F-14D. FLIGHT MANUAL (2004)

NAVAIR 01-F14AAD-1

2.24.6.1

Current Faults

message for 3 seconds followed by a flashing “CLR” mes-

sage for 4 seconds. Once all FLT fault codes are cleared, all

On the ground, currently existing faults are indicated

center segments “----” will be illuminated and the INC and

by a combination of caution/advisory lights and acronyms.

DEC pushbuttons can then be released. IBIT faults can only

After momentary depression of the INC or DEC pushbutton,

be cleared by the completion of a successful IBIT “PASS,”

the current fault codes will be listed following the “FAIL”

a power on reset, or loss of system power. Current FAIL faults

header. The current faults are logged in volatile memory and

can be cleared from the display by depressing MASTER

will be lost following removal of system power. If faults exist

RESET once the fault no longer exists or loss of system

when the aircraft goes weight off wheels, they are added to

power.

the in-flight fault listing.

2.24.6.2

In-Flight Detected Faults

In-flight faults will be logged and stored in chronologi-

cal order. During the normal startup or shutdown sequence,

Following an IBIT, a MASTER RESET will

depressing the INC or DEC pushbuttons will indicate if any

clearthe IBIT caution/advisorylight failureindi-

in-flight faults were logged. In-flight faults will be listed

cations, but will not clear the FAULT DISPLAY

following the “FLT” header. Each particular fault code will

IBIT codes. This does not indicate that the fail-

only be displayed once regardless of the number of failures

ures detected during IBIT are resolved. The

recorded during the flight, unless a POR is recorded in which

DFCS should not be considered fully opera-

case the sequence is allowed to repeat previously listed

tional. Onlythe successfulcompletionofanother

codes. In-flight faults remain in memory until manually

IBIT can verify proper system operation.

cleared by the pilot or ground crew. To avoid confusion,

in-flight faultsshouldbeclearedjustpriortoeachflight.Loss

2.25 LANDING GEAR SYSTEMS

of system power does not remove FLT faults from memory

to enable reference by maintenance personnel during post-

The aircraft has fully retractable, tricycle landing gear

flight troubleshooting.

operated by combined hydraulic pressure in the normal mode

of operation and a stored source of pressurized nitrogen for

2.24.6.3

IBIT Detected Faults

emergency extension. The landing gear retract forward so

that airloads and gravity assist on emergency extension. Air-

Successful completion of IBIT is indicated by the

oil shock struts with oil metering pins reduce landing loads

absence of caution/advisory lights and acronyms. A “PASS”

transmitted to the airframe, and the struts are fully extended

indication will also be displayed on the DCP. Depressing

with the gear in the wells. All landing gear doors remain open

MASTER RESET will then blank the display (or will return

with the gear extended. Design limit landing sink speed for

to alternating “IBIT” and “ARM” if all IBIT interlocks are

the aircraft is 1,520 feet per minute (nominal landing sink

still valid). IBIT detected faults are indicated by the appropri-

speed is about 650 feet per minute).

ate caution/advisory lights and acronyms, and a “NOGO”

indication on the DCP. After momentary depression of the

2.25.1

Landing Gear Handle

INC or DEC pushbutton, fault codes from the most recent

IBIT run will be listed following the “IBIT” header. The IBIT

The landing gear handle mechanically positions the

faults are logged in volatile memory and will be lost follow-

landing gear valve for normal operation. Pulling the handle

ing removal of system power.

mechanically selects emergency extension of the gear using

the pneumatic backup source. Both modes of gear operation

2.24.6.4

Clearing Fault Indications

can be accomplished without electrical power except for the

gear position indication, which requires dc essential No. 2

Simultaneous and continuous depression of the INC

bus power. Gear downlock actuators incorporate internal

and DEC pushbuttons for 7 sec will clear any logged FLT

mechanical finger locks that maintain the downlock inserted

fault codes. This will be indicated by a steady

“CLR”

CHANGE 1

2-132

NAVAIR 01-F14AAD-1

position in the absence of hydraulic pressure. The landing

gear handle contains other interlocks that are discussedunder

their respective systems such as weapons firing, jettison

systems, APC, maneuvering flaps, and ground power system

test panel.

z Illumination of indexer lights does not indi-

cate that the main landing gear are clear of the

Normal and emergency controls and displays associated

runway. Raising the gear before a positive rate

with operation of the landing gear are shown in Figure 2-72.

of climb is established will result in blown

main tires.

2.25.2

Main Landing Gear

z Illumination of indexer and approach lights is

Each main landing gear shock strut consists of anupper

not an indication of gear down and locked.

outer cylinder and a lower internal piston, which has a maxi-

mum stroke of 25 inches. A hard step (31,000 pounds

2.25.3

Nose Landing Gear

required for further compression) in the strut air curve pro-

vides a consistent 4-inch stroke remaining in the ground

The dual-wheel nose landing gear has a shock strut

static condition. A side-brace link is mechanically extended

consisting ofan outer cylinder anda lowerinternal pistonthat

from the inboard side of the strut outer cylinder to engage in

has a maximum stroke of 18 inches. During normal ground

a nacelle fitting and thus provides additional side load

operations, the strut is fully extended. Pilot control is pro-

support for ground operations.

vided to kneel the strut (4 inches stroke remaining) for cata-

pult operations (see Figure 2-77). During retraction, the fully

The path of the wheel assembly is controlled by the

extended nose strut is rotated forward by the retract actuator

drag brace as it folds (jackknifes upwards)during gear retrac-

into the well and enclosed by two forward and two aft doors.

tion and unfolds during extension. The fully extended shock

The forward doors are operated by a separate actuator that

strut and jackknifed drag brace retracts forward and rotates

also engages the gear uplock, whereas the two aft doors are

the wheel assembly 90

to lie flat in the wheelwell. Inboard,

mechanically linked to the shock strut. An uplock hook

outboard and aft main gear doors are individually actuated

actuator engages a roller on the lower piston to hold the gear

closed in sequence to provide fairing for the retracted gear.

and doors in the retracted position. During extension, the

An uplock hook on the shock strut engages a roller in the

telescoping drag brace compresses so that a downlock actua-

wheelwell to hold the gear in the retracted position. The main

tor mechanically locks the inner and outer barrel to form a

landing gear actuator on the inboard side of the shock strut

rigid member for transmission of loads to the airframe.

retracts and extends the gear assembly.

Note

The gear downlock actuator, mounted at the drag brace

knee pin, extends to prevent unlocking (jackknifing) of the

z There is no foolproof visual check of the nose

drag brace. Hydraulic pressure must be supplied to the

landing gear locked-down status. Neither the

downlock actuator in order to retract it against the spring

downlock mechanism, which is concealed in

action of the integral locking mechanism. A paint stripe

the fuselage nor insertion of the ground lock

across the drag brace knee pin provides an external visual

pin will provide a positive indication of

indication of the drag brace locked condition. A ground lock

gear-locked status. In flight, the pilot must

device clamps onto the downlock actuator rod for safetying

normally rely on his indicator. Visual deter-

the main gear.

mination of nose landing gear unlocked status

is assisted by a red band painted on the nose

Maximum strut extension and wheel alignment are

landing gear drag brace. If red is visible, the

controlled by torque arms that incorporate cam-operated

nosegear is not locked.

microswitches to detect a weight-on-wheels condition

(greater than

5 inches of strut compression). The single

z An additional sequencing switch in series with

split-type wheel assembly incorporates thermal fuse blow

the existing down-and-locked switch provides

plugs and a pressure relief device to prevent overinflation of

the pilot with a positive indication of nosegear

the tire.

2-133

ORIGINAL

NAVAIR 01-F14AAD-1

NOMENCLATURE

FUNCTION

HYD ISOL switch

FLT —

Combined system hydraulic pressure is shut off to the

landing gear, nosewheel steering and wheel brakes.

T.O./LDG — Switch is automatically placed in this position with gear

handle down. Combined hydraulic pressure is available to

all components.

DOWN LOCK ORIDE

Down —

Weight-on-wheels indication, prevents gear handle being

lever

retracted without pilot override (raising lever).

Up —

Weight-off-wheels indication, does not inhibit pilot raising

gear handle. Automatic operation by electrical solenoid.

LDG GEAR handle

Normal — Up and down overcenter action provides normal retraction

and extension by the combined hydraulic system.

Emergency — Down-push-turn-clockwise pull action provides emergency

extension of all gear by a compressed nitrogen charge.

Figure 2-72. Landing Gear Controls and Indicators (Sheet 2 of 2)

position. If the nose landing gear is unsafe in

locked in the down position with weight on wheels to

the down position because of premature

prevent inadvertent gear retraction. Pilot override of the

deployment of the nose landing gear locking

solenoid-operated handle lock can be effected by lifting the

pin, the nosegear indicator will indicate

downlock lever next to the gear handle. Vertical movement

unsafe and the transition light will illuminate.

of the gear handle causes a corresponding up and downselec-

tion of the landing gear with the combined hydraulic system

Maximum strut extension and wheel steering angle are

pressurized. Three flip-flop indicators provide a position

controlled by torque arms interconnecting the steering collar

display for each of the landing gear, and a gear transition light

and the lower piston (see Figure 2-77). The split-type wheel

on the control panel illuminates anytime the gear position

assembly incorporates a tire pressure relief device to prevent

and handle do not correspond. In addition, a WHEELS

overinflation of the tire. Additional hardware on the nose

warning light alerts the pilot if the landing gear is not down

landing gear include the launch bar, holdback fitting,

with flaps deflected greater than 10

and either or both

approach lights, nosewheel steering actuator, and taxi light.

throttles set for less than approximately 85-percent rpm.

The wheel axles incorporate recessed holes for attachment of

a universal tow bar with maximum steering angle of ± 120 .

z Unless attempting fast-cycle troubleshooting

for gear that indicates unsafe nosegear down,

transition light illuminated, wait for gear to

Restrict nosewheel deflection to ±90

to prevent

completely transition

(15

seconds with

structural damage to the nosewheel steering unit.

normal hydraulic pressure) before recycling

the landing gear handle. When fast cycling the

2.25.4

Landing Gear Normal Operation

gear handle, the pilot must immediately

return the gear handle to the down position to

The landing gear handle is mechanically connected to

avoid damaging the main landing gear doors

the landing gear valve that directs combined hydraulic fluid

and inducing a possible combined hydraulic

into the gear-up and gear-down lines and provides a path for

or brake system failure.

return flow. In the down position, the handle mechanically

sets the hydraulic isolation switch to provide hydraulic pres-

z Maximum landing gear tire speed is 190 knots.

sure for gear operation. The handle is electromechanically

2-135

ORIGINAL

NAVAIR 01-F14AAD-1

2.25.4.1

Landing Gear Handle Up

The emergency landing gear nitrogen bottle is located

in the nose wheelwell. Normal preflight bottle pressure is

Placement of the landing gear handle to UP actuates the

3,000 psi at 70_ F. Minimum bottle pressure for accomplish-

landing gear valve that ports hydraulic pressure to the down-

ing emergency extension to the down-and-locked position is

lock actuators, gear retract actuators, and, in sequence, to the

1,800 psi.

door and uplock actuators. The gear shock strut and door

uplocks are hydraulically operated into a mechanical over-

Pneumatic pressure is directed by separate lines to

center position. An UP indication is displayed on the gear

power open the gear door actuators in sequence, release the

position indicators when the gear are in the uplock and all

gear uplock actuators, pressurize the nosegear actuator to

doors closed.

extend the gear (main gear free fall), and pressurize the

downlock actuators. A normal gear-down indication is

2.25.4.2

Landing Gear Handle Down

achieved upon emergency gear extension. Following emer-

gency gear extension, nosewheel steering is disabled. Once

Placement of the LDG GEAR handle to DN actuates

the landing gear is extended by emergency means, it cannot

the gear control module to port hydraulic pressure to the door

be retracted while airborne and must be reset by maintenance

uplocks, door actuators, and the strut uplocks. The landing

personnel.

gear are hydraulically extended and assisted by gravity and

airloads. A gear-down symbol (wheel) is displayed on the

gear position indicators when the gear downlocks are in the

locked position. The gear transition light will go out when the

main gear side-brace links are engaged.

D Emergencyextensionofthelandinggearshall

Note

be logged in the Maintenance Action Form

(OPNAV Form 3760-2).

With the main gear downlock inserted but the

side-brace link not engaged, landing sink speed

D To facilitate in-flight refueling probe exten-

is restricted to 480 feet per minute. Minimize

sion when the landing gear has been blown

yaw and sideslips on touchdown and rollout.

down, raise the landing gear handle to give

priority to the refueling probe system.

2.25.5

Emergency Gear Extension

2.26 WHEELBRAKE SYSTEM

Although emergency gear extension can be initiated

with the landing gear control handle in any position, it is

The wheelbrake system provides power boost hydrau-

preferable that the LDG GEAR handle be placed in DN

lic control of the multiple disk-type main wheelbrakes using

before actuating the emergency extension system.

pressurized fluid in the landing gear down line from the

combined hydraulic system. Individual or collective wheel-

brake control can be modulated by depression of the rudder

toe pedals, or collective, unmodulated brake control is avail-

able with the parking brake. An antiskid system is provided

to operate electrohydraulically in conjunction with the

The landing gear handle must be held in the fully

normal wheelbraking mode. Wheelbrake controls are shown

extended emergency position for a minimum of

in Figure 2-73.

1 second to ensure complete actuation of the air

release valve. Approximately 55 pounds pull

Brake pedal and parking brake control motions are

force is required to fully actuate the emergency

mechanically transmitted to the power brake module

nitrogen bottle. The pulling motion should be

together with the antiskid valve. Separate hydraulic lines

rapid and continuous to ensure the air release

transmit normal and emergency fluid pressure from the

power brake module to the left and right wheelbrake

valve goes completely overcenter to the locked

position. The landing gear handle will be loose

assemblies. At each brake assembly, the normal and

emergency lines input fluid to the brake shuttle valve, which

(fore and aft) in its housing as an indication of

complete extension of the handle. An incomplete

applies brakes as a function of normal or emergency line fluid

pressure. Two wear-indicator pins on the brake piston

handle motion could cause partial porting of

gaseous fluid, initiating the emergency dump

housing measure lining wear for preflight inspection. For

new brakes, these pins extend approximately one-half inch

sequence. Interruption of handle motion without

completing the overcentering action of the valve

above the piston housing. When the pin is flush with the

could cause the extending gears to contact and

piston housing with the parking brake applied, the brake

damage the strut doors.

assembly is worn to the point of replacement.

CHANGE 1

2-136

NAVAIR 01-F14AAD-1

NOMENCLATURE

FUNCTION

Parking brake handle

Forward —

Parking brake released. Modulated braking action avail-

able with brake pedal depression.

Aft —

Parking brake set. No modulation of control, locks both

main wheel brakes.

BRAKES warning light

Indicates parking brake handle is pulled, antiskid has failed, or operation is

in auxiliary brake mode when brake pedals are depressed.

Brake Pedals

Press top of rudder pedals to command normal or auxiliary braking.

Hand pump

Recharges auxiliary and parking brake accumulators with gear handle

down. With REFUEL PROBE switch in FUS or ALL EXTD, provides

emergency extension or retraction of refueling probe regardless of gear

handle position.

BRAKE PRESSURE

Provides pilot indication of brake accumulator pressure remaining which is

gage

indicative of auxiliary and emergency brake cycles remaining.

ANTI SKID SPOILER BK

BOTH —

Antiskid activated. Spoiler brakes operative with weight

switch

on wheels and both throttles in IDLE.

OFF —

Antiskid deactivated, spoiler brakes inoperative.

SPOILER BK — Spoiler brakes operate with weight on wheels and both

throttles IDLE. Antiskid is deactivated.

Figure 2-73. Wheelbrake Controls and Indicators

2-137

ORIGINAL

NAVAIR 01-F14AAD-1

Four thermal relief plugs are mounted in each main

2.26.2

Normal Braking

wheel assembly to relieve tire pressure and thus avert a blow-

In the normal mode of operation, wheelbrake applica-

out because of hot brakes if the local wheel temperature

tion is modulated by brake pedal depression using pressur-

exceeds 428

ized fluid from the combined hydraulic system through the

brake module and through the normal brake line to the brake

The capacities of the wheelbrake assemblies are suffi-

assembly. In the normal mode of operation, the brake

cient to restrain the aircraft in a static condition on a dry

pressure gauge indication should continue to indicate a full

surface with MIL power set on both engines. The minimum

charge on the brake accumulators since this fluid energy is

hydroplaning speed for the main tires on a wet runway is

maintained by the combined hydraulic system. Normal

approximately 90 knots.

combined-system operations can result in pressure excur-

sions that will be trapped in the brake system. This can cause

2.26.1

Brake Characteristics

the brake pressure indicators to read beyond the full range of

the gauges. This will not affect system performance.

