The TRUs convert AC input to 28 Vdc.

charged by the battery charger.

They are self monitoring for over temperature. TRU fail

status is displayed on the UFD/EUFD.

2.99.1 Battery Power Control. Battery power is con-

2.101.1 Total DC Circuit Load Support. Either TRU,

trolled by the MSTR IGN switch (fig 2-88) on the pilot

in conjunction with either generator, is capable of support-

POWER lever quadrant. The ignition key is not required to

ing the total DC bus load.

actuate the MSTR IGN switch. With the MSTR IGN switch

in the OFF position, power is still applied to the refuel pan-

2.101.2 DC BUS Interface with Battery Electrical Sys-

el. When the refuel switch on the refuel panel is turned

tem. Primary DC bus 1 and DC bus 2 provide 28 Vdc to

ON, the aircraft may be refueled without the need to turn

battery bus 1 and bus 2 through two diodes. These diodes

the MSTR IGN switch to the BATT position. When the

isolate the battery electrical system from DC bus load

MSTR IGN switch is placed in the BATT position, the bat-

when TRU electrical power has failed.

tery is connected to the battery busses unless the busses

are being supplied by the Transformer Rectifier Unit

2.101.3 TRU Failure. In the event of a TRU failure, a

(TRU) power. If the battery busses are being powered by

caution message will be annunciated on the MPD and the

the output of the TRU, the battery is isolated and placed in

UFD/EUFD. The failed DC bus system is then transferred

the charge mode. In the event that TRU power is lost, the

to the remaining operational DC bus system.

battery will automatically be connected to the battery

busses (fig 2-83).

2.102 AC ELECTRICAL SYSTEM

2.99.2

Battery Power Indications. Any time the

The AC generation subsytem consists of two 45 KVA, 3

battery charger fails, an advisory message is displayed on

Phase, 4 wire, 115/200 Vac, 400Hz, air cooled, brushless

the UFD/EUFD, and battery charging is terminated. If the

AC generators and two Generator Control Units (GCU).

charger detects a battery fault an advisory message is dis-

Each generator normally supplies its respective AC bus,

played on the UFD/EUFD, and battery charging is termi-

AC bus 1 or AC bus 2. In the event of a generator failure

nated.

the remaining generator has sufficient power to supply all

AC loads, without the need to load shed. Each generator

2.100 BATTERY CHARGER

control unit provides voltage regulation, control and

protection for its associated generator. Protection is pro-

The battery charger is normally powered by 115 Vac sup-

vided against undervoltage, overvoltage, underfrequency,

plied from CBP 2. The charger automatically maintains

over current/short circuit, and differential/feeder fault op-

the battery in a full state of charge. Additionally the battery

eration.

[

BLK 1

Aircraft

equipped with

charger monitors the battery temperature and state of

7-511B11019-11 and -13 GCUs, underfrequency protec-

health. If it detects a battery fault, it is reported to the sys-

tion is inhibited in flight and undervoltage protection is in-

tem processor for display on the UFD/EUFD. The charger

hibited during underfrequency operation. Aircraft

also monitors itself for a fault and if a fault is detected that

equipped with 7 - 511B11019 - 15 - 17, - 19 and - 21 GCUs,

is also reported to the system processor for display on the

underfrequency protection is functional in flight and

UFD/EUFD. In the event of a battery or battery charger

undervoltage protection is inhibited during underfre-

failure, battery charging is automatically terminated and

quency operation. ] [

BLK 2

Underfrequency protection

can not be restarted until the fault has cleared. Finally the

is functional in flight and undervoltage protection is inhib-

battery charger controls the battery internal heater. When

ited during underfrequency operation.]

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 3

2-91

TM 1-1520-251-10

2.102.1 Generators. Each generator is capable of sup-

• [

BLK 2

Underfrequency (during ground opera-

plying power to all aircraft loads. Each generator is

tion, the AC generator output frequency drops to

equipped with a GCU which provides automatic start up

between 365 to 380 Hz for 1 to 3 seconds; during

and generator protection and fault detection for automatic

flight operation, AC generator frequency drops to

shutdown. If one generator fails, its load is automatically

306 Hz for 1 second).]

transferred to the other generator.

