F18. FLIGHT MANUAL (2008)

A1-F18EA-NFM-000

2.8.1.2

Normal Landing Gear Extension and Retraction. Normal landing gear extension and

retraction is electrically controlled by the LDG GEAR handle and uses hydraulic pressure from HYD

2A. With weight off the nose gear and the launch bar retracted, moving the LDG GEAR handle to the

UP position sends an electrical signal to the landing gear selector valves to initiate normal landing gear

retraction. Likewise, moving the LDG GEAR handle to the DN position sends an electrical signal to the

landing gear selector valves to initiate normal landing gear extension

If the launch bar does not return to the up and locked position after catapult launch or the nose gear

indicates WonW, the nose landing gear cannot be retracted. In either case, placing the LDG GEAR

handle UP will raise the main landing gear and leave the nose landing gear extended.

2.8.1.3

Emergency Landing Gear Extension. Emergency landing gear extension is mechanically

controlled by the LDG GEAR handle (front cockpit) or the EMERG LDG GEAR handle (rear cockpit

Lots 21 thru 25) and uses hydraulic pressure provided by the APU accumulator. Both handles are

mechanically connected to the landing gear emergency selector valves by a series of levers and cables.

Emergency extension is mechanically activated by rotating the LDG GEAR handle 90° clockwise and

pulling to detent (approximately 1.5 inches) or by pulling the EMERG LDG GEAR handle to the

detent.

Emergency landing gear extension opens the hydraulic arming valve and directs APU accumulator

pressure to the emergency selector valves. APU accumulator pressure is used to unlock the doors,

release the landing gear uplocks, and is applied to the drag brace locking actuator and sidebrace

downlock actuator. The nose landing gear extends by freefall aided by airloads and the drag brace

locking actuator. The main landing gear extends by freefall aided by the sidebrace downlock actuator.

Emergency extension can be performed with the LDG GEAR handle either UP or DN (DN is

recommended).

2.8.1.4

LDG GEAR Handle. The wheel-shaped LDG GEAR handle, located on the lower left main

instrument panel in the front cockpit, is used to control landing gear extension and retraction. A

downlock solenoid in the LDG GEAR handle assembly prevents gear retraction with WonW by

preventing movement of the handle from the DN position.

With WoffW and the launch bar retracted, electrically initiates normal

landing gear retraction.

Electrically initiates normal landing gear extension.

Emergency

Mechanically initiates emergency landing gear extension.

(Rotate handle 90°

clockwise and pull

to the detent)

2.8.1.5

DOWNLOCK ORIDE Button. The DOWNLOCK ORIDE button, located on the lower left

main instrument panel outboard of the LDG GEAR handle, is used to override the downlock solenoid.

If the downlock solenoid does not retract with WoffW (LDG GEAR handle cannot be moved from the

DN position), a failure has occurred in the downlock circuitry (Landing Gear Control Unit). If the

landing gear indicate three down and locked, cycling the landing gear handle is not recommended, as

proper landing gear functioning is questionable. However, if dictated by an emergency situation,

pressing and holding the DOWNLOCK ORIDE button will retract the mechanical stop and allow the

LDG GEAR handle to be moved to the UP position.

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The LDG GEAR handle must be in the full down position for the mechanical stop to properly engage

after landing (WonW).

If the DOWNLOCK ORIDE button is pressed or the mechanical stop is

not fully engaged, the LDG GEAR handle can be raised on the ground,

and the main landing gear will retract. The nose gear will not retract with

weight on the nose gear.

2.8.1.6

EMERG LDG GEAR Handle (F/A-18F). The EMERG LDG GEAR handle is located on the

lower left main instrument panel in the rear cockpit. Pulling the handle to the detent mechanically

initiates emergency landing gear extension.

2.8.1.7

Landing Gear Control Unit (LGCU). The LGCU monitors the position of the landing gear and

launch bar systems, provides cockpit indications of gear/launch bar position, and provides outputs to

various aircraft systems which are dependent on gear position (e.g., FCC A and B, the SMS, and the

SDC). The LGCU does not control landing gear extension and retraction.

The LGCU receives inputs from the LDG GEAR handle, the LAUNCH BAR switch, and the

following proximity switches: launch bar, landing gear uplocks, landing gear downlocks, planing links,

and WonW. The LGCU controls the red and green L BAR warning/advisory lights, the landing gear

position lights, the light in the gear handle, the landing gear warning tone, the downlock solenoid, and

all inputs to the FCCs and the SMS. The LGCU also performs a self-BIT and a functional check of all

proximity switches, providing MSP code input to the SDC.

2.8.1.8

Landing Gear Warning Light and Warning Tone. The landing gear warning light is a red light

located inside the LDG GEAR handle. The landing gear warning tone is a beeping tone heard in the

headset. The landing gear warning light and warning tone serve three purposes: to indicate a mismatch

between LDG GEAR handle position and actual gear position, to warn of a planing link failure, and to

provide a ″wheels warning.″

A steady warning light comes on whenever the landing gear is in transit and remains on until all

three gear are down and locked (LDG GEAR handle DN) or all gear doors are closed and locked (LDG

GEAR handle UP). If the landing gear is unsafe, the landing gear warning light remains on. The

warning tone is inhibited for 15 seconds to allow for normal landing gear extension and retraction. If

the warning light remains on for 15 seconds, the warning tone is annunciated to provide an aural

indication of unsafe landing gear position.

If a left or right planing link failure occurs with the landing gear down and locked (planing link

proximity switch not properly activated), the landing gear warning light will come on immediately

accompanied by the warning tone.

Lastly, when the LDG GEAR handle is UP, a flashing warning light accompanied by the warning

tone will be activated when airspeed is below 175 KCAS, altitude is less than 7,500 feet, and rate of

descent is greater than 250 fpm. This ″wheels warning″ is provided as a cue to check the position of the

landing gear at flight conditions where the LDG GEAR handle should normally be DN. The wheels

warning is also activated if calibrated airspeed and/or barometric altitude data are lost. In this case, the

standby airspeed and/or altitude indicators should be referenced prior to silencing the warning tone.

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2.8.1.8.1

WARN TONE SIL Button. The WARN TONE SIL button, located to the left of the LDG

GEAR handle, is used to silence the landing gear warning tone.

2.8.1.8.2

Landing Gear UNSAFE Light (F/A-18F). The landing gear UNSAFE light is a red light

located on the upper left main instrument panel in the rear cockpit. The light indicates a mismatch

between LDG GEAR handle position and actual gear position (e.g., gear in transit). The light does not

illuminate for a planing link failure, wheels warning, or loss of air data.

2.8.1.9

Landing Gear Position Lights. Three green landing gear position lights, located on the lower

left main instrument panel, are labeled NOSE, LEFT, and RIGHT. When the LDG GEAR handle is

DN, steady lights indicate that the corresponding landing gear is down and locked. The LEFT and

RIGHT landing gear position lights are also used to indicate a planing link failure. If the main landing

gear are down and locked but a planing link proximity switch is not properly activated, the

corresponding position light will flash.

A landing gear position of three down and locked is indicated by three steady green position lights

with the landing gear warning light out. Additionally, when illuminated inflight, the approach lights

provide an external indication that the landing gear is down and locked.

If a landing gear position light is out with the LDG GEAR handle DN and the landing gear warning

light out, a LT TEST should be performed to test the integrity of the position light bulb. If the bulb

tests bad, it is safe to assume that the gear is down and locked. During day operations, if all three

position lights appear to be out/dim, make sure the interior lights MODE switch is in the DAY position.

A landing gear position of three up and locked is indicated by the landing gear warning light out with

all three position lights out.

If one or more landing gear indicates unsafe, a visual inspection can only

confirm general position and obvious damage. There is no external

indication of a locked landing gear.

2.8.1.9.1

Landing Gear Position Lights (F/A-18F). Three green landing gear position lights, labeled

NOSE, LEFT, and RIGHT, are located on the upper left main instrument panel in the rear cockpit.

These lights have the same functionality as those in the front cockpit.

2.8.2 Nosewheel Steering System (NWS). The NWS system is used to provide directional control

and shimmy damping during ground operations. The NWS hydraulic power unit, attached to the nose

landing gear strut, is electrically controlled by commands from the FCCs and is hydraulically actuated

by pressure from HYD 2A (primary) or HYD 2B/ APU accumulator (backup). In the event of a HYD

2A failure, a pressure-biased shuttle valve routes HYD 2B pressure (if available), or APU accumulator

pressure to the NWS unit for backup operation. The FCCs accept input from the rudder pedals to

provide NWS commands.

The NWS system has two modes, NWS (low) and NWS HI. In the low mode (NWS cue in the HUD),

full rudder pedal deflection commands approximately 22.5° of nosewheel deflection. In the high mode

(NWS HI cue in the HUD), full rudder pedal deflection commands approximately 75° of nosewheel

deflection. The NWS system (low gain) incorporates a yaw rate feedback input from the FCCs, which

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is designed to suppress directional PIO tendencies by increasing directional damping during takeoff

and landing roll.

With loss of yaw rate information to the FCCs, directional PIO may occur

during aggressive ground tracking.

If the NWS system fails, the NWS caution is displayed and the NWS or NWS HI cue is removed

from the HUD. When failed, the NWS system reverts to a 360° free-swiveling mode.

2.8.2.1

NWS Engagement/Disengagement. With WonW, manual NWS engagement is provided by

actuation of the NWS/undesignate button. The method required to engage each of the two NWS

modes (low and high) is dependent on wing lock/unlock status.

With the wings spread and locked and NWS disengaged, the first momentary press and release of the

NWS button engages full-time NWS (low). NWS HI is engaged by subsequent press and hold of the

NWS button. With NWS disengaged, press and hold for greater than 1 second also engages NWS HI.

If the NWS button is released, the system reverts to NWS (low).

With the wings unlocked and NWS disengaged, the first momentary press and release of the NWS

button still engages full-time NWS (low). However, subsequent press and release engages full-time

NWS HI, providing hands-free NWS HI capability for operations in the carrier environment. If the

wings are subsequently spread and locked, NWS reverts to the low mode.

During landing, full-time NWS (low) is automatically engaged when the nose landing gear and at

least one main landing gear transition to WonW. If NWS is engaged with both HYD 2A and 2B failures,

the NWS or NWS HI cue will flash in the HUD as an indication that APU accumulator pressure is

depleting.

NWS is manually disengaged by pressing the paddle switch. NWS is automatically disengaged for

catapult launch, when the launch bar is extended. With the launch bar extended, NWS (low) can be

momentarily engaged to position the launch bar by press and hold of the NWS button. Additionally,

NWS is automatically disengaged when the nose landing gear transitions to WoffW during takeoff or

when power is removed from the FCCs.

2.8.2.2

Emergency High Gain NWS. With a FCS CH 2 or FCS CH 4 failure, normal nosewheel

steering is lost. Emergency high gain NWS can be regained by pulling the failed channel circuit

breaker, unlocking the wings, and momentarily pressing the nosewheel steering button.

When emergency high gain NWS mode is entered, NWS indications may

not be displayed on the HUD. As a result, inadvertent nosewheel steering

actuation may injure ground personnel.