Because carbon brakes contain solid disk-shaped

carbon rotors and stators, they cannot shingle. The thermal

characteristics prevent them from fusing together during or

following heavy braking.

z After heavy or repeated braking or if hot

Carbon brakes may produce a sudden increase in brake

brakes are suspected, allow a 5 to 10-minute

torque as brake pedal force is smoothly increased. This can

cooling period with the gear extended before

produce grabbing at low brake pedal force inputs. This grab-

retracting the gear.

bing is caused by excessive air in the combined hydraulic

z If heavy braking is used during landing or

system. Open-loop bleeding of the combined hydraulic sys-

taxiing followed by application of the parking

tem by maintenance personnel will reduce the amount of air

brake, normal brake operation may not be

in the system and should eliminate any associated grabbing.

available following release of the parking

If grabby brakes are experienced, smooth modulation to

brake if the brakes are still hot. Check for nor-

higher braking forces is easily accomplished after the initial

mal brake operation after releasing the park-

grabbing. The sudden increase in torque is most noticeable

ing brake and prior to commencing taxiing.

at moderate to slow taxi speeds. As groundspeed increases,

the kinetic energy of the aircraft increases and the effect of

2.26.3

Antiskid

the sudden torque increase is significantly reduced. Normal

braking technique should be used during normal rollout.

The antiskid system operates electrohydraulically in

conjunction with the normal mode of wheelbrake operation

The pilot must apply maximum pressure on the brake

to deliver maximum wheelbraking upon pilot command

pedals to hold the aircraft static at MIL. If carbon brakeshave

without causing a skid. Essential No. 2 bus dc power for

been heated up by a full-stop landing, and for about 45 min-

antiskid operation is supplied through the ANTI SKID/R

utes thereafter, they will probably not hold the aircraft static

AICS LKUP PWR circuit breaker (8E1) and controlled by

with military power set on both engines even with the parking

the ANTI SKID SPOILER BK switch (Figure 2-73). When

brake set. In this case, 75 to 100 pounds of pedal force will

energized, approximately 200 milliseconds are required for

hold the aircraft static with afterburner set on one engine and

antiskid system warmup. Individual wheel rotational veloc-

idle power set on the other. In all cases, holding the aircraft

ity is sensed by skid detectors mounted in the wheel hubs and

static at high power settings depends on adequate runway and

transmitted to the skid control box. The control box detects

tire conditions. Degraded conditions such as wet runways or

changes in wheel deceleration and reduces fluid pressure in

worn tires may result in tire skid at high power settings.

the normal brake lines to both wheels, simultaneously, to

prevent a skid.

With the antiskid system armed in flight, the touch-

down circuit in the control box prevents braking until weight

is on both main gear and the wheels have spun up, regardless

of brakepedal application. The antiskid system isinoperative

When the antiskid system becomes inoperative

at groundspeeds of less than 15 knots. During maximum-

knots during a maximum-effort stop,

effort antiskid braking, expect a rough, surging deceleration.

carbon brakes can lock the wheels and pedal

When the ANTI SKID SPOILER BK switch is in BOTH

pressure should be relaxed as the aircraft deceler-

during low-speed taxi (less than 10 knots for more than a few

ates through 15 knots during a maximum effort

seconds), subsequent acceleration of the aircraft through

antiskid stop.

approximately 15 knots will cause a temporary loss of brakes

ORIGINAL

2-138

NAVAIR 01-F14AAD-1

lasting from 2 to 10 seconds. Should this happen, use of the

electrical failure occurs in the antiskid system or if hydraulic

brakes can be regained instantly by turning antiskid OFF. To

pressure is withheld from either brake for greater than 1.2

preclude this possibility, antiskid must be OFF during taxi.

secondsbythecontrolbox,thesystemautomatically becomes

inoperative and illuminates the BRAKES warning light with

the ANTI SKID SPOILER BK switch in BOTH.

2.26.3.1

Antiskid Ground Test

z Failure of the weight-on-wheels switch

During ground operation, a self-test of the antiskid

results in continuous release signal with anti-

system can be initiated on the face of the control box with the

skid selected. Normal braking is available

system energized, parking brake handle released, and the

with antiskid off.

aircraft in a ground static condition. Before taxiing (chocks

z If the antiskid system fails, allowing antiskid

in place), but after releasing the parking brake and while the

to operate below 15 knots, place the ANTI

pilot presses the toe pedal brakes, the plane captain should

SKID SPOILER BK switch in OFF; otherwise

press the antiskid test pushbutton on the control box in the

the aircraft cannot be stopped using normal

nose wheelwell. Approximately 10 seconds is required for

braking.

self-test, which checks the operational status of the control

box, brake valve, and wheel sensors. Any discrepancies

detected will be displayed by the BIT flags on the face of the

control box (Figure 2-74).

Failure to release brakes prior to deselecting

A valid BIT test requires that three criteria be met: the

ANTI SKID may result in blown tires.

BIT flags on the face of the control box must check good,

the pilot must feel both brakes release during BIT test, and

The antiskid system is inoperative when the wheel-

the BRAKES warning light must not remain illuminated. A

brakes are in the auxiliary or parking modes of operation

flash of the BRAKES light coinciding with brake pedal

since the emergency brake lines bypass the brake valve. If an

thumps during the antiskid BIT check is acceptable.

Figure 2-74. Antiskid BIT Box

2-139

ORIGINAL

NAVAIR 01-F14AAD-1

Before initiating antiskid self-test by pressing the

D Even though braking action is available at

antiskid pushbutton on the control box, ensure

accumulator pressures less than 3,000 psi,

that the aircraft chocks are in place. Initiation of

braking force is proportional to pressure

antiskid self-test will release aircraft brakes.

remaining. Red band pressure (1,900 psi) is

sufficient to hold the brakes locked with the

2.26.4

Auxiliary Brake

aircraft stationary in all deck conditions;

Two different auxiliary brake systems are presently

however, rolling motion greatly increases

incorporated in the aircraft. Entry into the auxiliary brake

pressure requirements. Accumulator pressure

mode is the same for both systems. Transfer of normal brake

of up to 2,100 psi may be required to stop a

operation to the auxiliary mode is automatic without the

moving aircraft in a 4_ deck roll. In deck rolls

requirement for pilot action upon the loss of combined

greater than 6_, 3,000 psi may not be suffi-

hydraulic system pressure. Both auxiliary braking systems

cient to stop a moving aircraft.

have two brake accumulators that provide pressure for

D Complete loss of hydraulic fluid through the

auxiliary and parking brake modes of operation when

wheelbrake hydraulic lines will render

combined hydraulic system is not available. Accumulators

parking brake ineffective.

deliver

3,000 psi when fully charged by the combined

hydraulic system or hydraulic handpump (with the gear

2.26.5

BRAKES Warning Light

handle down only). When the combined hydraulic system

pressure decreases below 1,425 psi, the shuttle valve in the

The BRAKES warning light will illuminate whenever

power brake module shifts the brake system to the auxiliary

auxiliary brake pressure is applied to the brakes via the brake

brake mode.

pedals, indicating the combined hydraulic system pressure is

Approximately 13 to 14 full dual-brake applications

not available to the brakes and cautioning the pilot to monitor

are available in the auxiliary mode. Dual pneumatic BRAKE

brake application with the auxiliary brake pressure indicator.

PRESSURE gauges on the front cockpit center pedestal show

A postlight is installed above the BRAKE PRESSURE gauge

auxiliary and parking brake accumulator pressures. Full

to illuminate the dial.

capability operations of the brake accumulators in the

auxiliary modes of operation is predicated on the system

Note

serviced with a nitrogen precharge of 1,900

50 psi. The

green band of the dial indicates pneumatic pressure between

The postlight requires electrical power. Brake-

3,000 psi at the top of the band to 2,150 psi; the red band

riders on carrier night respot must use a flashlight

indicates pneumatic pressures between 2,150 and 1,900 psi

to check the cockpit brake pressure gauge.

at the bottomoftheband. Approximatelyfive auxiliarybrake

applications are available in the red band. Once the auxiliary

2.26.6

Parking Brake

braking system is depleted, braking must be accomplished by

the emergency/parking brake. Three applications of the

The parking brake mode provides a means for collec-

parking brake are available.

tive locking of the wheelbrakes to maintain a ground static

position during normal operations or during emergency con-

With either auxiliary brake system, additional braking

ditions. Aft movement of the parking brake handle provides

can be achieved only by pulling the parkingbrake handle aft.

for unmodulated porting of accumulator fluid pressure

If the shuttle valve in the power brake modules does not

through emergency lines to the shuttle valve at the wheel-

return to the normal position with combined hydraulic

brake assembly. In the parking brake mode, the brake pedals

pressure greater than 2,000 psi, the BRAKES warning light

have no effect on wheelbrake operation. Pushing the parking

will illuminate when a brake pedal is depressed. In this

brake handle forward releases wheelbrake pressure and the

instance the wheelbrake accumulators can be recharged only

power brake module reverts to the normal and auxiliary

by the hydraulic handpump with the landing gear handle

braking mode. When auxiliary mode braking action is no

down. Pilot manual isolation, or system automatic isolation

longer available by depression of the brake pedals, sufficient

of the combined hydraulic system, cuts off the supply of

accumulator fluid pressure remains for a minimum of three

combined hydraulic pressure to the power brake module so

parking brake applications.

that depression of the brake pedals will cause depletion of the

brakes’ accumulator charge.

CHANGE 1

2-140

NAVAIR 01−F14AAD−1

damping, and nosewheel centering. The power unit is located

on the lower portion of the nose landing gear strut outer

cylinder, which, through a ring gear, controls the directional

alignment and damping of the lower piston assembly.

For shipboard operations, before breaking down

and moving an aircraft without combined

Combined hydraulic system pressure is the motive power

hydraulic pressure, the parking brake handle

used for steering and centering. Electrical power is supplied

should be cycled and the toe brakes should be

from the essential dc bus with circuit protection by the NOSE

tested for effectiveness.

(The AUX and PARK

WHEEL STEER/AFCS circuit breaker (RC2) on the pilot right

brake needles should drop slightly after brake

knee panel. Hydraulic pressure is derived from the gear−down

application.) The brake pressure indicator should

line such that steering control is disabled subsequent to

then be pumped back up to the top of the green

emergency extension of the landing gear (Figure 2-75).

band with the cockpit handpump. The indicator

Note

should be maintained in the green band until the

If nosewheel steering is inoperative, the emer-

aircraft is secured. Full

3000 psi pressure is

gency gear extension air release valve may be

required if conditions are severe (greater than

tripped, which will prevent gear retraction.

4 degree roll, wet brakes, etc.).

2.27.1

Nosewheel Steering Control

Nosewheel steering control during ground operations

is energized by momentarily pressing the autopilot reference

and nosewheel steering pushbutton on the lower forward side

Normal brakes are not available with parking

of the pilot stick grip (see Figure 2-75). The system cannot

brake handle pulled. If parking brake accumula-

be engaged without weight on wheels. The system will

tor pressure is depleted, aircraft brakes are iso-

remain engaged until weight is off wheels, electrical power

lated from brake pedal master cylinders. Parking

is interrupted, or the pushbutton switch is pressed again.

Brake handle shall be pushed in to restore normal

Engagement of nosewheel steering is indicated by illumina-

brake operation.

tion of the NWS ENGA caution light. An automatic

In the absence of a pressurized combined hydraulic

nosewheel steering system disengage feature is provided. If

system, the wheelbrake accumulators can only be recharged

this feature has been activated by cycling the hook on deck

by the pilot hydraulic handpump with the landing gear handle

with the throttles at idle, then the nosewheel steering will be

in the down position.

disengaged and the NWS ENGA light extinguished when the

launch bar is lowered. The nosewheel steering automatic

disengage feature is deactivated if the nosewheel steering

button is depressed.

With the system engaged, nosewheel steering is

Complete loss of hydraulic fluid through the

controlled by rudder pedal position. Centering is unaffected

wheelbrake hydraulic lines will render parking

by directional trim displacement. Maximum steering author-

brake ineffective.

ity is 70_ either side of neutral, and the nosewheel can swivel

2.26.7

Wheel Antirotation

a maximum of 120_ about the centered position. With greater

weight on the nosewheel (wings forward, high gross weight,

During the initial phase of the landing gear retraction

cycle, pressurized fluid from the gear−up lines is directed to

etc.) the steering torque can only turn the nosewheel ± 5_

the power brake module to displace the normal metering

with the aircraft static. However, only a slight forward

valves to stop main wheel rotation before the wheels enter the

movement will provide the pilot with full−power steering

wells. This feature is not provided for the nosewheels.

authority. In a full pedal−deflection turn using nosewheel

steering, the aircraft pivots about a point between the main

gear such that the inboard main wheel rolls backward. Under

this condition, application of either main wheelbrake will

only serve to increase the radius of turn. Because of the

Illumination of indexer lights is not a positive indi-

outboard location of the engines, the application of thrust in

cation that the main landing gear is clear of the run-

tight turns should be made on the outboard engine to

way. Raising the gear before a positive rate of

efficiently complement the turning movement of the nose-

climb is established will result in blown main tires.

gear. Nosewheel centering is enabled by the same latching

relay that enables nosewheel steering automatic disengage-

2.27 NOSEWHEEL STEERING SYSTEM

ment with launch bar lowering. Therefore, if the nosewheel

The electrohydraulic nosewheel steering system provides

steering is automatically disengaged when the launch bar is

for on−deck aircraft directional control, nosewheel shimmy

lowered, the nosewheels will be hydraulically centered.

2−141

CHANGE 2

NAVAIR 01-F14AAD-1

2.27.2

Nosewheel Centering

2.28.1

Nose Strut Kneel

The nosewheel is automatically centered during gear

Prior to catapult hookup, the nose strut is compressed

retraction before the nosewheel enters the wheelwell. During

14 inches. Control of the nose strut kneel function isprovided

gear retraction with weight off wheels, hydraulic pressure

by the NOSE STRUT switch on the landing gear control

from the combined system bypasses the steering unit shutoff

panel (see Figure 2-76.) The three-position (EXTD, OFF,

valve to center the nosewheel independent of rudder pedal

and KNEEL) toggle switch is spring-loaded to return to the

movement. If the nosewheel iscocked beyond15

either side

center detent position of OFF. The position of the strut

of center after takeoff, the nosewheel is automatically

remains in the last commanded position independent of

prevented from retracting and the LAUNCH BAR advisory

electrical or hydraulic power interruptions. In both cases, the

light illuminates.

transfer control valve source of electrical power is the

essential No 2 bus and combined hydraulic system fluid is

During carrier arrestment, the nosewheel is centered

used as the transfer medium. With external electrical power

with weight on wheelsand hookdownwhenboth throttles are

on the aircraft, the combined hydraulic system must be

retarded to IDLE to prevent castoring during rollback. After

pressurized (>500 psi) before the control switch can com-

arrestment and rollback, the nosewheel will remain centered

mand a position change of the transfer control valve. The

until nosewheel steering is engaged.

control switch need only be held momentarily to effect a

change in transfer control valve position.

Selection of KNEEL releases hydraulic fluid from the

shock strut transfer cylinder to the combined hydraulic sys-

tem return line, causing the weight of the aircraft to compress

Nosewheel centering can contribute to launch bar

the shock strut 14 inches. Stroking of the nose strut causes the

misalignment in the catapult shuttle, which could

aircraft to rotate about the main wheels. The aircraft may be

result in premature launch bar separation during

taxied or towed in the strut-kneeled position except for the

launch. The nosewheel centering latching relay

nuisance trip of the launch bar at greater than 10

steering

must be deactivated by depressing thenosewheel

angle; this is the position used for taxiing onto the catapult

steering button after the hook check and prior to

and enhances accessibility to the forward fuselage compart-

entering the catapult. As this will also deactivate

ments during ground maintenance. Since the nose strut is

the nosewheel steering automatic disengagement

bottomed during the catapult launch stroke, the energy stored

function, the nosewheel steering must be manu-

in the last 4 inches of strut-piston stroke is released upon

ally disengaged when entering the catapult.

shuttle release at the end of the catapult stroke to impart a

noseup pitching moment to rotate the aircraft to the fly-away

2.27.3

Shimmy Damping

attitude without any control required by the pilot. All the

Shimmy damping is provided in the steering actuator.

stored energy is expended before the nosewheels leave the

Increased shimmy damping action is obtained with NWS

deck edge.

disengaged.