2.102.2 AC Power Indications. Anytime a generator is

a. Generator Drive. The generators are driven si-

failed, a caution message will be displayed on the UFD/

multaneously by the accessory gearbox. The accessory

EUFD and the MPD CAUTION page.

gearbox is driven by the APU or engines/main rotor.

2.102.3 AC Power Controls. Generator output is auto-

b. Generator Operation. During normal operation,

matically controlled by the GCU. In the event of an an-

each generator carries approximately 50% of the electri-

nunciation, the pilot may reset either generator with the

cal load. Each generator powers a three phase bus man-

GEN RST switch on the CHK OVSP TEST/GEN RST

aged by the EPMS. A bus interlock configuration prevents

panel (fig 2-89). A generator may be selected OFF

generators from operating in parallel or operating in paral-

through the SYS page.

lel with an external power source. Either generator is ca-

pable of supporting both AC busses.

c. Generator Electrical Output Manage-

ment. Generator output is managed by the associated

GCU, which regulates output voltage and provides protec-

tion against undervoltage, overvoltage, underfrequency,

over current/short circuit, and differential/feeder fault op-

eration. Underfrequency protection is inhibited in flight

and undervoltage protection is inhibited during underfre-

LBA0020

quency operation. The GCU connects its associated gen-

erator to the appropriate AC primary bus. The GCU dis-

Figure 2-89. Check Overspeed Test/Generator

connects and de-energizes its generator under the

Reset Panel

following adverse operating conditions:

• Overcurrent/short circuit

2.103 EXTERNAL/GROUND POWER

• Overvoltage (the AC generator output voltage is

greater than 125 Vac).

Ground power is applied to the helicopter by connecting

• Undervoltage (the AC generator output voltage

the power cable from the Aviation Ground Power Unit

drops below 100 Vac for longer than 200 ±50

(AGPU) to the external power receptacle located in the

milliseconds).

right hand aft fuselage (refer to Section XV). Ground pow-

er supplies power to all aircraft loads.

• [

BLK 1

Aircraft equipped with 7- 511B11019 - 11

and - 13 GCUs, underfrequency (during ground

2.103.1 External/Ground Power Controls. When the

operation, the AC generator output frequency

MSTR IGN switch is in the EXT power position, the exter-

drops to between 365 to 380 Hz for 1 to 3

nal power monitor automatically controls the application of

seconds).

power to the helicopter and monitors for degraded power

Aircraft equipped with 7 - 511B11019 - 15 - 17, - 19

supply.

and - 21 GCUs underfrequency (during ground

operation, the AC generator output frequency

drops to between 365 to 380 Hz for 1 to 3 seconds;

2.103.2 External/Ground Power Indications. The

during flight operation, AC generator frequency

UFD/EUFD will display an advisory when the external

drops to 306 Hz for 1 second). ]

power access door is open.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 3

2-93

TM 1-1520-251-10

Section XII. AUXILIARY POWER UNIT (APU)

2.104

INTRODUCTION

also provides, directly, pressurized air for the operation of

helicopter systems anytime the APU is ON (whether or not

the main engines are driving the transmission accessory

WARNING

section). The APU provides the means of engine starting

without the need for an AGPU. It is located just inboard of

• To prevent an accidental APU start,

the APU ECU PWR/START circuit

the right engine nacelle in the aft equipment bay. The APU

breaker located on the ELECTRICAL

consists of a gearbox, a compressor, and a turbine sec-

LOAD CENTER NO. 2 in the right hand

tion, together with associated fuel, lubrication, and electri-

EFAB shall be out when battery or ex-

cal systems.

ternal power is connected to the heli-

copter and unqualified personnel are

in or around the pilots crew station.