2.8.3 Wheel Brake System. The aircraft’s wheel brake system provides normal braking, anti-skid,

emergency braking, a parking brake, and main wheel anti-spin. Normal braking utilizes HYD 2A

pressure and is capable of functioning with a separate anti-skid system. The anti-skid system, when

enabled, provides maximum braking effectiveness on wet runways or during heavy braking by

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preventing wheel skid. When selected, emergency braking utilizes HYD 2B pressure, if available, or

brake and APU accumulator pressure to provide backup braking capability following a HYD 2A

failure. The anti-spin function stops main landing gear wheel rotation prior to landing gear retraction.

2.8.3.1

Wheel Brake Assembly. Each main landing gear wheel is fitted with hydraulically actuated

multiple disk brakes. There are two independent sets of brake lines running to each wheel brake

assembly: the normal brake line pressurized by HYD 2A and the emergency brake line pressurized by

HYD 2B or the brake and APU accumulators. See figure 2-17. Only one set of brake lines can be

pressurized at any given time. A shuttle valve on each wheel brake assembly switches from normal to

emergency brake pressure, depending on which is applied.

Each wheel brake assembly has a brake wear indicator pin, located on the inboard side of the wheel.

When the brakes are applied and the indicator pin is flush or below flush with the brake housing, the

brake pads require changing.

Each wheel assembly incorporates a fuse plug which is designed to melt and deflate the tire at

temperatures below those which would result in a catastrophic tire blowout.

2.8.3.2

Wheel Brake Operation. Each main wheel brake is controlled by a separate brake pedal,

integrated into the rudder/brake pedal mechanism. Pilot applied force to the top of each brake pedal

is transmitted by a series of cables and pulleys directly to the brake control hydraulic servovalves,

located in the nose wheelwell. The amount of hydraulic pressure applied to the wheel brakes by the

servovalves is directly proportional to brake pedal force. Dual brake pedal action provides symmetric

braking, while individual brake pedal action provides differential braking. In the F/A-18F (trainer

configuration), a second set of cables are routed to the servovalves from the rear cockpit brake pedals.

The servovalves are controlled by the pilot applying the most brake pedal force.

2.8.3.3

Normal Braking. Normal braking is enabled when HYD 2A is operable and the EMERG

BRK handle(s) are in the stowed position. The emergency brake valve is closed and the emergency

brake lines are unpressurized. During normal braking, HYD 2A pressure is applied through the left and

right servovalves proportional to the amount of pilot applied brake pedal force and is routed to the

main wheel brakes. When the ANTI SKID switch is ON, the anti-skid system modulates pilot applied

brake pressure in order to prevent wheel skid. When the ANTI SKID switch is OFF, the pilot must

regulate brake pedal force to prevent wheel skid.

2.8.3.4

Anti-skid System. The anti-skid system performs 4 basic functions which are designed to

maximize braking effectiveness during landing rollout: touchdown protection, wheel spin-up override,

skid control, and locked wheel protection. The anti-skid system is enabled when the ANTI SKID

switch is ON and the LDG GEAR handle is DN.

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Figure 2-17. Wheel Brake and Anti-skid System

Failure of either wheel speed sensor can lead to an anti−skid failure,

resulting in a complete loss of brakes. Placing the ANTI SKID switch to

the OFF position or pulling the EMERG BRK handle will bypass the

faulty system and restore braking ability. Judicious braking must be

used, as the anti−skid system is not available. Refer to the BRAKE

FAILURE/EMERGENCY BRAKES procedure.

The system contains two wheel speed sensors, an anti-skid control unit, and an anti-skid control

valve. The anti-skid control unit senses wheel speed and operates by electronically limiting the amount

of HYD 2A pressure that is applied to the wheel brakes through the anti-skid control valve and the

normal brake lines. Anti-skid is not available when emergency brakes are selected.

Touchdown protection delays initial brake application on landing by completely dumping brake

pressure until (1) weight is on the right main landing gear and wheel speed is over 50 knots or (2), if

a wet runway delays wheel spin-up, for 3 seconds after landing. This function prevents landing with

locked main wheels (tire blowouts) even if full brake pedal force is applied at touchdown.

Wheel spin-up override is activated at 50 knots wheel speed to allow normal braking if the right

WonW switch fails.

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Skid control is enabled when sensed wheel speed differs from what the anti-skid control unit

determines it should be (e.g., hydroplaning is detected). If the system detects wheel skid, anti-skid

limits the amount of HYD 2A pressure applied to both brakes as required to prevent skidding. If no

skid exists, full pilot-applied brake pressure is routed to the brakes.

If the speed of one wheel drops 40% below the other wheel, locked wheel protection dumps brake

pressure to both wheels until the speed of the slower wheel returns above 40% of the other. Locked

wheel protection is removed below 35 knots, so that full braking performance (including locking a tire)

is available for taxi and turning operations. Below 14 knots, anti-skid is completely disabled. Below 35

knots, judicious braking is required to avoid flat spotting tires.

Anti-skid protection is bypassed when the ANTI SKID switch is OFF. Normally limited by the

anti-skid system, 3,000 psi hydraulic brake pressure is available and regulated only by pilot brake pedal

forces. When using brakes at high speed without anti-skid protection, there is a very small margin

between effective braking and blown tires. Any force greater than approximately 55 to 60 lbs applied

to the pedals (6° to 7° of pedal rotation) will likely result in blown tires with either the ANTI SKID

switch OFF or emergency brakes selected. The use of normal (anti-skid off) brakes at high speed

should be done with extreme caution. If braking without anti-skid is needed at high speeds, initially

apply very light brake pedal pressure and gradually increase as required.

Use of brakes without anti-skid at high speed can result in blown tires

resulting in loss of directional control. If practical, rollout speed should be

as slow as possible before applying brake pedal pressure.

NOTE

Hot brakes and/or melted wheel assembly fuze plugs can be expected

any time maximum effort braking is used at heavy gross weights with

or without anti-skid, e.g., aborted takeoff or heavy weight landing

(above 46,000 lb GW) with high taxi brake usage.

2.8.3.4.1

ANTI SKID Switch. The ANTI SKID switch, located on the lower left main instrument

panel, is used to manually disable the anti-skid system, e.g., for carrier operations or following an

anti-skid failure (ANTISKID caution displayed). The switch is lever-locked in the OFF position.

Anti-skid system enabled for use with normal braking.

OFF

Anti-skid system disabled (SKID advisory displayed when the landing gear is

down).

2.8.3.4.2

Anti-skid BIT and the ANTISKID Caution. The anti-skid control unit performs two types of

BIT: initiated and periodic. IBIT is performed when power is initially applied to the anti-skid system:

(1) when the landing gear is lowered, (2) when the ANTI SKID switch is selected from OFF to ON

inflight, or (3) on the ground with the parking brake set. IBIT performs a complete test of the anti-skid

system 9 seconds after power is applied and runs for 4.5 seconds. With WonW, IBIT is inhibited with

the parking brake released, as wheel motion will cause a false BIT failure and brake pressure would be

dumped if brakes were applied. PBIT only performs a partial anti-skid test and runs whenever power

is applied and IBIT is not running.

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If an anti-skid failure is detected by either BIT, the ANTISKID caution will be displayed at BIT

completion. If an anti-skid failure is detected by PBIT, cycling the ANTI SKID switch will command

an IBIT and a more complete test of the system. When IBIT is running, the ANTISKID caution is

inhibited or is removed if previously displayed. If the ANTISKID caution returns after IBIT, the

ANTI SKID switch must be placed to OFF in order to isolate the failure and make sure that normal

braking (without anti-skid) is available.

For instance, assume the right wheel speed sensor has failed and an ANTISKID caution is displayed.

If the ANTI SKID switch is left ON during landing, touchdown protection circuitry will dump and

never restore brake pressure to both wheels. Normal braking will be lost. In this case, placing the ANTI

SKID switch to OFF will restore normal braking (without anti-skid), or pulling the EMERG BRK

handle will enable emergency braking (bypassing anti-skid).

• Do not cycle the ANTI SKID switch in response to an ANTISKID

caution immediately prior to landing. Cycling the ANTI SKID switch

removes the ANTISKID caution for up to 13.5 seconds as the system

performs IBIT even though the anti-skid system may still be failed

and, if the system is not failed, wheel motion at touchdown may cause

a false BIT failure and a dump of normal brake pressure when brakes

are applied.

• If the ANTI SKID switch is not placed to OFF with an ANTISKID

caution displayed, normal braking capability may be lost completely.

2.8.3.5

Emergency Braking. Emergency braking is enabled when either EMERG BRK handle is

pulled to detent. This action opens the emergency brake valve and applies backup hydraulic pressure

to the hydraulic servovalves. If available, backup pressure from HYD 2B is utilized through the aft

isolation and arming valves, which are open with WonW. If HYD 2 is failed completely, backup

pressure from both the brake and APU accumulators is used. Check valves are incorporated to prevent

the loss of accumulator pressure if HYD 2B is failed. With backup pressure applied, the servovalves

isolate HYD 2A pressure, if still available, so that the normal brake lines are unpressurized.

During emergency braking, backup pressure is applied through the left and right servovalves

proportional to the amount of pilot applied brake pedal force and is routed to the main wheel brakes

through the emergency brake lines. These lines bypass the anti-skid control valve, so the pilot must

regulate brake pedal force to prevent wheel skid.

Hydraulic accumulators and the brake accumulator pressure gauge are discussed in the Hydraulic

Power Supply System section.

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Figure 2-18. Emergency/Parking Brake Handle

2.8.3.5.1

EMERG BRK Handle (Front Cockpit). The EMERG BRK handle is combined with the

PARK BRK handle and is located on the lower left main instrument panel in the front cockpit. When

the handle is in the stowed, emergency position (horizontal), the ‘‘EMERG’’ label appears upright. See

figure 2-18. To select emergency brakes, the handle must be pulled to the detent while in the horizontal

position.

Stowed

Emergency brake valve closed. Normal braking selected.

(unmarked)

PULL

Emergency brake valve open. Emergency braking selected.

(to detent)

The position of the EMERG BRK handle is the only indication that

emergency braking is selected: no warning or caution is displayed. The

EMERG BRK handle(s) must be fully stowed in both cockpits to make

sure that normal braking with anti-skid is available.

Due to friction in the EMERG BRK handle mechanism, the handle may

not return to the fully stowed position unless positively pushed.

2.8.3.5.2

EMERG BRK Handle (Rear Cockpit Lots 21 thru 25). The EMERG BRK handle, located

on the lower left main instrument panel in the rear cockpit, is used to select emergency braking from

the rear cockpit. When stowed, the handle is oriented vertically.

Stowed

Emergency brake valve closed. Normal braking selected.

(unmarked)

PULL

Emergency brake valve open (to detent). Emergency braking selected.

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The position of the EMERG BRK handle is the only indication that

emergency braking is selected: no warning or caution is displayed. The

EMERG BRK handle(s) must be fully stowed in both cockpits to make

sure that normal braking with anti-skid is available.

Due to friction in the EMERG BRK handle mechanism, the handle may

not return to the fully stowed position unless positively pushed.

2.8.3.6

Parking Brake System. The parking brake is used to lock the main landing gear wheels when

the aircraft is parked. The parking brake is activated when the PARK BRK handle is rotated and

pulled to the locked position. This action places the emergency brake valve in the parking brake mode.

