F18. FLIGHT MANUAL (2008)

A1-F18EA-NFM-000

2.14 OXYGEN SYSTEMS

2.14.1 On Board Oxygen Generating System (OBOGS). OBOGS provides oxygen rich breathing gas

to the aircrew while either engine is operating. Engine bleed air is cooled and routed through the

OBOGS inlet air shutoff valve to the OBOGS concentrator. The breathing gas is routed from the

concentrator to a cockpit plenum, where the temperature is stabilized and a limited supply is stored for

peak flow demands. From the plenum, the breathing gas flows through the pilot services panel oxygen

disconnect, through the seat survival kit, to the aircrew regulators and masks.

A leak or break in the breathing gas system anywhere from the pilot

services panel oxygen disconnect to the mask will prevent either normal

OBOGS breathing gas or emergency oxygen from being delivered to the

mask in its intended concentration. Leaks and breaks may be difficult to

locate and/or verify. If hypoxic symptoms are experienced or an OBOGS

system degrade occurs, immediate descent below

10,000

feet cabin

altitude is required to prevent severe or incapacitating hypoxia.

The OBOGS concentrator is powered by the left 115 volt ac bus. Two molecular sieve beds in the

OBOGS concentrator remove most of the nitrogen from the engine bleed air. The nitrogen is dumped

overboard while the remaining output of oxygen rich breathing gas is supplied to the aircrew.

2.14.1.1 OBOGS Monitor. The CRU-99/A solid state oxygen monitor is located on the left side of the

seat bulkhead in the front cockpit and is powered by the left 28 volt dc bus. The monitor continuously

measures oxygen concentration in the OBOGS breathing gas and provides a discrete signal to activate

the OBOGS DEGD caution if the oxygen concentration falls below a predetermined level.

Loss of electrical power to the OBOGS monitor prevents reporting of

OBOGS DEGD conditions.

The monitor performs a power-up BIT during a 2 minute warm-up period and conducts a periodic

BIT every 60 seconds. No indication is provided if power-up or periodic BIT pass.

Preflight BIT of the monitor is accomplished by using either the pneumatic BIT plunger or the

electronic BIT pushbutton. Refer to figure 2-33. Pressing up and holding the pneumatic BIT plunger

for 15 to 65 seconds tests the operation of the OBOGS monitor by diverting cabin air into the monitor

to create a low oxygen concentration condition.

After initiating OBOGS pneumatic BIT, failure to ensure plunger is fully

extended may result in contamination of OBOGS plenum which may lead

to hypoxic conditions.

Momentarily pressing and releasing the electronic BIT pushbutton tests the monitor electronically.

Successful completion of either test activates the OBOGS DEGD caution, which clears automatically

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Figure 2-33. OBOGS Monitor

after BIT is complete and OBOGS resets to normal operation.

Successful OBOGS monitor BIT is required prior to flight. A BIT failure

indicates that there is no protection against inadequate oxygen concen-

tration or hypoxia due to a degraded OBOGS monitor. Good breathing

gas flow alone does not ensure adequate oxygen concentration.

2.14.1.1.1 OBOGS DEGD Caution. An OBOGS DEGD caution is set by the OBOGS monitor when

oxygen concentration is below a predetermined level. The OBOGS DEGD caution threshold is always

above cabin air conditions, in order to provide a physiological safety margin.

NOTE

Oxygen concentration may drop below the predetermined OBOGS

DEGD level if breathing gas flow is unlimited. Removing the mask

without placing the OXY flow knob to OFF, system leaks, and/or loose

aircrew hose connections can overwhelm system capacity and may

result in an OBOGS DEGD caution.

After a total loss of bleed air, OBOGS breathing gas flow will be available until the residual gas

within the system is depleted. The residual oxygen concentration may be sufficient to keep the OBOGS

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DEGD caution from illuminating, but loss of system pressure will ultimately lead to inadequate

pressure and an abrupt inability to breathe.

Low mask flow or increased breathing resistance may occur without an accompanying OBOGS

DEGD caution, which indicates a potential system degradation that may result in oxygen levels below

physiological requirements.

When cabin altitude is above 10,000

feet, OBOGS DEGD caution

procedures shall be executed for low mask flow, increased breathing

resistance, or OBOGS DEGD cautions of any duration.

A single brief appearance of the OBOGS DEGD caution immediately following placing the OBOGS

control switch to ON, turning the OXY FLOW knob to ON, or after donning or removing the mask is

normal. Certain malfunctions affecting the OBOGS system also may momentarily cause OBOGS

DEGD cautions. A momentary OBOGS DEGD caution may clear from the DDI so rapidly that the

aircrew may be unable to determine the cause of the MASTER CAUTION light/tone.

Repeated unexplained MASTER CAUTION lights/tones may be an

indication of OBOGS system degradation.

Even with the OBOGS DEGD caution displayed, OBOGS breathing gas under normal flow is of

higher quality (i.e., oxygen concentration, partial pressure and purity) than cabin air. Once cabin

altitude is below 10,000 feet, aircrew may elect to conserve emergency oxygen by resetting the

emergency oxygen release tab, then either removing the mask and breathing cabin air, or returning the

OXY FLOW knob and the OBOGS control switch to ON and breathing through the mask.

Pure oxygen accelerates recovery from hypoxia. Emergency oxygen shall

be used whenever hypoxic symptoms are recognized.

2.14.1.2 Breathing Regulator. The aircrew torso mounted breathing regulator reduces both normal

and emergency oxygen system operating pressures to breathing pressure levels. The regulator delivers

undiluted OBOGS breathing gas or emergency oxygen to the aircrew at positive pressure, the limits of

which increase automatically with altitude. It interfaces with the hose assembly, which connects with

the seat survival kit oxygen disconnect.

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2.14.1.3 OBOGS Control Switch. The OBOGS control switch, located on the left console in the front

cockpit, is used to control electrical power to the OBOGS concentrator and the OBOGS inlet air shutoff

valve.

ON Supplies electrical power and engine bleed air to the OBOGS concentrator.

OFF OBOGS system off.

2.14.1.4 OXY FLOW Knob. The OXY FLOW knob, located on the left console in both cockpits, is

used to control the supply of OBOGS breathing gas to each aircrew’s mask.

ON OBOGS flow supplied to the mask.

OFF OBOGS flow secured.

It is possible to place the OXY FLOW knob in an intermediate posi-

tion between the ON and OFF detents, which may result in a reduced

flow of breathing gas. The OXY FLOW knob should always be fully

rotated to the ON or OFF detent position.

2.14.2 Emergency Oxygen. Emergency gaseous oxygen is contained in a bottle in the seat survival

kit. The bottle is connected into the OBOGS supply hose as it passes through the kit. From this point,

the emergency oxygen and OBOGS breathing gas share a common path to the aircrew mask.

A leak or break in the breathing gas system anywhere from the pilot

services panel oxygen disconnect to the mask will prevent either normal

OBOGS breathing gas or emergency oxygen from being delivered to the

mask in its intended concentration. Leaks and breaks may be difficult to

locate and/or verify. If hypoxic symptoms are experienced or an OBOGS

system degrade occurs, immediate descent below

10,000

feet cabin

altitude is required to prevent severe or incapacitating hypoxia.

A pressure gauge is visible on the inside left front of the survival kit. The bottle provides

approximately 10−20 minutes of oxygen. Oxygen duration decreases with lower altitude.

Under less than optimum conditions (low altitude, heavy breathing,

loose−fitting mask, etc.), as few as 3 minutes of emergency oxygen may be

available.

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The emergency oxygen supply is activated automatically upon ejection. The emergency oxygen

supply may be activated manually by pulling the emergency oxygen green ring on the outside of the left

thigh. The emergency oxygen supply may be deactivated at aircrew discretion by pushing down on the

release tab immediately forward of the green ring.

With emergency oxygen selected, the OXY FLOW knob(s) shall be

placed to OFF. If not secured, OBOGS system pressure may prevent

emergency oxygen from reaching the breathing regulator. Additionally,

the OBOGS control switch should be placed to OFF to backup the OXY

FLOW knob.

2.15 FIRE DETECTION, FIRE EXTINGUISHING, AND BLEED AIR LEAK DETECTION SYSTEMS

The fire detection system contains dual-loop fire detectors and three FIRE warning lights. The fire

extinguishing system contains a READY/DISCH light and a fire extinguisher bottle. The two systems

provide engine bay, AMAD bay, and APU bay fire warning, engine and APU emergency shutdown, and

selective fire extinguishing capability. The fire extinguisher bottle is located in the aft fuselage

between the engines. The bottle contains a nontoxic gaseous agent which provides a one-shot

extinguishing capability.

Electrical power from the 28 vdc essential bus is required to operate the fire detection and

extinguishing systems. The systems can operate on battery power alone with the BATT switch ON.

A separate dry bay fire suppression (DBFS) system is incorporated to automatically detect and

extinguish a fire or explosion in the dry bays below fuel tanks 2, 3, and 4.

2.15.1 FIRE Lights. Two FIRE warning lights, one for each engine/AMAD bay, are located on the

upper left and right sides of the main instrument panel. The warning lights come on when a fire

condition is detected in the respective engine/AMAD bay. The left FIRE light indicates a fire condition

in the left engine/AMAD bay. The right FIRE light indicates a fire condition in the right engine/

AMAD bay.

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Each light is a pushbutton, which is guarded to prevent inadvertent actuation. Pushing the left or

right FIRE light arms the fire extinguisher bottle (FIRE EXTGH READY light on) and closes the

corresponding engine feed shutoff valve, the crossfeed valve, and the crosscooling valve.

Because the engine VEN/start pumps are fuel lubricated, pushing a FIRE

light prior to throttle OFF may damage the corresponding pump. To

reduce the likelihood of damage, FIRE lights should only be pressed as

directed by NATOPS (following throttle OFF for actual emergencies or

as specifically delineated for an FCF A profile).

If a FIRE light is pressed, the pushbutton stays in and approximately 1/8 inch of yellow and black

stripes is visible around the outer edges of the light.

2.15.2 APU FIRE Light. The APU FIRE warning light is located on the main instrument panel

inboard of the right FIRE light. The APU FIRE light comes on when a fire condition is detected in the

APU bay. The APU FIRE light is also a pushbutton. Pushing the APU FIRE light arms the fire

extinguisher bottle (FIRE EXTGH READY light on) and secures fuel to the APU. If the APU FIRE

light is depressed, the pushbutton stays in and approximately 1/8 inch of yellow and black stripes is

visible around the outer edges of the light.

2.15.3 FIRE Warning Voice Alerts. When the left, right, and/or APU FIRE lights are illuminated, the

ENGINE FIRE LEFT, ENGINE FIRE RIGHT, or APU FIRE voice alerts, respectively, are also

activated. If more than one FIRE light comes on at the same time, the voice alert priority is LEFT,

RIGHT, then APU.

2.15.3.1 FIRE Warning Lights (F/A-18F). The rear cockpit left, right, and APU FIRE warning lights

are advisory only. These lights are not pushbuttons, and they do not arm the fire extinguisher bottle

or shut down the engines or APU.

2.15.4 FIRE EXTGH READY/DISCH Light. The FIRE EXTGH READY/DISCH light is located on

the MASTER ARM panel on the left side of the main instrument panel. The top half of the light is

yellow and is labeled READY. The bottom half of the light is green and is labeled DISCH. When the

fire extinguisher bottle is armed (left, right, or APU FIRE light pressed), the yellow READY light is

illuminated.

The FIRE EXTGH READY/DISCH light is also a pushbutton. When the READY light is on,

pushing the light discharges the fire extinguisher bottle into the selected engine/AMAD/APU bay(s).

The FIRE EXTGH READY/DISCH light does not latch like the FIRE lights, which means a signal is

sent to discharge the fire extinguisher bottle ONLY when the light is held in the fully pressed position.

When the fire extinguisher bottle has discharged or pressure has been lost, the green DISCH light

comes on. Therefore, it is good practice with an engine/AMAD bay fire to hold the READY/DISCH

light pressed until the green DISCH light comes on. The fire extinguisher bottle should discharge

within 5 seconds.

For an inflight APU fire, discharge of the fire extinguisher bottle is delayed approximately 10

seconds from when the APU FIRE light is pushed. It is not necessary to hold the READY/DISCH light

for those 10 seconds. However, if the DISCH light does not come on 10 seconds after the APU FIRE

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and READY/DISCH lights have been pushed, the READY/DISCH light should be pushed and held

until the DISCH light does come on.

If more than one FIRE light is pressed, the fire extinguisher bottle may not discharge and, if it does,

the concentration of the extinguishing agent sent to the selected bays may be insufficient to extinguish

both fires.

2.15.5 APU Fire Extinguishing System. The APU fire extinguishing system is automatically acti-

vated with WonW and must be manually activated with WoffW. If an APU fire condition is detected

with WonW, the automatic function secures fuel to the APU, arms the fire extinguisher bottle, and

after 10 seconds discharges the fire bottle. Discharge of the bottle is delayed for 10 seconds to allow the

APU time to spool down before extinguishing agent is introduced.

Manual activation is accomplished by pushing the APU FIRE light and then the FIRE EXTGH

READY light. Like automatic activation, discharge of the bottle is delayed for 10 seconds after the

APU FIRE light is pushed. If an APU FIRE condition is detected with WonW, manual activation

should be performed to backup the automatic system.

Since the fire extinguishing system requires 28 vdc essential bus power,

the fire extinguisher bottle may not be discharged if the BATT switch is

turned off during the 10 second delay time.

2.15.6 Engine/AMAD Fire Extinguishing System. The engine/AMAD fire extinguishing system

must be manually activated. Manual activation is accomplished by lifting the guard and pressing the

corresponding FIRE light and then the FIRE EXTGH READY light. Pushing the FIRE light secures

fuel to the engine at the engine feed shutoff valve and isolates the left and right fuel systems by closing

the crossfeed and crosscooling valves. When the FIRE EXTGH READY light is pushed, the fire

extinguisher bottle is discharged without delay into the corresponding engine/AMAD bay.

Fire testing indicates that the probability of extinguishing a fire and

preventing relights is greatly increased by immediately discharging the

fire extinguisher.

2.15.7 FIRE Detection System Test. Each of the three FIRE warning lights contains four individual

light bulbs. Light bulb integrity can be tested during a LT TEST with ac power applied. If a

malfunction exists in a fire detection loop associated with the APU FIRE light, the APU FIRE voice

alert does not annunciate and none of the four individual bulbs in the light illuminate. If a malfunction

exists in a fire detection loop associated with either FIRE light, the corresponding ENGINE FIRE

LEFT/RIGHT voice alert does not annunciate and only the individual bulb (or bulbs) associated with

the malfunctioning sensor does not come on. Care must be taken to detect bulbs that are not on in the

FIRE lights during the loop test.

A successful test of the FIRE detection system should illuminate all four bulbs in each of the three

FIRE lights and should annunciate the ENGINE FIRE LEFT, ENGINE FIRE RIGHT, and APU

FIRE voice alerts.

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2.15.8 Bleed Air Leak Detection (BALD) System. The BALD system is designed to protect the

aircraft from damage resulting from a bleed air leak. The system contains a BALD controller and 11

detector sensing elements routed along the bleed air distribution lines (ducts, valves, and heat

exchangers). If a bleed air leak is detected, the system attempts to isolate the affected ducting by

automatically closing the appropriate bleed air shutoff valve(s). A leak detected upstream of the

secondary bleed air shutoff valve closes only the appropriate primary bleed air shutoff valve. A leak

detected downstream of the secondary bleed air shutoff valve closes all three valves (secondary and

both primaries).

When a leak is detected, the BALD system sends commands directly to the appropriate bleed air

shutoff valves, to the BLEED warning lights, and to the ACI, triggering the BLEED AIR LEFT

(RIGHT) voice alerts. When the bleed air shutoff valve(s) are commanded closed, the appropriate L

and/or R BLD OFF caution is displayed. The L BLEED and/or R BLEED warning lights extinguish

as soon as bleed air is removed from the leaking duct and may not be on long enough to be recognized

by the aircrew.