Note

Under certain launch conditions (high wind over

deck and light aircraft gross weights) the nose

strut will not be fully compressed during the cata-

If excessive nosewheel shimmy is encountered,

pult stroke. Subsequent nose rotation following

disengage nosewheel steering.

shuttle release will be at a less than normal rate.

Aircraft launch bulletins for the aircraft are

2.28 NOSEGEAR CATAPULT SYSTEM

written to ensure that catapult launch pressures

Catapult connection components on the nose landing

are sufficient to provide safe launch pitch rates

gear shock strut piston provide nosegear catapult capability.

and fly-away capability.

A launch bar attached to the forward face of the nosegear

steering collar guides the aircraft onto the catapult track and

Full extension of the nose strut after launch and weight

serves as the tow link that engages the catapult shuttle. A

off wheels provides a redundant and automatic transfer of the

holdback fitting secures the holdback restraint prior to

control valve to the extend position. With weight off wheels,

launch. The two-piston nose strut uses the stored energy

the NOSE STRUT switch is inoperative.

catapult principle to impart a positive pitch rotation move-

ment to the aircraft at shuttle release, thus providing for a

hands-off launch fly-away technique.

2-143

ORIGINAL

NAVAIR 01-F14AAD-1

NOMENCLATURE

FUNCTION

NOSE STRUT switch

EXTD — Hydraulic pressure causes strut to extend. Combined hydraulic

system must be pressurized before switch is activated on external

power. Launch bar is lifted into the up-lock position by torque arms

as strut extends 14 inches.

OFF — Spring-loaded return position.

KNEEL — Nose strut transfer control valve releases pressure in the shock

strut, which strokes 14 inches. Combined hydraulic system must

be pressurized before switch is active on external power. Launch

bar uplock can be released manually to allow bar to lower to deck,

or by turning nosewheel ± 10 .

LAUNCH BAR

Illuminates under the following conditions:

advisory light

Weight On Wheels

D Aircraft kneeled, throttles less than MIL (goes out when throttles are

advanced to MIL to provide lights out criterion for catapult launch).

D Launch bar not up and locked (normal operation)

Weight Off Wheels (inhibits nosegear retraction)

D Launch bar not up and locked

D Nosewheel not within ± 15

of center

D Nose strut not fully extended

LAUNCH BAR switch

ABORT — Enables pilot to disengage the launch bar from the catapult while

remaining at MIL power and in the kneel position.

NORM — Allows launch bar to be lowered.

Figure 2-76. Launch Bar Controls

ORIGINAL

2-144

NAVAIR 01-F14AAD-1

2.28.2

Launch Bar

deck-secured fittings. Prior to the application of single-

engine high power, the nose strut should be kneeled and slack

The launch bar is attached to the nosegear and serves

taken out of the holdback mechanism, otherwise dynamic

as the tow link for catapulting the aircraft (see Figure 2-77).

loads may exceed mechanism design strength conditions.

With the nose strut extended, the launch bar is held in the

retracted position. The launch bar can be lowered by kneeling

2.29 ARRESTING HOOK SYSTEM

the aircraft and turning the nosewheel greater than ±10

from

the centered position. The launch bar can also be lowered by

The arresting hook installation consists of a stinger

the deck crew with no pilot action after the aircraft has been

tailhook and associated control mechanism mounted to the

kneeled. A proximity sensing switch on the uplock detects

underside of the center fuselage. The hook shank is free to

the latch out of the locked position and illuminates the

pivot up and down at its attachment point. A pneumatic

LAUNCH BAR advisory light (see Figure 2-76). Ears on the

dashpot preloads the hook downto minimize hook bounce on

head of the launch bar engage under the lip of the catapult

contact with the deck. The hook shank is free to pivot left or

lead-in track and the head serves as a guide to steer the

right within a ± 26

sway angle with positive centering

nosewheel on the catapult track and engage the shuttle. For

action provided by a pneumatic damper housed inside the

an abort, the launch bar cannot be raised until the shuttle is

tailhook shank. The trail angle of the arresting hook provides

disengaged.

for hookpoint-deck contact even with the nose landing gear

strut fully compressed.

2.28.2.1

LAUNCH BAR Light

2.29.1

Arresting Hook Operation

The LAUNCH BAR advisory light is interlocked to go

off when both throttles are at MIL even though the launch bar

Normal operation of the arresting hook requires

position and mechanism remain unchanged; this action is

combined and flight hydraulic system pressure, dashpot

effected to establish a “lights out” criterion for launch. The

charged, and dc essential No. 2 electrical power. Because of

light circuit is disabled with nosegear up and locked. A

a redundant means of pilot control (electrical and mechani-

pilot-controlled LAUNCH BAR switch is installed that

cal), emergency extension of the arresting hook can be

enables the pilot to disengage the launch bar from the

accomplished without these sources of power.

catapult while remaining at MIL power and in the kneel

position. This switch is on the pilot left vertical console.

Note

Hookretraction requires electrical andcombined

hydraulic power.

2.29.1.1

Normal Operation

To avoid damage to the launch bar retract mech-

Normal operation

(Figure 2-78) on the pilot hook

anism, do not set the LAUNCH BAR switch to

control consists of a straight down-up movement of the

ABORT with the nosewheel deflected off center.

HOOK handle. This action actuates switches that provide

electrical command signals to the hook control valve. For

After the catapult launch stroke, extension of the strut

lowering the hook, the uplock is released and the lift cylinder

mechanically cams the launch bar up to the retracted-and-

is vented. Flight hydraulic pressure is the medium that

locked position. If the launch bar is not engaged in the uplock

disengages the hook uplock actuator. When flight hydraulic

with weight off wheels, the LAUNCH BAR advisory light

pressure drops below 2,100 psi with weight off wheels, the

will illuminate and nosegear retraction will be electrically

hook/auxiliary flap isolation relay circuit is energized. This

inhibited.

disables the arresting hook control valve and, therefore,

disallows normal hook extension. This condition remains

2.28.3

Holdback Fitting

until either the starboard engine-driven hydraulic pump

The holdback fitting is provided on the nose strut for

(flight) produces greater than 2,400 psi or weight on wheels

insertion of the holdback bar. Groundcrew must manually

is restored.

attach the bar before the aircraft is taxied into the catapult

lead track. The holdback bar is reusable and provides for

Note

repeated releases at a tow force of 76,000 pounds. Force

greater than this on launch causes the holdback bar to release

If emergency hook extension is inoperative in

the aircraft holdback fitting.

conjunction with a flight hydraulic failure,

cycling the HYD VALVE CONTR circuit

Single-engine, high-power turnup operations can use

breaker (8E5) with the hook handle down will

the holdback fitting to attach aircraft restraining hardware to

permit hook extension.

2-145

ORIGINAL

NAVAIR 01-F14AAD-1

2.29.1.2

Hook Retraction

3. Cockpit pressurization

For hook retraction, the control valve pressurizes the

4. Canopy seals

retract side of the lift cylinder and the lock side of the

actuator.

5. Windshield and canopy defogging

6. Windshield ant-ice

7. Anti-g suit inflation

Do not attempt to raise the hook when the hook

8. Wing airbag seals

is engaged in the arresting gear.

9. Gun-gas purging

When the arresting hook roller engages the uplock

10. Electronic equipment cooling and pressurization

mechanism, the lift cylinder is depressurized. On deck, hook

retraction time is approximately 3 seconds. The hook transi-

11. Temperature control of liquid coolant supplied to

tion light is illuminated as long as a discrepancy exists

APG-71radar control system, television camera set,

between the hook and cockpit handle positions. On-deck

and infrared search and track.

extension requires approximately 1 second. The transition

light will remain illuminated, unless the aircraft is kneeled,

2.30.1

ECS Air Sources

as contact with the deck precludes full hook extension.

2.30.1.1

Bleed Air

Note

The hook transition light may remain illuminated

The normal source of ECS air is ninth-stage bleed air

when the hook handle is lowered at airspeeds

from both engines. Through a series of manifolds and valves,

greater than300knotsbecauseofhookblowback.

this air is cooled and mixed to reduce temperature and pres-

sure to usable levels. The primary valves are the two engine

2.29.1.3

Emergency Hook Extension

bleed air shutoff valves, the dual pressure regulating and

shutoff valve, and the turbine compressor modulating and

The emergency control system lowers the hook by

shutoff valve, which are all controlled by the AIR SOURCE

mechanically (cable) tripping the uplock and venting the

selector pushbuttons: L ENG, R ENG and BOTH ENG

hook lift actuator pressure. Emergency extension of the hook

(Figure 2-79.)

may be initiated when the handle is in either UP or DN. In

either case, the hook handle is pulled aft (approximately

4 inches) and turned 90

counterclockwise. Rotation 90

2.30.1.2

RAM AIR Source

counterclockwise will lock the handle in the extended

position. With the handle locked, the hook will not retract

If either the RAM or OFF pushbutton is selected by the

regardless of the handle position (UP or DN).

pilot, the cooling turbine compressor is shut down and emer-

gency ram air can be used to ventilate the cockpits and pro-

Note

vide cooling air to the service and suit heat exchanger and

Afteremergency hookextension, thehook can be

those electronic subsystems requiring forced air cooling.

retracted airborne or on deck provided that the

However, if OFF is selected, pressurization to the service

handle is rotated

clockwise, pushed full

systems (canopy seal, anti-g suit, external fuel tank, wing

forward, and placed in UP. Combined and flight

airbag seal, OBOGS), and 400

F air supply to the windshield

hydraulic system pressures are required toretract

air defog and heating systems is lost. Selecting AIR

the hook while airborne. On deck, only combined

SOURCE RAM will provide air to the service systems and

hydraulicsystempressureisrequiredtoretractthe

400

F manifold air to the defog and heating systems.

hook.

2.30 ENVIRONMENTAL CONTROL SYSTEM

The ECS regulates the environment of flightcrew and

electronic equipment. The system provides temperature-

Selection of the AIR SOURCE pushbutton to

controlled, pressure-regulated air for the following systems.

RAM with a failure of the 400

F temperature

1. External drop tank pressurization

manifold will continue to circulate

400 F air

throughout the system surrounding aircraft

2. OBOGS

components and may cause a fire.

ORIGINAL

2-148

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

AIR SOURCE selector

RAM Ċ

Closes the bleed air flow modulator pressure regulator

pushbuttons

and shutoff valve, thereby securing the cooling bootstrap

turbine compressor. Inhibits gun firing. The RAM AIR

switch is enabled. Combined ram air and regulated 400_

F bleed air are available to the cockpits and air cooled

electronic equipment for temperature control. When

either BOTH ENG, L ENG or R ENG are selected, the

ram air door automatically closes.

L ENG Ċ

The left engine is the source of bleed air for the environĆ

mental control system and the right engine bleed air shutĆ

off valve is closed.

R ENG Ċ

The right engine is the source of bleed air for the environĆ

mental control system and the left engine bleed air shutoff

valve is closed.

BOTH ENG Ċ The right and left engine bleed air shutoff valves are open

and both supply bleed air to the environmental control

system. This is the normal position. Automatically closes

ram air door.

OFF Ċ

Both the left and right engine bleed air shutoff valves

and the dual pressure regulator valve are closed. Inhibits

gun firing. Pressurization and air conditioning are not

available. Enables the RAM AIR switch.

TEMP thumbwheel control

Selects cockpit and suit air temperature. It can be rotated through a 300_

arc (0 to 14) with mechanical stops at each end placarded COOL and

WARM. A midposition temperature (7) is approximately 70_ĂF in the

automatic mode. With the TEMP mode selector switch in AUTO the

temperature selected is automatically maintained by the modulating

temperature control valves. In MAN, the TEMP control thumbwheel must

be repositioned to maintain cockpit and suit air temperature. Air flow and

temperature will not change as a function of airspeed and altitude.

CABIN PRESS ALT

Displays cabin pressure altitude in 1,000−foot increments from

indicator

0 to 50,000 feet.

BLEED DUCT caution light

Indicates overheating (575° F or greater) along the high−temperature

bleed air duct routing forward of the engine fire wall past the primary heat

exchanger and then up to the right diverter area. An additional sensor,

detecting temperatures of 255° F or greater, senses from the right diverter

area, along the 400° F manifold and into the bootstrap turbine

compartment.

CABIN PRESS caution

Indicates cabin pressure is less than 5−psi absolute pressure or cockpit

light (RIO’s cockpit)

altitude is above 27,000 feet.

Figure 2−79. Air−Conditioning and Pressurization Controls and Indicators (Sheet 2 of 2)

ORIGINAL

2−150

NAVAIR 01-F14AAD-1

Interconnects inhibit gun firing with RAM or OFF

remainder is further cooled by the turbine compressor. Here

selected. The emergency ram-air door is on the lower right

the air is compressed, run through the secondary heat

side of the fuselage, inboard of the right glove. To activate the

exchanger, and then expanded in the turbine section,

ram air door, either the OFF or RAM AIR SOURCE

resulting in cold air that is mixed with 340_ F air to obtain

pushbutton must be depressed and the RAM AIR switch on

any temperature desired. The primary and secondary heat

the air-conditioning control panel must be moved to OPEN.

exchangers are between the left and right engine inlets and

the fuselage. At speeds above 0.25 Mach, ram air across the

heat exchangers is used for cooling. During ground opera-

tions and at airspeeds less than 0.25 Mach, airflow across the

heat exchanger is augmented by air-powered turbine fans.

Note

D Before opening the ram air door, reduce air-

With the system in MAN to increase airflow to

speed to 350 knots or 1.5 Mach, whichever is

forced-air-cooled equipment, place CANOPY

lower, to prevent ram air temperatures above

DEFOG-CABIN AIR control lever in CANOPY

110_ F from entering the system. After ram air

DEFOG.

flow is stabilized, airspeed may be varied as

The third heat exchanger is the service air-to-air heat

required for crew comfort or to increase flow

exchanger. This normally uses cold air from the cold-air

to electronic equipment.

manifold as a heat sink but can use emergency ram air if the

D With AIR SOURCE OFF selected, limit air-

cold-air manifold is not operating. Air from the service heat

speed to less than 300 knots/0.8 Mach to pre-

exchanger is used by the pressure suit, anti-g suit, canopy

vent damage to the deflated wing airbag seals.

seal, OBOGS, servo air, and for pressurization of wave-

guides, the radar liquid cooling loop tank, and the television

For maximum cockpit ram-air flow, the cockpit

camera set.

pressurization must be dumped. Pressing either L ENG, R

ENG or both ENG pushbuttons automatically closes the

2.30.2.1

Temperature Management

ram-air door if it is open.

The pilot can control cockpit temperature by selecting

2.30.1.3

External Air

either a manual (MAN) mode or automatic (AUTO) mode

with the TEMP mode selector switch (Figure 2-79). In the

The adapter for connecting a ground air-conditioning

AUTO mode, temperature (60_ F to 80_ F) is selected by

unit is under the fuselage, aft of the nose wheelwell. An

the pilot with the TEMP thumbwheel control. This desired

additional provision for connecting an external source of

temperature is maintained by a cabin temperature sensor in

servo air is in this same area.

the forward left side of the cockpit. In the MAN mode, the

External electrical power is automatically inhibited

TEMP thumbwheel control maintains airflow and tempera-

from AYK-141, IRST, TR1, TR2, and the CIU if external

ture. If cockpit inlet airflow temperature (in either AUTO or

air-conditioning is not connected to the aircraft. A pressure

MAN) exceeds 250_ F, a cockpit overtemperature switch

switch interrupts electrical power to the above forced-air-

closes the hot-air-modulating valve.

cooled equipment.