Backup hydraulic pressure from HYD 2B or the brake and APU accumulators is applied to the wheel

brake hydraulic servovalves and routed to the main wheel brakes through the emergency brake lines.

The PARK BRK caution will come on to alert the pilot that the parking brake is still set when both

throttles are advanced above about 80% N₂ rpm (INS on).

2.8.3.6.1

PARK BRK Handle. The PARK BRK handle is combined with the EMERG BRK handle

and is located on the lower left main instrument panel in the front cockpit. From the stowed

(horizontal) position, the handle must be rotated 90° counterclockwise and pulled to the locked

position, in order to activate the parking brake. When the handle is in the vertical position, the

‘‘PARK’’ label appears upright. See figure 2-18. If emergency brakes are selected, the handle must be

returned to the stowed position before the parking brake can be activated. Rotating the handle 45°

counterclockwise releases the lock and allows the handle to return to the stowed (horizontal) position.

Stowed

Parking brake released. Normal braking selected.

(unmarked)

TURN/

Parking brake set.

PULL

Several aircraft systems utilize parking brake activation to enable or disable logic. A set parking

brake is used to enable anti-skid BIT logic, GEN TIE logic, and INS alignment and is used to trigger

the PARK BRAKE caution.

2.8.3.7

Main Wheel Anti-Spin. The anti-spin function stops main landing gear wheel rotation prior

to landing gear retraction. When the LDG GEAR handle is moved to the UP position, main landing

gear retract pressure is supplied to the anti-skid control valve. Normal brake pressure is blocked and

this anti-spin pressure is routed to the wheel brakes through the normal brake lines. Unlock and

retraction of the main landing gear is delayed until anti-spin pressure is applied and main wheel

rotation has stopped.

2.8.4 Launch Bar System. The launch bar is electrically controlled, hydraulically extended, and

mechanically retracted. With weight on the nose gear, placing the LAUNCH BAR switch to EXTEND

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energizes the launch bar control valve and routes HYD 2A pressure to unlock, lower, and hold down the

launch bar. The green L BAR advisory light indicates that the launch bar has been extended.

With the launch bar extended, returning the LAUNCH BAR switch to RETRACT deenergizes the

launch bar control valve, isolates HYD 2A pressure, and allows dual retract springs to mechanically

return the launch bar to the up and locked position. A launch bar proximity switch is energized when

the launch bar is fully retracted.

When the launch bar is fully extended it is held against the deck by HYD 2A pressure. Deck load

springs allow vertical movement of the launch bar during taxi over the catapult shuttle. When the

aircraft is placed in tension on the catapult, the launch bar is held captive in the extended position by

the shuttle. Once in tension, the LAUNCH BAR switch should be placed to RETRACT in order to

remove HYD 2A pressure from the launch bar. When the LAUNCH BAR switch is placed to

RETRACT, the green L BAR light should go out.

Failure to place the LAUNCH BAR switch to RETRACT prior to

catapult launch may result in launch bar hydraulic seal failure and

possible loss of HYD 2A.

At the end of the catapult stroke, launch bar/shuttle separation occurs and allows the retract springs

to return the launch bar to the up and locked position. When engaged, the launch bar uplock prevents

the launch bar from dropping to the deck due to g-loads during landing.

If the launch bar does not return to the up and locked position after catapult launch (launch bar

proximity switch not energized), the nose landing gear cannot be retracted. In this case, placing the

LDG GEAR handle UP will raise the main landing gear and leave the nose landing gear extended.

2.8.4.1

LAUNCH BAR Switch. The LAUNCH BAR switch, located on the lower left main instrument

panel in the front cockpit, is used to control the position of the aircraft’s launch bar. The switch is

spring loaded to the RETRACT position and is electrically held in the EXTEND position only if

weight is on the nose gear.

RETRACT Launch bar control valve deenergized. Launch bar up.

EXTEND Launch bar control valve energized. Launch bar unlocked and extended by HYD

2A pressure. Green L BAR advisory light on.

2.8.4.2

LB Circuit Breaker. The LB circuit breaker is located on the left-hand circuit breaker panel

above the left console. The LB circuit breaker provides a secondary means to raise the launch bar

following a launch bar malfunction. When pulled, the circuit breaker manually deenergizes the launch

bar control valve, removing HYD 2A pressure and allowing the retract springs to raise the launch bar.

Typically, this action would be required only if (1) the LAUNCH BAR switch failed in the EXTEND

position with weight on the nose gear or (2) the nose gear failed WonW after launch and the pilot failed

to place the LAUNCH BAR switch to RETRACT.

2.8.4.3

L BAR Warning/Advisory Lights. Two L BAR lights, one green and one red, are located on

the left warning, caution, and advisory lights panel. Both lights are controlled by the landing gear

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control unit (LGCU), based on inputs from the LAUNCH BAR switch, the launch bar proximity

switch, and various landing gear proximity switches.

The green L BAR advisory light is used to indicate that the launch bar has been extended. The

LGCU illuminates the green L BAR light when all of the following conditions are met: weight on the

nose gear, the LAUNCH BAR switch in EXTEND, the launch bar not up (launch bar proximity switch

not energized), and the red L BAR warning light not on.

The red L BAR warning light is used to indicate failure of the launch bar to retract after catapult

launch or a failure in the launch bar control system (proximity switch failure). The LGCU illuminates

the red L BAR light when one of the following sets of conditions are met:

1. Launch bar not up and weight off the left main gear.

2. Launch bar not up and left main gear not down.

3. LAUNCH BAR switch in EXTEND and weight off the left main gear.

4. LAUNCH BAR switch in EXTEND and left main gear not down.

The first set of conditions is the primary L BAR warning, e.g., the launch bar does not retract fully

after catapult launch. The other three sets of conditions provide a backup L BAR warning if one or

more of the proximity switches which control launch bar functioning fail.

2.8.5 Arresting Hook System. The arresting hook is always down-loaded by a nitrogen-charged

accumulator (arresting hook snubber) contained in the arresting hook retract actuator. Arresting hook

extension is therefore accomplished by mechanically releasing the arresting hook uplatch mechanism

(HOOK handle down) and allowing snubber pressure and gravity to extend the hook. The hook should

extend in less than 2 seconds. At touchdown, the arresting hook snubber controls hook bounce and

provides a hold down force for arresting cable engagement.

Arresting hook retraction is accomplished by raising the HOOK handle. This electrically opens the

aft isolation valve and the arresting hook selector valve, routing HYD 2B pressure to the arresting hook

retract actuator. HYD 2B pressure overcomes the snubber down-load pressure and raises the hook. The

arresting hook uplatch mechanism captures and locks the hook in the up position. The hook should

retract in less than 4 seconds. If HYD 2B pressure is lost, the arresting hook cannot be retracted.

2.8.5.1

HOOK Handle. The HOOK handle, located on the lower right main instrument panel in the

front cockpit, is used to control arresting hook extension and retraction.

Retracts the arresting hook utilizing HYD 2B pressure.

(unmarked)

Down

Unlocks the arresting hook uplatch mechanism and allows arresting hook snubber

(unmarked)

pressure and gravity to extend the hook.

2.8.5.2

HOOK Light. The red HOOK light is located on the lower right main instrument panel

directly above the HOOK handle. The HOOK light comes on any time hook position does not agree

with HOOK handle position. The light comes on when the hook leaves the up and locked position and

remains on until the hook is fully extended (hook proximity switch energized). With WonW, the hook

will strike the ground before it reaches full extension, so the HOOK light will remain on.

In Lot 26 AND UP, there is a green HOOK light in the rear cockpit on the left warning, caution, and

advisory panel. This HOOK light illuminates when the hook is down.

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2.9 WING FOLD SYSTEM

The aircraft’s outer wing panels are designed to fold vertically to reduce the amount of deck space

occupied by the aircraft in the carrier environment. Each wing contains an independent wingfold

mechanism, which consists of two electric motors (one to lock/unlock the wings and one to spread/fold

the wings). During normal operation, the wings are spread, locked, unlocked, and folded in unison.

2.9.1 Wingfold Mechanism. Each wingfold mechanism contains a dc electric motor, which locks and

unlocks the wings, and an ac electric drive unit, which spreads and folds the wings. When the wings are

spread and locked, a locking bolt is electrically driven through the wingfold hinge, holding it in place.

When the wings are unlocked, a wing unlock flag (commonly called a beer can) protrudes from the

upper surface of the wing near the leading edge of the wingfold hinge, indicating that the locking bolt

is unstowed. The shaft of each beer can is painted red for easy identification. When the wings are

locked, the top of the beer can should be flush or near flush with the upper surface of the wing, and no

red should be showing.

Additionally, when the wings are folded, the ailerons are mechanically locked in the faired position

by a hook on the inboard aileron hinge, which engages an aileron locking pin. The aileron locking pin

is mechanically extended as the wings fold. The hook and locking pin are designed to prevent the

ailerons from blowing inboard over the TEFs when hydraulic power is not applied. If an aileron locking

pin should break, it is possible for the aileron to blow inward over the TEF. If this condition exists

during engine start, the TEF will retract into the aileron, damaging both surfaces.

If the wings are folded, note the position of the ailerons during the

preflight walk-around. If the aileron locking pins do not restrain the

ailerons in the faired position, make sure the ailerons are moved to a

faired or outboard position prior to engine start to preclude damaging the

ailerons and TEFs.

Each wingfold mechanism also contains a wing safety switch, which electrically prevents wingfold

movement. The safety switch is activated by a ″remove before flight″ pin inserted in the underside of

the wing near the wingfold hinge.

For ground crew operations (such as loading wingtip missiles), each wing can be manually unlocked,

folded, or spread. The beer cans can be manually extended by inserting a screwdriver into the wing

unlock motor (underside, leading edge). Once unlocked, the wings can be folded or spread by inserting

a speed handle into the electric drive unit (underside, trailing edge).

2.9.2 Wingfold Operation. With the wings folded, wing spread and lock is commanded by placing the

WINGFOLD switch to SPREAD. The SPREAD command is sent directly to the electric drive units to

spread the wings (there are no WonW or FCC interlocks). When each wing reaches the completely

spread position, power is removed to that electric drive unit, and that wing is automatically

commanded to lock. The WING UNLK caution will not be removed until both wings are locked (both

beer cans down). Once the wings are spread and locked, the ailerons will droop to the position

scheduled by the FCCs based on FLAP switch position.

The wings can be stopped in an intermediate position by placing the WINGFOLD switch to HOLD.

If the wings are spread, selecting HOLD unlocks the wings without folding, allowing full time NWS HI

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to be engaged for operations in the carrier environment. This function is useful when NWS HI is

desired but wingfold is not, e.g., dearming wingtip missiles after carrier arrestment.

With the wings spread and locked, wingfold is commanded by placing the WINGFOLD switch to

FOLD. In order for the wings to unlock, ground power must be applied or the aircraft must be WonW

(left main). The initial FOLD command electrically unlocks the wings (beer cans extended, WING

UNLK caution displayed) and fairs the ailerons. When the FCCs determine that (1) weight is on

wheels, (2) airspeed is less than 100 KCAS accelerating or 66 KCAS decelerating, (3) the ailerons are

faired, and (4) both wings are unlocked, the FOLD command is sent to the electric drive units to fold

the wings. When each wing reaches the completely folded position, power is removed to that electric

drive unit

2.9.3 WINGFOLD Switch. The WINGFOLD switch, located on the lower right main instrument

panel, is lever-locked in all three positions. The switch has a barrier guard to prevent inadvertent

actuation.