Automatic functioning of the BALD system may extinguish the L(R)

BLEED warning lights prior to aircrew recognition and may not trigger

the appropriate voice alerts. In this case, cycling the BLEED AIR knob to

remove the L and/or R BLD OFF cautions reintroduces hot bleed air to

the leaking duct. If the sensing element was damaged by the leak,

automatic shutdown and isolation capability may be lost. Extensive

damage and/or fire may result.

The BALD controller also sends a separate command to the SDC which sets an MSP code for the

appropriate sensing element.

A bleed air leak can be verified by MSP codes 953, 954, 955, 956, 957, 958, 959, 960 or 961 (code

determines leak location). An overpressure condition is indicated by MSP code 833 with no bleed air

leak codes.

Sustained high power, high speed operation at low to medium altitude

may result in dual BLEED warnings and loss of OBOGS, cabin pressur-

ization, and g-suit due to secondary bleed air pressure regulator bay

overheat.

2.15.8.1 Bleed Air Leak Detection System Test. The BALD system is tested by the FIRE test switch

in conjunction with the FIRE detection system. Actuation of the FIRE test switch tests the BALD

system sensors and circuitry. The BALD controller turns on the L and R BLEED warning lights,

annunciates the BLEED AIR LEFT/RIGHT voice alerts, and commands the bleed air shutoff valves

closed setting the L and R BLD OFF cautions. The warnings and cautions indicate that the test has

successfully passed.

The L BLEED and R BLEED warning lights go out when the FIRE test switch is released to NORM,

but the L and R BLD OFF cautions remain until the BLEED AIR knob is cycled through OFF to

NORM with ac power applied.

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2.15.8.2 FIRE Test Switch. The FIRE test switch, located on the forward left console, is used to

initiate a test of the FIRE detection and BALD systems. Operation of the FIRE test switch requires

28 vdc essential bus power. The switch is spring-loaded to the NORM position.

TEST A

Initiates a test of loop A of the fire detection and BALD systems.

NORM

Provides normal fire and bleed air leak detection.

TEST B

Initiates a test of loop B of the fire detection and BALD systems.

A successful test of the FIRE detection and BALD systems should illuminate all four bulbs in the

left, right, and APU FIRE lights, both L BLEED and R BLEED warning lights, and should annunciate

all of the following voice alerts in order: ″ENGINE FIRE LEFT, ENGINE FIRE RIGHT, APU FIRE,

BLEED AIR LEFT, BLEED AIR RIGHT″ (each repeated twice).

The BALD controller is sensitive to the duration of FIRE test switch actuation. If the switch is not

held in the TEST A or TEST B position for at least 2 seconds, the BALD controller may set a false

MSP code. Additionally, the BALD controller requires 3 seconds between TEST A and TEST B to

successfully reset. If the switch does not remain in NORM for at least 3 seconds, the BALD controller

may not successfully initiate the bleed air warnings and may set a false MSP code.

2.15.9 DBFS - Dry Bay Fire Suppression System. An active DBFS system is incorporated in the

center dry bays under fuel tanks 2, 3, and 4 to automatically extinguish any fires or explosions which

are detected in these areas. The system consists of fourteen optical fire detectors, seven extinguishing

units and a control unit. The control unit integrates system operation and performs BIT. If the DBFS

system fails BIT, a BIT advisory appears, and DBFS BIT status indicates DEGD.

The DBFS system is totally automatic and requires no aircrew action. The system is armed and

capable of suppressing a fire/explosion in the dry bays when the LDG GEAR handle is up and at least

one generator is on line. If a fire/explosion is detected, the controller discharges all extinguishers,

flooding all dry bays with an inert gas (BAY DISCH caution set). If a fire condition is still detected

after 3 seconds, a BAY FIRE caution is set to alert the aircrew that the fire was not extinguished. If

the fire condition ceases, the BAY FIRE caution resets. With the LDG GEAR handle down, the system

still gives a fire warning (BAY FIRE caution) but is not capable of extinguishing a dry bay fire.

2.16 ENTRANCE/EGRESS SYSTEMS

2.16.1 Canopy System. The cockpit is enclosed by a clamshell type canopy. The main components

of the canopy system are an electromechanical actuator, which provides powered and manual operation

of the canopy, and a cartridge actuated thruster with associated rocket motors, which provides

emergency jettison. When closed, the canopy is latched in place by three hooks on the bottom of each

side of the canopy frame and two forward indexer pins on the lower leading edge of the canopy frame.

When the canopy is closed, the latch hooks and indexer pins engage fittings along the canopy sill, and

the canopy actuator rotates the canopy actuation link over-center, locking the canopy. A mechanical

brake in the canopy actuator motor provides a redundant lock. An inflatable seal, installed around the

edge of the canopy frame, retains cockpit pressure when the canopy is locked. A rain seal is installed

outboard of the pressure seal to divert rain water away from the cockpit. See figure 2-34.

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A high voltage (100,000 volt) static electrical charge may build up in flight

and be stored in the canopy. If possible, make sure that ground crew

discharge the static electricity prior to egress. Otherwise, avoid direct

contact with the outside of the windscreen and canopy to prevent

electrical shock.

Taxiing with the canopy at an intermediate position can result in canopy

attach point damage and failure. Do not open or close the canopy with the

aircraft in motion.

2.16.1.1 Canopy Operation. During normal operation, the canopy is electrically actuated using

either the internal CANOPY switch or external canopy switch. The canopy actuator is powered by the

maintenance bus, which is powered directly by the battery in the absence of ac power (BATT switch

ON or OFF). On battery power, at least five open/close cycles should be available. With the canopy

open, the CANOPY switch must be held to lower the canopy to the rails, slide it approximately 1.5

inches forward, and lock it in place. With the canopy closed and locked, selecting OPEN on the

CANOPY switch (WonW) automatically unlocks and opens the canopy to the full up position. The

switch does not have to be held. Whether the canopy is opening or closing, selecting HOLD stops the

canopy at its present position.

If electrical power is not available, the canopy can be manually operated using either an internal or

external crank system. The canopy can be jettisoned using one of the internal CANOPY JETT

handles.

2.16.1.1.1 CANOPY Switch (Internal). The internal CANOPY switch is located beneath the right

canopy sill in the front cockpit. In LOT 26 and up, the rear cockpit CANOPY switch is located on the

lower right portion of the instrument panel. The CANOPY switch is spring loaded to the HOLD

position and is solenoid-held in the OPEN position only with WonW. The solenoid can be overridden

at any time by placing the switch to HOLD. With WoffW, the switch must be held in the OPEN

position to raise the canopy. Opposing position control commands between the front cockpit and rear

cockpit switches result in a fail-safe OPEN command.

OPEN

Unlocks and/or raises the canopy.

HOLD

Stops the canopy at any point during the open or close cycle.

Lowers and, if held, closes and locks the canopy (CANOPY caution out when

closed and locked).

2.16.1.1.2 Canopy Switch (External). The external canopy switch is located inside the external

power receptacle door (door 9) on the left side of the aircraft below the canopy and LEX. The switch

provides electrical operation of the canopy from outside the cockpit. The switch has the same positions

and operates identically to the internal CANOPY switch, except that the OPEN position is not

solenoid held. After AFC 366, the switch is guarded to prevent inadvertant actuation.

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2.16.1.1.3 Manual Canopy Handcrank (Internal). The internal manual canopy handcrank is stowed

in a clip beneath the left canopy sill. The canopy can be manually opened or closed by inserting the

handcrank into the crank socket immediately above the stowage clip. Approximately 70 counterclock-

wise turns are required to fully open the canopy. Clockwise cranking closes the canopy. A cable is

provided to prevent loss of the handle if dropped.

2.16.1.1.4 Manual Canopy Actuation Fitting (External). The external manual canopy actuation

fitting, a 3/8 inch drive socket on the left side of the aircraft below the canopy, is used to manually

operate the canopy. Inserting a 3/8 inch drive tool in the socket and then turning counterclockwise

approximately 35 turns opens the canopy. Turning the drive tool clockwise closes the canopy.

2.16.1.2 Canopy Jettison System. For canopy jettison, a cartridge initiated thruster is utilized to

unlatch the canopy by moving it 1.5 inches rearward, after which two canopy frame mounted rocket

motors fire to rotate the canopy up and aft, clear of the ejection seat path. The thruster, which provides

attachment for the canopy actuator link during normal canopy operation, is activated by pulling the

ejection seat firing handle or internal canopy jettison handle(s) (figure 2-34). The canopy can be

jettisoned closed, open, or in any intermediate position.

2.16.1.2.1 CANOPY JETT Handle. The CANOPY JETT handle is black and yellow striped and is

located on the left inboard canopy sill just aft of the instrument panel in the front cockpit. Pressing the

button on the tip of the handle unlocks the handle. Pulling the handle aft initiates the canopy jettison

sequence. A ‘‘remove before flight’’ pin is used to manually secure the CANOPY JETT handle between

flights.

2.16.1.2.2 CANOPY JETT Handle (F/A-18F). The rear cockpit CANOPY JETT handle is black and

yellow striped and is located on the left console. Pressing the button on the forward tip of the handle

unlocks the handle. Pulling the handle up initiates the canopy jettison sequence. A REMOVE

BEFORE FLIGHT pin is used to manually secure the CANOPY JETT handle between flights.

2.16.2 Boarding Ladder. A five-step boarding ladder (figure 2-35), stowed under the left LEX,

provides access to the cockpit and the top of the aircraft. Ladder extension and retraction can be

accomplished only from outside the cockpit, either manually or by the ladder remote release button

(after AFC 366) or switch (before AFC 366).

The ladder is extended manually by releasing the latch on the stow assist handle on the ladder’s left

rail (allowing it to drop slightly) and while supporting the ladder, rotating the stow assist handle to

vertical (releasing the remaining two mechanical uplocks on the underside of the LEX). The ladder

rotates down to the extended position. The stow assist handle is then secured. The drag brace locks

when extended to its full length to provide longitudinal stability for the ladder. Lateral stability is

provided by the V-shaped side brace attached to the side of the fuselage.

The LEX is narrow and highly sloped. Use caution to avoid loss of

footing.

NOTE

The ladder is not visible from the cockpit.

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Normal cockpit egress is accomplished by grasping the canopy sill firmly with both hands, leaning

outboard and, using the ladder marking decals as a guide, stepping over the LEX toward the ladder.

The first step is approximately 15 inches below the leading edge of the LEX.

The ladder is stowed by detaching the rigid side brace connection from the fuselage. Pulling the

collar on the drag brace down permits the telescoping drag brace to unlock and compress as the

boarding ladder is rotated up and aft to the stowed position. The latches are manually engaged and

locked by pushing them full up until locked flush with the forward beam. If necessary, the stow assist

handle can be used to assist in stowing the ladder by releasing the handle and pushing the ladder to

the stowed position and pushing the stow assist handle to the closed (stowed) position and releasing it.

With electrical power on the aircraft, a LADDER caution comes on whenever the proximity switch in

the aft portion of the ladder well is not actuated. With the ladder stowed and the 3 latches locked, the

LADDER caution goes out.

The ladder is extended remotely by opening the ground power receptacle door and holding the

ladder remote release switch (lower switch) to the down (DEPLOY) position. The three latches are

opened by a battery powered actuator. The ladder drops to the open position while being restrained

from free falling by a dampening strut (drop time approximately 3 to 4 seconds). All other procedures

for securing the ladder are the same as manual opening.

2.16.2.1 Ladder Remote Release Switch (Before AFC 366). The external ladder remote release

switch is located inside the external power receptacle (door 9) on the left side of the aircraft below the

canopy and LEX. The switch is lever locked and spring loaded to the UP position. Switch positions are:

Removes electrical power from the unlock actuators which allows the ladder to lock

when it is manually stowed.

DEPLOY Applies electrical power to the three unlock actuators which unlocks the actuators

so the ladder can free fall. The switch must be pulled out and down and held in

DEPLOY until the actuators are unlocked.

Due to the close proximity of the ladder remote release switch and the

external canopy switch, positive switch identification is required to

prevent inadvertently lowering the canopy and injuring personnel egress-

ing the aircraft.

2.16.2.2 Ladder Remote Release Button (After AFC 366). The guarded external ladder remote

release button is located inside the external power receptacle (door 9) on the left side of the aircraft

below the canopy and LEX. Pressing and holding the button applies electrical power to the three

unlock actuators which unlocks the actuators so the ladder can free fall.

2.16.3 Ejection Seat. The SJU-17 (V)1/A, 2/A, and 9/A, and SJU-17A (V)1/A, 2/A, and 9/A NACES

(Navy Aircrew Common Ejection Seat) are ballistic catapult/rocket systems that provide the pilot with

a quick, safe, and positive means of escape from the aircraft. See Ejection Seat, foldout section, for

ejection seat illustrations. The seat system includes an initiation system which, after jettisoning the

canopy and positioning the occupant for ejection, fires the telescopic seat catapult. Canopy breakers on

the top of the seat give capability of ejecting through the canopy. As the seat departs the aircraft and

the catapult reaches the end of the stroke a rocket motor on the bottom of the seat is fired. The thrust

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ORIGINAL

A1-F18EA-NFM-000

Figure 2-35. Boarding Ladder

of the rocket motor sustains the thrust of the catapult to eject the seat to a height sufficient for

parachute deployment even if ejection is initiated at zero speed, zero altitude in a substantially level

attitude.

NOTE

Safe escape is provided for most combinations of aircraft altitude,

speed, attitude, and flight path within the envelope of 0 to 600 KCAS

airspeed and 0 to 50,000 feet.

2.16.3.1 SJU-17(V) 1/A, 2/A, and 9/A, and SJU-17A (V)1/A, 2/A, and 9/A NACES Seat. Timing of

all events after rocket motor initiation is controlled by the electronic sequencer which utilizes altitude,

acceleration, and airspeed information to automatically control drogue and parachute deployment and

seat/man separation throughout the ejection seat’s operational envelope. In the event of partial or total

failure of the electronic sequencer, a 4-second mechanical delay initiates a barostatic release unit which

frees the occupant from the seat and deploys the parachute between 14,000 and 16,000 feet MSL if the

ejection occurred in or above this altitude range. The emergency barostatic release unit operates

immediately after the 4-second delay if the ejection occurred below 14,000 feet MSL. An emergency

restraint release (manual override) system provides a backup in the event of failure of the barostatic

release unit. The seat is stabilized and the forward speed retarded by a drogue chute attached to the

top and bottom of the seat. The parachute deployment rocket is automatically fired to withdraw the

parachute from deployment bag. Full canopy inflation is inhibited until the g forces are sufficiently

reduced to minimize opening shock. There are 5 modes of operation. See figure 2-36 for parameters

that determine the mode of operation and the corresponding parachute deployment and drogue chute

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A1-F18EA-NFM-000

Figure 2-36. SJU-17 and SJU-17A Ejection Modes

release times. At high altitude the drogue chute deploys to decelerate and stabilize the seat. The seat

falls drogue retarded to 18,000 feet MSL where the drogue is released, the main parachute is deployed,

and seat/man separation occurs. At medium altitude, (between 18,000 and 8,000 feet MSL), and at low

altitude (below 8,000 feet MSL) parachute deployment is automatically delayed from 0.45 to 2.90

seconds (depending upon airspeed and altitude) after first seat motion to allow the drogue chute to

decelerate and stabilize the seat.

The main parachute is a 21 foot aeroconical canopy type, stored in a headbox container on top of the

ejection seat. The parachute is steerable and contains water deflation pockets which aid in dumping air

from the canopy after landing in water. The seat drogue chute is stored in a separate container on top

of the drogue deployment catapult. The seat contains controls for adjusting seat height and for locking

and unlocking the inertia reel shoulder restraint straps. A survival kit is installed in the seat pan.

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A1-F18EA-NFM-000

2.16.3.2 SEAWARS - SEAWATER Activated Release System. SEAWARS is a seawater activated

system that automatically releases the parachute from the crew member. When the sensing-release

units are immersed in seawater, cartridges are fired which allow the crew member to separate from the

parachute.

2.16.3.3 Ejection Control Handle. The ejection control handle, located between the crewman’s legs

on the front of the seat pan, is the only means by which ejection is initiated. The handle, molded in the

shape of a loop, can be grasped by one or two hands. To initiate ejection, a 20 to 40 pound pull removes

the handle from its housing, and a continued pull of 30 to 60 pounds is required to pull both sears from

the dual initiators. Either of the initiators can fire the seat. After ejection, the handle remains attached

to the seat. The ejection control handle safes the ejection seat safe/armed handle.