The conditioned air entering the cockpit is divided

forward and aft, with 50 percent of the air going to each

2.30.2

Cockpit Air-Conditioning

cockpit. A CANOPY air diffuser lever on the right console

ECS manifolding consists of:

in each cockpit individually controls the percentage of

airflow through the cockpit diffusers and the canopy defog

1. The high-temperature (bleed air) manifold

nozzles. When the lever is in CABIN AIR (full aft), 70

2. The 400_ manifold

percent of the air is directed through the cockpit diffusers and

30 percent through the canopy defog nozzles. In DEFOG, 100

3. The cold-air manifold.

percent of the air is directed through the canopy defog

High-temperature engine bleed air is routed through

nozzles.

the primary heat exchanger. The cooled output of this heat

exchanger is split and a portion is mixed with hot engine

2.30.2.2

Vent Airflow Thumbwheel

bleed air to a temperature of approximately 340_ F; the

This control has no function.

2-151

CHANGE 1

NAVAIR 01−F14AAD−1

2.30.2.3

Anti−G Suit

2.30.3.1.2

Ground Operation

Each anti−g suit is connected to the aircraft pressurizaĆ

During ground operation with electrical power, exterĆ

tion system by an anti−g suit hose that delivers pressurized air

nal air−conditioning, and servo air available to the aircraft

to the suit control valve and then to the suit through a

and the GND CLG switch in RADAR, the cockpit low−flow

composite disconnect. Below 1.5g, the suit remains deflated.

sensor is overridden. The OFF position of the GND CLG

A spring−balanced anti−g valve automatically opens when

switch enables cockpit air priority. With engines running on

g forces exceed 1.5g. Operation of the anti−g suit valve may

the ground, select OFF on the ground cooling switch.

be checked by depressing the test button marked G VALVE

on each flightcrew’s left console.

2.30.3.2

Cockpit Air Priority Function

2.30.3

Electronic Equipment Cooling

The cockpit air priority function is operational during

all engine−on operations (FO−14). It provides the cockpit with

Ambient cooled equipment in the electronic bays is

priority over the radar liquid−cooling loop in the event there

cooled by the air exhausted from the cockpits. Equipment

is a shortage of conditioned air. On engine power the GND

incapable of being cooled by free convection is cooled from

CLG switch (Figure 2Ć80.) should always be in OFF and the

the cold−air manifold.

canopy locked to enable the cockpit air priority function.

A schematic of the radar and electronic equipment

There is no indication to the flightcrew that the cockpit

cooling is shown in FO−14. Controls and lights are shown in

priority action is taking place unless it progresses to the point

Figure 2Ć80.

that the SENSOR COND advisory light illuminates. Even

then, it is only one of several problems that could have

2.30.3.1

Radar Liquid Cooling

triggered the light.

Radar equipment is cooled by liquid coolant (FO−14).

The heat is rejected in the ram air heat exchanger. This is

2.30.4

Pressurization

accomplished by circulating coolant fluid through the

electronics and ram−air heat exchanger and/or the radar heat

2.30.4.1

Cockpit Pressurization

exchanger. The cooling loop is also used for automatic

From sea level to 8,000 feet altitude the cockpit is

warmup of the radar using 400_F manifold.

unpressurized. Between altitudes of 8,000 feet to 23,000 feet

The radar liquid cooling loop incorporates a separate

the system maintains a constant cockpit pressure altitude of

ram−air liquid−heat exchanger. A ram−air door is located

8,000 feet. At altitudes above

23,000 feet, the cockpit

under the right glove, forward of the primary heat exchanger

pressure regulator maintains constant 5−psi pressure differenĆ

inlet. There are no cockpit controls for this ram−air door. It

tial greater than ambient pressures. An illustration of the

is controlled by the radar controller and is independent of the

cabin pressure schedule is shown in Figure 2Ć81.

air−conditioning and pressurization system. The radar system

ram−air heat exchanger automatically maintains the liquid

2.30.4.1.1

Cockpit Pressure Indicators

temperature within operating limits when ram air is used for

A cockpit pressure altimeter (Figure 2Ć79) is provided

cooling.

for the pilot. The rear cockpit has a CABIN PRESS light on

the CAUTION and ADVISORY panel. The CABIN PRESS,

2.30.3.1.1

Controls and Lights

light illuminates when cockpit pressure drops below 5 psi

Figure 2Ć80 shows the controls and lights associated

absolute pressure or cockpit altitude is above 27,000 feet.

with the radar cooling loop. The radar cooling loop is

activated by the RADAR COOLING switch on the RIO left

2.30.4.1.2

Cockpit Pressure Malfunctions

outboard console. In ON, the radar cooling loop is activated

If the cockpit pressure regulator malfunctions, the

for airborne operation. A temperature sensor in the heat

cockpit safety valve will open to prevent a cockpit pressure

exchanger outlet illuminates the SENSOR COND advisory

differential from exceeding a positive 5.5−psi or a negative

light when the liquid temperature goes above 104_F. In

differential of 0.25 psi. The cockpit pressure regulator and

addition, a pressure switch in the radar pump illuminates the

the safety valve are pneumatically operated and function

SENSOR COND advisory light when pump output pressure

independently through separate pressure sensing lines.

is too low.

If the coolant pump temperature rises to 230 ± 5_F, the

thermal switch opens, shutting down the pump to prevent

pump failure and illuminating the SENSOR COND advisory

light.

ORIGINAL

2−152

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

GND CLG switch

APG−71 Ċ Cockpit low flow sensor is overridden.

CAUTION

• Servo air required to actuate servo operated valves.

• Use RADAR only when engines are shut down.

OFF Ċ Cockpit low flow interlock is operational. OFF shall be selected

when engines are operating.

C & D HOT caution light

Indicates DD or PTID overheat condition.

Figure 2−80. Avionic Equipment Liquid Cooling Controls and Lights (Sheet 2 of 2)

2.30.4.1.3

Cockpit Pressure Dump

D Extended operations in clear air with the

windshield air on may cause windshield

Cockpit pressurization can be dumped by the pilot by

cracking and discoloration.

selecting DUMP with the CABIN PRESS switch. When

2.30.6

Gun−Gas Purging

DUMP is selected the safety valve is immediately opened

and the cockpit is depressurized.

External airflow is used to ventilate the gun compartĆ

ment for gun−gas purging. A flush air inlet on the fuselage

2.30.4.2

Canopy Seal Pressurization

gun bump and an aft louvered door containing a FOD screen

provide a continual flow of air to purge gun gases.

Pressurized air from the air−conditioning system is

ducted through the cockpit to the canopy seal. The seal is

2.30.7

Degraded ECS Operation

automatically inflated when the canopy actuator is moved to

There are various temperature and pressure safeguard

the closed position. A check valve in the canopy pressure

systems that cause the ECS system to shut down if an unsafe

regulating valve prevents the loss of canopy seal pressurizaĆ

situation is detected. A complete failure of the dual valve will

tion if the conditioned air manifold is depressurized. Initial

cause it to shut down the pressurization and air−conditioning

movement of the canopy actuator automatically deflates the

system. Should that fail to close, a pressure switch will

seal.

close both engine bleed air shutoff valves if an overpressure

(155 psi) situation exists in the outlet of the primary heat

2.30.5

Windshield Air and Anti−Ice

exchanger. A shutdown of the bleed air supply duct, either

automatically or pilot−selected AIR SOURCE OFF pushĆ

Compressor bleed air at approximately 340_ĂF and at

button, will cause total ECS air shutdown.

high pressure is directed over the outside of the windshield

through a fixed−area nozzle. This blast of hot air over the

windshield will evaporate rain and ice and prevent its further

accumulation. It is activated by selecting ON with the

WSHLD AIR switch. A temperature overheat sensor at the

base of the windshield protects the windshield from overĆ

Failure of the left or right weight−on−wheels

heating. When the sensor detects overheating (300_ĂF), a

switches to the in−flight mode can cause loss of

signal closes the pressure regulating valve and illuminates

engine ejector air to the IDGs and hydraulic heat

the WSHLD HOT advisory light on the pilot CAUTION

exchangers causing thermal disconnect and/or

ADVISORY light panel (Figure 2Ć82).

heat damage to the generators and aircraft

hydraulic systems.

Note

After an automatic shutdown of the system, the

pilot should select either OFF or RAM AIR

D Selecting WSHLD AIR ON prior to entering

SOURCE to enable the emergency ram−air door

rain or icing conditions may cause windshield

and then hold ram air switch to OPEN for approxĆ

cracking because of the rapid cooling effects

imately 50 seconds to provide ram−air cooling to

of precipitation.

electronic equipment and to the cabin.

ORIGINAL

2−154

NAVAIR 01−F14AAD−1

Figure 2Ć81.ĄCabin Pressure Schedule

Note

1. No cockpit airflow.

2. RIO COOLING AIR advisory light illuminated.

D Loss of electrical power with bleed air still

operating will result in smoke entering the

3. RIO SENSOR COND advisory light illuminated.

cockpit through the ECS when the aircraft is

on the deck. In flight only cold air will be supĆ

4. If ram−air cooling is not selected, extended flight

plied to the cabin and suit. Icing of the water

with AIR SOURCE OFF could cause an overĆ

separator may occur, causing reduced flow to

heating condition of the converter interface unit and

the cabin. Since the ECS panel is dependent

a subsequent loss of primary attitude and navigaĆ

on electrical power, selector pushbuttons will

tional indications

(i.e., multifunction displays,

be inoperative.

HUD, NAVAIDs).

D Retarding throttles to IDLE above 30,000 feet

The pilot should press the AIR SOURCE RAM

may result in a considerable reduction in ECS

pushbutton and set the RAM AIR switch to OPEN to open the

airflow, leading to a loss of cockpit pressurĆ

ram−air door to provide forced−air cooling to the electronic

ization, SENSOR COND light, and/or

equipment and to the cabin.

COOLING AIR light.

ECS duct failures may be indicated by diminishing

If the 400_ manifold reaches 475_ F, a 400_ĂF shutoff

cabin cooling airflow and/or cabin pressurization with or

valve closes, stopping the flow of unconditioned engine

without COOLING AIR advisory light illumination. Duct

bleed air to the 400_ F manifold. If either compressor inlet

failures may additionally be indicated by pressurization loss

or turbine inlet temperature becomes excessive, the refrigĆ

to the service systems and airflow loss to rain removal, defog,

eration unit will shut down. Cockpit indications will be as

and heating systems. This cannot be verified if the system is

follows:

2−155

ORIGINAL

NAVAIR 01−F14AAD−1

not in use. Selection of AIR SOURCE OFF and RAM AIR

concentrator and the monitor continue to function as long as

OPEN is appropriate when any indication of duct failure

the OBOGS switch is in the ON position. The monitor will

exists. ECS malfunctions that are not caused by duct failure

automatically shift back to the OBOGS supply source when

are usually indicated by loss of temperature control without

it detects adequate concentrator output.

a cabin or system airflow/pressurization degradation. Failure

Power to the monitor is provided by 28 Vdc OBOGS

of the 400_F modulating valve or duct should not cause

control power from the essential dc bus No. 1 when the

illumination of the cooling air light. Any duct failure in this

OBOGS master switch is in the ON position. The sensor in

area associated with the COOLING AIR light is strictly

the monitor is heated for proper operation. Upon initial

coincidental. However, the duct failure between the primary

selection of the OBOGS master switch to ON, the OBOGS

heat exchanger and the turbine compressor assembly, or

is powered and functioning but the monitor will not be

between the secondary heat exchanger and the turbine

accurately detecting oxygen concentration until the sensor is

compressor assembly, could cause degraded cooling airflow

warmed up. This can take up to 2 minutes, depending on the

and a COOLING AIR light to illuminate.

ambient temperature. The OBOGS light will not be illumiĆ

Actuation of the overtemperature switch results in

nated during the warmup period. The pilot may test operation

cycling of the 400_F valve. During this period the heating

of the monitor via the press−to−vent TEST button. The button

capacity of the 400_ F manifold would be degraded.

actuates a valve that must be held for up to 1 minute to vent

oxygen sensor. Laboratory testing has demonstrated that the

2.31 OXYGEN SYSTEM

test can normally be completed in approximately 15 seconds.

Once vented, the monitor will sense insufficient oxygen,

Breathing oxygen is provided to each crewmember by

illuminating the cockpit caution lights and shifting the

the OBOGS. A backup oxygen system provides a supply of

oxygen supply source to BOS. The monitor will automatiĆ

gaseous oxygen sufficient for a maximum range descent in

cally shift back to OBOGS operation and extinguish the

the event of a failure of the OBOGS. In addition, emergency

caution light after release of the TEST button. Testing has

oxygen is available to each crewmember through a high−

demonstrated this occurs within 5 to 7 seconds, but may take

pressure, gaseous oxygen bottle located in the ejection seat

up to 20 seconds.

survival kit.

2.31.1

On−Board Oxygen Generating System

(OBOGS)

The OBOGS provides 95−percent pure pressure− and

The aircrew will not have any indication of a failĆ

temperature−regulated oxygen to each crewmember. The

ure of the monitor. If the aircrew suspects the

system includes an oxygen concentrator, an oxygen monitor,

onset of hypoxia at any time, immediately select

and two regulators. Controls and indicators for the OBOGS

BACKUP. The monitor may be tested once the

are shown in Figure 2Ć83.

aircraft has descended to a cabin altitude of

The oxygen concentrator is in the right side of the

10,000 feet or less and the ON position on the

fuselage adjacent to and beneath the forward cockpit.

OBOGS master switch has been reselected.

Filtered and cooled ECS service air is directed to the oxygen

The OBOGS regulators are chest−mounted, pressure−

concentrator when ON is selected on the OBOGS master

demand type through which pressure−and temperature−

switch on the pilot cockpit panel. A molecular sieve in the

regulated oxygen is provided to each crewman. Pressure

concentrator removes the nitrogen from the compressed air,

breathing is activated above 34,000−foot cabin altitude.

leaving a breathing gas equivalent in concentration to

95−percent oxygen at 34,000 feet. The oxygen concentrator

When the OBOGS master switch is on, filtered, cooled

receives 115−Vac motor power from the pilot ac essential bus

engine bleed air is directed to the oxygen concentrator where

No. 1 and heater power from the ac right main bus. OBOGS

a molecular sieve removes the nitrogen from the compressed

28−Vdc control power is provided by essential dc bus No. 1

air, leaving a breathing gas consisting of 95−percent oxygen.

via the OBOGS CONTR circuit breaker (7A1).

The oxygen monitor checks system operation to ensure that

a sufficient concentration of oxygen is being generated,

The oxygen monitor is on the pilot right console. It

provides a cockpit indication, and brings the backup gaseous

constantly monitors the oxygen concentrator output to ensure

supply on line as required. A test button on the monitor

a sufficient concentration of oxygen is being generated.

enables the pilot to verify that the monitor and the backup

When the monitor detects an oxygen partial pressure less

oxygen system are functioning. When pressed, the OBOGS

than 182 mm Hg, it generates an alarm signal that illuminates

advisory light illuminates indicating the system is in backup.

the OBOGS caution lights, shuts off output from the

concentrator and enables the backup oxygen system. The

2−157

ORIGINAL

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

OBOGS master switch

BACKUP Ċ Deenergizes oxygen concentrator and Process Air Shutoff

Valve. Enables Backup Oxygen.

ON Ċ

Applies power to oxygen generator and oxygen monitor.

Opens solenoid valve providing ECS Service Air to Oxygen

Concentrator.

OFF Ċ

Removes power from OBOGS and BOS. Process air shutoff

valve closes.

VENT AIRFLOW

NOTă FUNCTIONAL.

OXYGEN SUPPLY valve

ON Ċ

Opens oxygen supply permitting OBOGS or BOS oxygen flow

to crewmember.

OFF Ċ

Secures OBOGS and BOS oxygen flow to crewmember.

OBOGS caution light

Illuminates when OBOGS has failed or OBOGS master switch is in OFF

or BACKUP

B/U OXY LOW caution

Illuminates when pressure remaining in BOS assembly oxygen cylinder

lights

is below 200 psi.

BACKUP OXY PRESS

Indicates pressure remaining in BOS assembly oxygen cylinder.

Indicator

OXYGEN MONITOR

Provides functional test of the oxygen monitor, BOS, and OBOGS control

TEST Button

systems.