FOLD

Unlocks the wings (WING UNLK caution displayed), fairs the ailerons, and, when

(& unlock)

allowed by the FCCs, folds the wings.

HOLD

Stops wing movement in an intermediate position. If spread, unlocks the wings.

(& unlock)

SPREAD Spreads and locks the wings. (WING UNLK caution removed when both wings are

(& lock)

locked).

Ensure the WINGFOLD switch is lever-locked in the SPREAD position

during takeoff checks. If the wings are commanded to unlock or fold

during a catapult shot, the wings will unlock, the ailerons will fair, the

wings may fold partially, and the aircraft will settle.

2.9.4 Wingfold Overheat Cutout Protection. The wingfold electric drive units are designed to meet

the following duty cycle requirements: two (2) fold-spread cycles followed by a twelve (12) minute

cooldown period. If wingfold operation exceeds this duty cycle, overheat cutout protection may

shutdown wingfold operation to prevent actuator damage. Once overheat cutout protection has been

activated, normal wingfold operation is not restored until actuator temperature drops within limits;

however, the wings can still be unlocked, folded, or spread manually.

2.10 FCS - FLIGHT CONTROL SYSTEM

The flight control system (FCS) is a fly-by-wire, full authority control augmentation system (CAS).

The FCS provides four basic functions: aircraft stability, aircraft control, departure resistance, and

structural loads management. Since the basic airframe is statically neutral to slightly unstable, a

primary function of the FCS is to maintain aircraft stability at all flight conditions. The FCS also

provides full authority control of the aircraft by implementing the basic flight control laws which

determine aircraft response to pilot inputs. Pilot inputs from the stick and rudder pedals send

electrical commands to two quad-redundant, digital flight control computers (FCC A and FCC B).

There is no mechanical linkage between the stick and rudder pedals and the flight control surfaces.

FCC software determines what commands are sent to the various flight control surfaces to exercise

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pitch, roll, and yaw control of the aircraft. Additionally, the FCS provides departure resistance by

either refusing to accept or by tailoring pilot inputs that would otherwise lead to an aircraft departure.

Lastly, the FCS provides structural loads management by limiting g-available to prevent an aircraft

overstress or by retracting flight control surfaces at airspeeds that would otherwise exceed the

structural limits of the airframe. See figure 2-19 for a functional diagram of the flight control system.

2.10.1 Flight Control Surfaces. The aircraft has 12 primary flight control surfaces including leading

edge flaps (LEFs), trailing edge flaps (TEFs), ailerons, twin rudders, horizontal stabilators, and

spoilers. LEFs, TEFs, ailerons, and stabilators can be moved both symmetrically or differentially for

pitch and roll control. Flight control surface deflection limits are shown in figure 2-20.

Pitch control is accomplished with symmetric stabilators and, in some conditions, with rudder toe-in

or rudder flare. Roll control is accomplished with combinations of ailerons, differential stabilators,

differential LEFs, and differential TEFs dependent on flight condition and CAS operating mode. The

twin rudders deflect symmetrically for directional control. There is no dedicated speedbrake surface.

Instead, a ″speedbrake function″ is provided by partial deflection of several of the primary flight

control surfaces.

Hydraulic power to all flight control surface actuators is supplied by HYD 1 and HYD 2. Stabilator

and TEF actuators are powered simultaneously by one HYD circuit from each system. All other

actuators are powered by a single primary HYD circuit, with backup hydraulic power available through

a hydro-mechanical switching valve. See the Hydraulic System section, specifically the Hydraulic Flow

Diagram, to determine which HYD circuits power each flight control surface actuator.

Surface

Deflection limits *

Aileron

25° TEU to 42° TED

Rudder

40° left or right

Stabilator

24° TEU to 20° TED

LEF

5° LEU to 34° LED

TEF

8° TEU to 40° TED

LEX Spoilers

0° or 60° TEU

* Tolerance ±1°, or ±3° for spoilers.

Figure 2-19. FCS Surface Deflections

2.10.1.1 Spoilers. The spoilers are mounted on top of the fuselage near the aft end of the LEX. The

spoilers are controlled by the FCCs and have two fixed positions: 0° (down) or 60° TEU. The 60° TEU

position is activated by the speedbrake function or when more than 15° TED stabilator is commanded

(forward stick) above 22° AOA to aid in recovery from high AOA.

2.10.2 FCCs - Flight Control Computers. Two flight control computers (FCC A and FCC B) provide

the computations which implement the aircraft’s flight control laws. A four-channel architecture is

used to provide FCS redundancy. Each FCC contains two individual central processing units (CPUs),

which each run one channel of the FCS. CH 1 and CH 2 are resident in FCC A, with CH 3 and CH 4

in FCC B.

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Most inputs to the FCCs (rate gyros, accelerometers, air data sensors, stick and rudder pedal

position sensors) are quad-redundant, one input for each channel. Each of the four CPUs runs

independent and parallel flight control computations. Sensor inputs as well as CPU outputs are

continuously monitored by the FCCs for agreement. When there is disagreement, the erroneous signal

is discarded, if possible.

Rate and acceleration data are provided by two independent Attitude and Heading Reference Sets

(AHRS), one for each FCC. Each AHRS has two sets of ring laser rate gyros and two sets of

accelerometers, which provide four independent sources of pitch, roll, and yaw rate information, and

four independent sources of normal and lateral acceleration. The AHRS units have the capability to

provide attitude, heading, and longitudinal acceleration data, but it is not currently utilized. The

physical rate and acceleration sensors in each AHRS channel are not aligned with the aircraft’s pitch,

roll, and yaw axis. This raw sensor data is converted to the aircraft’s pitch, roll, and yaw axis by

microprocessors internal to each AHRS. As a result of this architecture, a single rate gyro failure in one

channel results in all three axis rates being unusable in that same channel (CAS P, R, Y in one channel

Xd out). Similarly, if any of the accelerometers fail, all acceleration data from that AHRS channel is

unusable (N ACC and L ACC in one channel will both be Xd out).

FCC channel outputs are transmitted to the appropriate flight control actuators and to other aircraft

systems such as the MCs. While FCC computations run in all four channels, all flight control actuators

are not commanded in all four channels. The stabilators and TEF actuators do receive command

signals from all four FCC channels. However, each aileron, rudder, spoiler, and LEF actuator only

receives command signals from two FCC channels, one from FCC A and one from FCC B. The

2-channel actuators on the left side of the aircraft receive inputs from CH 1 and CH 4 while the

2-channel actuators on the right side receive inputs from CH 2 and CH 3. This channel distribution can

be seen on the FCS format.

2.10.2.1 FCC Temperature Monitoring. FCC A contains a thermocouple which monitors the

temperature within the computer and provides a signal to FCC CH 1 and CH 2. If an over-temperature

condition is detected, the FCS HOT caution and caution light come on, and the ″Flight computer hot,

Flight computer hot″ voice alert annunciates. Additionally, FCC A indicates OVRHT on the BIT

status line. In this case, placing the AV COOL switch to EMERG provides emergency ram air cooling

to FCC A and the right TR via a dedicated FCS ram air scoop. FCC B also contains a thermocouple,

but does not set the FCS HOT cautions. The only indication of an over-temperature condition in FCC

B is a BIT status indication of OVRHT.

2.10.2.2 AV COOL Switch. The AV COOL switch is located on the lower right main instrument panel

outboard of the caution light panel.

NORM

FCS ram air scoop retracted.

EMERG Deploys the FCS ram air scoop for emergency ram air cooling of FCC A, the right

TR, and other essential avionics.

Once deployed, the FCS ram air scoop cannot be retracted inflight.

2.10.3 FCS Redundancy and Survivability. Hydraulic redundancy is provided by distributing flight

control actuators among the four HYD circuits. This arrangement minimizes the probability of losing

multiple actuators due to catastrophic damage to any single actuator or its hydraulic lines. Following

a single HYD system failure, the other HYD system is capable of powering the entire FCS. Loss of

HYD 1 or HYD 2 in up and away flight does not affect aircraft control. However, in the takeoff and

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landing configuration, small but controllable roll and/or yaw excursions may be expected as hydraulic

switching valves cycle to their backup circuits.

The primary electrical power source for each FCC channel is a dedicated output from one of two

permanent magnet generators (PMGs). See the Electrical System section for FCC Electrical Redun-

dancy. Should a power interruption occur to any single FCC channel, the FCC power supply

automatically switches to a ″keep alive″ circuit connected directly to the maintenance bus for 7 to 10

seconds. This makes sure that the FCCs have uninterrupted power to maintain full operation during

all predictable electrical bus switching transients.

For survivability, wiring for one channel from each computer is routed through the upper part of the

aircraft with wiring for the other through the lower part of the aircraft. This routing minimizes the

possibility of loss of any one flight control surface due to system failures or battle damage. If a

stabilator actuator fails due to multiple FCS or hydraulic failures, the FCS automatically reconfigures

to maintain 3-axis control and acceptable handling qualities by using the remaining surfaces. There is

no mechanical FCS reversion mode.

2.10.4 CAS Operating Modes. The control augmentation system (CAS) operates in two basic modes:

Powered Approach (PA) and Up-AUTO (UA). Mode selection is controlled by FLAP switch position

and airspeed. With the FLAP switch in HALF or FULL and with airspeed below approximately 240

KCAS, CAS implements flight control laws tailored for the takeoff and landing configuration (PA).

With the FLAP switch in AUTO, CAS implements flight control laws tailored for up and away flight

(UA). If the FLAP switch is left in HALF or FULL, the aircraft automatically transitions from PA to

UA when airspeed increases above approximately 240 KCAS. This is known as ″auto flap retract.″ In

this case, the amber FLAPS light comes on to alert the pilot to check FLAP switch position. The flight

control laws utilized in each mode are tailored to provide maximum maneuverability while maintaining

predictable handling qualities and departure resistance.

2.10.4.1 FLAP Switch. The FLAP switch, located on the lower left main instrument panel, is used to

select the CAS operating mode and to position the TEFs and aileron droop for takeoff and landing.

AUTO

Selects UA operating mode for up and away flight.

HALF

Selects PA operating mode for the takeoff and landing configuration. Sets TEF

deflection and aileron droop to 30° TED (WonW or at approach speed).

FULL

Selects PA operating mode for the takeoff and landing configuration. Sets TEF

deflection and aileron droop to 40° TED (WonW or at approach speed).

2.10.4.2 Flap Position Lights. Three flap position lights, two green and one amber, are located on the

lower left main instrument panel. The green HALF and FULL flap lights are used to indicate FLAP

switch position and are not indications of actual TEF/aileron position. The FCS format should be

referenced to determine actual LEF, TEF, and aileron position.

FLAPS

FLAP switch in HALF or FULL and airspeed above 240 KCAS (auto flap retract),

(amber)

abnormal flap condition (any flap is off or lacks hydraulic pressure), spin detected

by Spin Recovery System, or GAIN switch in ORIDE.

HALF

FLAP switch in HALF and airspeed below 240 KCAS.

(green)

FULL

FLAP switch in FULL and airspeed below 240 KCAS.