2.16.3.4 Ejection Seat SAFE/ARMED Handle. To prevent inadvertant seat ejection, an ejection seat

safe/armed handle is provided. The handle, forward on the right seat armrest, safeties the seat when

rotated up and forward, and arms the seat when it is rotated aft and down. The safe/armed handle is

locked when placed to either of these two positions and the handle must be unlocked by squeezing a

locking lever within the handle cutout before changing positions. When in the armed position the

visible portion of the handle (from the occupant’s vantage point) is colored yellow and black with the

word ARMED showing. In the safe position, the visible portion is colored white with the word SAFE

showing. The seat is safe only when the word SAFE is entirely visible on the inboard side of the

SAFE/ARM handle and the handle is locked in the detent. Placing the handle to the SAFE position

causes a pin to be inserted into the ejection firing mechanism to prevent withdrawal of the sears from

the dual seat initiators.

2.16.3.4.1 CK SEAT Caution. The CK SEAT caution light is located on the caution light panel and

repeats the DDI CHECK SEAT caution. The caution comes on when the right throttle is at MIL or

above, weight is on wheels, and the ejection seat is not armed.

2.16.3.5 Shoulder Harness Inertia Reel. Pilot shoulder harness restraint is provided by a dual strap

shoulder harness inertia reel mounted in the seat below the parachute container. The dual inertia reel

shoulder straps connect to the parachute risers which in turn are buckled to the seat occupant’s upper

harness. The inertia reel locks when the reel senses excessive strap velocity. Manual locking and

unlocking of the reel is controlled by the shoulder harness lock/unlock handle on the left side of the

seat bucket. During ejection a pyrotechnic cartridge is fired to retract the shoulder harness to position

the seat occupant for ejection.

2.16.3.6 Shoulder Harness Lock/Unlock Handle. The shoulder harness lock/unlock handle on the

left side of the seat bucket has two positions. To operate, the handle must be pulled up against spring

pressure, moved to the desired position, and released.

FORWARD The inertia reel prevents the reel straps from being extended and ratchets any

(locked)

slack in the straps back into the reel.

AFT

The reel allows the pilot to lean forward, but the inertia portion of the reel contin-

(unlocked)

ues to protect by locking the reel when it senses excessive strap velocity. Once

locked, the pilot can normally lean forward again after a slight release in pressure

on the reel straps.

2.16.3.7 Leg Restraint System. A leg restraint system is located on the front of the ejection seat. The

function of the system is to secure the occupant’s legs to the seat during ejection. The system consists

of two adjustable leg garters, a restraint line, and a snubber box for each leg. One garter is worn on the

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ORIGINAL

A1-F18EA-NFM-000

Figure 2-37. Leg Restraint System

thigh and one on the lower leg. The restraint lines are routed through the garter rings and the snubber

box as shown in figure 2-37. One end of each restraint line is secured to the cockpit floor and the other,

after being routed through the snubber box and both garter rings, is secured to the seat just outboard

of the snubber box by a releaseable pin. During ejection, the slack in each line is taken up and the

tension builds up to finally separate the lines at the tension rings in the leg lines. At man/seat

separation, the pins on the other end of the lines are released by the time release mechanism. The pins

are also released when the manual override handle is pulled. Both the lower garter and thigh garter

contain a quick release buckle which disconnects the ring through which the leg restraint line runs,

permitting the pilot to egress from the aircraft wearing both upper and lower garters. In addition, toe

clips are installed on the tops of the rudder pedals to prevent contact between the toes and the

instrument panel during ejection.

2.16.3.7.1 Leg Restraint Snubber Release Tabs. The leg restraint lines are adjusted to give the pilot

more leg movement by pulling inboard the leg restraint snubber release tabs (figure 2-37) and

simultaneously pulling the leg restraint lines forward through the snubber box.

2.16.3.8 Seat Survival Kit (SKU-10/A). The SKU-10/A survival kit is used with the SJU-17 and

SJU-17A ejection seats. This survival kit, which fits into the seat bucket, is a contoured rigid platform

which contains an emergency oxygen system and a fabric survival rucksack (figure 2-38). A cushion on

top of the platform provides a seat for the aircrew.

The rigid platform forms a hard protective cover to the survival package and oxygen system and is

retained in position in the seat bucket by brackets at the front and lugs secured in the lower harness

locks at the rear. Attached to the lugs are two adjustable lap belts with integral quick release fittings.

A flexible oxygen and communication hose is installed in the left aft side of the upper kit to provide

a connection to the aircrew for aircraft oxygen and communication. An emergency oxygen cylinder,

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A1-F18EA-NFM-000

pressure reducer, and associated pipe work are mounted on the underside of the platform. A green

manual emergency oxygen operating handle is mounted on the left side of the platform and a pressure

gage is on the inside face of the left leg support. The emergency oxygen can be activated manually by

pulling the green emergency oxygen handle upwards. The green emergency oxygen handle can be reset,

shutting off the flow of emergency oxygen, by pushing downward on the button on the front end of the

emergency oxygen handle assembly. The emergency oxygen is automatically activated during ejection

by a lanyard connected between the floor and the survival kit. An AN/URT-33A locator beacon is

located in a cutout in the left leg support. The beacon is actuated during ejection by a lanyard

connected to the emergency oxygen lanyard.

The survival rucksack is retained to the underside of the rigid platform by 5 fabric straps and a

double cone and pin release system. The package accommodates a life raft and survival aids. Two

yellow manual deployment handles are mounted on the aft surface of the kit. Pulling either handle

enables the aircrew to deploy the raft and survival package after man/seat separation. The life raft

inflates automatically on survival package deployment and is attached to the survival package with a

line. If the survival kit must be deployed after water entry, a snatch pull on the red manual activation

handle near the CO₂ bottle is required to inflate the life raft.

2.16.3.9 Manual Override Handle. A manual override handle permits releasing the pilot’s lower

harness restraints and the leg restraint lines for emergency egress and permits resuming part of the

ejection sequence (man/seat separation and main parachute deployment) in the event of sequencing

failure during ejection. The manual override handle, on the right side of seat bucket and just aft of the

ejection seat safe/arm handle, is actuated by pressing a thumb button on the forward part of the handle

and rotating the handle up and aft. If the manual override handle is actuated on the ground or in the

air before ejection, survival kit attachment lugs and leg restraint lines are released, the inertia reel is

unlocked, and the ejection seat safe/armed handle automatically rotates to the SAFE position. During

ground emergency egress, after the manual override handle is pulled and the parachute riser fittings

are released, the pilot is free to evacuate the aircraft with the survival kit still attached. If the manual

override handle is actuated after ejection but before man/seat separation occurs, the following events

take place: release of survival kit attachment lugs, negative-g strap (SJU-17 (V)1/A, 2/A, and 9/A only),

leg restraint lines, and inertia reel straps; firing of the manual override initiator cartridge; firing of the

barostatic release unit; and firing of the parachute deployment rocket, which deploys the parachute.

The ejection seat safe/armed handle automatically rotates to the SAFE position whenever the manual

override handle is actuated.

Pulling the manual override handle automatically rotates the ejection

seat safe/armed handle to the SAFE position, releases the survival kit

attachment lugs and leg restraint lines, and unlocks the inertia reel. If

this is done inflight, the aircrew will be unable to eject.

2.16.3.10 Seat Bucket Position Switch. The seat bucket position switch is on the left side of the seat

bucket, forward of the shoulder harness lock/unlock handle. The forward switch position lowers the

seat bucket, the aft position raises the seat bucket. The center off position, to which the switch is spring

loaded, stops the seat bucket. The maximum vertical travel of the seat bucket is 5.1 inches for the

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ORIGINAL

A1-F18EA-NFM-000

SJU-17 (V)1/A, 2/A, and 9/A, and 6.1 inches for the SJU-17A (V)1/A, 2/A, and 9/A. The actuator

should not be operated over 1 minute during any 8 minute period.

• To prevent increased risk of thigh slap or leg contact injuries, aircrew

with a buttock-to-knee length greater than 25.5 inches should not use

either of the two forward backpad positions. Aircrew with buttock-to-

knee length between 24.6 and 25.5 inches should not use the full

forward backpad position.

• Actuation of the seat bucket position switch with the leg restraints

under the seat bucket may result in an inadvertent ejection.

Actuation of seat bucket position switch with lap belts, shoulder harness,

and/or leg restraints outside or under seat bucket may damage ejection

seat, leg restraints, and/or Koch fittings.

2.16.3.11 Backpad Adjustment Mechanism (SJU-17A (V)1/A, 2/A, and 9/A). The backpad adjust-

ment mechanism handle is on the seat bucket adjacent to the top left hand side of the backpad and is

connected to the backpad by a linkage. The backpad has three positions, full-forward, middle, and full-

aft, which give a total forward/aft adjustment of 1.6 inches. When the handle is in the full-up position,

the backpad is full-aft, and when the handle is full-down, the backpad is full-forward. To move the

backpad, the adjustment handle is moved within a quadrant until a spring-loaded plunger engages in

one of the three detent positions in the quadrant. Set the backpad for personal comfort and best access

to flight controls during initial strap-in and prior to flight.

2.16.4 Ejection Seat System (F/A-18F). The ejection seats in the F/A-18F are ejected at opposite

divergent angles to one another. The rear seat diverges to the left while the forward seat diverges to the

right. The amount of divergence is influenced by the weight of the aircrew and the speed of the ejection.

The heavier the aircrew and the faster the speed the less the resulting divergent angle. In addition, a

sequencing system is installed to allow dual ejection initiated from either cockpit or single (aft) seat

ejection initiated from the rear cockpit. A command selector valve is installed in the rear cockpit to

control whether ejection from the rear cockpit is dual or single.

2.16.4.1 EJECTION MODE Handle (F/A-18F). The EJECTION MODE handle is located on the right

side of the main instrument panel in the rear cockpit. The EJECTION MODE handle is used to select

the desired ejection sequence to be initiated from the rear cockpit, or provide for single ejection for solo

flight. Positioning is accomplished by pulling out while turning to the desired position. The SOLO

position requires the use of a collar to hold the handle in that position. To release from AFT

INITIATE, pull then turn clockwise.

NORM

Single rear seat ejection when initiated from the rear cockpit. Dual ejection (rear

(vertical)

seat first) when initiated from the front cockpit.

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A1-F18EA-NFM-000

AFT

Dual ejection (rear seat first) when initiated from either cockpit.

INITIATE

(horizontal)

SOLO

Front seat ejection only when initiated from the front cockpit. Front seat ejection

(45° CCW)

is immediate. Rear seat ejection only when initiated from the rear seat. Rear seat

ejection is immediate.

• SOLO mode shall NOT be selected when both seats are occupied. If

SOLO mode is selected when both seats are occupied, simultaneous

ejection initiation may result in a collision between seats.

• SOLO mode shall be selected when the aircraft is being flown solo.

Alternate selection when flying solo results in ejection of unoccupied

seat and possible collision with the front cockpit seat.

When selecting NORM or SOLO from AFT INITIATE, the handle must

be pulled before rotation or damage to the command selector valve may

result.

2.16.4.2 SEAT CAUT MODE Switch (F/A-18F). The SEAT CAUT MODE switch is located in the

rear cockpit above the command selector valve. The switch position changes the operation of the CK

SEAT caution for solo or dual flight.

NORM

CK SEAT caution is activated by either seat remaining safed. Switch is spring

loaded to this position.

SOLO

CK SEAT caution is activated only by the front seat remaining safed. Switch must

be pinned to remain in this position.

2.16.4.3 CK SEAT Caution (F/A-18F). The CK SEAT caution light is located on the caution light

panel, and repeats the DDI CHECK SEAT caution display. The caution is displayed when the right

throttle is at MIL or above, weight is on wheels, and the cockpit ejection seat is not armed with the rear

cockpit NORM/SOLO switch set to SOLO or either ejection seat is not armed with the NORM/SOLO

switch set to NORM.

2.17 EMERGENCY EQUIPMENT

2.17.1 Jettison Systems. The jettison systems consist of the emergency jettison system and the

selective jettison system.

2.17.1.1 Emergency Jettison. Emergency jettison is performed by pushing the EMERG JETT

button with either the LDG GEAR handle UP or with WoffW. When activated, the emergency jettison

system jettisons all stores, launchers, and racks from the BRU-32 racks on the six wing pylon stations

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ORIGINAL

A1-F18EA-NFM-000

(2, 3, 4, 8, 9, and 10) and the centerline station (6). Emergency jettison is sequential by station pairs:

3 and 9, 2 and 10, 4 and 8, then 6. There is a 100 msec, ±25 msec, delay before the first set of stations

is jettisoned and in between each subsequent set.

2.17.1.2 EMERG JETT Button. The EMERG JETT button is located on the left side of the main

instrument panel and is black and yellow striped. The button must be pressed and held during the

entire jettison sequence. The EMERG JETT button is used to initiate emergency jettison with the

LDG GEAR handle UP or with WoffW.

The EMERG JETT button must be pressed for 500 msec to make sure all

stores are jettisoned.

If the EMERG JETT button has been pushed on the ground prior to

takeoff and remains stuck in, emergency jettison is activated as soon as

the aircraft goes WoffW. The only cockpit indication of this condition is

SMS BIT status DEGD and MSP 082 (Emergency Jettison Switch Failed

On).

2.17.1.3 Selective Jettison. Selective jettison is performed using the SELECT JETT knob, in

conjunction with the JETT STATION SELECT buttons, and jettisons stores in a safe condition. The

stores or the launchers/racks (with any attached stores) can be jettisoned from the centerline and wing

stations, and the missiles can be jettisoned from the fuselage stations.

Selective jettison requires ARM conditions satisfied and all the landing gear up and locked. ARM

conditions are satisfied with WoffW, LDG GEAR handle UP, MASTER ARM switch in ARM, and

SIM mode unboxed. ARM status can be confirmed on the STORES page. All the landing gear up and

locked can be confirmed by the absence of the LDG GEAR handle warning light/landing gear warning

tone with the LDG GEAR handle UP.

Selective jettison of the centerline and wing stations requires station(s) selection by the JETT

STATION SELECT buttons and STORES or RACK/LCHR selection by the SELECT JETT knob.

Selective jettison of a fuselage station missile requires R FUS MSL or L FUS MSL selection by the

SELECT JETT knob.

With all the requirements met, selective jettison is performed by pressing the JETT center

pushbutton in the SELECT JETT knob.

2.17.1.3.1 JETT STATION SELECT Buttons. The JETT STATION SELECT buttons are on the left

edge of the instrument panel below the emergency jettison button. The buttons are labeled CTR, LI,

RI, LM, RM, LO and RO. Pressing a button turns on an internal light and selects a weapon station for

jettison. The JETT STATION SELECT buttons are also used in the backup A/G weapon delivery

modes for weapon selection; refer to A1-F/A-18EA-TAC (Series).

2.17.1.3.2 SELECT JETT Knob. The SELECT JETT knob on the left vertical panel has rotary

positions L FUS MSL, SAFE, R FUS MSL, RACK/LCHR, and STORES. L FUS MSL and R FUS

MSL select either fuselage missile for jettison. The RACK/LCHR and STORES positions select what

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ORIGINAL

A1-F18EA-NFM-000

is to be jettisoned from the weapon station(s) selected by the JETT STATION SELECT buttons. The

JETT center pushbutton activates the jettison circuits provided ARM conditions are satisfied and all

the landing gear are up and locked. The SAFE position prevents any selective jettison.

2.17.1.3.3 AUX REL Switch. The AUX REL switch, on the lower instrument panel, is used to enable

jettison of hung stores or store and rack/launcher combinations from BRU-32 racks on stations 2, 3, 4,

6, 8, 9, and 10. A need to use the AUX REL switch is indicated by a hung indication on the DDI after

a selective jettison or a normal weapons release is attempted. Placing the switch to ENABLE selects

the auxiliary release function. AUX RELEASE has the same requirements as those listed above for

selective jettison. Jettison is initiated by selecting RACK/LCHR or STORES on the selective jettison

knob, selecting the hung store station by pressing the appropriate station jettison select button, and

pressing the JETT center pushbutton of the selective jettison knob. The SMS provides a jettison signal

to fire the auxiliary cartridge in the BRU-32 rack on which the hung store or store and rack/launcher

combination is loaded. After the cartridge is fired, the store or rack/launcher is gravity dropped with

the store in a safe condition. This switch is also used with some weapons for a second normal release

attempt after these weapons have been hung during a first normal release attempt. Refer to

A1-F18EA-TAC-Series for these weapons and procedures.