• The APU ECU testing performed

2.104.1 Fuel System. The aft fuel cell provides all fuel

shows that certain transmitters can

for APU operation and is discussed in more detail in Sec-

preclude the APU from being started

tion IV of this chapter. The APU fuel control automatically

or will cause it to be shut down. Addi-

regulates fuel flow. The APU burns approximately 175 lb

tionally, it is possible that uncomman-

of fuel per hour.

ded overspeeds could occur without

overspeed protection. The aircrew

should exercise caution when starting

the APU in any location where the APU

2.104.2 Lubrication System. The APU has a self con-

has not been previously operated, es-

tained oil system. An oil filler cap is located on the left side

pecially in the vicinity of high powered

of the unit. Oil cooling is provided by airflow over cooling

transmitters.

fins at the compressor inlet. The oil level sight gage is lo-

cated on the right side and is an integral part of the oil

The APU (fig 2-90) indirectly provides hydraulic pressure

and electrical power for the operation of helicopter sys-

sump. The sight gage can be inspected through a door in

tems whether the engines are operating or not. The APU

the bottom of the right engine nacelle.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-94

Change 2

TM 1-1520-251-10

APU ELECTRONIC CONTROL UNIT

APU

APU EXHAUST DUCT

LBA0391

Figure 2-90.

APU Location

2.104.3 Electrical System. The APU electrical system re-

OFF. The ECU also transmits the shutdown signal auto-

quires DC power. This power is normally delivered by the he-

matically for fault detection of any of the monitored func-

licopter battery when the MSTR IGN switch is set to BATT.

tions. An advisory is sent from the low oil pressure switch

when automatic shutdown occurs. The PTO clutch actua-

2.104.4 APU Control Switch. The APU control switch is a

tion is also controlled by the APU ECU. Normally, PTO

momentary button switch located on the pilot POWER lever

clutch engagement occurs at 60% APU speed if the main

quadrant (fig 2-88). This switch applies dc power to energize

transmission temperature is above 0 °F; however, if the

the APU boost pump, open the APU fuel shutoff valve, and

main transmission temperature is below 0 °F, it will en-

sends a signal to the APU ECU to commence the start se-

gage at 95% APU speed.

quence. The APU momentary button switch turns the APU

ON and OFF, and has a hinged cover to prevent inadvertent

2.104.6 Main Transmission Low Oil Temperature Sen-

actuation. When the APU is turned ON and the APU speed

sor. The main transmission low oil temperature sensor

reaches 95%, the ON legend in the button illuminates and an

provides transmission oil temperature information to the

advisory is displayed on the UFD/EUFD.

system processor for cold start of the APU.

2.104.5 APU Electronic Control Unit (ECU). The APU

2.104.7 Utility Hydraulic Accumulator. The accumu-

ECU provides for automatic start and operation of the

lator, located on the deck of the aft equipment bay below

APU and Power Takeoff clutch (PTO) engagement. The

the APU, provides hydraulic pressure to the APU hydrau-

start valve opens when the APU switch is pressed to ON

lic starter through a solenoid operated hydraulic start

and closes automatically at 60% APU speed. The ECU

valve. The utility hydraulic accumulator is discussed in

monitors the APU for loss of thermocouple, overtempera-

more detail in Section VI of this chapter.

ture, overspeed, overcurrent, low oil pressure, percent

rpm, and exhaust gas temperature. The APU ECU trans-

2.104.8 APU Advisories. The APU advisories are dis-

mits a shutdown signal to the APU whenever the APU but-

played on the UFD/EUFD and MPD and are listed in Table

ton switch on the pilot POWER lever quadrant is set to

2-15.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 2

2-95

TM 1-1520-251-10

Section XIII. LIGHTING

2.105 LIGHTING EQUIPMENT

The aircraft lighting system consists of exterior and interi-

or lighting and their associated controls. The main func-

NOTE

tion of the aircraft exterior lighting is to provide visual in-

dications to other aircraft for navigation and safety

The search/landing light, crew station flood

reasons. The aircraft interior lighting provides crewmem-

lights, signal lights, and utility lights are the

bers with Night Vision Goggle (NVG) compatible lighting.

only lights available under battery power.