(green)

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2.10.5 Control Augmentation System (CAS).

2.10.5.1 Pitch CAS. Pitch CAS (P CAS) utilizes normal acceleration, pitch rate, and AOA feedback,

each scheduled based on aircraft flight conditions, to tailor aircraft response to pilot stick inputs and

to provide stabilator actuator commands. P CAS operates by comparing aircraft response to the pilot’s

longitudinal stick input, driving the stabilator actuators symmetrically until the difference is reduced

to zero.

In UA, with neutral longitudinal stick, comparing pilot input to aircraft response has the effect of

constantly trimming the aircraft to steady-state, hands-off

1g flight, essentially removing the

requirement for manual trim. In maneuvering flight, P CAS modifies aircraft response to stick inputs

creating the effect of changing stick forces to provide pilot cueing. Actual stick forces for a given stick

displacement do not change with flight condition. At high airspeeds, P CAS is a g-command system

requiring 3.5 pounds of stick-force-per-g. At medium airspeeds, P CAS acts as a hybrid pitch rate and

g-command system. Pitch rate feedback is used to increase apparent stick-force-per-g (heavier stick

forces) to cue the pilot that airspeed is decreasing and less g is available. At low airspeed, P CAS is

primarily an AOA command system using AOA feedback above 22° AOA to provide increasing stick

forces with increasing AOA. With large forward stick inputs, P CAS augments nose-down pitch rates

by flaring the rudders and raising the spoilers.

In PA, AOA and pitch rate feedbacks are used to augment inherent airframe pitch damping and

stability. P CAS nulls the difference between the commanded AOA and actual AOA. With neutral

longitudinal stick, P CAS maintains trim AOA. Unlike UA, pitch trim is required in PA to trim the

aircraft on-speed. Rudder toe-in is used to improve longitudinal stability and to aid aircraft rotation

during takeoff or bolter. Rudder toe-in is a function of AOA. At 0° AOA or with WonW, the rudders

are toed-in 40°. Rudder toe-in decreases linearly to 0° of toe at 12° AOA. Additional AOA feedback is

provided above 12° AOA which increases stick forces with increasing AOA to provide stall warning.

Pitch rate feedback helps maintain tight pitch attitude control during turns.

2.10.5.2 Roll CAS. Roll CAS (R CAS) schedules aileron, differential LEF, differential TEF, and

differential stabilator commands in response to lateral stick inputs to achieve the desired roll

characteristics. Roll rate feedback, scheduled based on aircraft flight conditions, is used to augment

inherent airframe roll damping. Differential LEFs and TEFs are only used in UA. The LEFs deflect

differentially up to 5° when below 25,000 feet and above Mach 0.6. Differential TEFs are not used

above 10° AOA or below -5° AOA. At high airspeeds, aileron, differential stabilator and differential

TEF travel are reduced to provide consistent roll rate response and to aid in preventing structural

loads exceedances. At low airspeeds, aileron and differential stabilator travel are reduced with

increasing AOA to minimize adverse yaw. Differential stabilator may also be limited due to pitch

commands which have priority over lateral commands.

With clean wing or A/A missile loadings (no wing tanks), maximum roll rate is limited to

approximately 225°/second. With A/G store or external fuel tank codes set in the armament computer

for any wing station and the pylon rack hooks closed for those stations, maximum roll rate is limited

to approximately 150°/second to avoid exceeding pylon structural load limits. If all stores are shown as

HUNG, roll rate limiting is removed; however, an R-LIM OFF caution appears on the DDI.

R CAS incorporates two features to reduce pitch-roll inertial coupling induced departures. Based on

pitch rate and Mach number, the first feature reduces the roll command when the pilot applies an

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excessive combined lateral/longitudinal stick input. The second feature limits the roll command when

the aircraft is already rolling and longitudinal stick is moved rapidly. This second feature is removed

at low altitude and high speed since available pitch rate does not result in significant pitch-roll inertial

coupling.

2.10.5.3 Yaw CAS. Yaw CAS (Y CAS) uses yaw rate and lateral acceleration feedback to provide

directional axis damping and to augment pilot commands to the twin rudder actuators. A rolling-

surface-to-rudder interconnect (RSRI) adjusted by roll-rate-to-rudder crossfeed (scheduled with

AOA), and lateral acceleration feedback are used to minimize sideslip for roll coordination. To provide

departure resistance and enhanced maneuverability at high AOA, directional stability is augmented

utilizing INS pitch and roll attitudes along with the FCS sensors to synthesize sideslip and sideslip rate

feedback to the ailerons and differential stabilators. These lateral surfaces are used in this sense as

directional controllers by taking advantage of the strong yawing moments they produce at high AOA.

Below 13° AOA, rudder pedal deflections provide yaw by symmetric rudder deflection. At 25° AOA

and above, rudder pedal deflections no longer provide yaw control inputs but instead act entirely as a

roll controller (identical to lateral stick input) by commanding aileron and differential stabilator with

the RSRI commanding the required rudder deflection for roll coordination. Rudder pedal inputs are

summed with lateral stick inputs and this combined input is limited to a value equal to a maximum

lateral stick input. Therefore, applying pedal opposite to lateral stick cancels lateral stick inputs

proportional to the pedal input, e.g., full opposite pedal cancels a full lateral stick command resulting

in zero roll rate. Between 13° and 25° AOA, rudder pedal deflection gradually changes from pure yaw

control to pure roll control. This method of control provides enhanced departure resistance at high

AOA.

Some traditional directional control capability is returned at low airspeed and high AOA only when

the pilot applies lateral stick and rudder in the same direction. This feature starts becoming effective

only at airspeeds below approximately 225 KCAS, from 20° to 40° AOA, but is most effective at

approximately 170 KCAS and 34° AOA. Enabling this feature outside of these conditions would

compromise departure resistance. When this feature is enabled, the sum of lateral stick and rudder

pedal command is no longer limited to a value equal to a full lateral stick input. The excess roll

command is fed to the directional axis to command sideslip. For example, adding full rudder pedal with

a full lateral stick input provides a maximum roll and yaw command. Alternatively, adding lateral stick

to an existing full rudder pedal input has the same effect. The resulting aircraft motion is a highly

controllable nose-high to nose-low reversal.

At high airspeeds, symmetric rudder deflection is reduced and the rudders are toed in to avoid

exceeding vertical tail structural limits.

In PA mode, synthesized sideslip rate feedback augments aerodynamic directional damping and

stability.

2.10.5.4 Flap Scheduling. In UA, LEFs, TEFs, and aileron droop are scheduled as a function of AOA

and air data to optimize cruise and turn performance, to improve high AOA characteristics, and to

provide load alleviation (when required). In general, LEFs start to deflect as AOA increases above

approximately 3°, reaching full deflection (34° LED) by about 25° AOA. In general, TEFs start to

deflect above 2 to 3° AOA, are at full scheduled deflection (approximately 10 to 12° TED) from

approximately 6 to 15° AOA, and begin to retract as AOA increases further. In other words, TEFs are

deflected in the heart of the maneuvering envelope to produce more lift and are retracted at high AOA.

In UA, aileron droop is scheduled to 50% of TEF deflection at low AOA and to 0° at high AOA.

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In UA, flap scheduling is altered slightly based on the presence of wing tanks. With wing tanks

installed, TEF deflection is slightly lower at most flight conditions. LEFs and TEFs typically begin to

deflect at slightly slower Mach but follow the same trends as those mentioned above.

In PA, LEFs are scheduled as a function of AOA to maximize lift. TEFs are scheduled as a function

of airspeed for load alleviation but should be at maximum scheduled deflection at approach speed. In

PA, aileron droop is scheduled with TEF deflection. Following field takeoff or catapult launch,

TEF/aileron droop is latched for 10 seconds after the transition to WoffW. This feature is designed to

improve catapult launch characteristics by ensuring the flaps do not retract immediately after launch.

However, if approximately 190 KCAS is exceeded prior to expiration of the 10 second timer, the TEFs

and aileron droop do begin to retract for loads alleviation. LEF, TEF, and aileron droop scheduling are

shown in figure

2-21.

FCS Mode

Configuration

Status

LEF Position

TEF Position AIL Droop

WonW

3° LED

2° TED

1° TED

50% of TEF

Scheduled

No Wing Tanks

Scheduled

(<10° AOA),

WoffW

with

with M,AOA

0°

M,AOA,Alt

(>15° AOA)

WonW

3° LED

4° TED

2° TED

50% of TEF

Scheduled

Wing Tanks

Scheduled

(<10° AOA),

WoffW

with

with M,AOA

0°

M,AOA,Alt

(>15° AOA)

WonW

15° LED

30° TED

Flaps HALF

Scheduled

30° TED

WoffW

with AOA

(on-speed)

WonW

15° LED

40° TED

Flaps FULL

Scheduled

40° TED

WoffW

with AOA

(on-speed)

Figure 2-21. Flap Schedules

2.10.6 Speedbrake Function. The aircraft is not fitted with independent speedbrake surfaces. A

″speedbrake function″ is provided to increase drag by partial deflection of several of the aircraft’s

primary flight control surfaces: ailerons, rudders, TEFs, and spoilers. The stabilators are commanded

to counter pitch transients during speedbrake extension and retraction. The full speedbrake function

can only be commanded in UA.

At subsonic speeds in UA, the speedbrake function flares the rudders and symmetrically raises the

ailerons TEU to approximately 95% of the capability of each surface at the given flight conditions.

This makes sure approximately 5% of surface authority is available for yaw and roll control. If needed,

rudder and aileron priority is given to yaw and roll commands. TEFs are also symmetrically lowered

to further increase drag and to counter the loss of lift caused by deflecting the ailerons TEU. The

spoilers are raised to the full up 60° position only when the speedbrake command reaches 75%. At

subsonic speeds, the stabilator is used to offset any pitch transients that occur due to the deflection of

all speedbrake surfaces except the spoiler. Delaying spoiler deflection until 75% allows the pilot to use

partial speedbrakes for speed modulation, while avoiding minor spoiler induced pitch transients.

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At supersonic speeds in UA, speedbrake surface deflections are changed. The rudders are not

deflected above 1.05 M due to vertical tail loads. The ailerons and TEFs are not deflected above 1.1 M

due to a lack of effectiveness. The spoilers are therefore deflected immediately upon speedbrake

actuation, since they are the only effective surface at these conditions. At supersonic speeds, the

stabilator is used to counter spoiler deployment. The speedbrake function is completely disabled above

1.5 IMN.

In UA, the speedbrake function is ramped out above 16° AOA or below -9° AOA to preserve

lateral-directional stability and between -3.0 to -1.5g for airframe loads.

In PA, the speedbrake function is disabled with WoffW. With WonW and the FLAP switch in HALF

or FULL, the speedbrake function only deploys the spoilers. While the spoilers can be deployed during

landing rollout or aborted takeoff, the drag increase is minimal, and rollout distance is not appreciably

decreased. With WonW and the FLAP switch in AUTO, full extension of the speedbrake function

commands 20° of rudder flare, 23° of TEU aileron, 7° of TED TEF, 60° of spoiler, and a 2° TED

stabilator change.

2.10.6.1 Speedbrake Switch. The speedbrake switch, located on the inboard side of the right throttle

grip, is used to enable/disable the speedbrake function, e.g., extend/retract the speedbrake surfaces.