2.17.2 Warnings/Cautions/Advisories. The warning/caution/advisory system provides visual indi-

cations of normal aircraft operation and system malfunctions affecting safe operation of the aircraft.

The lights are on various system instruments and control panels in the cockpit. The red warning lights

normally indicate a systems malfunction that could be a severe hazard to further flight, and may

require immediate action. Caution lights and displays normally, but not always, indicate malfunctions

that require attention but not immediate action. After the malfunction has been corrected, warning

and caution lights and caution displays go out. Advisory lights and displays indicate safe or normal

conditions and supply information for routine purposes. Warning, caution and advisory displays are

NVG compatible. Caution and advisory displays appear on the left or right DDI and the MPCD,

depending on the number of displays in operation. The advisory displays start at the bottom of the

display and are preceded by ADV. The caution displays, in larger characters than the advisory displays,

appear immediately above the advisory displays. The caution lights, located on the caution lights panel

and the instrument panel, are yellow lights. The advisory lights, scattered throughout the cockpit(s),

are green. Lights that have been lit on the caution lights panel flash when overheated to prevent light

damage.

2.17.2.1 MASTER CAUTION Light. A yellow MASTER CAUTION light, on the upper left part of the

instrument panel, comes on when any of the caution lights or caution displays come on. The MASTER

CAUTION light goes out when it is pressed (reset). An audio tone is initiated whenever the MASTER

CAUTION light comes on. The tone is of 0.8 second duration and consists of a 0.25 second sound

followed by a 0.15 second sound of higher pitch, followed by one repetition of these sounds. The tone

does not repeat unless the original condition causing the tone clears and recurs 5 seconds after the first

tone, regardless of whether or not the MASTER CAUTION is reset. Additional cautions sound the

tone, regardless of whether or not the MASTER CAUTION is reset, providing about 5 seconds have

elapsed since the previous caution. Pressing the MASTER CAUTION when it is unlighted causes the

uncorrected caution and advisory displays to reposition to the left and to a lower level, provided there

is available space vacated by corrected caution and advisory displays. To restack the cautions and

advisories when the MASTER CAUTION is lighted, the MASTER CAUTION must be pressed twice:

first, to turn off the MASTER CAUTION light and second, to reposition the caution and advisory

displays. A reset MASTER CAUTION light (and tone) comes on if there is at least one uncorrected

caution present when weight is on the wheels and both throttles are moved beyond approximately 80%

rpm if both throttles were below 80% for at least 60 seconds.

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A1-F18EA-NFM-000

2.17.2.2 MASTER CAUTION Light (F/A-18F). A yellow MASTER CAUTION light, on the upper

instrument panel comes on whenever the MASTER CAUTION light in the front cockpit comes on.

The rear cockpit MASTER CAUTION light goes out whenever the front cockpit MASTER CAUTION

is reset.

2.17.2.3 Dimming and Test Functions. There are no provisions for testing the caution and advisory

displays and each DDI contains its own display dimming controls. The warning/caution/advisory lights

are dimmed by the warning/caution lights knob and are tested by the lights test switch. The following

lights can be dimmed by the warning/caution lights knob, but once in the dimmed lighting range

cannot be varied in intensity: MASTER CAUTION light, landing gear handle warning, L BAR

warning, HOOK warning, L BLEED warning, R BLEED warning, APU FIRE warning, left and right

engine FIRE warning.

2.17.3 Voice Alert System. For certain critical warnings and cautions, voice alert transmissions are

sent to the aircrew’s headset. The message is repeated twice; for example, APU FIRE, APU FIRE. The

voice alert requires no reset action on the pilot’s part and the alert is not repeated unless the original

condition ceases for 5 seconds or more and then recurs. For cautions with voice alert, the voice alert

replaces the master caution tone; however, the master caution tone backs up the voice alert system and

provides a tone if the voice alert system malfunctions. FIRE, APU FIRE, L BLEED, and R BLEED

warning lights are not backed up by the master caution tone. Voice alert is the only audio warning for

these problems. With dual generator failure, the following voice alert warnings operate from battery

power: APU FIRE, ENGINE FIRE LEFT (RIGHT), and BLEED AIR LEFT (RIGHT). All voice alert

cautions, and the master caution tone are inoperative on battery power during dual generator failure.

Once a voice alert has been activated, it cannot be interrupted by a higher priority voice alert. All

voice alerts play until completed. The primary radar low altitude warning (WHOOP, WHOOP), is

repeated at the lowest priority until reset or disabled by the pilot. With MC OFP 20X AND UP OR

H3E AND UP, the BINGO voice alert is repeated every 30 seconds until the BINGO setting is

adjusted.

NOTE

With an MC1 failure (LOTs 21-24), the voice alert does not sound

when the aircraft descends below the altitude set by the primary low

altitude warning setting.

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A1-F18EA-NFM-000

CAUTION

VOICE ALERT

IFF 4

MODE 4 REPLY

FCS

FLIGHT CONTROLS

FCS HOT

FLIGHT COMPUTER HOT

L (R) OVRSPD

L (R) EGT HIGH

L (R) FLAMEOUT

ENGINE LEFT (RIGHT)

L (R) OIL PR

L (R) STALL

L (R) ENG

L (R) ENG VIB

FUEL LO

FUEL LOW

BINGO

WARNING

VOICE ALERT

ALTITUDE

L (R) BLEED AIR

BLEED AIR LEFT (RIGHT)

L (R) FIRE

ENGINE FIRE LEFT (RIGHT)

APU FIRE

2.17.4

Terrain Awareness Warning System (TAWS) (MC OFP 18EA and H2E AND UP). The terrain

awareness warning system alerts the aircrew of a controlled flight into terrain (CFIT) condition during

all mission phases. The system operates any time that the navigation mission computer (MC1) and

TAMMAC digital mapping set (DMS) are functional. TAWS functions as a safety backup system and

not as a performance aid. TAWS has been designed to eliminate false warnings, minimize nuisance

warnings, and generate consistent aircrew response in all aircraft master modes. Five possible voice

warnings are provided to indicate the correct initial response to an impending CFIT condition, and a

visual cue is provided to indicate the recovery direction of pull, or in some instances, to command an

increase in turn rate. All TAWS warnings should be treated as though an imminent flight into terrain

condition exists. Pilot response to a TAWS warning should be instinctive and immediate.

TAWS uses data from the following inputs: FCC, INS, RADALT, GPS, and digital terrain elevation

data (DTED). DTED resides in the DMS as part of TAMMAC and is used to provide the

forward-prediction capability that protects against flight into rising terrain. The TAWS option is

reached by pressing MENU-HSI-DATA-A/C as shown in figure 2-39. The TAWS option boxes

automatically at start-up.

When a DMS is not installed in the aircraft or is not operational, protection from CFIT events is

provided by the Ground Proximity Warning System (GPWS). BIT may be initiated on the DMS by

pressing the appropriate pushtile of the BIT display. The BIT can take up to 185 seconds to complete.

During the BIT, TAWS is not operational. Therefore, the GPWS algorithm is used to determine the

presence of possible CFIT events. There is no capability for pilot selection of GPWS if DMS is

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Figure 2-39. HSI-DATA-A/C Controls

operational. The GPWS algorithm runs continuously with outputs being overwritten if TAWS is

operational. This prevents erroneous values during an unexpected transition from TAWS to GPWS.

2.17.4.1 TAWS Modes. TAWS has two operational modes: TAWS-with-DTED, and TAWS- without-

DTED. These modes switch automatically depending upon the available sensor data and flight phase.

When the aircraft position (latitude and longitude) is accurately known and DTED for the local area

has been loaded onto the DMS (during the theater load process), TAWS is in the TAWS-with-DTED

mode and provides protection against varying terrain ahead of the aircraft. When the aircraft position

is not accurately known, DTED for the local area is unavailable, or TAWS determines the aircraft is

in a landing phase, TAWS transitions to the TAWS-without-DTED mode and provides protection

against flight into level or descending terrain as GPWS does.

When operating over the ocean, DTED does not exist and TAWS will be in the TAWS-without-

DTED mode. However, there is no degradation in protection because the ocean is relatively flat. As the

aircraft approaches the coast or islands, DTED may be available (depending upon the theater load)

and TAWS will automatically switch back to the TAWS-with-DTED mode.

Operation of TAWS in the TAWS-without-DTED mode is still an improvement over GPWS as

TAWS incorporates a more robust performance model and additional input sensor redundancy.

2.17.4.2 TAWS Operation. TAWS incorporates signal processing that determines a best estimate of

aircraft position and altitude (AGL and MSL). TAWS protection algorithm continuously computes

two recovery trajectories: Vertical Recovery Trajectory (VRT) and Oblique Recovery Trajectory

(ORT). VRT is the standard GPWS-like recovery: roll to wings-level, if needed, and pull to recover.

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ORT assumes that you maintain the current bank angle and pull to recover (increase turn rate). Both

computed trajectories include the following assumptions:

a. Pilot Response Time is the time from issuance of a TAWS warning to the time that the pilot

actually initiates recovery. Pilot Response Time is set at 1.3 seconds.

b. Roll Recovery Phase is the time necessary to roll the aircraft to near wings-level. This assumes

at least ½ lateral stick will be used for bank angles less than 70° and at least ¾ lateral stick

will be used for bank angles ≥ 70°.

c. G-Onset Phase is the time required to pull to the target recovery g. The target recovery g is

80% of the instantaneous g available, or 5g, whichever is less. The g-onset phase assumes that

rapid aft stick motion will be used (full deflection within ¾ second). In addition, TAWS

assumes that throttles will be moved to MAX if below corner speed and to IDLE if above

corner speed.

d. Dive Recovery Phase is the remainder of the trajectory until terrain clearance is achieved.

TAWS assumes a terrain clearance of 50 ft.

When TAWS senses that the aircraft is in the landing configuration, the recovery assumptions must

change since the desire is to land. TAWS defines the landing phase as below 500 ft AGL, less than 200

KCAS, landing gear down and locked, and more than one minute since a waveoff or takeoff. In the

landing phase, TAWS protects against landings of greater than the structural limit of the landing gear

(1584 fpm). To allow this, TAWS switches to TAWS-without-DTED and provides a warning when the

landing is predicted to exceed the structural limit of the landing gear.

TAWS provides protection against gear-up landings. When the aircraft is below 200 KCAS, below

150 ft AGL, more than one minute since waveoff or takeoff, and the landing gear is not down and

locked, a TAWS warning is provided.

2.17.4.3

TAWS Warnings. TAWS provides clear, unambiguous, and directive aural and visual cues

to the aircrew. Aural warnings provide the aircrew with a wake-up call and correct initial response

while visual warnings provide the aircrew with correct follow-on recovery information.

2.17.4.3.1 Voice Warnings. TAWS uses the ACI to provide aural cues to the aircrew. The aural cues

are distinct from any other cues that the aircrew may receive. The TAWS voice alert warnings are:

″Roll−Left...Roll−Left″,

″Roll−Right

Roll−Right″,

″Pull−Up...Pull−Up″,

″Power...Power″, and

″Check Gear″. Each of these warnings is issued at a level 3−6 dB above the present voice alerts. The

TAWS voice warnings provide a wake−up call to the aircrew and indicate the most appropriate initial

response for the given aircraft state, not necessarily the only required response. The aural cue repeats

until the warning condition is cleared. TAWS aural warnings have priority over all current aural tones.

Note that TAWS requires a -1018 or greater ACI load (or MIDS equivalent) which is capable of

generating the ″Roll Left/Right″ warnings. Earlier ACI loads were only capable of generating ″Roll

Out″ warnings and are not desired for use with TAWS.

A ″Roll Right...Roll Right″ warning is issued when a roll to the right is the correct initial response.

A ″Roll Left...Roll Left″ warning is issued when a roll to the left is the correct initial response.

A ″Power...Power″ warning is issued when the roll requirement conditions have not been met and

adding power is the correct initial response. This occurs when the aircraft is below 200 KCAS, the AOA

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Figure 2-40. TAWS HUD Visual Recovery Warning - Pull Up (VRT)

is above 8.5° for PA configuration (or 18° AOA for up and away configuration) and the throttle is not

already at MAX. The correct response to this warning is to select MAX afterburner.

A ″Pull Up...Pull Up″ warning is issued when the above conditions have not been met and pulling up

is the correct response or when the ORT is the recovery trajectory.

When a warning is given to protect against a gear-up landing, the following aural cues may be heard:

″Pull Up...Pull Up″ followed two seconds later by ″Check Gear″ when the gear handle is in the UP

position and a gear-up landing condition has been assessed (repeated every 4 seconds).

″Check Gear″ repeated every 8 seconds when the gear handle is down and a gear up landing condition

has been assessed.

2.17.4.3.2

Visual Warnings. A visual recovery arrow is provided in the center of the HUD and HUD

format on the DDI. The recovery arrow indicates the direction of recovery. The visual warning is

displayed when a CFIT condition is present and is removed when the CFIT condition is cleared.

TAWS visual recovery cues are designed to be used in conjunction with TAWS voice warnings.

There are several voice warning/visual recovery cue combinations. When the arrow points UP in the

HUD (i.e., along the lift vector), a longitudinal pull is the correct response and an aural ″Pull Up...Pull

Up″ is heard. This is a VRT recovery if the aircraft is close to wings level, or it is an ORT (increased

turn rate) recovery if the aircraft is banked such that the TAWS algorithm assessed that an increased

turn rate would provide the quickest recovery from an impending CFIT condition. Figures 2-40 and

2-41 depict these two situations. Both situations require a longitudinal pull as the correct response,

however, the first case (VRT) depicts a dive recovery while the second case (ORT) depicts a recovery

requiring an increase in turn rate by increasing g when already in an established angle of bank.

When the arrow points anywhere other than UP in the HUD (i.e., not along the lift vector, but

perpendicular to the horizon), it may be accompanied by either a ″Roll Left (Right)″...″Roll Left

(Right)″ or ″Pull Up...Pull Up″ voice warning. The voice warning indicates the correct initial response,

then the aircrew should roll or pull as required to place or maintain the TAWS recovery arrow straight

up in the HUD (i.e., along the lift vector). For example, if a ″Roll Left...Roll Left″ voice warning is

issued with an accompanying HUD recovery arrow displayed in the HUD that is perpendicular to the

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Figure 2-41. TAWS HUD Visual Recovery Warning - Pull Up (ORT)

horizon, the correct response is to roll left to align the lift vector with the HUD recovery arrow and then

perform a dive recovery. If a ″Pull Up...Pull Up″ voice warning is issued with an accompanying HUD

recovery arrow displayed in the HUD that is perpendicular to the horizon, the correct response is to

apply g along the current lift vector and then, referencing the HUD recovery arrow, roll to align the lift

vector with the HUD recovery arrow and perform a dive recovery. Figure 2-41 depicts a situation in

which a Roll or Pull Up aural warning could be issued. If a ″Roll Right...Roll Right″ aural warning was

issued, a roll to the right would be the correct initial response and then a dive recovery would be

continued with a longitudinal pull. If a ″Pull Up...Pull Up″ aural warning was issued, a longitudinal pull

would be the correct initial response and then a roll to the right to align the lift vector with the HUD

recovery arrow followed by a longitudinal pull for a dive recovery would be the follow-on recovery

procedure.

2.17.4.4 ACI Configuration Check. There are different ACI configurations in the F/A-18 aircraft.

Two TAWS software loads exist: -1016 and -1018. The -1016 ACIs can command these four aural cues:

″Roll Out...Roll Out″, ″Pull Up...Pull Up″, ″Check Gear″, and ″Power...Power″. The -1018 and greater

ACIs added logic for replacing the ″Roll Out...Roll Out″ aural cue with ″Roll Left...Roll Left″ and ″Roll

Right...Roll Right″. Due to the variety of possible combinations, on cold start power-up, the MC

commands a ″Roll Left...Roll Left″ to the ACI to determine if it is in the -1018 configuration. If there

is no response, the MC commands a ″Roll Out...Roll Out″ to the ACI to determine if the ACI is in the

-1016 configuration. If no response is received the second time the MC attempts to command a ″Roll

Out...Roll Out″, the TAWS/GPWS voice warnings will not be heard, however the visual arrow will still

be present when a warning is issued.