Figure 2−83. Oxygen System Controls and Indicators (Sheet 2 of 2)

2.31.2

Backup Oxygen System (BOS)

The BOS assembly consists of an oxygen cylinder,

pressure gauge, pressure regulator, fill port, pressure transĆ

The BOS consists of a BOS assembly, BOS controller,

ducer, low−pressure switch, manual shutoff valve, and quick

B/U OXY LOW caution light, and a BACKUP OXY PRESS

disconnect on a palletized assembly that is removable for

indicator. This system was designed to provide only enough

servicing and maintenance. A 200−cubic−inch, high−pressure

oxygen for maximum−range descent. In the event of an

cylinder containing 500 to 590 liters of gaseous oxygen at

OBOGS failure, the aircrew must take immediate action to

1,800 to 2,100 psi, respectively, provides a backup oxygen

conserve backup oxygen.

supply to the OBOGS. The BOS assembly is located on the

Switching to the backup system can be accomplished

right forward side of the fuselage, just below the forward end

three ways:

of the pilot cockpit.

1. Automatically upon monitor detection of an

The BOS controller enables flow from the BOS assemĆ

OBOGS failure or loss of OBOGS control power

bly via a diaphragm valve. This diaphragm valve is conĆ

2. Manually via direct selection of BACKUP on the

trolled by two solenoid valves and an aneroid valve. The BOS

OBOGS master switch

controller is in the BOS assembly compartment. Power for

automatic operation of the BOS controller is provided by

3. Automatically with total loss of electrical power or

28 Vdc essential bus No. 1 via the OBOGS CONTR circuit

selection of OFF on the OXYGEN system master

breaker (7A1). Alternate power is provided via the BOS

switch, when the aircraft is above 10,000 feet MSL.

CONTR/B/U OXY LOW circuit breaker (7A4) for automatic

Backup oxygen cannot be disabled above 10,000 feet

activation of backup oxygen in the event of a failure of the

MSL by turning the OXYGEN system master switch

OBOGS control relay and when BACKUP is manually

off. Therefore, the individual OXYGEN SUPPLY valves

selected.

(Figure 2Ć83.) in both cockpits must be used to turn off

oxygen flow to the personnel regulators.

2−159

ORIGINAL

NAVAIR 01−F14AAD−1

The B/U OXY LOW caution light is actuated by the

The pitot−static system is composed of two separate

BOS assembly low−pressure switch when the BACK UP

systems with individual pitot−static probes, one on each side

OXY PRESS gauge reads less than 200 psi, or when BOS

of the forward fuselage.

CONTR/B/U OXY LOW power is lost. Figure 2Ć84 provides

The left pitot pressure (P_T) probe supplies the pilot

backup oxygen breathing time for two crewmembers for

standby airspeed indicator and the left AICS programmer.

various cabin altitudes based upon BOS oxygen cylinder

The right pitot pressure (P_T) probe supplies the RIO standby

pressure.

airspeed indicator, the right AICS programmer, and the

CADC with airspeed indications. An electrical P_T input from

2.31.3

BOS Pressure Indicator

the left AICS programmer is supplied to the CADC backup

The BACK UP OXY PRESS indicator (Figure 2Ć83.),

channel as airspeed indications for wing sweep.

on the right side of the pilot right knee panel, shows the

The left and right forward static ports

(P_S1) are

pressure in the BOS assembly oxygen cylinder. The indicator

manifolded to provide static pressure to the pilot standby

will not function unless the BOS manual shutoff valve on the

airspeed indicator, standby altimeter, vertical speed indicaĆ

BOS assembly is open.

tor, and the CADC. Static pressure from the right aft (P_S2)

static ports supply the RIO standby airspeed indicator,

2.31.4

Emergency Oxygen Supply

standby altimeter, and the right AICS programmer. The static

The 50−cubic−inch oxygen cylinder in the survival kit

pressure from the left aft (P_S2) static ports supply the static

of each ejection seat provides a limited supply of gaseous

pressure to the left AICS P_S sensor. An electrical P_S input

oxygen. This oxygen cylinder can be manually activated in

from the left AICS programmer is supplied to the CADC

the event of a failure of the OBOGS and depletion of the

backup channel for wing sweep. Static pressure from the left

backup supply. The cylinder is charged to 1,800 to 2,100 psi

aft (Ps₂) static ports supply the left AICS programmer.

and a pressure gauge is visible on the inside face of the

The CADC and AICS programmers provide Mach

left−thigh support. Flow from the emergency cylinder is

number information to the digital flight control computers

routed through a pressure reducer and a shuttle valve, then

(DFCCs). The alpha computer, angle of attack probe (used

follows the path of the normal oxygen system, flowing

for displaying angle of attack to the pilot), and the AICS

through the oxygen regulator to the face mask. The supply of

programmers provide angle of attack information to the

oxygen available in the emergency cylinder is adequate for

DFCCs. Electrical interfaces are shown on Figure 2Ć85.

up to 8 to 10 minutes, depending upon altitude. The manual

actuation handle is a green ring under the left side of the

Note

survival kit cushion.

D With the in−flight refueling probe extended,

the pilot and RIO standby altimeters and airĆ

speed indicators show erroneous readings

because of changes in airflow around the

pitot−static probes.

Turn the OXYGEN supply valve to OFF before

pulling the emergency oxygen manual actuating

D The RUDDER AUTH caution light may illuĆ

handle if contamination of the normal system is

minate when the in−flight refueling probe is

suspected. Failure to do so will inhibit seatpan

extended. Press the MASTER RESET button

shuttle valve operation, preventing flow of emerĆ

to reset the light.

gency oxygen.

2.32.1

Pitot−Static Heat

Note

Each pitot−static probe is equipped with electrical

Flow of oxygen from the emergency cylinder can

heating elements to prevent icing. Pitot−static heat is

be stopped by reseating the manual actuation

controlled by the pilot through the ANTI−ICE switch on the

handle.

pilot right console. In AUTO/OFF, pitot probe heat is

available only with weight off wheels. ORIDE/ON activates

2.32 PITOT−STATIC SYSTEM

the probe heat elements independently of the weight−on−

wheels switch and illuminates the INLET ICE caution light

The pitot−static pressure system supplies impact (pitot)

on the CAUTION ADVISORY panel. OFF/OFF removes

and atmospheric (static) pressure to the pilot and RIO flight

heat from the probes.

instruments, to the CADC, and to the engine AICS programĆ

mers. Some systems require static pressure only; others

require static and pitot pressure (see Figure 2Ć85).

ORIGINAL

2−160

NAVAIR 01−F14AAD−1

BACK−UP OXYGEN PRESSURE

CABIN ALTITUDE

2000

1600

1200

800

400

200

35 & ABOVE

100

2.9

2.5

2.2

SL.

1.8

Minutes remaining based on two−man consumption.

Duration data should be used as a guide

Consumption rate based on 13.1 liters per minute per man.

Figure 2Ć84.ĄBackup Oxygen Duration Chart

1. Calligraphic or stroke writing is displayed on the

HUD, MFDs, and the DD.

2. Raster video (for example, radar, and television)

The ANTI−ICE switch should normally be in

generated internally (VDI formats) or provided by

AUTO/OFF during takeoff and landing. Engine

an external sensor, with or without a stroke overlay,

anti−icing has adverse effects on engine stall marĆ

is displayed on the MFDs and the DD.

gin.

3. Alphanumeric data is displayed on the multistatus

indicator and the radio frequency and radio frequenĆ

2.33

CONTROL AND DISPLAY SYSTEM

cy/control indicators.

The control and display system (Figure 2Ć86) provides

the crew with navigation, aircraft status, and flight tactical

Displays presented on the HUD and MFDs are

information. The control and displays system consists of two

identified as formats. The formats are categorized as display

display processors

(DP1 and DP2), three multifunction

format groups.

displays (pilot center MFD1, pilot right MFD2, and RIO

MFD3), and a heads−up display system, cockpit television

HUD format groups consist of takeoff/landing/

sensor, HUD−VIDEO panel, pilot displays control panel, and

navigation (TLN), air−to−air (A/A), air−to−ground (A/G), and

a multistatus indicator.

multimode formats that can be overlaid on the other three.

The HUD also displays a manual reticle and a test pattern.

The control and display system also sends display

information to the digital display, the radio frequency

MFD display format groups are shown in Figure 2Ć87.

indicator, radio frequency/control indicator, and the mission

The HUD and MFD VDI format groups are basically the

video recorder.

same; however, HUD symbology is scaled to be overlaid on

The data entry unit is a remote terminal that communiĆ

the real world, and certain differences, such as symbol

cates with the mission computers via the multiplex buses.

location, addition, and deletion occur between the HUD and

MFD VDI formats. MFDs also display repeats of the PTID

2.33.1

Display Types

and DD as well as TCS and CCTVS video.

The following types of display information are proĆ

vided by the MFD system:

2−161

ORIGINAL

NAVAIR 01-F14AAD-1

DISPLAY FORMAT

FORMATS WITHIN GROUP

NAV Align Formats

CV MAN DATA

CV Ships Inertial Navigation System (SINS) DATA

IFA

Standard Attitude Heading Reference System (SAHRS)

(Norm Mag SHDG)

SAHRS CV

Inertial Navigation System (INS) UPDATE Format

Continuous Update Formats

NAV AID CORRECTIONS

NAV AID ENABLED

NAV AID OPTIONS

SURFACE WAYPOINT POSITION Format

Stores Management System (SMS) Format

SPIN INDICATOR Format

ENGINE MONITOR Format

On Board Checkout (OBC) Formats

OBC BASIC

OBC Groups: CD, CNI, FLT, NAV, AUX, TAC, EW,

SMST, and SNSR

OBC Failed Data: CADC, CIU, SAHRS, DINS, GPS,

DEU, DP1, DP2, MC1, MC2, MDL/DSS, APC, EMSP1,

EMSP2, IFX, SMS, SWITCHES, RWR, and RDR/TCS

MAINT Format

Failure History Format (FHF)

Cooperative Support Software (CSS) Format

Missile Status Readout Formats

MISSILE SUBSYSTEM 1

MISSILE SUBSYSTEM 2

Electronic Counter Measures (ECM) Format

Recon Formats

RECON DATA

RECON WPT DATA1

RECON WPT DATA2

Tactical Situation Display (TSD)

TSD MENU

TSD PRIORITY

TSD DECLUTTER 1

TSD DECLUTTER 2

TSD COMMAND

TSD REPLY

TSD TARGET MODIFIER

JTIDS Data Readouts (JDR)

OWN AIRCRAFT PPLI

AIR PPLI

NON-AIR PPLI

INDIRECT PPLI

Infrared Search and Track (IRST) Formats

IRSTS NORMAL

IRSTS CSCAN

IRSTS SUMMARY

Figure 2-87. Display Format Groups (Sheet 2 of 2)

CHANGE 1

2-168

NAVAIR 01−F14AAD−1

2.33.2

Display Processors

is on the PDCP and the MFD power switches are on each

MFD as a part of the DAY/AUTO/NIGHT switch.

Two display processors

(DP1 and DP2) drive the

display system. The DPs receive various signal inputs from

After a short warmup (under 2 minutes), the default

the aircraft systems. These signals are processed and

formats appear on the displays. The default formats, with

converted to display information for the HUD, MSI, MFDs,

weight on wheels, are as follows:

DD, RFI, RFCI, and the mission video recorder.

1. HUD Ċ TLN basic

2.33.2.1

Normal Operation

2. MFD Ċ VDI TLN basic

During normal operation, DP1 drives the HUD and

MFD1, while DP2 drives MFD2 and MFD3. Should either

3. MFD2 Ċ OBC basic

DP fail, the mission computer commands backup operation,

4. MFD3 Ċ OWN A/C basic.

where the remaining DP provides limited functions.

If the mission computers are not in communication,

2.33.2.2

DP Backup Operation

test patterns will appear on all four displays.

During backup operation, the remaining DP drives the

Format selection for the HUD is made by use of the

HUD, and MFD1. Should one of these three displays be OFF

MODE pushbuttons on the PDCP and by the type of steering

or subsequently selected off, then MFD2 will operate. If both

selected. MFD format families are selected by pressing the

stroke generators in the remaining DP are in use, an MFD

pushbutton adjacent to a menu legend or by cursor designaĆ

format that is normally produced by stroke writing may be

tion of the legend. Every MFD format (except repeats) has

generated in raster. With the following exception, either DP

MENU select as the center pushbutton on the lower edge of

can perform any display function: Mission video record is not

the display. Also appearing on all formats for immediate

performed during backup operation.

selection are SMS to the left of MENU and ECM to the right

of MENU. Other selections vary according to format

2.33.2.3

Data Failure Modes

requirements. When a repeat format (HUD, DD, or PTID) is

being displayed on the MFD, no legends are available for

In addition to the backup mode, there are other failure

format selection. To change formats from a repeat, press any

modes. Some examples are as follows.

pushbutton. This returns MENU1 to the MFD, permitting

If the DPs fail to receive pitch and roll data, the

other format selections to be made. Cursor designation of

message PITCH/ROLL FAIL will appear on the MFDs and

legends cannot be used with repeat displays.

all pitch/roll−related symbols are removed from the displays.

The symbols are returned if pitch and roll information is

2.33.4

Heads−Up Display (HUD)

restored.

The HUD (Figure 2Ć88) provides a combination of

If the DPs lose communication with the MCS, a manual

real−world cues and flight direction symbology, projected

reticle will appear on the HUD and the MFDs will display

directly on a combining glass assembly. The flight informaĆ

only the message DP−MC COMM FAIL and MENU1. The

tion on an optical combiner is projected in the pilot forward

lighted MODE pushbuttons also turn off with a loss of MCS

field of view. The display is focused at infinity, thereby

communication. Should communications be restored, the

creating the illusion that the symbols are superimposed on the

DP−MC COMM FAIL message is removed and the MODE

real world (and so that visual cues received from outside the

buttons are lighted again. If the MC performed a cold start or

aircraft are not obscured). The pilot usually steers based on

a system reset, default formats are presented on the displays.

interpretation of the visually observed real world. The HUD

can be selected to be the primary flight reference for all flight

2.33.3

System Operation

regimes displaying navigation and weapon delivery informaĆ

tion. The HUD symbol brightness control is on the HUD; all

The display system requires 115 V, 400 Hz electrical

other HUD controls are on the PDCP.

power DP1, HUD, and MFD1 receive power from ac

essential No. 2 bus and DP2, MFD2, and MFD3 are on the ac

2.33.4.1

Pilot Displays Control Panel (PDCP)

left main bus. All displays and DPs are electrically protected

by circuit breakers. There are no power switches for the DPs.

The PDCP on the pilot right console (Figure 2Ć89)

Each of the displays has a power switch that is normally

provides control of the mode and display presentation of the

turned off at the conclusion of flight. The HUD power switch

2−169

ORIGINAL

NAVAIR 01−F14AAD−1

Figure 2Ć88.ĄHeads−Up Display

HUD, VDI, ECM, and TCS formats. Display informaĆ

an MFD, or is being recorded, bright white flashes of video

tion is dependent on the mode selected with the A/A, A/G,

will be displayed or recorded. This is normal for BIT

and TLN pushbuttons.

operation.

2.33.4.2

Color Cockpit Television Sensor

2.33.5

Multistatus Indicator (MSI)

(CCTVS)

The MSI is an LCD panel on the lower center

The CCTVS is an electro−optical system that images

instrument panel below the center MFD

(MFD1)

symbology present on the HUD combiner and outside world

(Figure 2Ć86). The MSI displays the weapon type and status

information as well. The unit consists of a video sensor head

on each store station.

on the HUD and an electronic unit in the HUD−VIDEO panel

The lower row displays weapon status: ready,

(Figure 2Ć86)). The sensor signal can be fed to the mission

degraded, ready/selected or degraded/selected. The selected

video recorder and can be displayed on the MFDs. Operation

symbol never appears alone; it is always superimposed over

of the CCTVS is controlled by the VIDEO CONTROL

the ready or degraded symbol. Figure 2Ć90 provides a

switch on the HUD−VIDEO control panel.

representative display of available MSI symbols along with

their meanings.

2.33.4.3

HUD−VIDEO Control Panel

The upper row of the display identifies the weapon.