The forward and center positions are detented, while the aft position is spring-loaded back to center

and must be held. In the F/A-18F (trainer configuration), the rear cockpit speedbrake switch has

override priority over the front cockpit switch.

Forward

Retracts speedbrake surfaces (full retraction in 2 seconds).

(unmarked)

Center

Stops speedbrake surfaces at an intermediate position.

(unmarked)

Aft

Extends speedbrake surfaces (full extension in 2 seconds).

(unmarked)

NOTE

• If the speedbrake switch is held or fails in the aft position for more

than 5 minutes, the speedbrake switch is declared failed, and the FCS

caution is set. If the switch is failed or is held in the aft position when

the FCS RESET button is pushed, the speedbrake surfaces are

retracted, Xs are set on the DEGD row of the FCS page, and the

speedbrake function is disabled for the remainder of the flight.

• If the front cockpit switch is held in the aft position during any FCS

RESET attempt, the speedbrake switch is declared failed; the speed-

brake surfaces are retracted; and the speedbrake function is disabled

for the remainder of the flight. This allows the speedbrake surfaces to

be retracted before the 5 minute timer expires, if the front cockpit

switch is stuck in the aft position.

• In the F/A-18F (trainer configuration), if the rear cockpit switch fails

in the aft position, the

5 minute timer must expire before the

speedbrake surfaces can be retracted with the FCS RESET button.

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2.10.6.2 SPD BRK Light. The green SPD BRK light is located on the left warning, caution, and

advisory panel on the main instrument panel. The SPD BRK light comes on anytime the speedbrake

surfaces are not fully retracted.

2.10.7 G-Limiter Considerations. In order to understand what protection the aircraft’s g-limiter

provides, pilots must understand the difference between ″design limit-g″ and ″reference load factor (Nz

REF).″ See the Acceleration Limitations chart in the Operating Limitations chapter for a plot of Nz

REF versus gross weight, for g-limiter specifics, and for gross weight related g-restrictions.

2.10.7.1 Design Limit-g. The aircraft was designed to sustain a limit-g of +7.5g or -3.0g (symmetric)

only at or below its fighter design gross weight of 42,097 lb. At higher gross weights, design limit-g is

reduced to keep from exceeding the structural limitations of the airframe. An ″overstress″ is defined as

a g-level that exceeds the design limit-g at the aircraft’s current gross weight. Above 42,097 lb gross

weight, design limit-g is reduced by the aircraft’s relative gross weight (42,097/GW), such that the

positive design limit is +7.5g * (42,097/GW) and the negative design limit is -0.4 * (positive limit-g). At

the aircraft’s maximum gross weight (66,000 lb), design limit-g is only +4.8g or -1.9g.

Due to the increased airframe and pylon loads that accompany high-g rolling maneuvers, the aircraft

also has a design limit-g for abrupt full-stick rolls (FSR). Abrupt FSRs are defined as full lateral stick

in less than 1 second. The positive design FSR limit is +6.0g below 42,097 lb GW and 80% of the

symmetric design limit-g above 42,097 lb. The negative design FSR limit is -1.0g at all gross weights.

At 66,000 lb GW, the positive design FSR limit is only +3.8g.

2.10.7.2 Reference Load Factor (Nz REF). Reference load factor (Nz REF) is the value that the MC

uses to set the g-limiter when outside of the transonic g-bucket (described below). With increasing

gross weight, Nz REF is the same as design limit-g until the gross weight where +5.5g (-2.2g) is

available (57,405 lb GW). Above 57,405 lb GW, Nz REF is held fixed at +5.5g (-2.2g) in order to assure

that the pilot always has those g-levels available even if they would result in an overstress. Since the

g-limiter may not prevent an overstress at gross weights above 57,405 lb, the pilot must be responsible

for preventing an overstress in this gross weight region.

2.10.7.3 G-Limiter. The g-limiter essentially limits the amount of positive and negative g that can be

commanded by the pilot at a particular gross weight in order to prevent an aircraft overstress. Once the

pilot reaches the stick displacement required to attain the Nz REF g-limit, further stick inputs do not

increase g. This is commonly called ″being on the limiter.″ Once the stick is relaxed to the limit

displacement, g-control below Nz REF is regained. The g-limiter functions to maintain both the

positive and negative Nz REF limits.

During abrupt longitudinal stick inputs, g-limiter overshoots are not uncommon. G-limiter over-

shoots of up to +0.5g or -0.2g are allowed and do not constitute an over-g. An ″over-g″ is defined as a

g-level which exceeds the overshoot thresholds and sets MSP code 811 (positive exceedance) or 925

(negative exceedance). An over-g condition requires a postflight inspection to determine if an

″overstress″ occurred.

For rolling maneuvers commenced above the positive FSR limit, the g-limiter also provides some

protection. In this region, the g-limiter attempts to reduce commanded-g towards the positive FSR

limit to prevent an overstress. However, if the rate of lateral stick input exceeds the capability of the

g-limiter, an actual rolling overstress may result without setting an 811 MSP code (set only if the

symmetric over-g threshold is exceeded). The g-limiter treats rolling maneuvers with less than ¾ inch

lateral stick as symmetric maneuvers.

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A G-LIM 7.5G caution, accompanied by the FLIGHT CONTROLS, FLIGHT CONTROLS voice

alert, is set for any of the following: FUEL XFER, CAUT DEGD, MC2, SMS failure, or an invalid fuel

quantity. A G-LIM 7.5G caution indicates that the positive symmetrical command limit has been set

to +7.5g regardless of gross weight or stores loading. If the G-LIM 7.5G caution is set, the pilot must

limit commanded g-level to prevent an overstress.

Very high g-onset rates are possible with rapid aft stick movement, with

or without g-limiter override. A very high g-onset rate can cause imme-

diate loss of consciousness (G-LOC) without the usual warning symptoms

of tunnel vision, greyout, and blackout. The effects of G-LOC may last 20

seconds or longer after the g level is reduced to near 1.0g.

2.10.7.4 G-Bucket. Due to the aerodynamic phenomenon known as transonic pitch-up, the g-limiter

incorporates a g-bucket designed to prevent an aircraft positive over-g during transonic deceleration.

In the g-bucket, the g-limiter reduces the positive command g-limit below Nz REF (figure 2-22). This

reduction is a maximum of 1.0g above 20,000 feet, and 1.7g below 15,000 feet. For example, if Nz REF

is +7.5g and altitude is ≤15,000 feet, the g-limiter only allows +5.8g to be commanded while in the

g-bucket. The symmetrical command limit is never reduced below +4.5g.

NOTE

• G-bucket reduction reduces maximum commandable-g.

• Magnitude of transonic pitch-up increases as rate of Mach change

increases. High drag loadings with idle power settings generally have

the largest transonic pitch-ups. High drag loadings (e.g., A/G stores)

have a g-bucket that extends into a lower Mach range.

• Largest measured transonic pitch-up was 2.2g for <15,000 feet. This

magnitude pitch-up was seen on both A/A and A/G loadings.

The 0.2g deep mini-g-bucket extension in the Mach 0.85 to Mach 0.94 range was added in the FCC

OFP to protect against over-g in that region.

The Mach range for the deeper part of the g-bucket is dependent on external stores configuration.

The deeper part of the g-bucket is entered at Mach 0.905 accelerating (with at least one wing tank or

A/G wing store) or at Mach 0.941 accelerating with no wing tanks or A/G stores. Regardless of stores

loading, when decelerating, the g-bucket is entered at Mach 1.045 and is exited at Mach 0.83.

Full stick roll (FSR) limits are reduced in the g-bucket to 80% of (Nz REF minus no more than a

1.0g reduction). For example, if Nz REF is +7.5g, Mach is 0.95, and altitude is ≤15,000 feet, the

g-limiter sets the FSR limit to 80% of (Nz REF minus 1.0) even though the bucket depth is 1.7g.

If the pilot wants to have maximum-g available during a turning maneuver (e.g., the merge), or avoid

the deeper part of the g-bucket, IMN should be 0.90 or less (with at least one wing tank or A/G store),

or 0.93 or less (clean or with A/A stores). Note that even though the 0.2g mini-g-bucket might be active,

flight testing has shown the g-level will be at or slightly above Nz REF.

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2.10.7.5 G-Limiter Override. A g-limiter override feature can be enabled to allow a 33% increase in

the command g-limit for emergency use (allows a 10g command at 7.5g Nz REF). G-limiter override is

selected by momentarily pressing the paddle switch when the stick is near the full aft limit. When

g-limiter override is selected, a G-LIM OVRD caution is set along with a 927 MSP code. Override is not

disengaged until the stick is returned to near the neutral position.

2.10.7.6 Roll Rate Limiting. Roll rate limiting is enabled in R CAS when external wing tanks or A/G

stores are mounted on wing pylons (hooks closed). If any A/G store indicates HUNG, a R-LIM OFF

caution is set and roll rate limiting is removed. In this case, higher than normal roll rates are possible

and may exceed the structural limitations of the airframe/pylons if pilot-imposed lateral stick limits

are not applied.

2.10.8 Air Data Function. The air data function is provided by the FCCs and not a separate

computer. The FCCs receive input from pitot-static sensors, total temperature sensors, the angle of

attack probes, the standby altimeter barometric setting, and the mission computers. The FCC air data

function applies appropriate source error corrections to the air data sensor inputs and calculates

accurate true altitude, airspeed, Mach number, AOA, and outside air temperature (OAT). Computed

air data is used internally by the FCC control augmentation system (CAS) and is also supplied to the

MCs for IFF altitude reporting, weapon system calculations, and landing gear wheels warning, to the

FADECs for engine control, and to the ECS controller for ECS scheduling and fuel tank pressurization

and vent.

2.10.8.1 Pitot-Static/Total Temperature Probes. Two combined pitot-static/total temperature

probes are mounted on the left and right forward fuselage. Each probe contains one pitot pressure

source, two static pressure sources, and a total temperature sensor. One of the static pressure sources

from each probe is connected together and pneumatically averaged. This average static source is

provided to the left and right pressure transmitter sets along with the corresponding pitot pressure

source. The left pressure transmitter set provides pitot-static input to FCC CH 1 and 4 with the right

providing input to FCC CH 2 and 3. The FCC air data function corrects sensed pitot and static

pressures for position error to provide accurate true air data for FCC calculations, MC calculations, and

display in the HUD. The pitot pressure source and the second static pressure source from the left probe

are used to drive the standby flight instruments (altitude, airspeed, and VSI). The second static

pressure source from the right probe is unused.

Each pitot-static probe also contains an integral total temperature sensor. Each total temperature

sensor converts sensed temperature to an electrical signal. The output of the left total temperature

sensor is sent to FCC CH 2, with the right to FCC CH 4. The FCCs use total temperature to calculate

OAT. Each pitot-static/total temperature probe is electrically heated to prevent icing.

2.10.8.2 AOA Probes. Two AOA probes are mounted on the left and right forward fuselage. Each

probe mechanically measures local AOA by aligning with the airstream. An integral AOA transmitter

set converts the mechanical input to a two-channel electrical signal which is sent to the FCCs. The left

AOA transmitter set provides input to FCC CH 1 and 4 with the right to FCC CH 2 and 3. The FCC

air data function corrects the sensed local AOA to a true AOA and provides the output to the MC for

display on the HUD. FCS CH 4 supplies the AOA signal which drives the AOA indexer lights and the

approach lights.