2.17.5 GPWS - Ground Proximity Warning System. GPWS is designed to backup the pilot by

providing an alert of impending controlled flight into terrain (CFIT). GPWS provides warnings of

potentially unsafe maneuvering flight conditions such as excessive bank angles, excessive sink rates,

gear up landings, floor altitude violations, and altitude loss during recovery. The system is operational

as long as MC1, radar altimeter, and air data systems are ON and functional. The GPWS algorithm

operates in the backg1round of the OFP with no cockpit indications until an actual CFIT warning is

required. The system provides distinctive aural and visual warning cues only, to alert and direct

recovery from an impending CFIT condition. The pilot maintains full control of the aircraft for

recovery.

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Figure 2-42. TAWS HUD Visual Recovery Warning (VRT)

2.17.5.1 GPWS Sensors/Modes. GPWS is a look-down system with no forward-looking capability.

GPWS uses the radar altimeter as the primary source of terrain clearance information and the FCC air

data function, GPS, and INS as backup altitude sources when radar altitude is invalid. Radar altitude

is considered invalid by GPWS above 4,950 feet AGL or at a pitch or angle of bank greater than 50°.

With valid radar altitude data, GPWS calculates terrain slope from inputs from the INS and the radar

altimeter. Over descending terrain, GPWS assumes the terrain descends indefinitely (until the system

senses a change in terrain slope). This mechanization allows for maximum protection while minimizing

nuisance warnings.

For the first 5 seconds after radar altitude becomes invalid (as indicated by a flashing ″B″ in the

HUD or RALT Xd out on the UFCD), GPWS provides no CFIT protection. After 5 seconds, the system

enters ″COAST″ mode for a period of up to 2 minutes. While in COAST mode, GPWS calculates an

estimate of the aircraft current height above terrain. COAST mode can only be enabled while the

aircraft is not transonic and was over flat terrain (defined as slope less than 2°). CFIT warnings can still

be generated while in COAST mode. If the aircraft was transonic or was not over flat terrain when

radar altitude data was lost, GPWS transitions into the BYPASS mode. In the BYPASS mode, no

CFIT warnings are generated. Full protection is resumed from both modes when valid radar altitude

data is restored.

2.17.5.2 Altitude Required For Recovery Calculations. GPWS calculations for altitude required for

recovery include the loss of altitude due to persistency timers, pilot reaction time, time to roll wings

level, target g-onset rate, and steady state dive recovery time. GPWS pilot reaction time varies

depending on flight conditions but is a minimum of 0.5 second in the GPWS LAT envelope (±30° AOB,

0 to 30° dive, 450 to 560 KCAS). Pilot reaction time is reduced in the GPWS LAT envelope, where pilot

situational awareness is typically good, in order to reduce false warnings. Time to roll wings level is

based on a ½ to ¾ lateral stick displacement roll at 1g. Target g-onset rate is 80% of the available

g-onset rate up to (1) 5g/sec (less than 400 KCAS or greater than 30° AOB) or (2) 6g/sec (greater than

400 KCAS and less than 30° AOB). Steady state dive recovery time is based on a target sustained-g of

80% of g-available up to (1) 5g (less than 400 KCAS or greater than 30° AOB) or (2) 6g (greater than

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400 KCAS and less than 30° AOB). Regardless of which category applies, these g-onset

rates and sustained-g levels require an aggressive pilot response.

2.17.5.3 CFIT Protection Provided.

Above 150 feet AGL -

Above 150 feet AGL, GPWS continuously calculates the altitude required to recover. A CFIT

warning is issued if the altitude required to recover plus a variable safety buffer and an added terrain

clearance altitude is greater than the current altitude above terrain. The terrain clearance altitude

varies between 30, 50, and 90 feet, depending on flight conditions.

Below 150 feet AGL -

Below 150 feet AGL, GPWS transitions to provide warnings of CFIT conditions related to takeoff

and landing. These warnings are based on (1) the time since a WoffW transition (takeoff or T&G) or

a waveoff and then (2) a combination of landing gear position, airspeed, altitude, and sink rate. GPWS

defines a waveoff as 1000 fpm rate of climb for more than 5 seconds while below both 500 feet AGL and

200 KCAS. If the following sets of conditions are valid for greater than 0.3 seconds when the aircraft

altitude is less than 150 feet, a CFIT warning is provided. The CFIT warning is cancelled when the

condition no longer exists for 0.3 seconds.

1. Less than 60 seconds after WoffW or a waveoff:

a. Floor Altitude - less than 90 feet AGL and greater than 250 KCAS.

b. Takeoff Sink Rate - less than 150 feet AGL, less than 250 KCAS, greater than 300 fpm sink.

2. More than 60 seconds after WoffW or a waveoff:

a. Floor Altitude - less than 90 feet AGL and greater than 200 KCAS.

b. Check Gear - less than 150 feet AGL, less than 200 KCAS, descending, and landing gear not

down.

c. Landing Sink Rate - less than 150 feet AGL, less than 200 KCAS, landing gear down, and an

excessive sink rate. The allowable sink schedule varies from a maximum of 2,040 fpm to a

minimum of 1,488 fpm based on altitude and GW.

d. Bank Angle - less than 150 feet AGL, less than 200 KCAS, greater than 45° AOB for one

second.

Below 150 feet AGL, GPWS does not directly account for the recovery

capabilities of the aircraft. Therefore, recovery may not be possible

following a warning under extreme flight conditions.

2.17.5.4 GPWS Warning Cues. GPWS provides distinctive, clear, unambiguous and directive visual

and aural cues to the aircrew for each potential CFIT condition.

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Figure 2-43. GPWS HUD Roll Warning Cues

2.17.5.4.1 GPWS HUD Recovery Arrow. The GPWS visual warning cue is a steady arrow located in

the center of the HUD. See figure 2-43. The HUD recovery arrow is always perpendicular to the horizon

and points in the direction of pull required for recovery. The HUD recovery cue is displayed

simultaneously with all voice warnings except CHECK GEAR. The HUD recovery arrow remains

displayed until GPWS calculates that a CFIT condition no longer exists.

2.17.5.4.2 GPWS Voice Commands. Refer to figure 2-44 for GPWS aural warning cues.

Voice commands automatically transition to the appropriate command for the current stage of

recovery (e.g., ROLL OUT transitions to PULL UP when AOB becomes less than 45°). The voice

commands are terminated when the appropriate recovery maneuver is initiated (e.g., a PULL UP is

initiated within 0.5 g of the GPWS calculated target-g).

• In addition to following the voice commands, additional pilot action

may be required to avoid an unrecoverable situation (e.g., aft stick

with a POWER call or power addition/subtraction with a PULL UP

call.)

• GPWS voice alerts are delayed if other voice alerts are currently being

transmitted.

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GPWS Warning

Aural Cue

Repetition Rate

Condition

Excessive bank angle

″ROLL LEFT (RIGHT), ROLL LEFT (RIGHT)″

2 seconds

Excessive take-off

″POWER, POWER″

2 seconds

sink rate

Excessive landing

″POWER, POWER″

2 seconds

sink rate

Gear-up landing

″CHECK GEAR″

8 seconds

ALDR or floor

″POWER, POWER″ for airspeed <210 KCAS and

2 seconds

altitude

AOB ≤45°

″ROLL LEFT (RIGHT), ROLL LEFT (RIGHT)″

2 seconds

for AOB >45°

″PULL UP, PULL UP″ for all other flight

2 seconds

conditions

Figure 2-44. GPWS Aural Cues

2.17.5.5 Areas of Limited CFIT Protection. Areas where CFIT protection is considered limited are

as follows:

1. In the COAST mode (5 to 120 seconds outside the valid RALT envelope).

2. Over rising terrain of greater than 2° slope (GPWS is inhibited to prevent nuisance warnings).

3. Within the GPWS LAT envelope where allowable pilot reaction times have been reduced (±30°

AOB, 0 to 30° dive, 450 to 560 KCAS).

4. Below 150 feet AGL in the landing phase (less than 200 KCAS) where warnings are designed only

to prevent hard landings.

At certain high speed, high gross weight conditions, overriding the

g-limiter may be required for recovery from dives greater than 50° and

will likely be required for dives between 10 and 25°.

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2.17.5.6 Areas of No CFIT Protection. Areas of no protection are as follows:

1. Loss of air data or RALT, INS, or MC1 failed or off (non-AMCD aircraft), or either MC1 or MC2

failed or off (AMCD aircraft).

2. Less than 6 seconds after WonW.

3. Less than 5 seconds or greater than 120 seconds outside the valid RALT envelope.

4. Transonic flight (0.95 to Mach 1.04) outside the valid RALT envelope.

5. For 1.5 seconds after a break X is displayed.

6. After a waveoff until exceeding 1,000 fpm for 5 seconds.

7. Dives greater than 50° after 2 minutes above 5,000 feet AGL.

2.18 INSTRUMENTS

Refer to foldout section for cockpit instrument panel illustration. For instruments that are an

integral part of an aircraft system, refer to that system description in this section.

2.18.1 Standby Attitude Reference Indicator. The standby attitude reference indicator is a self-

contained electrically driven gyro-horizon type instrument. It is normally powered by the right 115

volts ac bus. If this power fails it is automatically powered by an inverter operating off the essential 28

volts dc bus. An OFF flag appears if both power sources fail or the gyro is caged. During caging the gyro

initially cages to 4° pitch and 0° roll regardless of aircraft attitude. After 3 to 5 minutes, the indicator

reads 0° pitch and 0° roll. Power should be applied for at least 1 minute before caging. The indicator

displays roll through 360°. Pitch display is limited by mechanical stops at approximately 90° climb and

80° dive. As the aircraft reaches either stop, the gyro tumbles 180° in roll. A needle and ball are at the

bottom of the instrument. A one needle width turn is 90° per minute.

2.18.2 Standby Airspeed Indicator. The standby airspeed indicator displays airspeed from 60 to 850

KIAS. It operates directly from left pitot and static pressure.

2.18.3 Standby Altimeter. The standby altimeter is a counter-pointer type. The counter drum

indicates altitude in thousands of feet from 00 to 99. The long pointer indicates altitude in 50-foot

increments with one full revolution each 1,000 feet. A knob and window permit setting the altimeter

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Figure 2-45. Angle of Attack Indexer

to the desired barometric setting. This setting is also used by the flight control computers. The standby

altimeter operates directly from left static pressure.

2.18.4 Standby Rate of Climb Indicator. The standby rate of climb indicator displays vertical speed

on a scale from 0 to ±6,000 fpm and operates directly from left static pressure.

2.18.5 Standby Magnetic Compass. A conventional aircraft magnetic compass is mounted on the

right windshield arch in the front cockpit and in the rear cockpit in Lots 21 thru 25.

2.18.6 Angle Of Attack Indexer. The angle of attack indexer is mounted to the left of the HUD. It

displays approach angle of attack (AOA) with lighted symbols; corresponding AOA indications are

shown on the HUD (see figure 2-45). The indexer operates with the landing gear down and locked and

weight off the gear. The lighted symbol(s) flash if the arresting hook is up and the hook bypass switch,

on the left vertical panel, is in CARRIER. The symbols will not flash with the arresting hook up and

the hook bypass switch in FIELD. The switch is solenoid held to FIELD and automatically goes to

CARRIER when the arresting hook is lowered or aircraft power is removed. The AOA indexer knob on

the HUD controls dimming of the symbols. All symbols light when the lights test switch on the interior

lights control panel is held to TEST.

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2.19 AVIONICS SUBSYSTEM

The avionics subsystem combines the integration and automation needed for operability with the

redundancy required to ensure flight safety and mission success. Key features of the system include

highly integrated controls and displays, inertial navigation set with carrier alignment capability, and

extensive built in test capability. The avionics subsystems operate under the control of two mission

computers with primary data transfer between the mission computers and the other avionics

equipment via the mux buses.

2.19.1 Mission Computer System. The mission computer system consists of two digital computers

(MC1 and MC2) which are high speed, stored program, programmable, general purpose computers

with core memory. Computer de-selection is made with the MC switch on the MC/HYD ISOL panel.

With non-AMCD aircraft, the two mission computers interconnect with the primary avionics

equipment on the avionics multiplex (mux) buses. MC1, referred to as the navigation computer,

performs processing for navigation, control/display management, aircraft built in test (BIT), status

monitoring operations and backup for MC2. MC2, referred to as the weapon delivery computer,

performs processing for air-to-air combat, air-to-ground attack, and tactical control/display.

There are six avionics mux bus channels with redundant paths (X and Y) for each channel.

The mission computer:

1. Computes and controls the data sent to the cockpit displays,

2. Computes missile launch and weapon release commands,

3. Provides mode control and options for various avionics systems,

4. Generates BIT initiate signals to and equipment operational status from various avionics

systems.

With AMCD aircraft, the front and rear DDIs are driven directly by the MC over a high speed

interface bus, not by avionics mux bus commands. The HUD is driven directly by redundant

connection to either MC. MC1 drives the front and rear LDDIs and HUD while MC2 drives the front

and rear RDDIs and HUD. When an MC is off or non-functional, the displays driven by that MC show

a green square in the center of the display. Each MC provides the same level of functionality in the

single MC backup mode of operation.

On AMCD aircraft, with both MCs inoperative and the left generator operative, the SDC provides

a limited HUD format on the front MPCD/UFCD, prevents the FADEC and ECS controller from going

into default mode operation, and provides left/right ATS cautions when necessary. See Chapter 25

Backup/Degraded Operations for a description of SDC Backup Mode.

AMCD II mission computers are used on aircraft with the aft cockpit 8 x 10 display installed. The

8 x 10 display receives its primary signal through MC2, and is inoperative with computers prior to

AMCDII. The AMCD II computer provides digital video color capability to the 8 x 10 display via the

Fiber Channel Network Switch (FCNS) and High Speed Video Network (HSVN).

The computers receive inputs for navigational data and steering command computations from the

inertial navigation system, electronic flight control system, multipurpose display group, TACAN, and

backup attitude and the navigation system. The computers control display symbology and information

presented to the pilot by the multipurpose display group.

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2.19.1.1 MC Switch. The MC switch, located on the outboard edge of the aft left console, is used to

manually turn OFF either of the two mission computers, MC1 or MC2.

1 OFF Removes power to MC1.

NORM Both MC1 and MC2 are powered with ac electrical power available.

2 OFF Removes power to MC2.

On non-AMCD aircraft with WonW, both engines will spool up from ground idle to flight idle, if

power is removed from MC1 (MC1 switch in 1 OFF).

2.19.1.2 Mission Data Entry. Mission data (date and flight number) can be manually loaded into the

mission computer for data recorder documentation. Data is entered by performing the following:

1. On the DDI - Press MENU, MUMI, then ID.

2. On the UFCD - Enter Julian Date (DATE).

3. On the UFCD - Enter Flight Information (FLT).

Mission data can be manually loaded into the mission computer through the Memory Unit Mission

Initialization (MUMI) display or automatically loaded into the mission computer through the Data

Storage Set (DSS). The DSS consists of the Memory Unit (MU) and the Memory Unit Mount (MUM)

and provides memory storage for aircraft parameters, maintenance data, and avionics initialization

data. The DSS receives, stores, retrieves, and transmits data with the mission computer.

2.19.1.2.1 Mission Initialization. The MU provides the capability to load the following mission

initialization files: HARM, RADAR, MU ID, TACAN, WYPT/OAP, Combined Interrogator Tran-

sponder (CIT), Sequential Steering (S/S), data link/ID, Overlay Controlled Stores (OCS), bomb wind

data, and ALR-67 display resets. The S/S file can have a 15 point sequence consisting of Geographic

Reference Points (GEOREF), GPS waypoints, and almanac data initialization files. Loading is done at

aircraft power up or when MUX communication is lost for more than 1 second and regained. If MUX

communication is not regained, a MU LOAD caution is displayed and an AV MUX error message is

displayed on the MUMI display. Manual loading may be done using the MUMI display.