2.106 EXTERIOR LIGHTING EQUIPMENT

Exterior lighting equipment (fig 2-91) consists of formation

lights, navigation lights, anticollision lights, a search/land-

ing light, and an inspection and maintenance light.

NAVIGATION

LIGHT (WHITE)

FORMATION

LIGHT (GREEN)

FWD INSPECTION AND

MAINTENANCE LIGHT

RECEPTACLE

SEARCH/LANDING

LIGHT (WHITE)

RETRACTED POSITION

FWD INSPECTION AND

MAINTENANCE LIGHT

RECEPTACLE

(INSIDE AVIONICS BAY)

FORMATION LIGHT (GREEN)

ANTICOLLISION LIGHT

(RED OR WHITE)

(TYPICAL BOTH SIDES)

NAVIGATION LIGHT

(LEFT SIDE RED,

RIGHT SIDE GREEN)

FORMATION LIGHT (GREEN)

LBA0545

Figure 2-91. Exterior Lighting Equipment

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-96

TM 1-1520-251-10

2.106.1 Formation Lights. The green formation lights

in a fairing under the forward section of the right Extended

provide visual orientation information regarding position of

Forward Avionics Bay (EFAB), just forward of the landing

the helicopter to adjacent aircraft in formation flight. They

gear attachment. The searchlight control switches (fig

are located on the upper surface of each wing, the upper

2-92) are located on the flight control grip of the collective

centerline of the aft fuselage, and on the upper surface of

control stick. Setting the ON/OFF/STOW switch to ON

the vertical stabilizer. The formation (FORM) lights rotary

supplies 28 vdc to the searchlight. The searchlight posi-

control knob on the pilot Exterior/Interior Lighting (EXT LT/

tion switch is a four position thumbforce control switch

INTR LT) panel varies the intensity of the formation lights

used to manually control searchlight motion.

from OFF to BRT (bright). For description and operation

of pilot and CPG lighting control panels, refer to paragraph

2.107.

2.106.2 Navigation Lights. The navigation lights indi-

SEARCH/LANDING

LIGHT

cate aircraft position and direction to other aircraft during

ON/OFF/STOW

flight. One green light is located on the right engine na-

SWITCH

celle and one red light is located on the left engine nacelle.

One white aft navigation light is located on the top aft side

FLIGHT

CONTROL

of the vertical stabilizer. The navigation (NAV) lights

GRIP

switch on the pilot EXT LT/INT LT panel has three posi-

tions, BRT, OFF, and DIM.

2.106.3 Anti-collision Lights. The anti-collision lights

provide a visual anti-collision warning during all phases of

SEARCH/LANDING

flight. High intensity red and white anti-collision strobe

LIGHT

POSITION

lights are located on each engine nacelle. The anti-colli-

SWITCH

sion (ANTI COL) lights switch on the pilot EXT LT/INT LT

LBA0480

panel has three positions: WHT, OFF and RED.

Figure 2-92. Search/Landing Light Control

CAUTION

Switches

• The searchlight can reach tempera-

2.106.5 Inspection/Maintenance Light. The inspec-

tures capable of igniting combustible/

tion/maintenance light is stored in the left equipment stow-

flammable materials. Do not land in

age compartment. Two plug-in electrical receptacle loca-

areas such as high grassy meadows

tions and the light’s 50 ft long cord facilitates inspection

with the searchlight ON.

and maintenance at all points on the helicopter. One re-

• Burning of the aircraft skin is possible

ceptacle is located adjacent to the CPG station on the un-

if the searchlight is over - extended.

derside of the right EFAB, forward of the searchlight. The

second receptacle is located in the right aft avionics bay.

• R/H EFAB door should not be left open

Operation of the inspection/maintenance light is con-

with the searchlight on.

trolled by an OFF/BRT rheostat switch which is integral

with the light.