It is possible to damage and jam an AOA probe such that it continues to send signals to the FCCs.

FCC software is designed to minimize flying qualities degradation in the event of a stuck/jammed AOA

probe. The FCCs incorporate an AOA estimator which is used to identify the good AOA probe if one

is damaged. If an AOA probe split is transient, the estimator is used to identify the good probe and no

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cautions are set. If the AOA probe split persists, an FCS caution is set, AOA is Xd in all four channels,

and the estimator is used for FCC calculations. See the HUD Symbology Degrades with Air Data

Function Failure paragraph and Part V for more details on AOA failures. Each AOA probe and AOA

probe cover are electrically heated to prevent icing.

2.10.8.3 PITOT ANTI ICE Switch. The PITOT ANTI ICE switch is located on the ECS panel on the

right console. This switch is used to power the electric heaters for the pitot-static/total temperature

probes, the AOA probes, and the AOA probe covers. All heaters are thermostatically controlled to

prevent damage to their corresponding sensors. With WonW, the thermostat set points are reduced to

prevent damage when cooling airflow is not provided.

Pitot and AOA heaters on manually (WonW or WoffW).

AUTO

Pitot and AOA heaters on automatically with WoffW. Heaters off with WonW.

Failure of both AOA probe heaters in icing conditions may cause a sharp

uncommanded nose down attitude, uncontrollable by normal stick forces

or paddle switch actuation.

2.10.9 Flight Controls.

2.10.9.1 Stick. A traditional center mounted control stick is used to provide pitch and roll inputs to

the FCS. Since there is no mechanical linkage between the stick and the FCCs or the flight control

surfaces, stick feel is provided by two feel-spring assemblies and two eddy current dampers. The feel

spring assemblies provide a linear stick force versus stick displacement gradient in each axis. Two

4-channel position sensors, one in each axis, measure stick displacement and send longitudinal and

lateral stick commands to the FCCs proportional to stick displacement. Stick force and displacement

are listed in figure 2-23 for full stick travel. The eddy current dampers provide stick motion damping

in each axis. Additionally, the control stick is mass balanced to minimize longitudinal stick movement

resulting from accelerations normally experienced during catapult launch.

In the F/A-18F (trainer configuration), a control stick is also fitted in the rear cockpit and is

mechanically linked to the one in the front cockpit.

Displacement Force

Direction

(in)

(lbs)

Forward

2.5

Stick

Aft

5.0

Left/Right

3.0

Pedal

Left/Right

1.0

100

Figure 2-23. Stick and Pedal Travel Limits

2.10.9.2 Rudder Pedals. Two rudder pedals (left and right) are used to provide directional inputs to

the FCS for yaw/roll control inflight or NWS control with WonW. Since there is no mechanical linkage

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between the rudder pedals and the FCCs or the flight control surfaces, rudder pedal feel is provided

by two feel-spring assemblies. The feel spring assemblies provide a linear pedal force versus

displacement gradient. Two 4-channel position sensors, one on each pedal, measure pedal displace-

ment and send directional commands to the FCCs proportional to pedal displacement. Rudder force

and displacement are listed in figure 2-23 for full pedal travel. The rudder pedals are also used to

provide NWS commands and to actuate toe-operated wheel brakes.

In the F/A-18F (trainer configuration), two rudder pedals are also fitted in the rear cockpit but are

not mechanically linked to the rudder pedals in the front cockpit. Pedal inputs from either cockpit are

summed together and transmitted to the FCCs. A half pedal input from the front cockpit and a half

pedal input from the rear cockpit results in a full rudder pedal command to the FCCs. Similarly,

opposing rudder pedal inputs in each cockpit cancel each other.

2.10.9.2.1 RUD PED ADJ Lever. A RUD PED ADJ lever, located on the center pedestal in each

cockpit, is spring loaded to the up and locked position. When the lever is held down, the rudder pedals

are unlocked and can be moved forward and aft in ½ inch increments. Both pedals are spring loaded

to move aft and must be pushed forward to the desired position. Releasing the RUD PED ADJ lever

locks the pedals in the new position.

• Restrain the rudder pedals during adjustment. Unrestrained rudder

pedals may damage the rudder pedal mechanism.

• Ensure the rudder pedals are locked in position after adjustment.

Failure to lock the rudder pedals may result in uncommanded forward

rudder pedal movement inflight.

2.10.9.3 Stick Grip FCS Controls. The FCS controls located on the stick grip include the pitch and

roll trim switch, the NWS button, and the autopilot/NWS disengage switch. See figure 2-24. In the

F/A-18F (trainer configuration), the FCS controls on the rear cockpit stick grip are identical to those

in the front cockpit.

2.10.9.3.1 Pitch and Roll Trim Switch. The pitch and roll trim switch is located on the top right of

the stick grip. Movement of the pitch and roll trim switch electrically biases the FCCs and does not

reposition the stick.

Forward

Trims nose-down.

Aft

Trims nose-up.

Left

Trims left-wing-down.

Right

Trims right-wing-down.

Pitch and roll trim inputs can be made incrementally or held for faster trim rates. In UA, little if any

pitch trim is required due to the automatic trimming function provided by P CAS. In PA, pitch trim

is required to trim for on-speed AOA. Lateral trim is typically only required immediately after takeoff

or following changes in lateral weight asymmetry (fuel and/or stores). Pitch trim is not monitored for

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Figure 2-24. Stick Grip FCS Controls

runaway trim. However, roll trim is monitored for a stuck switch. If the trim switch is held or is stuck

in the left or right position for more than 40 seconds, the FCS caution is set and the roll trim function

of the switch is disabled for the remainder of the flight. Roll trim can be faded to zero by pushing and

holding the FCS RESET button or TO/TRIM button for approximately 4 seconds.

2.10.9.3.2 NWS Button. The undesignate/nosewheel steering button is located on the front of the

stick grip. The NWS button is used to engage NWS modes, as described in the NWS System

paragraphs in the Utility Hydraulic Functions section. The undesignate function of the NWS button

is described in the Weapon Systems Controls section.

2.10.9.3.3 Paddle Switch. The autopilot/NWS disengage switch, commonly called the ″paddle

switch,″ is located on the lower front of the stick grip. The paddle switch is used to disengage NWS with

WonW, to disengage all autopilot modes with WoffW, and to enable g-limiter override with WoffW. To

enable g-limiter override, the paddle switch must be momentarily pressed with the stick near the aft

limit.

2.10.9.4 RUD TRIM Knob. The RUD TRIM knob is located on the FCS panel on the left console in

the front cockpit only. Movement of the RUD TRIM knob electrically biases the FCCs and does not

reposition the rudder pedals. Rudder trim authority is ±10° and ±22.5° of rudder surface deflection in

UA and PA, respectively. PA rudder trim authority is set to allow zero pedal forces during a HALF flap,

single engine approach. Yaw trim is zeroed by mechanically centering the RUD TRIM knob when the

T/O TRIM button is pushed with either WonW or WoffW.

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2.10.9.5 T/O TRIM Button. The T/O trim button is located in the center of the RUD TRIM knob on

the FCS panel on the left console. With WonW, holding the T/O TRIM button pressed drives pitch

trim to 4° TEU stabilator, roll trim to neutral, and yaw trim to neutral by mechanically centering the

RUD TRIM knob. Depending on initial trim position the T/O TRIM button may need to be pressed

for up to 4 seconds. When these takeoff trim settings are reached, the TRIM advisory is displayed on

the LDDI for as long as the T/O TRIM button is held depressed. With WoffW, pressing the T/O TRIM

button for as long as 4 seconds drives roll trim to neutral, centers the RUD TRIM knob, but does not

affect pitch trim.

2.10.10 Yaw Rate Warning Tone. In UA (flaps AUTO), a yaw rate warning tone is provided to alert

the pilot of excessive yaw rate that may lead to an aircraft departure. The yaw rate warning tone is

generated by the FCCs and is initiated at 40°/second yaw rate with a 1 Hz pulse rate. The tone pulse

rate increases linearly as yaw rate approaches 60°/second, where the pulse rate remains constant at 10

Hz. There is no yaw rate warning tone in PA.

2.10.11 AOA Warning Tone. An AOA warning tone is provided to alert the pilot of excessive AOA

that may lead to aircraft settle and/or departure.

With flaps HALF or FULL, the AOA warning tone is triggered at 14° AOA with a 1 Hz pulse rate.

The tone pulse rate increases linearly as AOA approaches 35°, where the pulse rate remains constant

at 10 Hz.

With MC OFP H3E AND UP, with flaps AUTO, the AOA warning tone is triggered when the AOA

limits corresponding to the FLY lateral weight asymmetry are exceeded. A 500 ft-lb buffer is applied

to the lateral weight asymmetry threshold when triggering the tone to account for fuel slosh. If there

is an AOA failure, or more than one fuel quantity is invalid and/or weapon station indicates HUNG on

stations 2 − 10 (FLY value removed), the tone will not be triggered and the AOA TONE caution will

be displayed.

2.10.12 Spin Recovery System. The aircraft incorporates an automatic spin detection and recovery

system. A spin is declared when both of the following conditions are met: (1) airspeed is below

approximately 120 ±15 KCAS and (2) the yaw rate threshold is exceeded. The yaw rate threshold is

exceeded, for example, if a 15 to 20°/second yaw rate persists for approximately 15 seconds or a 50 to

60°/second yaw rate persists for approximately 2 seconds. For cases where the pilot is intentionally

commanding a high AOA roll (e.g., pirouette) the yaw rate threshold persistence is increased from 15

seconds to 25 seconds.

Immediately following a low-speed maneuver less than 77 KCAS (e.g., tail-slide), airspeed limits for

SPIN logic are opened to 180 KCAS or 12 seconds, whichever comes first. During this time, the normal

acceleration feedback gain is removed to avoid excess coupling, and spin mode arrows will be displayed

to aid recovery if yaw rate exceeds the threshold.

Once a spin has been detected, the spin recovery system places the SPIN MODE recovery displays

on both DDIs (figure 2-25), illuminates the amber FLAPS light, and drives the LEFs to 34° LED and

the TEFs to 4° TED. The displayed spin recovery arrow always indicates the proper direction for

anti-spin lateral stick inputs whether the spin is upright or inverted. Anti-spin lateral stick inputs are

aileron-into for upright spins and aileron-opposite for inverted spins. When lateral stick is placed with

the arrow, automatic spin recovery mode (ASRM) is engaged. With ASRM engaged, all CAS feedback

and control surface interconnects are removed, providing full aileron, rudder, and stabilator authority

for spin recovery. If the stick is neutral or is moved in the wrong direction, the SPIN MODE formats

remain displayed, the LEFs and TEFs remain deflected, but the FCS remains in CAS, and ASRM is

not engaged.

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Figure 2-25. SPIN Recovery Display

If ASRM is engaged during spin recovery, the spin arrow is removed and the FCS automatically

reverts to CAS when either of the following conditions are met: (1) airspeed is above approximately 245

KCAS or (2) the yaw rate threshold is no longer exceeded. After recovery, LEFs and TEFs return to

normal scheduling, and the SPIN MODE formats are replaced with the MENU page after 2 seconds.