2.19.1.2.2 MUMI - Memory Unit Mission Initialization Format. The MUMI format (see figure 2-46)

is accessible from the SUPT MENU and with WonW provides a visual indication of mission

initialization files loaded from the MU. If the MU directory indicates that no user files are present, the

MU ID displays NO IDENT. When the MU directory indicates a user file is present, MC1 displays the

option. When the option is selected and the file is being read by MC1, the option is boxed. If the read

is successful, the file is loaded and the option is unboxed. When a file is present and errors have

resulted from reading the file, the following occurs:

1. The MU ID displays NO IDENT.

2. The applicable load error is displayed (HARM, RDR, TCN, WYPT, S/S, OCS, GPS WYPT, GPS

ALM, ALR 67, WIND, DL13, or CIT).

3. MC1 sends the appropriate maintenance code to the SDC.

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Figure 2-46. MUMI Display

4. If WonW, an MU LOAD caution is displayed on the DDI.

2.19.1.2.3 CRYPTO Switch. Setting the intercommunications amplifier control CRYPTO switch to

the ZERO position sends an erase signal to the MU. This causes the MU to erase all data stored

between predetermined memory locations.

2.19.1.2.4 Erase and Hold Data. The erase controller (EC) within MC1 provides the capability to

automatically or manually erase, or inhibit erasing, of classified data contained in the MU, SMS, MC1,

and MC2.

NOTE

With AMCD aircraft, during an ERASE, the DDI controlled by the

MC undergoing the erase will flash STANDBY, then briefly display a

green square, then flash STANDBY until the erase is complete.

When the EC determines classified mission initialization files have been read from the MU, the EC

classified data management system is activated. When activated, MC1:

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1. Displays the HOLD and ERASE options on the MUMI display.

2. Displays the CDATA advisory.

3. Sends applicable maintenance code(s) to the SDC.

2.19.1.2.5 Automatic Erase. The MU, SMS, MC1, and MC2 automatically erase classified data

when all of the following criteria are met.

1. Airspeed is less than 80 KCAS.

2. Left and right engine THA less than 29°.

3. Transition from WoffW to WonW.

4. Pilot does not select erase inhibit (HOLD) or MC SUSPEND options.

NOTE

Automatic erase can be inhibited by selecting the HOLD pushbutton

option.

5. HOLD boxed with MU displayed prevents automatic erase of the MU.

6. HOLD boxed with ALL displayed prevents automatic erase of all units (MU, armament

computer, MC1, and MC2).

The EC commands the MU and the armament computer, then MC1 and MC2 to erase. The MC

ERASE IN XX SEC countdown timer starts (60 seconds). During the countdown, an MC SUSPEND

pushbutton option is displayed. The MC SUSPEND option may be toggled between boxed (selected)

and unboxed (deselected). When the timer reaches zero and the MU and armament computer have

finished erasing, the decision to continue erasing the remainder of MC1 and MC2 depends on the MC

SUSPEND option being deselected (unboxed); when deselected, the remaining erase of MC2 and MC1

is completed.

NOTE

With AMCD aircraft, during an ERASE, the DDI controlled by the

MC undergoing the erase will flash STANDBY, then briefly display a

green square, then flash STANDBY until the erase is complete.

Automatic erase is also initiated by pilot ejection. The state of the HOLD options is ignored during

pilot ejection.

2.19.1.2.6 Manual Erase. Manual erase is a two pushbutton process and is initiated by pressing the

ERASE pushbutton on the MUMI display. When the ERASE pushbutton is pressed, the option to

proceed with the erasure (ERASE) and the option to cancel the erase (CNX) replaces the HOLD and

ERASE options. Selecting the second ERASE option initiates erasure. While erase is in progress,

ERASE is boxed and erasing proceeds the same as automatic erase. When erasing is complete, the

ERASE pushbutton unboxes. While erase is in progress one of the following is displayed on the MUMI

display:

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1. ERASING - erasing of unit is in progress.

2. COMPLETE - erasing of unit is complete.

3. FAILED - unit failed to erase.

NOTE

With AMCD aircraft, during an ERASE, the DDI controlled by the

MC undergoing the erase will flash STANDBY, then briefly display a

green square, then flash STANDBY until the erase is complete.

When erase fails, the MC1 retains the MUMI ERASE and HOLD pushbutton options and displays

the ERASE FAIL caution on the DDI. When erasing is complete, MC1 removes the ERASE, HOLD,

and MC SUSPEND pushbutton options from the MUMI display, removes the CDATA advisory from

the display, and resets the applicable maintenance code(s).

2.19.1.2.7 Backup Erase Controller. If MC1 fails, backup erase capability is provided by MC2 by

providing an ERASE option on the HSI format. When the ERASE pushbutton is depressed, the option

to proceed with the erasure (ERASE) and the option to cancel the erasure (CNX) are provided.

Selecting the second ERASE option initiates erasure. While ERASE is in progress, the ERASE option

is boxed; however, additional cuing is not provided. In backup mode, pilot ejection is the only

automatic erase provided. With AMCD aircraft, the MUMI display for the MC being erased disappears

while the erase is in progress.

2.19.2 Master Modes. There are three master modes of operation: navigation (NAV), air-to-air

(A/A), and air-to-ground (A/G). Controls, displays, and the avionics equipment operation are tailored

as a function of the master mode selected. The navigation master mode is entered automatically when

power is applied to the aircraft, when the air-to-air or air-to-ground modes are deselected, when the

landing gear is lowered, when the SPIN mode activates, or when the aircraft has WonW and the THA

is greater than 27°. The A/A master mode is entered either by pressing the A/A master mode button

alongside the left DDI or by selecting an A/A weapon with the A/A weapon select switch on the control

stick. The A/G master mode is selected by pressing the A/G master mode button. The selection is

performed by the stores management set (SMS), and the SMS identifies the selected master mode to

the mission computer.

2.19.2.1 Steering Information. The sources of steering information available in the NAV master

mode are waypoint, TACAN, instrument landing system, and data link. The data link modes available

in the NAV master mode are vector and automatic carrier landing. TACAN and waypoint steering are

mutually exclusive; selecting one automatically deselects the other. Data link, ILS, and TACAN (or

waypoint) steering can be provided simultaneously. The ACL mode is selectable only in the NAV

master mode and the vector mode is available in all master modes. Steering information is used by the

Automatic Flight Control System to provide coupled steering options.

2.19.3 Cockpit Controls and Displays. The cockpit controls and displays which are used for

navigation operation are on the multipurpose display group.

2.19.4 Multipurpose Display Group. The multipurpose display group consists of the right and left

digital display indicators (DDIs), the multipurpose color display (MPCD), 8 x 10 display, the digital

map set (DMS), the head-up display (HUD), the CRS (course) set switch, the up front control display

(UFCD) and the HDG/TK (heading/ground track) set switch. The multipurpose display group

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A1-F18EA-NFM-000

presents navigation, attack, and aircraft attitude displays to the pilot. The multipurpose display group

converts information received from the mission computer system to symbology for display on the DDIs,

the MPCD, the UFCD, and the HUD. The HUD camera records the outside world and HUD

symbology. The left and right DDIs, 8 x 10 display and the MPCD contain pushbuttons for display

selection and selection of various equipment operating modes. The UFCD is an active matrix liquid

crystal display with an infrared (IR) touchscreen for operator inputs. Refer to Part VII for the

operation of each component.

2.19.4.1 CRS Set Switch. The course set switch, located on the main instrument panel on the video

record panel, manually sets the desired course on the HSI display.

2.19.4.2 HDG/TK Set Switch. The heading/ground track set switch located on the main instrument

panel on the video record panel, manually sets the heading marker on the desired heading/ground track

on the HSI display.

2.19.4.3 DDIs. The left and right DDI (LDDI/RDDI) are physically and functionally interchange-

able, giving the ability to display desired information on either indicator. With non-AMCD aircraft, the

DDIs also provide symbology to the HUD. The left indicator is used primarily for stores status, built

in test status, engine monitor, caution, and advisory displays. The right indicator is normally used for

radar and weapon video displays.

(Non-AMCD aircraft) The DDIs are NVG compatible and display three colors (red, yellow, and

green) for stroke information. A monochrome version of the digital map can be selected on the left DDI.

Either DDI can provide raster generation for the HUD.

NOTE

It is possible that a transient condition may cause the displays to

blank or provide an erroneous display on the left or right DDI, or

HUD. The problem may be cleared by manually cycling the power to

the left or right DDI.

(AMCD Aircraft) The DDIs have full color capability in all display modes and are NVG compatible.

2.19.4.3.1 Brightness Selector Knob - Non-AMCD Aircraft. Placing this rotary knob to OFF

prevents the DDI from operating. Placing the knob to NIGHT provides a lower brightness control

range and no automatic contrast control. Turning the knob to DAY provides a higher brightness

control range.

2.19.4.3.2 BRT Control Knob - Non-AMCD Aircraft. This knob varies the intensity of the DDI

presentation.

2.19.4.3.3 Brightness Knob - AMCD Aircraft. Placing this rotary knob to OFF prevents the DDI

from operating. When turned on, rotating the knob clockwise increases display brightness, while

rotating the knob counterclockwise decreases display brightness.

2.19.4.3.4 GAIN Control - AMCD Aircraft. This three-position rocker switch affects the existing

gray-scale, shifting the backg1round brightness up or down, with no impact on displayed symbology.

The center position is off, while the up arrow increases backg1round brightness and the down arrow

decreases backg1round brightness. The up and down arrows have momentary and scroll functionality,

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A1-F18EA-NFM-000

depending on how long they are held. Current values are temporarily displayed when switches are

pressed.

2.19.4.3.5 CONT Control. This knob (non-AMCD aircraft), or three-position rocker switch (AMCD

aircraft), varies the contrast between symbology and the dark backg1round on any level of brightness.

With AMCD aircraft, when the contrast rocker is pressed, a graphical and numeric representation of

the current setting is momentarily displayed on the DDI.

2.19.4.3.6 DDI Pushbuttons. There are 20 pushbuttons on each DDI which are used to select the

function and the mode for proper indicator display.

2.19.4.3.7 MENU Formats. There are two MENU formats (figure 2-47), TAC (tactical) and SUPT

(support) through which display selections can be made. The two menu formats can appear on either

the DDI, the UFCD, 8 x 10 display or MPCD.

The TAC MENU is indicated by the word TAC appearing just above the MENU option. The TAC

MENU is selected by actuating the MENU option on any format other than the TAC MENU format.

The SUPT MENU is indicated by the word SUPT appearing just above the MENU option. The SUPT

MENU is selected by actuating the MENU option on the TAC MENU. With MC OFP H2E AND UP,

the current time of day, consisting of two minute digits and two second digits appear in place of the

MENU legend (cold start default). Time of day is also turned on/off with the TISM option on the

Engine Format. MENU pushbutton functionality is unchanged.

The TAC MENU allows selection of the following formats: AZ/EL, HUD, RDR ATTK, STORES,

FLIR, NFLIR, DL13, HARM, A/G WPN, JSOW, SA, IMAGE, CAS, EW, and TGT DATA. The SUPT

MENU allows selection of the following formats: ADI, HSI, HMD, NETS, GPS, MIDS, ROE/IFF

PROG, BIT, CTT, MUMI, CHKLST, ENG, FCS, UFC BU, FPAS, and FUEL.

Some of the options on the MENU formats are conditional and are not always displayed. NFLR,

FLIR, LST, and CAM are listed only if the equipment is communicating with the mission computer.

HARM DSPLY is displayed when HARM is on board and CLC communicating. A/G missile display

(WEDL DSPLY, MAV DSPLY, etc.) is displayed when the MC has determined from the armament

control processor set that a weapon station has been selected which contains one of these weapons.

In Non-AMCD Aircraft, if the navigation computer (MC1) is not on, only HSI is displayed on the

SUPT menu. If the weapon delivery computer (MC2) is not on, the SUPT menu remains unchanged

and the TAC menu does not display the STORES, FLIR/LST/CAM or NFLR, AWW-13, HARM, or

A/G displays. If both mission computers are off or not communicating with the display, the DDI,

MPCD, and UFCD display only a flashing STANDBY in the center of the screen.

In AMCD Aircraft, if either mission computer is off or failed (MC Backup Mode), the options

available on the displays are the same regardless of which computer is failed. MC1 is no longer referred

to as the navigation computer and MC2 is no longer referred to as the weapons computer. If MC1 is

failed or not communicating, the left DDIs in both cockpits will either flash STANDBY or have a green

square present in the middle of the displays. If MC2 is failed or not communicating with the display,

the right DDIs in both cockpits will either flash STANDBY or have a green square present in the

middle of the displays. The 8 x 10 display will not be capable of displaying format symbology or video.

The following options are available in MC Backup mode: HUD, EW, FCS, MIDS, MUMI, ENG, UFC

BU, FUEL, ADI, and HSI. If A/A master mode is selected, the RDR ATTK, SA, and AZ/EL formats

are available as well. No A/G displays are supported in MC Backup mode. If both mission computers

are off or failed, a backup HUD format (driven by the SDC) is displayed on the UFCD and MPCD in

the front cockpit.

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A1-F18EA-NFM-000

Figure 2-47. MENU Format

2.19.4.3.8 Electronic Attitude Display Indicator (EADI). The electronic attitude display indicator is

available for display on the left or right DDI as an alternative to the attitude display on the HUD

(figure 2-48). The EADI display is selected by selecting the ADI option on the SUPT MENU. The pitch

ladder is displayed in 10° increments. A small circle is displayed on the ball to represent the zenith and

a circle with an inscribed cross is displayed to represent the nadir. A turn indicator which displays FCS

yaw rate is provided below the ball. A standard rate turn (3°/second) is indicated when the lower box

is displaced so that it is under one of the end boxes.

Selecting the INS or STBY options at the bottom of the display determines the source of attitude

information used to generate the display. Upon power-up with WonW, the EADI attitude initializes to

STBY (STBY boxed), thus using the standby attitude reference indicator for attitude source

information. Selecting the INS option (INS boxed) uses attitude information provided by the INS.

Selection of INS or STBY on the EADI does not change the source of attitude data for the HUD.

Airspeed and altitude are displayed in boxes at the top left and right. Altitude source is displayed

to the right of the altitude box and the vertical velocity is displayed above the altitude box. When ILS

is selected the deviation needles are displayed in reference to the waterline symbol. The ILS needles

are yellow when COLOR is selected on the Attack display.

2.19.4.3.9 Shaped Attitude Display Indicator. When the ADI option is selected on the SUPT menu,

a shaped attitude display indicator is available on the MPCD, UFCD, or DDI and 8 x 10 display

(AMCD aircraft). The pitch ladder is displayed in 10° increments, with ±30° in pitch being displayed

when the gear is up. With the gear down, the pitch ladder is displayed in 5° increments with ±15° in

pitch being displayed. A turn indicator is provided below the ball. A standard rate turn (3°/second) is

indicated when the lower box is displaced so that it is under one of the end boxes.

On the MPCD and 8 x 10 display, the shaped ADI is black above the horizon to represent sky, and

green below the horizon to represent the ground. On the UFCD, the shaped ADI is unshaded above the

horizon (sky) and shaded green below the horizon (ground).

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ORIGINAL

A1-F18EA-NFM-000

Figure 2-48. Electronic Attitude Display Indicator

Selecting the INS or STBY options at the bottom of the display determines the source of attitude

information used to generate the display. Upon power-up with WonW, the ADI attitude initializes to

STBY (STBY boxed), thus using the standby attitude reference indicator for attitude source

information. Selecting the INS option (INS boxed) uses attitude information provided by the INS.

Selection of INS or STBY on the ADI does not change the source of attitude data for the HUD.

Airspeed and altitude are displayed in boxes at the top, altitude source is displayed to the right of

the altitude box and vertical velocity is displayed above the altitude box. When ILS is selected the

deviation needles are displayed in reference to the waterline symbol. The ILS needles are yellow when

COLOR is selected on the attack display.

2.19.4.4 MPCD - Multipurpose Color Display. The MPCD is an NVG compatible digital display

capable of providing any MENU selectable format except the video on the A/G radar display (figure

2-50). The MPCD drives itself using information received on the MUX.