Operation of the CCTVS is controlled by the HUD−

Two dashed lines at a store station indicate that the missile at

VIDEO control panel (Figure 2Ć86)). The panel contains the

that station has failed or is hung. A blank display on a station

two−position VIDEO CONTROL toggle switch, a BIT

indicates no weapon is loaded or the weapon loaded is not

button, a green GO light, and a yellow NO GO light. Setting

recognized.

the VIDEO CONTROL switch to ON provides power to the

CCTVS; selecting OFF removes power. Depressing the BIT

There are no controls on the MSI. Power to the MSI is

button initiates a CCTVS self−test. A good test results in a

provided by the HUD subsystem. The MCS must be

momentary flash of the yellow NO GO light followed by a

transmitting data for a display to be presented. Selecting

steady green GO light. A failure results in a steady yellow NO

TEST on the HUD PWR switch causes all LCD segments on

GO light. During BIT, if CCTVS video is being displayed on

the MSI to be displayed.

ORIGINAL

2−170

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

TCS FOV (Television

NAR Ċ Selects TCS narrow field of view for display on pilot’s MFD1.

Camera Field of View)

WIDE Ċ Selects TCS wide field of view for display on pilot’s MFD1.

ECM switch

ORIDE Ċ Enables ECM display to override whatever is being displayed on

MFD2 for as long as the threat is being reported.

OFF Ċ ECM display override not enabled.

ELEV LEAD Control

A continuous rotary control that provides a range of elevation positions for the

HUD manual reticle with the 0 mr setting coincident with the armament datum

line (ADL). Clockwise rotation increases elevation lead.

HUD/VDI ALT source

BARO Ċ Selects barometric altimeter as source for display of altitude on

switch

HUD and VDI.

RDR Ċ

Selects radar altimeter as source for display of altitude on HUD

and VDI. Radar altitude is displayed as follows:

• Below 5000 feet AGL

• Radar altitude valid

• AOB ≤ 45_

HUD PWR switch

TEST Ċ

(Momentary) Presents an intersecting vertical and horizontal line at

(lever lock)

the center of the HUD field of view, and illuminates all segments of

the multistatus indicator (MSI).

ON Ċ

Provides power to HUD and MSI.

OFF Ċ

Removes power from HUD and MSI.

FORMAT switch

ANLG Ċ

Selects analog dial format for HUD display of airspeed and

altitude.

BOTH Ċ Selects a combination of analog dial and digital readout for HUD

display of airspeed and altitude.

DGTL Ċ

Selects digital readout format for HUD display of airspeed and

altitude.

DECLUTTER switch

NORM Ċ Normal display symbology is presented.

LVL 1 Ċ

Depending on MODE selected, the following symbols are

removed:

TLN ć GEAR UP

(AOA bracket and target pointer/AON are not displayed)

•

Vertical velocity

•

AOB scale

•

Peak G

Figure 2−89. Pilot Displays Control Panel (Sheet 2 of 3)

ORIGINAL

2−172

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

DECLUTTER switch

TLNć GEAR DOWN

(continued)

(Target pointer/AON and Mach are not displayed)

• Peak G (displayed as required in normal mode only)

• AOB scale

• Radaraltitude

A/A

• Radar altitude readout

A/A and A/G

(Vertical velocity, AOB

scal

e, and AOA bracket

are not displayed)

• AOA

eadout

• Potential flight path marker (PFPM)

LVL 2 − Depending on MODE selected, the following additional symbols are

removed:

TLN ć GEAR UP

• AOA

• Mach

• Nav range

• PFPM

• Radar altitude readout

• Digital boxes

• Clock/Timer readout

• Steering Mode

/WPT #/

Range

TLN ć GEAR DOWN

• AOA

• Digital boxes

• PFPM

• Vertical velocity

• Clock/Timer readout

• Steering Mode/WPT #/Range

A/A

• Nav range

A/G

(Closure and target pointer/AON are not displayed)

• Radar altitude readout

A/A and A/G

(AOB scale, AOA bracket, and vertical velocity are not displayed)

• Mach number

• Peak G

• Digital boxes

• Heading sca

• Gho st FPM

• Clock/Timer readout

• Steering Mode/WPT #/Range

CAGE ENBL/DSBL

Momentary contact pushbutton used to enable/disable HUD CAGE option.

Pushbutton

Caging restricts pitch ladder and flight path marker symbols in azimuth to the

center of the HUD display.

Figure 2−89. Pilot Displays Control Panel (Sheet 3 of 3)

2−173

ORIGINAL

NAVAIR 01−F14AAD−1

Figure 2Ć90.ĄMultistatus Indicator Symbols/Meanings

An MFD displays tactical and flight command situaĆ

Normally the pilot uses the MFD below the HUD

tions, navigation, and discrete information either separately

on the aircraft centerline as the primary in−the−cockpit flight

or simultaneously with radar and TV data. There is also a

instrument.

power/brightness select switch above the display screen

Attitude information is displayed on the MFD VDI

(Figure 2Ć91).

format by an aircraft reticle, a horizon line, and a calligraphic

pitch ladder. The aircraft reticle is fixed at the center of the

2.33.6

Multifunction Displays (MFD)

display, and the horizon line and pitch ladder move about it

The three identical MFDs are CRT displays with 20

in accordance with the aircraft pitch and roll attitudes.

pushbuttons around the perimeter of the display screen. The

The flight parameters displayed include magnetic

MFD pushbuttons with adjacent legends are used for menu

heading, data link

(D/L), commanded airspeed

(Mach

selection, data entry/readout, and system test and/or status

number), airspeed, altitude, and vertical velocity.

indications. The three programmable MFDs, two in the pilot

instrument panel and one in the RIO instrument panel

Note

provide display flexibility such that either crewmember is

able to select any display available, allowing the pilot and

If pitch or roll data is not updated within 240

RIO to monitor and back up each other. The HUD format may

milliseconds, the pitch ladder and roll marker

be repeated on any MFD by depressing pushbutton No. 11

will be blanked and the horizon, sky, and ground

from the MENU1 format.

plane will darken.

Multifunction pushbuttons with adjacent CRT legends

2.33.7

Cursor Controls

located around the perimeter of the MFD are used for menu

selection, data entry/readout, and system test and/or status

Both the pilot and RIO have cursor controls

indications. An MFD displays tactical and flight command

(Figure 2Ć92) that permit the remote selection of MFD

situations, navigation, and discrete information either sepaĆ

pushbutton options as well as symbol and spot hooking. A

rately or simultaneously with radar and TV data.

symbol is hooked when the cursor is placed over a format

ORIGINAL

2−174

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

Pushbuttons

20 momentary contact pushbuttons that provide for selection of display,

operating modes, and system parameters. A selected legend is normally

enclosed by a rectangular box. A dashed rectangular box indicates that a

legend has been selected but is not available.

CONT control

Varies the amplitude of the shades of gray. Effects are most visible when

viewing video or raster graphics.

BRT (brightness)

Varies intensity of overall display. As brightness is decreased, fewer shades

control

of gray are discernable.

Figure 2−91. Multifunction Display (Sheet 2 of 2)

symbol and cursor designate is activated. Hooking is used to

2.33.7.3

Cursor Hooking Functions

set waypoints on the HSD waypoint format and to select

Spot, symbol, and MFD pushbutton hooks can be

tracks and other symbols, for the purpose of obtaining

performed by the pilot on the HUD or MFD by activation

information, or identifying symbols of interest on the TSD

of the TDC or on the PTID or MFD by the RIO through use

format. The cursor symbol is a small circle inside a larger

of the sensor hand control.

circle when displayed on the MFDs and a circle with four tic

marks extending from the circle inward at 0_, 90_, 180_, and

Normal symbol hooking is accomplished by placing

270_ when displayed on the HUD.

the cursor over the desired symbol using either the TDC or

the SHC and activating the appropriate cursor designate

2.33.7.1

Throttle Designator Controller (TDC)

switch. The hooked symbol brightens and the previously

hooked symbol returns to normal intensity. Symbol hooks are

The pilot controls cursor position with the throttle desĆ

used to display additional information about those symbols

ignator controller. The TDC is a circular disk that is a comĆ

or to designate tracks for functions that are format dependent.

bination fourway force sensor and momentary switch on the

Only HUD, PTID, TSD, and IRST normal format support

outboard throttle grip. Finger pressure on the outer edges of

symbol hooking.

the control will move the cursor in the direction selected.

MFD pushbutton hooks permit remote activation of

When cursor movement exceeds the limit of a display that is

MFD pushbutton functions through the TDC or SHC. They

adjacent to another display (e.g., the right edge of MFD1 or

are accomplished by positioning the cursor over the desired

the bottom of the HUD), the cursor will move to the adjacent

MFD menu choice and activating the cursor designate

display. If the cursor symbol reaches a display limit that is not

switch.

adjacent to another display (e.g., the right edge of MFD2),

the cursor remains at that limit. Depressing and releasing the

2.33.8

Displays, Formats, and Symbology

TDC designates the cursor position.

The paragraphs that follow describe the HUD and MFD

2.33.7.2

RIO Cursor Control

displays. Sample formats from format families are illusĆ

trated, symbols associated with these families are identified

The RIO cursor control is on the sensor hand control.

and defined, and format selection is described.

It consists of a four−position select switch, a two−position

Many symbols are common to more than one format

(half−full action) trigger switch, and a handgrip. When the

family. Once a symbol has been defined for a format family,

top, bottom, or right edge of the select switch is pressed, the

the definition is not repeated when describing other format

DD, PTID, or MFD3, respectively, is selected for cursor disĆ

families. Certain features, such as changes in scaling

play. Pressing the left edge toggles sensor control between

between formats, that are obvious when viewing the display

radar and infrared. The cursor symbol becomes visible when

are not covered.

the trigger switch is pressed to the half−action position. Full

trigger depression designates the cursor position. Cursor

All symbols available to a format are illustrated;

symbol movement is controlled by handgrip movements.

however, they will rarely be displayed at the same time. Not

all formats are illustrated. Where only minor differences

exist, they will be noted. Formats that contain only alphaĆ

numerics are described but are not illustrated.

ORIGINAL

2−176

NAVAIR 01−F14AAD−1

2.33.8.1

Warning, Caution, Advisory Indicators

Note

The large cross that appears on the HUD when

Warnings are displayed on the lower center of the HUD

the HUD PWR switch is set to TEST is generated

viewing area. These warnings are: L FIRE, R FIRE, L

by the HUD, independent of the DP, and is used

STALL, R STALL, and RDC SPD. The CLSN advisory is

to check HUD operation.

also displayed on the HUD. If there are more than two warnĆ

ings, then they will scroll up at the rate of one warning per

The HUD and MFD test patterns also momentarily

second.

appear during IBIT and following a system reset. Both test

patterns are written in stroke and are used to check stroke

On the MFDs, warning/caution/advisory indications

accuracy.

are shown in a viewing window that appears on all formats

except repeats. This window is displayed in the upper left of

The MFD/KROMA test pattern (a future−growth color

the MFD and is referred to as the CAW (Caution Advisory

display) includes an MFD TEST legend, used to select the

Warning) window or CAW box. The message window allows

MFD RASTER test pattern.

up to four CAWs to be displayed at one time. If more than four

The MFD RASTER test pattern allows for testing of

CAWs are to be displayed, they scroll up from the bottom of

individual pushbuttons. When a button is pressed, a solid−line

the window at a rate of one per second. Warning, caution,

box appears around the PRESS legend; pressing the button

advisory legends are independent of format and may be

again removes the box. The diamond and blinking breakĆ

directed to a specific crewmember. Figure 2Ć93 lists specific

away symbol are used to check RASTER accuracy. Numerics

CAWs and the crewmember to whom they are directed.

0 through 7 check RASTER shades of gray. Selecting EXIT

returns the display to MFD/KROMA test pattern.

When warning, caution, or advisories are displayed,

pressing the pushbutton above the CAW window (PB6) will

2.33.8.3

HUD Formats

remove the window and replace it with a boxed CAW legend.

Pressing the CAW pushbutton when the legend is boxed

HUD format category (TLN, A/A, A/G) is normally

returns the window and indications to the display and

selected by use of MODE buttons on the PDCP. However,

removes the box from the legend.

air−to−air formats are selected automatically if the pilot

selects a weapon using the weapon select switch on the stick

Note

grip; selects RDR PLM/PAL, IR PLM/PAL (all with gear up)

with the sensor mode switch; lifts the ACM guard; or if VSL

If a repeat format is on MFD1, the CAW window

HI/VSL LO is selected with the sensor mode switch or DD.

is shifted to MFD2 in its current state, open or

Air−to−ground formats are automatically selected when an

closed (acknowledged). New CAWs continue to

air−to−ground weapon is selected on the SMS format. The

be displayed on MFD2 until the repeat format is

HUD default format is the TLN basic format (Figure 2Ć95).

removed from MFD1. If a repeat format is disĆ

This format is displayed on power−up and if DP1 experiences

played on both MFD1 and MFD2 or on MFD3,

a cold start (power outage of over 1 second).

receipt of CAW data removes the repeat format

from MFD1 and/or MFD3 and displays a new

The amount of information displayed on HUD formats

format with the CAW in the appropriate window.

is pilot selectable by means of the FORMAT and DECLUTĆ

Receipt of a data−link advisory removes the

TER switches on the PDCP. Symbols are also added or

repeat format from MFD1 (if appropriate) and

deleted by the mission computer depending on aircraft status,

MFD3 and displays the menu format with data−

steering mode, and weapon selection. When the FORMAT

link advisories.

(DD and PTID displays are

switch is set to BOTH, airspeed and altitude information are

repeats on MFD1; HUD, DD, and PTID displays

displayed as boxed digital readouts with analog dials. In the

are repeats on MFD2 and MFD3).

ANLG position, the boxes are removed from the digital

readouts. In the DGTL position, only the boxed digital

2.33.8.2

Test Patterns

readout is presented and the analog dials are removed.

The test patterns (Figure 2Ć94) appear on the HUD and

The position of the HUD/VDI ALT switch on the PDCP

MFD when the display system is turned on with the MCs off

selects the type of altitude data that is to be displayed, either

during ground tests and are generated by the DP.

radar or barometric. If radar is selected and a valid

ORIGINAL

2−178

NAVAIR 01−F14AAD−1

ACRONYM

TYPE

AIRCREW

DISPLAY

FUNCTION

L STALL

BOTH

HUD/MFD

Warns of left engine stall.

R STALL

BOTH

HUD/MFD

Warns of right engine stall.

L FIRE

BOTH

HUD/MFD

Warns of fire in left engine.

R FIRE

BOTH

HUD/MFD

Warns of fire in right engine.

RDC SPEED

PILOT

HUD/MFD

Safe Mach number exceeded for current

position of flaps.

W/S

PILOT

MFD

Indicates failure of wingsweep system.

L N2 OSP

PILOT

MFD

Indicates overspeed of left rotor N₂.

R N2 OSP

PILOT

MFD

Indicates overspeed of right rotor N₂.

L N1 OSP

PILOT

MFD

Indicates overspeed of left rotor N₁.

R N1 OSP

PILOT

MFD

Indicates overspeed of right rotor N₁.

L TBT OT

PILOT

MFD

Indicates overtemp of left turbine blade.

R TBT OT

PILOT

MFD

Indicates overtemp of right turbine blade.

L FLMOUT

PILOT

MFD

Indicates left engine flameout.

R FLMOUT

PILOT

MFD

Indicates right engine flameout.

L IGV SD

PILOT

MFD

Indicates left inlet guide vane adjust schedule

is not correct.

R IGV SD

PILOT

MFD

Indicates right inlet guide vane adjust schedule

is not correct.

A/P REF

PILOT

MFD

Indicates autopilot mode is selected but not

engaged.

CLSN

PILOT

HUD

Indicates collision course steering to target has

been selected.

IFF ZERO

RIO

MFD

Indicates the identification friend or foe

transponder is not operating correctly.

Figure 2Ć93.ĄWarning, Caution, Advisory Functions (Sheet 1 of 3)

2−179

ORIGINAL

NAVAIR 01−F14AAD−1

ACRONYM

TYPE

AIRCREW

DISPLAY

FUNCTION

AAI ZERO

RIO

MFD

Indicates the air to air intercept interrogator is

not operating correctly.

SDU ALM

RIO

MFD

Indicates the JTIDS Secure Data Unit is not

operating properly or does not contain valid

JTIDS crypto keys. Under certain conditions the

display of this alarm is normal.

ASPJ HOT

RIO

MFD

Indicates an overtemp condition of the airborne

self−protection jammer.