NOTE

2.107 LIGHTING CONTROLS

Searchlight motion is inhibited for 60 sec-

onds after the ON/OFF/STOW switch is

Aircraft exterior lighting (excluding the searchlight) is con-

placed in the STOW position.

trolled via the EXT LT section of the pilot EXT LT/INTR LT

panel (fig 2-93). Aircraft interior lighting can be controlled

2.106.4 Search/landing Light. The retractable search-

independently for each cockpit by use of the INTR LT sec-

light provides omnidirectional search and landing light

tion of the pilot EXT LT/INTR LT panel or the CPG INTR

during low visibility conditions. The searchlight is located

LT panel.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-97

TM 1-1520-251-10

NOTE

The Lighting Distribution Unit in the pilot

crewstation contains variable resistors

ÑÑÑ

(available to maintenance personnel) to al-

low balancing of individual panels in the

ÑÑÑ

standby instruments lighting system for uni-

ÑÑÑ ÑÑÑ

form intensity within the crewstation

ÑÑÑ

throughout the STBY INST control dimming

ÑÑÑ

range.

ÑÑÑ

2.108 INTERIOR LIGHTING EQUIPMENT

ÑÑ

ÑÑÑ

ÑÑÑ ÑÑ

The interior lighting system includes four subsystems

ondary/emergency, signal, and standby instrument light-

ing. Each cockpit also has one utility light.

LBA0025

2.108.1 Primary Lighting. Primary lighting consists of

integrally illuminated lightplates, display bezels and key-

pads. The PRIMARY variable resistor rotary control al-

lows for continuous adjustment of the crewstation primary

lighting system brightness from OFF to BRT. Turning the

Figure 2-93.

Pilot EXT LT/INTR LT Panel

PRIMARY variable resistor from OFF towards BRT in-

forms the Lighting System Controller that the cockpit is in

a “night” condition and allows for the SIGNAL RST posi-

tion to function for resetting to a dim mode. The FLOOD

2.107.1 INTR LT Panel. The INTR LT section of the pi-

control must be OFF or less than the midpoint of its range

lot panel controls the lighting in the pilot’s station and con-

to allow SIGNAL RST. The PRIMARY lights control is on

sists of four rotary control knobs and a PRESS-TO-TEST

the pilot EXT LT/INTR LT panel and on the CPG INTR LT

button. The CPG INTR LT panel (fig 2-94) controls the

panel.

lighting in the CPG’s station and consists of three rotary

control knobs and a PRESS-TO-TEST button.

2.108.2 Secondary/Emergency Lighting. Secondary/

emergency lighting consists of floodlights to provide crew

compartment illumination and is powered by the battery

bus as an alternate to primary lighting. The FLOOD vari-

able resistor rotary control allows for continuous adjust-

ment of the floodlighting system brightness from OFF to

ÑÑÑ

ÑÑ

BRT. Turning the FLOOD control clockwise past the mid-

point of the control range informs the Lighting System

ÑÑÑ

Controller that the cockpit is in a “bright” condition and re-

ÑÑÑ

verts the SIGNAL lights to BRT. Turning the FLOOD con-

trol counterclockwise past the midpoint of the control

range allows for the SIGNAL lights to reset to a dim mode.

The FLOOD lights control is on the pilot EXT LT/INTR LT

ÑÑÑ

panel and on the CPG INTR LT panel.

ÑÑÑ

2.108.3

Standby Instrument Lighting. Standby

instrument lighting consists of integral illumination of the

LBA0026

four standby flight instruments in the pilot station. The

STBY INST variable resistor rotary control allows for con-

tinuous adjustment of the standby instruments lighting

system brightness from OFF to BRT. The STBY INST

Figure 2-94. CPG INTR LT Panel

lights control is on the pilot EXT LT/INTR LT panel.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-98

TM 1-1520-251-10

2.108.4 Signal Lighting. Signal lighting includes warn-

manually turn the control to RST for night conditions.

ing, caution, and advisory indicators/switches. The SIG-

Turning the FLOOD control clockwise past the midpoint of

NAL variable resistor rotary control allows for continuous

the control range informs the Lighting System Controller

adjustment from DIM to BRT for the SIGNAL lighting. The

that the cockpit is in a “bright” condition and reverts the

rotary control knob for SIGNAL lighting includes a RST

SIGNAL lights to the bright mode. Turning the FLOOD

position to allow for reset to dim mode for night conditions.