The BRT knob controls the overall video and symbology brightness and also acts as the power

control for the MPCD and UFCD; a detented OFF position at the extreme counterclockwise position.

The CONT knob adjusts the video contrast of the MPCD display. The day/night mode is controlled

by the day/night/NVG switch on the interior lights panel.

2.19.4.4.1 Standby Indication. A flashing STANDBY indication is provided in the center of the

display on initial power-up and when there is invalid mux communication with the MC. If the

STANDBY condition persists for a few seconds, cycling power to the MPCD may return the MPCD

and UFCD to normal operation.

2.19.4.4.2 Brightness and Contrast Controls. The brightness control is used to adjust the overall

brightness of the MPCD display surface, allowing symbology and video to be adjusted. The full

counterclockwise position of the brightness control knob shuts off power to both the MPCD and the

UFCD. Selecting ON powers both the MPCD and the UFCD.

The contrast control is used to adjust video contrast.

2.19.4.4.3 Symbology Rocker Switch. The symbology (SYM) brightness rocker switch is used to

adjust the brightness of the symbology without affecting the brightness of the MPCD video.

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ORIGINAL

A1-F18EA-NFM-000

Symbology change feedback in provided in the left center of the display by a digit from 0 to 9. This

feedback is provided while the rocker switch is pressed and for 5 seconds after the rocker switch is

released.

2.19.4.4.4 DDI Formats on MPCD. Every display format is available on the L or R DDI. A/G radar

symbology and monochrome video is available on the MPCD. A/G video is not available.

2.19.4.4.5

MPCD Formats. With non-AMCD aircraft, if two different map formats are provided in

the front and rear cockpits, only one displays the map; the other provides only symbology. The priority

of the format which displays the map is HSI, followed by SA.

With AMCD aircraft, if two different formats are requesting MAP video underlay, only one displays

the video underlay; the other provides only symbology. MAP video display priority is determined by

two things: 1) which display surface is requesting the MAP underlay and 2) which format is requesting

the MAP underlay. Highest display surface priority is given to the MPCD followed by the DDIs. The

SA format is given the highest priority, followed by the HSI format, for formats requesting the MAP

underlay. Once the MAP is displayed for any format, (SA, HSI), any display surface with that same

format selected will also have the MAP underlay displayed. The MAP underlay is only supported on

the SA or HSI formats.

2.19.4.4.5A 8 x 10 Display. On aircraft LOT 26 AND UP, the 8 x 10 display replaces the aft MPCD.

The 8 x 10 display has a full color Active Matrix Liquid Crystal Display surface and 30 pushbuttons

used for operator inputs. Only 20 buttons are functional, and they duplicate the MPCD/DDI menu

layouts. The top row outboard buttons and top four buttons on each side are not used. The brightness

control switch adjusts the video brightness. The full counterclockwise position, a detented OFF

position, removes power from the 8 x 10 display and the aft cockpit UFCD. Two rocker switches

provide Gain and Contrast control. A numeric value displayed near the switches is used to adjust the

gain and/or contrast to a desired level. Pushbuttons are backlit, controlled by the INST PNL knob on

the aft cockpit interior lights panel. The following formats and related sublevel formats are available:

TAC Menu, SUPT Menu, HSI, EW, SA, TFLIR and BIT. Figure 2-49 shows symbology placement and

map coverage for the 8 x 10 display. See figure FO - 4, foldout section for aft cockpit arrangement with

the 8 x 10 display installed.

The 8 x 10 display is connected directly to MC2. The computer recognizes whether an 8 x 10 display

or an MPCD is installed and uses the corresponding software that matches the configuration at

startup. This affects the UFCD controls and functions in addition to the 8 x 10 display. When MC2 is

offline or during initialization, the 8 x 10 display does not show a format or process bezel inputs. A

flashing STANDBY indication is provided in the center of the 8 x 10 display when MC2 is not

communicating properly with the 8 x 10 display. A standby display pattern (approximately a 1- by

1-inch green square) is present when the Image Transfer Bus (ITB) has lost symbology, video, or a

combination, and the selected format is affected. When interface between MC2 and 8 x 10 display is

lost, only symbology is displayed (via ITB). There is no pushbutton feedback, video control, or BIT

reporting when interface is degraded. The day/night mode of the 8 x 10 display (and rear UFCD) is set

via the DAY/NITE/NVG switch on the forward cockpit interior lights panel through MC2.

In the Situational Awareness (SA) format, the cursor (captains bars) may be slewed over the entire

8 x 10 display surface. Because this area is larger than the DDIs/MPCD, if another display is showing

the same format and the cursor on the 8 x 10 display is slewed beyond the center area, the cursor on

the DDI/MPCD is limited. When this happens, the cursor on the DDI/MPCD displays flashes

continuously. When a SA format has DC priority assigned, the initial cursor position is determined by

the cockpit that initiates the slewing action.

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A1-F18EA-NFM-000

2.19.4.5 HSI Display Symbology. Basic HSI symbology such as the compass rose, ground track

pointer, lubber line (for magnetic heading), true airspeed readout, ADF bearing pointer, groundspeed

readout, and aircraft symbol are not described. These symbols are shown in figure 2-50.

Display format lines, just below the top row of pushbutton labels, indicate the formats being

displayed on the left, center, right and UFCD displays in both the cockpit (left side of HSI display) and

rear cockpit (right side of HSI display).

Radar target and GEOREF symbols are described in A1-F18EA-TAC-Series. The following

paragraphs describe unique navigation symbology. Refer to part VII for a description of how these

symbols are integrated with the navigation system:

1. Waypoint/OAP data. Data for the current steer to waypoint/OAP is displayed on the upper right

corner of the HSI. Waypoint/OAP data consists of bearing, range, and TTG (time-to-go) up to

8:59:59 based on distance and ground speed. When a waypoint/OAP or offset to the OAP is

designated (becomes a target), this data relates to the target. When a waypoint is a waypoint that

was transferred from the GPS, an ID code is displayed under the waypoint data.

2. TACAN data. TACAN data is displayed on the upper left corner of the HSI. TACAN data

consists of bearing, range (slant range), TTG (based on distance and present ground speed), and the

station identifier.

3. Waypoint/OAP symbology. Waypoint/OAP symbology consists of a waypoint/OAP symbol and a

bearing pointer and tail. The waypoint/OAP symbol indicates the position of the selected

waypoint/OAP relative to the aircraft symbol. The waypoint/OAP bearing pointer and tail are

displayed inside the compass rose and indicate bearing to the selected waypoint/OAP. Waypoint/

OAP symbology is displayed whether or not waypoint/OAP steering is selected. When the selected

waypoint/OAP is outside the HSI range scale, the waypoint/OAP symbol does not appear, but the

bearing pointer and tail appears. When a waypoint/OAP is designated, the waypoint/OAP symbol

and circle inside the pointer change to a diamond shape. The offset symbol appears when steering

is to an OAP. The offset symbol indicates the position of the offset relative to the OAP.

4. TACAN symbology. TACAN symbology consists of a TACAN symbol, and TACAN bearing

pointer and tail. The TACAN symbol indicates the position of the TACAN station relative to the

aircraft symbol. The TACAN bearing pointer and tail are located outside of the compass rose and

indicate bearing to the TACAN station. When the TACAN station is outside the HSI range scale, the

TACAN symbol does not appear but the bearing pointer and tail appear. When TACAN range

becomes invalid the TACAN symbol is not displayed.

NOTE

TACAN symbology displayed inside of the compass rose is filtered to

prevent excessive movement of TACAN symbols due to RF

interference. However, the ″fly−to″ needle displayed in the HUD with

TACAN steering selected is not filtered and represents the raw data

received by the TACAN. As a result, for brief periods of time, the

HUD and HSI may display conflicting information regarding aircraft

position with respect to the selected TACAN course line. Aircrew

should use HUD displayed TACAN information when conducting

TACAN approaches.

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A1-F18EA-NFM-000

5. Heading select/ground select marker and readout. The heading select/ground select marker is

maneuvered along the periphery of the compass rose using the HDG/TK switch. The digital readout

of the selected heading is located on the lower left corner of the HSI. The heading select/ground

select marker and digital readout are part of the heading select/ground track select mode of the

autopilot.

Course line arrow and readout. The course line arrow indicates the selected course to the

waypoint/OAP or TACAN station. The course is selected using the CRS switch. The digital readout

of the selected course is displayed on the lower right corner of the HSI. The course line arrow is not

displayed when TACAN range is invalid.

6. TDC assignment symbol. The TDC assignment diamond is displayed on the upper right corner

of the HSI. This symbol indicates that the TDC is assigned to the HSI. The TDC assignment

diamond indicates control is assigned to both cockpits. Other symbols indicate cockpit (d) or rear

cockpit (e) TDC control and SLEW control. SLEW is done by actuating the sensor control switch

AFT while in the NAV or A/G master mode. The word SLEW is displayed in the TDC assignment

diamond position when the SLEW option is active.

7. Coupled steering symbology. CPL and the source of the steering information is displayed on

either side of the aircraft symbol in the center of the HSI display whenever the flight control system

is coupled in azimuth to a steering source. Steering source can be WYPT, TCN, or SEQ#. The couple

cue flashes for 10 seconds and is removed if the steering signal is lost or becomes invalid.

8. Sequential steering lines. The sequential steering lines are displayed when a sequence is entered

and when one of the sequence options (SEQ1, SEQ2, SEQ3, or SEQL) is boxed. The sequential

steering lines are available for display in all HSI modes and range scales. Sequential steering lines

are not displayed at power up with WonW and are removed when: magnetic heading is invalid,

aircraft position is invalid, or map slew is selected.

9. Time of day. Zulu time of day (ZTOD) or local time of day (LTOD) are displayed on the lower left

corner of the HSI. For aircraft that pass the FIRAMS real time clock power up BIT, ZTOD does not

need to be entered. For aircraft that do not pass the FIRAMS real time clock power up BIT, ZTOD

must be entered.

10. Groundspeed required. Groundspeed required appears below the current groundspeed readout.

Groundspeed required indicates the groundspeed required to a target based on entered ZTOD, time

on target (TOT), and the target.

11. Elapsed time (ET)/countdown (CD) time. ET and CD time are displayed on the lower right

corner of the HSI, however, only one of the timers can be displayed at a time. ET or CD timer must

be selected to be displayed. ET initializes to zero minutes and seconds and CD time initializes to six

minutes and zero seconds.

12. Aircraft heading. Aircraft heading is indicated on the compass rose. Aircraft heading and bearing

data can be selected as either magnetic or true. With true heading selected, the letter T appears

below the lubber line and the word TRUE appears below the selected scale readout. There is no

indication when magnetic heading is selected.

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ORIGINAL

A1-F18EA-NFM-000

Figure 2-51. HUD Controls

2.19.4.6 HUD - Head-Up Display. The HUD is on the center main instrument panel. The HUD is

used as the primary flight instrument, weapon status, and weapon delivery display for the aircraft

under all conditions. (Non-AMCD aircraft) The HUD receives attack, navigation, situation, and

steering control information from the left or right DDI symbol generators (under mission computer

control) and projects symbology on the combining glass for head-up viewing. (AMCD aircraft) The

HUD receives all symbology from MC1 or MC2. The HUD is electrically interfaced with the UFCD.

The HUD has NVG compatible raster display capability to allow it to display NFLR video. The

controls for the HUD are below the UFCD and are described in the following paragraphs. See figure

2-51.

2.19.4.6.1 HUD Symbology Reject Switch. The HUD reject switch is located on the HUD control

panel on the center main instrument panel. This switch is used to control the amount of symbology

displayed on the HUD.

NORM Displays full HUD symbology.

REJ 1

Removes the Mach, g, and peak-g indications, the bank angle scale and pointer, the air-

speed and altitude boxes, energy caret (landing gear down), and the ground speed

required cue.

REJ 2

Removes REJ 1 symbology and the heading scale, current heading caret, command

heading marker, NAV/TCN range, and the ET, CD, LTOD or ZTOD timer.

2.19.4.6.2 HUD Symbology BRT Control Knob. This knob is used to turn on the HUD and then

varies the display intensity.

2.19.4.6.3 HUD Symbology Brightness Selector Switch. This is a two-position toggle switch with

positions of DAY and NIGHT. Placing the switch to DAY provides maximum symbol brightness in

conjunction with the HUD symbology brightness control. With the switch set to NIGHT, a reduced

symbol brightness is provided in conjunction with the HUD symbology brightness control.

2.19.4.6.4 Black Level Control Knob. The black level control knob, located on the HUD control

panel, adjusts the NFLR video plus or minus ½ a shade of gray per increment when rotated.

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2.19.4.6.5 HUD Video Control Switch. The video control switch, located on the HUD control panel,

enables NFLR video display on the HUD with selectable polarity (white hot/black hot).

W/B Selects white hot/black hot polarity.

VID

Displays NFLR video in the HUD, if available.

OFF NFLR video off.

2.19.4.6.6 BAL Control Knob. The balance control, located on the HUD control panel, adjusts the

stroke brightness relative to the raster brightness. Rotating the switch from 12 o’clock towards the VID

position holds the brightness of the video (as set by the brightness control switch) and reduces the

brightness of the stroke symbology. The opposite is true when rotating the switch toward the SYM

position.

2.19.4.6.7 AOA Indexer Control Knob. This knob controls the brightness of the indexer lights.

2.19.4.6.8 ALT Selector Switch. The ALT selector switch, located on the HUD control panel, is used

to select the primary altitude source for display on the HUD and for use in the mission computer

(weapon systems calculations).

BARO Selects barometric altitude.

RDR Selects radar altitude.

2.19.4.6.9 ATT Selector Switch. The ATT selector switch, located on the HUD control panel, is used

to select the primary attitude source used for display in the HUD and in MC and FCC computations.

INS

Functions identically to the AUTO position.

AUTO Selects filtered INS data as the primary attitude source. The INS automatically reverts

to gyro mode, using unfiltered data if its processor fails. The MC automatically selects

the standby attitude reference indicator for attitude information if the INS fails com-

pletely.

STBY Selects the standby attitude reference indicator. The FCCs no longer use INS data and

the HIAOA advisory is displayed.

2.19.4.6.10 HUD Symbology. The following paragraphs describe HUD symbology as related to basic

navigation, steering (direct, great circle, courseline, and ILS), navigation target designation, advisories

and landing, see figure 2-52. Refer to part VII for a description of how these symbols are integrated into

the navigation system. Also, refer to section VII for unique ACL data link symbology. Refer to

A1-F18EA-TAC-Series, for symbology concerning the A/A and A/G master modes, weapons, RWR and

the data link vector mode.

1. Heading. The aircraft magnetic/true heading is indicated by the moving 30° heading scale. The

actual aircraft heading is directly above the caret/T symbol. The moving heading scale provides

trend information during turns. As the aircraft turns right, the scale moves from right to left.

Magnetic or true heading may be selected. Magnetic heading is indicated by a caret below the

heading scale. True heading selection is indicated by a T appearing below the current heading.

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2. Airspeed. Calibrated airspeed from the FCC is provided in the box on the left side of the HUD.

The tops of the airspeed and altitude boxes are positioned at the aircraft waterline, which is 4° up

from the optical center of the HUD.

HUD airspeed should normally read less than 50 knots while sitting still

on the ground. A reading of more than 50 knots on the ground may

indicate an Air Data failure.

3. Altitude. The altitude presented in the box on the right side of the HUD may be either barometric

altitude or radar altitude depending on the setting of the altitude switch on the HUD control panel.

When the altitude switch is in the BARO position, barometric altitude is displayed. When the

altitude switch is in the RDR position, radar altitude is displayed and is identified by an R next to

the altitude. If the radar altitude is invalid, barometric altitude is displayed and a B next to the

altitude flashes to indicate that barometric altitude is being displayed rather than radar altitude. An

X displayed next to the barometric altitude indicates that the altitude value may be inaccurate. The

ten thousand and thousand digits are 150% size numbers. The hundred, ten, and unit digits are

120% size numbers, except that below 1,000 feet they are 150% size.