JTID HOT

RIO

MFD

Indicates an overtemp condition of the JTIDS

R/T.

RWR

RIO

MFD

Indicates the radar warning receiver is not

operating correctly.

FWD ASPJ

RIO

MFD

Indicates the forward ASPJ is not operating

correctly.

AFT ASPJ

RIO

MFD

Indicates the aft ASPJ is not operating corĆ

rectly.

AFT CG

BOTH

MFD

Indicates that stores station status has shifted

center of gravity to preclude landing without

correction.

MC 1

RIO

MFD

Indicates mission computer 1 is not operating

correctly.

MC 2

RIO

MFD

Indicates mission computer 2 is not operating

correctly.

CIU

RIO

MFD

Indicates the computer interface unit is not

operating correctly.

MC1 HOT

RIO

MFD

Indicates an overtemp condition of the mission

computer #1.

MC2 HOT

RIO

MFD

Indicates an overtemp condition of the mission

computer #2.

INS

RIO

MFD

Indicates the inertial navigation system is not

operating correctly.

IMU

RIO

MFD

Indicates the inertial measurement unit is not

operating correctly.

GPS FAIL

RIO

MFD

Indicates the MAGR unit is failed or turned off.

POSITION

RIO

MFD

Position differs. Will be displayed if:

D GPS and INS differ by more than 4 nm.

D GPS and SAHRS differ by more than 11 nm.

D SAHRS and INS differ by more than 13 nm.

Figure 2−93. Warning, Caution, Advisory Functions (Sheet 2 of 3)

ORIGINAL

2−180

NAVAIR 01−F14AAD−1

ACRONYM

TYPE

AIRCREW

DISPLAY

FUNCTION

GPS QUAL

RIO

MFD

Degradation of GPS occurred so GPS output

is invalid.

VELOCITY

RIO

MFD

Feet/second.

CIU HOT

RIO

MFD

Indicates an overtemp condition of the CIU.

DP1 HOT

RIO

MFD

Indicates an overtemp condition of display

processor 1.

DP2 HOT

RIO

MFD

Indicates an overtemp condition of display

processor 2.

SMS HOT

RIO

MFD

Indicates an overtemp condition of the stores

management system.

RDR HOT

RIO

MFD

Indicates an overtemp condition of the radar

system.

HUD HOT

PILOT

MFD

Indicates an overtemp condition of the HUD.

RWR HOT

RIO

MFD

Indicates an overtemp condition of the radar

warning receiver.

DSS HOT

RIO

MFD

Indicates an overtemp condition of the data

storage system.

DEU HOT

RIO

MFD

Indicates an overtemp condition of the data

entry unit.

MPS HOT

RIO

MFD

Indicates an overtemp condition of the missile

power supply.

IRSTS HOT

RIO

MFD

Indicates an overtemp condition of the infrared

search and track system.

TARPS

RIO

MFD

Indicates the tactical air reconnaissance pod

system is not operating correctly.

IPF

RIO

MFD

Indicates a failure in the JTIDS R/F output

detected by the JTIDS Interference Protection

Feature.

JTID

RIO

MFD

Indicates the joint tactical information distribution

system is not operating correctly.

SAHR HOT

RIO

MFD

Indicates an overtemp condition of the standard

attitude heading reference set.

Figure 2−93. Warning, Caution, Advisory Functions (Sheet 3 of 3)

2−181

ORIGINAL

NAVAIR 01−F14AAD−1

SYMBOL

FUNCTION

Water line

Indicates fuselage reference line (FRL). Displayed when attitude information

is not valid. Also displayed when gear down or the flight path marker is at,

or beyond the HUD’s full field of view.

Heading pointer

Actual aircraft heading is displayed below the stationary heading pointer.

Pitch/flight path ladder

Ladder displays aircraft climb/dive angle and roll angle. Aircraft vertical flight

path angle is indicated by the position of the flight path marker on the pitch/

flight path ladder. Positive pitch lines are solid and negative pitch lines are

dashed. To aid in determining flight path angle when it is changing rapidly,

the pitch lines are angled toward the horizon at an angle half that of the

flight path angle. For example, in a 40_ climb, the pitch lines are angled 20_

toward the horizon. "Up" appears at +90_ and "down" appears at −90_.

Ghost flightpath marker

Displayed at the true velocity vector position when the flight path marker is

caged and the true velocity vector position differs from the caged position in

azimuth. When the true flight path marker position is actually outside the

HUD total field of view, the symbol will be pegged at the edge of the total

FOV and flash.

Radar altitude indicator

Displays radar altitude when the aircraft is below 5000 feet AOL and bank

angle is less than 45_. If RDR is selected as the altimeter source and valid

radar altitude exists, the radar altitude is displayed within the dial, replacing

the barometric altitude. An R is displayed to the right of dial to indicate radar

altitude. If BARO is selected and a valid radar altitude exists, radar altitude

is displayed above the altitude dial or box.

Altitude analog dial

The HUD analog altimeter consists of ten dots encircling the altitude readĆ

out. Each dot indicates altitude in hundreds of feet with the zero mark

located at the top center of the dial.

Altitude pointer

An analog pointer indicating altitude moves uniformly around the inside of

the altitude dial based on indicated altitude. Increasing altitude is indicated

by clockwise rotation of this pointer.

Digital altitude readout

Digital barometric, radar, or GPS altitude is displayed depending on the

source of the data. When the ALT switch is in the BARO position, baroĆ

metric altitude is displayed. When the ALT switch in the RDR position and

if aircraft altitude is 5000 feet or lower, radar altitude is displayed within the

dial and is identified by an R to the right of the least significant digit. If the

radar altitude becomes invalid by exceeding 5000 feet or 45_ AOB, baroĆ

metric altitude is substituted and a B will flash to indicate that barometric

altitude is being displayed rather than radar altitude. A G is only displayed

when SCADC altitude becomes invalid and the GPS altitude is used. The

G acronym will flash if RADALT is selected and both SCADC and radar

altitude are invalid.

Vertical velocity readout

The vertical velocity readout consists of a maximum of five digits for a posiĆ

tive vertical velocity indication and a maximum of four digits with a leading

minus sign for a negative vertical velocity indication. If the limit 32,999 or

−9,999 is exceeded, a minus sign with four X’s (−XXXX) is displayed. It is

displayed below the six o’clock dot of the altitude dial.

Figure 2−95. HUD TLN Basic Format (Sheet 2 of 4)

ORIGINAL

2−184

NAVAIR 01−F14AAD−1

SYMBOL

FUNCTION

Negative three degree

Indicates the negative 3 degree position on the pitch ladder.

marks

Barometric pressure

The barometric pressure setting used by the display system and the weapon

setting

system is the value set on the pilot’s barometric altimeter.

The setting will be displayed for up to 5 seconds on the HUD and VDI

in the TLN mode when the setting is changed. At 18,000 feet, if it is off,

the symbol comes on and blinks for 5 seconds.

Navigation source and

WPT, GPS or TCN; waypoint number, range.

range readout

Clock/timer readout

Derived from the GPS clock. Default time is ZTOD. Options selectable from

the HSD ZTOD pushbutton are ZTOD, TTG, TTGT, TREL, ETA, ET and CD.

Bank scale

Provides indication of bank angle to ±45_. Tick marks are provided at 0_,

±10_, ±20_, ±30_ (slightly larger) and ±45_.

Bank angle pointer

Moving pointer provides indication of aircraft bank angle. At bank angles in

excess of ±45_, the pointer will be pegged at ±50_ and will flash.

Warning/caution/

The warnings L STALL, R STALL, L FIRE, R FIRE, and RDC SPEED and the

advisory readout

CLSN advisory will appear on the HUD in the steady condition. Up to two

indications may be displayed at any one time. When more than two indicaĆ

tions are present, they scroll up from the bottom at the rate of one

per second.

Mach number

Indicates speed of the aircraft in mach.

Peak aircraft g

Peak Aircraft g is displayed on the HUD as follows;

TLN Gear Down: If aircraft g falls below +0.0 or exceeds +2.0.

TLN Gear Up, A/A, A/G: If aircraft g falls below −2.0 or exceeds +4.5.

Peak g indication is displayed until a declutter mode is cycled.

Aircraft g

Aircraft g is displayed on the HUD as follows;

TLN Gear Down: If aircraft g falls below +0.5 or exceeds +1.5.

TLN Gear Up, A/A, A/G: If aircraft g falls below −2.0 or exceeds +4.5.

Angle of attack

Indicates angle of attack in units.

Note

When the TLN gear down format is displayed, the AOA readout is

removed when AOA is between 14 and 16 units. If AOA is greater than

14 units and decreasing, the readout remains off until AOA decreases

below 13 units. If AOA is less than 16 units and increasing, the readout

remains off until AOA increases above 17 units.

Airspeed pointer

An analog pointer indicating airspeed moves uniformly around the inside

of the airspeed dial based on indicated airspeed. Increasing airspeed is

indicated by clockwise rotation of the pointer.

Airspeed dial

The HUD analog airspeed dial consists of ten dots encircling the airspeed

readout. Each dot indicates airspeed in tens of knots with the zero mark

located at the top center of the dial.

Extended horizon line

Represents the horizon with respect to the aircraft and changes orientation

with any change in aircraft pitch or roll.

Figure 2−95. HUD TLN Basic Format (Sheet 3 of 4)

2−185

ORIGINAL

NAVAIR 01−F14AAD−1

SYMBOL

FUNCTION

Digital airspeed readout

Provides digital readout of calibrated airspeed.

Flight path marker

The flight path marker is displaced in azimuth and elevation to present

computed flight path. Aircraft vertical flight path angle is indicated by the

position of the flight path marker on the pitch/flight path ladder. In the caged

mode, the flight path marker is caged in azimuth and the true flight path

marker position is indicated by the display of the ghost flight path marker

when the true position is more than 2_ from the caged position. The flight path

marker can be caged or uncaged by alternately pressing the CAGE/SEAM

switch. On selection of TLN or A/A, the flight path marker is initially caged;

selection of A/G presents the uncaged mode initially.

Heading scale

Aircraft magnetic heading is indicated by the moving 360_ heading scale.

In TLN, the major divisions are numbered every 10 degrees. In A/A, the major

divisions are numbered every 20 degrees.

Note

When the aircraft is in the TLN mode with the gear handle down, the

heading scale remains 2 degrees above the position of the flight path

marker. The lowest point of the heading scale, including the numbers,

will never rise above the normal (gear up) position. The heading scale is

occluded by the altitude and airspeed dials and readout.

Angle of attack

The AOA bracket is a pitch related variable that indicates the deviation of the

bracket

current AOA from a desired value and is vertically referenced to the left wing

of the flight path marker symbol. The center of the bracket represents the optiĆ

mum AOA. The bracket moves lower with respect to the flight path marker as

AOA increases and it moves higher as AOA decreases.

Potential flight path

Indicates the acceleration along the flight path marker. Provides a graphical

marker (PFPM)

representation of the ability to change the flight path angle by varying the

thrust acceleration and/or angle of attack. Deceleration is indicated by the

PFPM below the flight path marker and acceleration by the PFPM above the

flight path marker.

Figure 2−95. HUD TLN Basic Format (Sheet 4 of 4)

radar altitude exists (altitude < 5,000 feet and AOB < 45_),

2.33.8.3.1

Takeoff/Landing/Navigation Formats

radar altitude is displayed in the center of the altitude dial. If

TLN formats are categorized by the selected steering

the aircraft’s altitude exceeds 5,000 feet or the radar altitude

mode and landing gear position. TLN basic, the HUD default

becomes invalid, the system automatically substitutes baroĆ

format, does not display steering information. Refer to

metric altitude and a B" will flash to the right of the analog

Figure 2−95 for the location and description of TLN basic

dial to indicate radar altitude is not being used. Switching

symbology.

HUD/VDI ALT to BARO removes the flashing B."

Steering mode selection is made through MFD pushĆ

The symbols removed by the DECLUTTER switch

button or cursor designate action on the VDI AWL formats.

vary with formats and are discussed in the applicable paraĆ

TCN, DEST and GPS steering mode selections are also

graphs. Refer to Figure 2Ć96 for declutter information in

available on the HSD by boxing the TACAN data buffer or

TLN−GD, TLN−GU, A/A, and A/G modes.

waypoint data buffer. Making a steering mode selection

changes TLN basic to TLN TCN (TACAN), MAN (manual),

DEST (destination), GPS, D/L (data link), or AWL (all−

weather landing).

ORIGINAL

2−186

NAVAIR 01−F14AAD−1

Steering mode is identified on the HUD by a legend in the

2.33.8.3.2

Air−to−Air Formats

data readout display area. Steering modes are described in

A/A formats (Figure 2Ć99 and Figure 2Ć100) are preĆ

Chapter 20.

sented when the pilot selects the A/A pushbutton on the pilot

displays control panel, when a weapon is selected, the radar

TCN selection adds a course−steering arrow and

hot modes are selected, or when the ACM guard is lifted. The

course−deviation dots. Distance to the TACAN staĆ

A/A formats provide target acquisition, weapon status, and

tion is displayed to the right of the TCN legend.

shoot prompts as well as primary flight information. Target

MAN steering selection adds a commanded heading

data and the selection legends A/A, PH, SP, SW, and G are

marker to the heading scale. The commanded headĆ

displayed. Quantity of the selected weapons is also shown.

ing marker also appears on destination, datalink,

When GUN is selected, the quantity number indicates rounds

and AWL formats.

remaining in hundreds. A large X" through a weapon

selection legend indicates that the master arm switch is

DEST steering selection adds the waypoint destinaĆ

SAFE.

tion range to TLN basic format.

Refer to the Supplemental NATOPS Flight Manual,

NAVAIR 01−F14AAD−1A, for a description of air−to−air

D/L selection displays the range to the data−link

attack.

destination. A large flashing X" will appear in the

center of the display when a data−link waveoff

2.33.8.3.3

Sensor Mode Indications

command is received.

Radar modes are indicated on the HUD via alphaĆ

AWL steering selection provides for the display of

numerics. The radar mode alphanumerics are removed when

ACL and ILS, ACL only, ILS only, or no ACL and

the radar is off or in the computer mode. An X" overlays the

ILS glidepath situation displays. The display of the

mode indication if the IRST is failed (Figure 2Ć99). The radar

HUD flight director glideslope and centerline steerĆ

mode alphanumerics are as follows:

ing can also be independently controlled. Selections

are made via MFD pushbutton activation on the VDI

Hot range while search (HRWS)

AWL format. A large, flashing X" will appear in

the center of the display when a waveoff command

Manual rapid lock−on (MRL)

is received. Distance to the TACAN station is disĆ

played as is the TCN legend.

Pilot automatic lock−on (PAL)

When the landing gear handle is placed in the down

Pilot lock−on mode (PLM)

position, the HUD cage/uncage function is enabled on the

CAGE/SEAM switch located on the inboard throttle, the

Pulse Doppler search (PDS)

system transitions to TLN−GD mode, and all weapon

selections are cleared. In TLN−GD mode, the Mach number

Pulse Doppler single−target track (PDSTT)

is removed and aircraft g is displayed if the g’s fall below

+0.5 or exceed +1.5; peak g is displayed in normal declutter

Pulse search (PS)

mode if aircraft g falls below 0.0 or exceeds +2.0; the horizon

line is extended across the HUD field of view and a flying

Pulse single−target track (PSTT)

W" (waterline) symbol is added at the fuselage reference

Range while search (RWS)

line.

10. Range while search velocity (RWSV)

Note

11. Sniff (SNIFF)

The waterline symbol is also added in other HUD

formats when the flight path marker is at or

12. Standby (STBY)

beyond the HUD field of view or when altitude

data is lost.

13. Track while scan automatic (TWSA)

Figure 2Ć97 shows the symbols that are added during

14. Track while scan manual (TWSM)

TACAN and AWL flight director steering modes, landing

gear down, with digital or analog display selection. Refer to

15. Vertical scan lock−on high (VSLHI)

Figure 2Ć98 for a description of the symbols that are available

for TLN formats.

16. Vertical scan lock−on low (VSLLO).

ORIGINAL

2−188

NAVAIR 01−F14AAD−1

SYMBOL

FUNCTION

Command heading

This symbol indicates the heading required to achieve the selected course.

marker

Course selection may be manual, data link commanded, or waypoint destination.