control counterclockwise past the midpoint of the control

The warning lights have a fixed daytime brightness (28

range allows for the SIGNAL lights to reset to a dim mode.

vdc) and a fixed nighttime brightness (14 vdc). The cau-

The PRESS TO TEST button allows all SIGNAL lighting

tion and advisory lights have a fixed daytime brightness

lamps to be tested at once. The SIGNAL lights control and

(28 vdc) and a variable nighttime brightness (14 vdc

PRESS TO TEST button are on the pilot EXT LT/INTR LT

max.). The SIGNAL variable resistor controls the signal

panel and on the CPG INTR LT panel.

lighting brightness for the advisory lights from BRT to

DIM, but does not extinguish the lights. Initial aircraft pow-

2.108.5 Utility Light. The utility light is stowed (via a

er-up sets the SIGNAL lighting to full daytime brightness.

bayonet base) in the left side of each crewstation. The

If the PRIMARY lights control is ON, and the FLOOD light-

hand-held light has a coiled cord and is powered by the

ing control is less than the midpoint position, then turning

battery bus. The utility light has a built-in variable rheostat

the SIGNAL control clockwise to the momentary RST

control for continuous adjustment of the light from OFF to

position results in reset to the dim mode for night condi-

BRT, and a press-to-hold BRT button. The beam can be

tions. In the event of a power interrupt, the signal lighting

adjusted from SPOT to FLOOD. The color of the illumina-

reverts to the daytime brightness; the crewmember must

tion can be adjusted to WHITE or NVG GREEN.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-99

TM 1-1520-251-10

Section XIV. FLIGHT INSTRUMENTS

2.109 INTRODUCTION

• T4

PERF (Performance) page button

• T6

A/C UTIL (Utility) page button

Instruments discussed in this section are, for the most

part, those that directly measure flight performance. Flight

• L5

-W- (Waterline) button

instruments are available on the IHADSS and the MPD

• B6

FLT SET button

FLT page. The CPG can present flight symbology, HOD/

HDD displays, and video displays on the ORT.

2.110.1 MPD FLT Page Buttons. The FLT page (figs

2-95 and 2-96) provides the following button selections:

2.110 MPD FLT PAGE

a. ENG Page Button. Pressing the ENG page but-

The FLT page (fig 2-95) presents flight instrument

ton presents the Engine instrument page.

symbology required for flight from the MPD. The majority

of the buttons and some of the graphics are presented

b. FLT Page Button (On second level only). Press-

only when the FLT SET button is selected,

ing (deselecting) the FLT page button returns the MPD to

the top-level format.

c. FUEL Page Button. Pressing the FUEL page

button presents the Fuel page.

d. PERF Page Button. Pressing the PERF page

button presents the Performance page.

e. A/C UTIL Page Button. Pressing the A/C UTIL

button presents the Aircraft Utility page.

-W- Button. The -W- button allows the crew to

align the attitude indicator and HMD horizon line to the

waterline symbol in the pitch axis. The word BIAS is pres-

ented below the waterline button when a bias has been

applied using the up or down bias buttons. Deselecting

the - W - button clears any manually applied bias from the

attitude indicator and HMD horizon line.

g. FLT SET button. Selection of the FLT SET but-

ton presents the control buttons necessary to set or enter

LBA0580A

data on the FLT page. Deselecting the FLT SET button

blanks these control buttons and additional graphics, and

Figure 2-95. FLT Page

returns the format to the FLT page.

• T1

ENG (Engine) page button

2.110.2 MPD FLT Page graphics. The FLT Page pro-

• T3

FUEL page button

vides the following graphics (fig 2-96):

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

2-100

TM 1-1520-251-10

4. Lubber Line. The lubber line is aligned to the

centerline of the aircraft. It is the reference for both heading

and angle of bank.