4. Barometric setting. The barometric setting used by the air data function in the FCC is the value

set in the standby altimeter. When the barometer setting is changed on the standby altimeter, the

barometric setting is presented below the altitude on the HUD to provide a head-up baro-set

capability. The display remains for 5 seconds after the change is made. In addition, the baro-set

value is displayed and flashed for 5 seconds when the aircraft descends below 10,000 feet at an

airspeed less than 300 KCAS.

5. Angle of Attack. True angle of attack in degrees is displayed at the left center of the HUD. AOA

values displayed on the HUD are filtered and may slightly lag actual true AOA. Therefore, it may

be possible to trigger the AOA tone slightly prior to seeing the applicable limit AOA in the HUD.

The HUD AOA is generally driven by the FCS using the AOA probes. However, around 42° AOA

(and -9°), the HUD reverts to an MC-computed AOA based on INS data and winds. The HUD AOA

flashes indicating that the displayed HUD AOA may be inaccurate because the INS computed winds

have not been updated in the past three minutes. Low AOA, small bank angles, and small rates are

required to update the wind. With MC OFP H3E AND UP, the pitch trim AOA value is displayed

next to the ATC HUD advisory location while trimming and for two seconds after trimming with

WoffW and flaps HALF or FULL. The value is displayed with or without ATC engaged but is not

displayed with autopilot engaged.

6. Mach number. The aircraft Mach number is displayed immediately below the angle of attack.

7. Aircraft g. Normal acceleration of the aircraft is displayed immediately below the Mach number.

8. Peak aircraft g. A peak positive g indication is displayed on the HUD below the normal g when

a threshold of 4.0g is exceeded. The peak positive g display can be removed by cycling the clutter

reject switch to one of the reject positions.

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With a VVEL advisory displayed, sustained climbs and descents, such as

penetration from the marshal stack, can result in uncued (no cautions)

vertical velocity errors and a possible inaccurate velocity vector position.

Error magnitudes increase at slower airspeeds and lower altitudes. Errors

of up to 3° (actual flightpath 3° below the displayed velocity vector) have

been observed in the landing configuration. Three minutes of level flight

may be required to allow the INS to correct the vertical velocity errors.

NOTE

The VVEL advisory will be displayed if a GPS is not installed or if

masking prevents GPS positioning of a sufficient quality to ″aid″ the

INS vertical velocity loop.

The position of the velocity vector is limited to an 8° radius circle centered at the HUD optical

center. If the velocity vector reaches this limit during high angle of attack flight or large yaw and/or

drift angles, it flashes rapidly to indicate that it does not accurately indicate flight path (velocity vector

is HUD limited).

With GPS operating, if the INS velocity data becomes unreliable, the mission computer utilizes GPS

information. If INS velocity data becomes unreliable the mission computer utilizes FCC air data

function information and the last available wind data to compute the velocity vector and this degraded

velocity vector is indicated by a slow flashing of the symbol. In the NAV master mode, the velocity

vector may be caged to the vertical center line of the HUD by the cage/uncage switch on the throttle.

When it is caged, a ghost velocity vector is displayed at the true velocity vector position if that position

is more than 2° from the caged position. The flight path/pitch ladder and steering information are

referenced to the caged position. The ghost velocity vector flashes when limited. The flight path/pitch

ladder is referenced to the waterline symbol when the velocity vector is caged.

The velocity vector is automatically fixed at the horizon with WonW and ground speed less than 80

knots, and its status cannot be changed until WoffW. However, the velocity vector’s selected

caged/uncaged status does not change during touch-and-go landings if the ground speed remains above

80 knots.

11. Flight path/pitch ladder. The vertical flight path angle of the aircraft is indicated by the position

of the velocity vector on the flight path/pitch ladder. The horizon and flight path/pitch angle lines

represent the horizon and each 5° of angle between ±90°. Positive pitch lines are solid and are above

the horizon line. Negative pitch lines are dashed and are below the horizon line. The outer segments

of the lines point toward the horizon. Each line is numbered and the numbers rotate with the lines

so that inverted flight can easily be determined. 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, the 50° pitch line is angled 25° toward the horizon. In level flight, the pitch lines

are not angled. The zenith is indicated by a circle and the nadir is indicated by a circle with an X

in it. Aircraft pitch angle can be determined by comparing the tops of the altitude and airspeed

boxes (which represent the aircraft waterline) with the pitch ladder when the wings are level, but the

flight path/pitch ladder normally rotates about the velocity vector and determination of pitch angle

may be difficult at high roll angles.

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12. Vertical velocity readout. This value is displayed above the altitude box and indicates vertical

velocity in feet per minute. This is displayed in the NAV master mode only. Descent is indicated by

a minus sign.

13. HUD landing symbology. When any two landing gear are down, the Mach number, g, and peak

g are deleted and an AOA bracket, extended horizon bar, waterline symbol, and energy caret appear.

The center of the AOA bracket represents the optimum approach AOA. The bracket moves lower

with respect to the velocity vector as AOA increases and moves higher as AOA decreases. When the

energy state of the aircraft is in equilibrium, the energy caret points to the ″right wing″ of the

velocity vector and the aircraft neither accelerates or decelerates. With an energy deficit, the energy

caret moves lower with respect to the velocity vector and the aircraft decelerates; with excess energy,

the energy caret moves higher and the aircraft accelerates.

14. Waypoint/OAP, mark point, TACAN, or target data. Waypoint/OAP and mark data consists of

range (horizontal), and the steer-to point identifier (W, O, or M) and number located on the lower

right corner of the HUD. TACAN data consists of slant range and a Morse code identifier located

on the lower right corner of the HUD. When a steer-to point is designated, range remains displayed

and the steer-to point identifier changes to TGT.

15. Coupled steering symbology. While coupled steering is engaged CPL SEQ#, CPL WYPT, CPL

TCN, CPL BNK, CPL ASL, CPL HDG, or CPL P/R appears on the right side of the HUD display

above the navigation data.

16. ILS symbology. When ILS steering is selected, an azimuth deviation bar (localizer) and elevation

deviation bar (glideslope) appear on the HUD.

17. ZTOD, LTOD, ET, and CD time. The ZTOD, LTOD, ET, or CD time is displayed on the lower

left corner of the HUD. These timers are mutually exclusive. Only one timer is available for display

on the HUD at a time. When the FIRAMS real time clock power up BIT passes, ZTOD does not need

to be entered, but when the FIRAMS real time clock power up BIT does not pass, ZTOD must be

entered. ET initializes to zero minutes and seconds. CD initializes to 6 minutes and zero seconds.

18. Command heading marker. When waypoint/OAP or TACAN direct great circle steering is

selected, the command heading marker is displayed just below the heading scale.

19. Steering arrow and dots. When waypoint/OAP or TACAN course line steering is selected, the

steering arrow and dots appear on the HUD.

20. Required ground speed cue. When steering is engaged to the target in a sequence, the required

ground speed cue appears under the airspeed box.

21. Target designation symbology. When a target is designated, a target designation symbol

(diamond) appears below the heading scale indicating target heading. Another target designation

symbol (diamond) appears indicating the target line of sight (LOS).

2.19.4.6.11 HUD Symbology Degrades. The avionics suite has built in redundancy with two mission

computers for data management and two DDIs (LOTs 21-24), or two MCs (LOTs 25 and up), for

symbol generation. Likewise, if the attitude select switch is in the AUTO or INS position, back up data

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sources are automatically selected to provide HUD symbology when failures are detected. Refer to

figure 2-53, for the HUD displays discussed below.

HUD Symbology Degrades with INS Failure. When a failure occurs in the INS, HUD

bank angle, velocity vector, pitch ladder, and heading indications can be expected to be

impacted. With GPS operating, the mission computer utilizes GPS information for the velocity

vector. If INS attitude is valid but INS velocities are not valid the mission computer

automatically uses the INS attitude and GPS velocities to position a non-flashing velocity

vector. With a degradation of the air data function (probe or pressure transmitter set damage

or failure) calibrated airspeed, barometric altitude, indicated Mach number, and vertical

velocity indications may be impacted.

When the INS experiences a total shutdown (dump) with the attitude select switch in AUTO

or INS, or if the attitude switch is deliberately placed in standby, a stationary waterline symbol

replaces the velocity vector indicating that the standby attitude reference indicator is now

providing attitude data. This failure is normally accompanied by the MASTER CAUTION

light, tone, and INS ATT caution. Place the attitude select switch in the STBY position,

crosscheck the HUD against standby instruments, and attempt an in-flight alignment.

Due to the tendency of the standby attitude reference indicator to precess, it is suggested that

flying in instrument meteorological conditions (IMC) using the ARI as a primary attitude

reference be minimized. A partial IFA (In-Flight Alignment) is always recommended whenever

possible to recover the INS attitude platform.

HUD Symbology Degrades with Air Data Function Failure. An air data function

failure in the FCC results in loss of associated data from the HUD display as shown in figure

2-53. Such a failure also inhibits operation of cruise flight Automatic Throttle Control and

disables the altitude signal used for IFF altitude reporting. An air data function failure may

affect cabin air flow and cabin air temperature.

The pressure transmitter set can produce erroneous signals without cautions or advisories if

the pitot tube or AOA probes receive damage. As the air data function degrades, loss of some

or all of the following data from the HUD may occur:

(1) Calibrated airspeed or barometric altitude. The loss of calibrated airspeed and/or baro-

metric altitude data results in activation of the landing gear handle warning light and tone

with the gear UP. Aircrew action is to reference the applicable standby airspeed or altitude

indicator and then silence the tone.

(2) Angle of Attack. Loss of AOA in three or more FCC CHs causes AOA to be removed from

the HUD.

(3) Vertical velocity indicator. Pilot action on loss of the vertical velocity indication is to check

that the aircraft is in the NAV master mode and to reference the standby vertical velocity

indicator.

(4) Mach number. Pilot action on loss of the Mach number indication is to reference the

standby airspeed indicator.

If an AOA probe becomes jammed (does not move), the FCC continues to receive valid signals until

the pilot executes a maneuver that causes the reading between the AOA probes to differ more than 15°

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in UA or 5.5 to 15° in PA, depending on sideslip. HUD displayed airspeed may be inaccurate without

annunciation if a pitot tube is damaged.

A jammed, blocked, or damaged pitot tube/AOA probe may not be

annunciated if system errors are not large enough to set a caution. Be

alert for unannunciated pitot static and AOA errors during flight in icing

conditions or if damage is suspected after a bird strike or IFR basket

impact during inflight refueling.

Air data inputs from the MC are used by the INS to help smooth or dampen pitch ladder and

velocity vector position. A complete air data function failure does not immediately affect the pitch

ladder/velocity vector, but these displays eventually degrade. If subtle damage to the AOA probe is

suspected, the pilot should make a crosscheck of airspeed with a wingman if possible. The standby

airspeed indicator receives signals from the left pitot static probe, so it is accurate if only the right

probe is damaged. AOA checks with a wingman should be made in landing configuration if a jammed

AOA probe is suspected. Crosschecking in cruise configuration may give a satisfactory crosscheck, but

the probe may be bent in such a way that AOA anomalies are accentuated on landing configuration.

Landing with automatic throttle control (ATC) may be affected. If damage is suspected, ATC during

landing is not recommended.

When AOA is declared invalid (e.g., AOA Four Channel failure), the HUD AOA display and AOA

bracket are removed and the AOA indexer lights and approach lights are inoperative. GAIN ORIDE

provides fixed gains to the FCS and allows the pilot to select, through the FCS status display, either

the left or right probe. The center (INS) AOA value allows the pilot to compare AOA values to select

the undamaged probe. Once selected, this probe drives the HUD AOA display, AOA bracket, AOA

indexer, and approach lights. If the incorrect probe is selected, the information provided to the pilot

and LSO may be in error but has no impact on the flight control system as the gains are fixed. Notify

the LSO that a single probe has been selected.

2.19.4.6.12 HUD Advisory Data Symbology. The displays in figure 2-52 show some of the advisories

that can appear on the HUD in the NAV master mode. The advisories are associated with nose wheel

steering and approach power compensator. Although the advisories are shown on the gear down

display, most of them can appear on the basic HUD display. Refer to Part VII for description of data

link system and advisories.

The automatic throttle control/nosewheel steering advisories are displayed above the distance

display whenever the ATC or the NWS is engaged. If the ATC is disengaged by any means other than

actuation of the ATC engage/disengage switch, the advisory is flashed for 10 seconds before it is

removed from the display or, if a pilot attempt to engage ATC is not successful, ATC is flashed for 10

seconds and removed.

2.19.4.6.13 HUD BIT Checks. The HUD has two methods of built-in tests: manually initiated and

automatic test. Refer to BIT-Status Monitoring Subsystem for the procedures and displays used for

the HUD BIT checks.

2.19.4.7 CRS Set Switch. The course set switch manually sets the desired course on the HSI display.

2.19.4.8 HDG/TK Set Switch. The heading/ground track set switch manually sets the heading

marker on the desired heading/ground track on the HSI display.

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2.19.5 Up Front Control Display (UFCD). The UFCD is on the main instrument panel below the

HUD in the front cockpit. In the rear cockpit, the UFCD is located below the MPCD in LOTs 21 thru

25, and above the MPCD or 8 x 10 display in LOT 26 AND UP. The UFCD is an active matrix liquid

crystal display with an IR touchscreen used for data entry inputs consisting of digits or NEWS, and

control of the CNI systems (autopilot modes, IFF, TACAN, ILS, data link, radar beacon, UHF/VHF

radios, and ADF), see figure 2-55. In addition, the touchscreen can be used as a multi-function display

for display formats, including video. The UFCD is used in conjunction with the two DDIs, the 8 x 10

display and the MPCD to enter navigation, sensor, and weapon delivery data. UFCD option selections

and inputs are transmitted directly to the MPCD and on to the mission computers. (The mission

computers pass these inputs to the control converter (CC) for CNI equipment control). The UFCD is

NVG compatible. In the F/A-18F, the front and rear cockpit UFCDs operate independently. When

different formats are being displayed, the only data common to both UFCDs is radio channel and

frequency information. Both cockpit UFCDs present the results of changes to radio channels or

frequencies at the same time, regardless of which cockpit performed the change. The cockpit not

performing data entry does not see touch highlights as the data is entered, only the result of the data

entry. If both pilot and WSO enter digits on the keypad for the same option, both entries are accepted,

with the second entry overwriting the first. When pilot and WSO are on the same data entry or CNI

format, asterisks are provided in the top left and right of the scratch pad.

(LOT 26 AND UP) The aft UFCD is electrically controlled through the 8 x 10 display Off/On/

Brightness knob when it is installed. The aft UFCD flashes STANDBY when the MC1 communication

to the UFCD is disrupted. A standby display pattern in the the center of the UFCD display surface,

similar to the DDI pattern, indicates degraded image processing from MC1. When the Mono video

connection from the 8 x 10 display is lost or degraded, the aft UFCD will not display anything. When

video synchronization is lost or degraded, the aft UFCD may display symbology that is not coordinated

with the video. If the interface between MC1 and the aft UFCD is lost, the touch screen capability and

brightness/contrast control will not work. When MC1 is inoperative, the aft UFCD is not capable of

displaying format symbology or analog video.

When the pilot or WSO touch a keypad option on the touch screen a highlight appears indicating

that the option has been selected. Figure 2-55 shows an example of a selected option. When a new

format is selected on the touchpad, the highlight does not remain on the new format.

Keypad options use a first finger in mechanization. Only one option can be selected at a time. If two

or more selections are attempted at one time, none of them are considered valid.

Some formats initialize with data in the scratchpad, e.g., the COMM sublevel of the CNI format

initializes with the comm frequency in the scratchpad. When data entry is started, the digits in the

scratchpad are blanked, allowing data entry. If a comm frequency is being entered the decimal remains

displayed, allowing frequency entry in relation to the decimal point to be viewed.

The UFCD uses a double clear mechanism. The first selection of the CLR keypad option removes the

last digit that was entered in the scratchpad. The second selection of the CLR option removes all the

digits which have been entered in the scratchpad. If the scratchpad is flashing due to the 10 second

timer (no entries in the previous 10 seconds), pressing CLR stops the flashing and performs the

previously described CLR function. If the scratchpad is flashing due to an error, pressing CLR stops

the flashing and removes all digits from the scratchpad.

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F18. FLIGHT MANUAL (2008) - page 19