Where commanded heading is beyond display scale limits, the symbol will be

pegged at the nearest edge to the commanded heading. This symbol does not

appear on the basic or TACAN formats.

Waterline

Indicates fuselage reference line (FRL). Displayed when attitude information is not

valid. Also displayed when gear down or TLN Gear Up, or the flight path marker is at

or beyond the HUD’s full field of view.

Breakaway, waveoff

A large flashing X will appear in both D/L and AWL steering modes if a

WAVEOFF command has been received.

ILS precision course

Consists of two independent needles (vertical and horizontal) which form a cross

needles

pointer. The horizontal needle responds to ILS glide slope error and the vertical

needle responds to ILS localizer error. Null/center indications are provided to enable

the pilot to null the error and keep the vertical and horizontal needles centered.

AWL legend

This message indicates that the all weather landing steering mode has been

selected.

Range

Depending on the format, this message will indicate either the range to the TACAN

station, data link destination or distance to waypoint destination. The legends TCN,

D/L, or WPT may also appear. When in the manual steering mode no range appears

but the MAN legend is displayed.

Extended horizon line

Indicates the horizon with respect to the aircraft with landing gear down. Changes

orientation with any change in aircraft pitch and roll.

Course arrow and

The course arrow represents the selected course to the TACAN station. Two dots

deviation dots

will appear on the side of the flight path marker toward the course arrow and perpenĆ

dicular to the arrow. The dot closest to the flight path marker represents a half scale

deflection of 4_ off course, while the outermost dot represents full scale deflection

of 8_ off course. When the aircraft crosses the selected course, the arrow moves to

the opposite side of the flight path marker and the dots would appear on that side.

For deviations of more than 9_, the arrow pegs. If the arrow is centered on course,

the dots disappear. Flight path marker centered over the course arrow indicates

being on course. For TACAN bearings aft of ±90_, the arrow will be dashed.

Angle−of−attack

The AOA bracket is a pitch−related variable that indicates the deviation of the current

bracket

AOA from a desired value and is vertically referenced to the left wing of the flight

path marker symbol. The center of the bracket represents the optimum AOA.

The bracket moves lower with respect to the flight path marker as AOA increases

and it moves higher as AOA decreases.

Flight director

The flight director symbol provides glide slope and centerline steering information

computed by the mission computer using navigation system parameters and

Data Link information from the SPN−42/46 ACLS system. The box with the three

dots will provide the pilot with optimal glide path intercept and following when the

flight path marker is inside the flight director box and the three dots are aligned with

the wings and the tail of the flight path marker. The same procedures are used

whether the flight path marker is caged or uncaged. The flight director symbol is

removed from the HUD when the FLT DIR pushbutton on the VDI is unboxed.

ACL steering indicator

Provides ACL steering commands driven by the ASW−27C data link.

Figure 2−97. HUD Added Symbology (Sheet 2 of 2)

ORIGINAL

2−190

NAVAIR 01−F14AAD−1

FORMAT

SYMBOL

DATA

BASIC

AWL

LINK

DESTINATION

MANUAL

TACAN

Aircraft G Readout

(On all formats except GEAR DOWN & DECLUTTERć2)

Airspeed Dial

(On all formats except DIGITAL)

Airspeed Readout Box

(On all formats except ANALOG or DECLUTTERć2)

Airspeed Readout

(On all formats)

Altitude Dial

(On all formats except DIGITAL)

Altitude Readout Box

(On all formats except ANALOG or DECLUTTERć2)

Altitude Readout

(On all formats)

AnglećofćAttack Readout

(On all formats except DECLUTTERć2)

Bank Scale

(On all formats except DECLUTTERć1 and 2)

Baro Setting Readout (5 sec)

Ghost flight path marker

Extended Horizon Line

(On all formats when GEAR DOWN)

Heading Scale

Horizon

(On all formats when GEAR UP)

Reference Markers

Mach Readout

(On all formats except GEAR DOWN and GEAR UP DECLUTTERć2)

Peak A/C G Readout

(On all formats except DECLUTTERć1 and 2)

Pitch Ladder ć TLN

Radar Altitude Readout

(On all formats except GEAR DOWN DECLUTTERć1 & 2 and GEAR UP

DECLUTTERć2)

Flight Path Marker

Vertical Velocity Readout

(On all formats except DECLUTTERć1 & 2 in GEAR UP and DECLUTTERć2

in GEAR DOWN)

Clock/Timer Readout

(On all formats except DECLUTTERć2)

Figure 2Ć98.ĄHUD Symbology Available on TLN Formats (Sheet 1 of 2)

2−191

ORIGINAL

NAVAIR 01−F14AAD−1

FORMAT

SYMBOL

DATA

BASIC

AWL

LINK

DESTINATION

MANUAL

TACAN

Waterline

(On all formats or when flight path marker pegged or altitude data invalid)

Altitude Source ć B, R, or G

HUD Cursor

Potential Flight Path Marker

(On all formats except DECLUTTERć2)

Angle of Attack Bracket

(On GEAR DOWN only. All formats)

IRST Pointer

TCS Pointer

Caution/Advisory/Warning

Breakaway Symbol

Command Heading Marker

HUD Steering Legend ć AWL

HUD Steering Legend ć TCN

HUD Steering Legend ć D/L

HUD Steering Legend ć MAN

HUD Steering Legend ć WPT

ILS Precision Course Needles

Range Readout

ACL Steering Indicator Tadpole

Flight Director

Course Arrow & Deviation Dots

Notes:

+ indicates that the symbol is available for display on the selected format.

o indicates that the symbol is not available for display on the selected format.

Figure 2−98. HUD Symbology Available on TLN Formats (Sheet 2 of 2)

ORIGINAL

2−192

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

Radar target

Indicates the radar line of sight (LOS) to the target. Symbol is displayed on all

designator and target

A/A weapon modes when radar is tracking a target. The symbol is positionable

aspect

over the total HUD field of view (FOV). When the target moves beyond the

FOV limits of the HUD, the symbol will peg and flash. In STT, target aspect is

represented by a pointer which points in the direction of the aspect angle.

Zero target aspect is straight down.

Target pointer

Indicates the direction of the target designator (TD) box position on the HUD.

The target pointer is present when pointing to a TD box under the following

conditions: 1) FONO 1 track is outside the IFOV (Instantaneous Field of View).

2) PTID hooked track is outside the IFOV with no FONO 1 present. 3) Closest

TMA is outside the IFOV and there is no FONO 1 PTID hooked track.

Angle−off−the−nose

The angle−off−the−nose (AON) indicator defines the angle between the FRL

indicator

and the target line of sight that the target pointer is pointing to, in the plane

described by the FRL and the target pointer. When the target designator is

being pointed to by the target pointer, a three digit readout is displayed

indicating the AON of that target. The AON indicator is not earth stabilized.

The AON readout is centered below the origin of the target pointer and is

given in degrees.

Steering tee

Provides azimuth steering only, in search mode. Provides elevation and

azimuth steering in track mode.

IRST symbol

Indicates IRST target location. Up to four displayed.

Target range

Range of closest radar target in nautical miles and tenths. Numeric is disĆ

played only when range is valid.

Target altitude

Altitude of closest radar target in thousands of feet. Displayed only when

information is valid.

Target closing velocity

Displays closing rate to radar target. A minus sign indicates an opening

velocity.

Target ID

Target ID display.

Navigation data

Displays steering mode, waypoint selected or range (DEST or D/L).

Clock/timer readout

TTG is selected from the HSD ZTOD pushbutton, and indicates time−to−go to

read designated waypoint.

IRST mode indicator

Displays current IRST mode.

Radar mode indicator

Displays current radar mode.

Weapon select legend

Displays missile type and quantity, if selected, or gun and rounds remaining,

in hundreds. If no weapon is selected, displays A/A.

Master arm safe cue

A large X through the A/A or weapon select legend indicates the Master Arm

Switch is in SAFE.

TCS pointer

Indicates TCS track location.

NCTR

Indicates non−cooperative target recognition is available.

Refer to NAVAIR 01−F14AAD−1A.

Figure 2−99. HUD A/A Search Formats (Sheet 2 of 2)

ORIGINAL

2−194

NAVAIR 01−F14AAD−1

FORMAT

PHOENIX

SPARROW

SIDEWINDER

GUN

BASIC

TRACK

SYMBOL

MMGS

BACKUP

Aircraft G Readout

Airspeed Dial

(On all formats except DIGITAL)

Airspeed Readout Box

(On all formats except ANALOG & DECLUTTERć2)

Airspeed Readout

(On all formats)

Altitude Dial

(On all formats except DIGITAL)

Altitude Readout Box

(On all formats except ANALOG & DECLUTTERć2)

Altitude Readout

(On all formats)

Angle of Attack Readout

(On all formats except DECLUTTERć2)

Baro Setting Readout

Clock/Timer Readout

(On all formats except DECLUTTERć2)

Ghost Flight Path Marker

(On all formats except DECLUTTERć2)

Heading Scale

(On all formats except DECLUTTER 1 & 2)

Horizon

Reference Markers

Mach Readout

(On all formats except DECLUTTERć2)

Peak A/C G Readout

(On all formats except DECLUTTERć2)

Pitch Ladder

(On all formats except DECLUTTERć1 & 2)

Radar Altitude Readout

(On all formats except DECLUTTERć1 & 2)

Flight Path Marker

Waterline

(On all formats when flight path marker pegged, gear is down, or altitude information is

invalid.)

Altitude Source ć B or R

HUD Cursor

IRST Pointer

TCS Pointer

Caution/Advisory/Warning

Breakaway Symbol

Command Heading Marker

Select Legends,

A/A

PH#

SP#

SW#

Weapon ć Qty

Figure 2Ć100.ĄHUD Symbology Available on A/A Formats (Sheet 1 of 2)

2−195

ORIGINAL

NAVAIR 01−F14AAD−1

FORMAT

PHOENIX

SPARROW

SIDEWINDER

GUN

TRACK

SYMBOL

BASIC

MMGS

BACKUP

Master Arm Switch Safe

Cue

Target Range ć RNG, #

Target Range Indicator

Waypoint Select

Steering Tee

Target Designator

Target Closing Velocity

Target Altitude

Target ID

TACAN Digital Readout

Flood Illumination Pattern

Sidewinder Seeker Circle

SHOOT Cue

Reticle

Reticle A

Reticle B

Target Range Tape

Target Lead Cue

BATR Symbol

Gun Mode

Indication ć MAN

Reticle Depression ć #

A/A Gun/Backup Mode

Reticle

Notes:

+ indicates that the symbol is available for display on the selected format.

o indicates that the symbol is not available for display on the selected format.

Figure 2−100. HUD Symbology Available on A/A Formats (Sheet 2 of 2)

ORIGINAL

2−196

NAVAIR 01−F14AAD−1

IRST modes are indicated on the HUD via alphaĆ

format (Figure 2Ć104) allowing gun aiming by displaying the

numerics. The IRST mode alphanumerics are removed when

A/A gun/backup mode reticle. The reticle depression angle,

the IRST is failed. The IRST alphanumerics are as follows:

adjusted by the ELEV LEAD knob on the PDCP is shown in

the lower right corner of the HUD along with the MAN

1. Cooldown (COOL)

gun−mode indication.

2. Hot IR (HOTIR)

2.33.9

MFD Formats

3. Pilot automatic lock−on (PAL)

Initial turn on or a cold start (defined as a system reset

or a MCS power outage of at least 300 milliseconds) causes

4. Pilot lock−on mode (PLM)

the following default formats to be displayed: With gear

down and weight on wheels, VDI/TLN basic on MFD1, OBC

5. Single−target track (STT)

basic on MFD2, and OWN A/C on MFD3; with gear down

and weight off wheels, VDI/TLN basic on MFD1, HSD on

6. Standby (STBY)

MFD2, and HSD on MFD3; and with gear up and weight off

wheels, TSD on MFD1, HSD on MFD2, and HSD on MFD3.

7. Track while scan automatic (TWSA)

The actual format displayed on MFD3 depends on the

navigation mode selected and the conditions existing at the

8. Track while scan manual (TWSM).

time. If the NAV MODE switch is at OFF, the OWN A/C

basic format is displayed.

2.33.8.3.4

Air−to−Ground Formats

MFD3 may also act as a controller of the DEU in that,

Pushbutton selection on the PDCP or selection of an

when certain formats are being displayed on MFD3, the DEU

air−to−ground weapon places the A/G basic format on the

is commanded to display corresponding slaved formats.

HUD (Figure

2Ć101 and Figure 2Ć102). The A/G basic

Refer to Figure 2Ć105 for a listing of MFD3/DEU slaved

format can display waypoint and TACAN information. A/G

control conditions.

DECLUTTER and ANLG and DGTL displays are similar to

A/A formats. Refer to the Supplemental NATOPS Flight

With the exception of high−priority formats (ECM and

Manual, NAVAIR 01−F14AAD−1A.

spin), which appear when required, most MFD formats are

selectable by means of MFD pushbutton or cursor designaĆ

2.33.8.4

Overlay Symbology

tion. The actual format that will appear may depend on other

factors, however, such as master mode selection, aircraft

Symbology (Figure 2Ć103) may be overlaid on disĆ

state (TLN, A/A, or A/G), steering mode selection, and the

played HUD formats when additional information is

alignment condition. The MENU legend appears on every

required. These include RECON, TWS, and IRST TWS.

MFD format except for HUD, DD, and PTID repeats. These

RECON, used with the TARPS pod, is selected as an

repeat formats do not display selections; pressing any

overlay from the MFD RECON formats. This overlay adds

pushbutton when in a repeat mode will place MENU1 on the

the RECON command heading marker, command ground−

MFD. The MENU legend is located above the center

track line, RECON steering symbol, target−designator hexaĆ

pushbutton on the lower edge of the MFDs. Also appearing

gon, and camera selection legend.

on every MFD format are the SMS and ECM pushbutton

legends (Figure 2Ć106).

Radar track while scan adds up to four radar target

diamonds that indicate the four closest targets. Size of the

Repeated depressions of the ECM or SMS pushbuttons

symbols indicates relative proximity (i.e., the largest is the

toggle between these formats and the previously selected

closest). The four symbols are of preset sizes, not scaled to

display. This permits the crew to quickly check ECM or SMS

reflect actual distances.

conditions without having to reselect previous formats.

The infrared search and tracking system TWS adds up

Selecting MENU places menu 1 on the display. The

to four triangular IRST symbols to existing formats. Unlike

legend reads MENU1 and is enclosed by a rectangular box.

TWS, these symbols are all the same size. Both IRST and

Selecting MENU1 when it is boxed places menu 2 on the

TWS symbols are added automatically when a target is being

display with the legend MENU2 displayed in the box. The

tracked.

MENU pushbutton toggles between MENU1 and MENU2.

Menu selection changes the pushbutton legend but does not

2.33.8.5

Manual Reticle

alter the display being presented.

If the mission computer loses communication with

both DPs, the DP driving the HUD provides a manual reticle

2−197

ORIGINAL

NAVAIR 01−F14AAD−1

FORMAT

SYMBOL

BASIC

CCIP

MANUAL

Aircraft G Readout

Airspeed Dial

(On all formats except DIGITAL)

Airspeed Readout Box

(On all formats except ANALOG & DECLUTTER − 2)

Airspeed Readout

(On all formats)

Altitude Dial

(On all formats except DIGITAL)

Altitude Readout Box

(On all formats except ANALOG & DECLUTTER − 2)

Altitude Readout

(On all formats)

Angle of Attack Readout

(On all formats except DECLUTTER − 1 & 2)

Baro Setting Readout

Clock/Timer Readout

(On all formats except DECLUTTER − 2)

Ghost flight path marker

(On all formats except DECLUTTER − 2)

Heading Scale

(On all formats except DECLUTTER − 2)

Horizon

Reference Markers

Mach Readouts

(On all formats except DECLUTTER − 2)

Peak A/C G Readout

(On all formats except DECLUTTER − 2)

Pitch Ladder

(On all formats)

Radar Altitude Indicator

(On all formats except DECLUTTER − 2)

Flight Path Marker

Altitude Source−B or R

HUD Cursor

Figure 2Ć102.ĄHUD Symbology Available on A/G Formats (Sheet 1 of 2)

2−199

ORIGINAL

F-14D. FLIGHT MANUAL (2004) - page 4