5. Command or Bobup Heading. The com-

mand heading chevron symbol along the bottom of the

magnetic heading scale indicates the heading to the next

navigation waypoint selected through the navigation sub-

system. The symbol is presented when there is a valid next

waypoint and the heading scale is presented. Upon selec-

tion of the cyclic symbol select bobup mode, the symbol

represents the heading at the initiation of the bobup sym-

bol. The symbol remains at the bobup heading until the bo-

bup mode is disengaged or the heading becomes invalid.

6. ADF Bearing. The ADF bearing is presented

along the bottom of the magnetic heading scale, represent-

ing the instantaneous bearing to the tuned ADF Non - Di-

rectional Beacon (NDB). It is displayed only when a valid

bearing signal is received by the ADF system and the

heading scale is presented.

LBA5199A

7. Alternate

(Pilot or CPG) Sensor Bear-

ing. The alternate crewmembers sensor bearing is pres-

ented as a triangle along the bottom of the heading scale. It

Figure 2-96. FLT Page Symbols

is only presented when the heading scale is presented and

the opposite crewmembers sight is valid. This allows one

crewmember to see the other crewmember’s selected sen-

Heading Scale. The heading scale is cen-

sor’s line of sight azimuth. The symbol is not presented in a

tered in the top area of the page. The moving scale has a

crew station when the other crewmembers selected sight is

total range of 360°. It displays an instantaneous 180° in 10°

FCR.

increments, and the major cardinal points are labeled N, S,

E, and W. Current magnetic heading is located in the center

8. Barometric or Inertial Altitude. Barometric

of the heading scale as a 3 digit data field. The heading

altitude is presented in feet in the upper right area of the

scale is presented when the EGI has determined its true

page. The total range of barometric altitude is from - 2300

heading, and has a position for determining magnetic vari-

to 20,000 ft, in 10 ft increments. Barometric altitude is not

ance.

presented when the ADS has identified internal continuous

BIT faults, when there is no last stored altimeter settings, or

2. FCR

Centerline Bearing. The FCR cent-

when barometric pressure data is out of valid range. The

erline bearing is displayed along the bottom of the magnet-

inertial altitude (calculated MSL altitude) is WHITE in color

ic heading scale. It represents the azimuth of the current

and presented in feet with the WHITE text “INRTL” beneath

FCR scan centerline

(or line of bearing) when the

it when barometric altitude is not available, or when the EGI

centerline is within the displayed portion of the heading

has identified internal continuous BIT faults, and when in-

scale. The FCR centerline bearing is presented when the

heading scale is presented and the requirements for pre-

ertial altitude data is valid. Inertial altitude presented is the

calculated MSL altitude using inertial altitude information.

sentation of the FCR footprint on the TSD Page are met.

The total range of inertial altitude is from - 2300 to 20,000 ft,

3. Bank Angle. The bank angle indicator pro-

in 10 ft increments. The System Processor computes the

vides an analog indication of the aircraft’s current bank

displayed INRTL MSL altitude, using the Inertial altitude

angle in degrees. The bank angle indicator consists of a

and present position information from the primary INU.

fixed curved tape presenting a minimum of 30° to the left

Maximum allowable error is 70 ft.

and 30° to the right of the 0° bank in 5° increments. The

bank angle triangle is aligned and fixed to the roll move-

9. Radar Altitude HI and LO. The radar altitude

ment of the attitude indicator. As the bank angle increases,

HI and LO indicators are displayed just above and below

the bank angle tick marks expand to show a larger portion

the radar altitude digital readout. The YELLOW HI and

of the compass rose. When the angle of bank exceeds

RED LO indications can be set to present at 1 to 1428 ft. If

”20°, the entire symbol is shown WHITE in color. The bank

the HI or LO indications are set to 0, they are not presented

angle is presented if the attitude indicator is presented.

to the crew.

Use or disclosure of this information is subject to the restriction(s) on the title page of this document.

Change 4

2-101

HELICOPTER, ATTACK, AH-64D LONGBOW APACHE. TECHNICAL MANUAL (2002) - page 2