NOTE

During highly oscillatory spins or spins that transition from upright to

inverted or from inverted to upright, the SPIN MODE displays may

disappear momentarily.

2.10.12.1 Spin Recovery Displays. When the spin recovery system detects an upright left spin or an

inverted right spin, a left spin arrow appears on both DDIs to indicate the proper direction of the

anti-spin lateral stick input.

SPIN MODE

STICK

LEFT

When the spin recovery system detects an upright right spin or an inverted left spin, a right spin

arrow appears on both DDIs to indicate the proper direction of the anti-spin lateral stick input.

SPIN MODE

STICK

RIGHT

When lateral stick is placed in the direction of the arrow, the word ″ENGAGED″ appears below the

words ″SPIN MODE″ on both DDIs to indicate that ASRM has been successfully engaged.

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SPIN MODE

is replaced by

SPIN MODE

ENGAGED

When the SPIN MODE formats appear on the DDIs, airspeed is always displayed in the upper left

corner, with altitude in the upper right and AOA in the lower center. See figure 2-23.

2.10.12.2 SPIN Switch. The SPIN switch is located on the right side of the main instrument panel.

The switch is guarded to prevent actuation. The SPIN switch was designed to allow for activation of

a manual spin recovery mode (MSRM). However, with all CAS feedback and control surface

interconnects removed, flight in MSRM will result in a departure and, once departed, will prevent

departure and/or spin recovery. SPIN arrow logic and ASRM functionality have been optimized and

thoroughly flight tested to produce accurate spin mode detection and positive spin recovery.

RCVY

Prohibited.

NORM

ASRM available when a spin is detected.

Selection of manual spin recovery mode (SPIN switch in RCVY) seri-

ously degrades controllability, will prevent recovery from any departure

or spin, and is prohibited.

2.10.13 Stabilator Failure Control Law Reconfiguration. Stabilator reconfiguration consists of

additional control laws which augment baseline CAS control laws to compensate for the complete loss

of a single stabilator. Stabilator reconfiguration is automatically enabled following the detection of a

complete stabilator failure (3 or more FCS Xs in a single stabilator or a dual HYD circuit failure - HYD

1B/2A or 1A/2B). If hydraulics are intact, the failed stabilator is driven to 2° TEU and locked.

Following a dual HYD circuit failure, the failed stabilator must be driven to the locked position by

aiding airloads. If unaiding airloads are applied, actuator mechanization prevents the stabilator from

moving further away from the locked position.

The reconfigured control laws are designed to compensate for the loss of the pitch and roll

contribution of the failed stabilator. In the pitch axis, pitch commands to the remaining stabilator are

doubled to produce more pitching moment. In PA, rudder toe-in and rudder flare are also used to aid

the pitching moment capability of the remaining stabilator. In the roll axis, differential stabilator

commands are disabled. A stabilator-to-rolling-surface interconnect is used to compensate for the roll

generated by single stabilator movement. In PA, this interconnect is stabilator to aileron. In UA, it is

stabilator to aileron and differential TEF. At high airspeed in UA, differential LEFs are also used. In

the yaw axis, the baseline differential stabilator portion of the RSRI continues to be used to counter

the yaw generated by single stabilator movement.

2.10.14 GAIN ORIDE. GAIN ORIDE allows the pilot to select a set of fixed CAS gains when an FCS

malfunction prevents normal CAS gain scheduling (e.g., loss of AOA or pitot-static data). With the

GAIN switch in ORIDE, the FCCs use fixed values for speed, altitude, and AOA depending on the

position of the FLAP switch. These fixed gains cause the LEFs, TEFs, and aileron droop to be driven

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to the fixed positions shown in figure 2-26. GAIN ORIDE should generally provide acceptable handling

qualities at flight conditions which approximate the fixed gains. At flight conditions that deviate from

the fixed gains, a slight degradation in handling qualities should be expected. Refer to chapter 11 for

details. The aircraft stalls at a lower than nominal AOA since the LEFs are fixed. Transition to or from

the landing configuration should be performed at 180 KCAS. For best results, maintain on-speed AOA

during the approach and landing.

Flight with GAIN ORIDE selected is prohibited above 10° AOA or above 350 KCAS (flaps AUTO),

200 KCAS (flaps HALF), or 190 KCAS (flaps FULL) to ensure control system stability and to reduce

the potential for departure. When GAIN ORIDE is selected, the amber FLAPS light comes on along

with either the CRUIS advisory (flaps AUTO) or the LAND advisory (flaps HALF or FULL). Alpha

tone is disabled in GAIN ORIDE.

Fixed Gains

LEF

TEF

AIL Droop

FLAP Switch

(°LED)

(°TED)

Mach

KTAS

Feet

°AOA

AUTO

0.80

459

39,000

3.5

HALF

0.23

151

500

8.1

FULL

0.21

139

500

Figure 2-26. GAIN ORIDE Flap Positions and Gain Schedules

2.10.14.1 GAIN Switch. The GAIN switch, located on the FCS panel on the left console, is used to

select GAIN ORIDE. The switch is guarded in the NORM position to prevent inadvertent actuation.

ORIDE

Selects fixed CAS gains according to FLAP switch position.

NORM

Selects normal CAS gain scheduling.

2.10.15 FCS Failures. The FCS detects failures through three types of BIT: initiated (IBIT), periodic

(PBIT), and maintenance (MBIT). FCS IBIT is performed during Before Taxi Checks to run a

thorough test of the system prior to flight. PBIT is a less thorough test of the FCS and runs

continuously when other BITs are not running. MBIT is typically run by maintenance personnel and

is the most comprehensive test of the FCS.

FCS failures are annunciated by any or all of the following indications: the FCS caution, the FCES

caution light, the ″Flight controls, Flight controls″ voice alert, FCS format Xs, and/or BIT Logic

Inspection (BLIN) codes. FCS format Xs and BLIN codes identify the location and type of failure.

However, not all FCS related components/functions are covered by Xs on the FCS format matrix. For

such components/functions, valid BLIN codes may be the only indication of the location of the failure.

Therefore, until the nature of the failure is determined, BLIN codes that appear without Xs should be

treated with the same level of concern as those that do.

Typically, BLIN codes that have three digits or less are generated by PBIT, e.g., 341. Four digit and

five digit BLIN codes are generated by FCS IBIT, e.g., 4573 and 10165.

2.10.15.1 FCES Caution Light. The Flight Control Electronic Set (FCES) caution light is located on

the lower right caution lights panel. The primary purpose of the FCES caution light is to alert aircrew

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A1-F18EA-NFM-000

of critical FCS related failures when MC1 is failed. When MC1 is failed, the normal DDI FCS related

cautions are not generated and the FCS format is not available for troubleshooting. When MC1 is

operative, the FCES caution light is merely a secondary indication of an FCS related failure. The

specific FCS cautions which also trigger the FCES caution light are listed in figure 2-27.

2.10.15.2 FCS RESET Button. The FCS RESET button is located on the FCS panel on the left

console. This button is used to perform several FCS related functions. Following detection of FCS

related hardware and/or software failures (e.g., FCS Xs and/or BLIN codes), pressing the FCS RESET

button commands a reset of FCC failure detection circuitry. If the FCS related failure was momentary

and no longer exists, an FCS RESET (a) restores the failed actuator/component, (b) removes all FCS

failure indications (FCS caution, FCES caution light, and Xs; preflight BLIN codes only), and (c)

displays the RSET advisory for 10 seconds to indicate a successful reset. If the failure remains (a) the

failed actuator/component is not restored, (b) the FCS failure indications return, and (c) the RSET

advisory is displayed for 10 seconds to indicate an unsuccessful reset. In other words, the FCS

RESET button does not fix a detected failure; it merely allows components to be

restored and failure indications to be removed, if and only if the failure no longer exists.

Prior to takeoff (cycle to WoffW), a successful FCS RESET automatically clears all BLIN codes.

Inflight or post-flight, however, BLIN codes are not automatically cleared with a successful FCS

RESET in order to preserve this data for maintenance troubleshooting. Inflight and post-flight BLIN

codes can be cleared, if desired, by pushing the FCS RESET button simultaneously with the paddle

switch.

Additionally, the FCS RESET button is used in conjunction with the FCS BIT consent switch to

enter the FCS exerciser mode.

2.10.15.3 FCS Exerciser Mode. The FCS exerciser mode is incorporated to aid hydraulic system

warming during cold weather starts. The exerciser mode allows hydraulic fluid and hydraulic seals to

warm towards normal operating temperatures without making large surface movements. Large surface

movements with a cold hydraulic system can result in hydraulic seal damage, leaks, and loss of fluid.

On the ground, the FCS exerciser mode is initiated by simultaneously holding the FCS BIT consent

switch in the ON position while pressing the FCS RESET button. When initiated, the mode cycles the

stabilators, flaps, ailerons, and rudders through 20% of full travel for 10 cycles in 20 seconds. The

operation can be stopped prior to 20 seconds by pressing the paddle switch.

During cold weather starts, avoid activating any hydraulic actuated system for two minutes after

both engines are online. This allows hydraulic fluid to warm both systems and prevents hydraulic seal

damage and potential hydraulic leaks. If the aircraft has not flown within 4 hours with ambient

temperatures below -18°C (0°F), up to three selections of the FCS exerciser mode may be required in

order to obtain a successful FCS RESET (after the initial 2 minute warmup).

In standard or warm conditions, do not initiate the FCS exerciser mode

multiple times in an attempt to get a successful FCS RESET. In such

conditions, multiple initiations may excessively elevate hydraulic system

temperatures, increasing actuator and hydraulic pump seal wear and

potentially decreasing component life.

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2.10.15.4 FCS BIT Consent Switch. The FCS BIT consent switch is located above the right console

beneath the right canopy sill. The switch is used in conjunction with the FCS BIT option or the FCS

RESET button to initiate FCS IBIT or the FCS exerciser mode, respectively. See the FCS Initiated

BIT (IBIT) section at the end of chapter 2 for details.

When held (for at least 2 seconds) during selection of the FCS option, initiates

FCS IBIT. When held during a press of the FCS RESET button, initiates FCS

exerciser mode.

OFF

FCS IBIT and FCS exerciser mode not selected.

2.10.15.5 FCS Related Cautions. FCS related cautions shown in figure 2-27

are described in the

Warning/Caution/Advisory Displays in Part V.

Associated Cockpit Indications

Caution

″Flight Controls,

Master Caution

Flight Controls″

Light

Tone

FCES Light

Voice Alert

AOA

Air Data

ATC Fail

AUTO PILOT

P CAS

R CAS

Y CAS

CHECK TRIM

CK FLAPS

FC AIR DAT

FCS

Note 2

FCS HOT

Note 1

FLAPS OFF

FLAP SCHED

G-LIM 7.5G

G-LIM OVRD

HYD 5000

NWS

R-LIM OFF

RIG

S/W CONFIG

NOTES

1. The FLIGHT COMPUTER HOT, FLIGHT COMPUTER HOT voice alert and FCS HOT light on the

caution lights panel are activated when the FCS HOT caution is set.

2. Also displayed when any aileron, stabilator, or rudder actuator failed off (Xd out and a ″bold X″ over

the surface position on FCS Status Display).

Figure 2-27. FCS Related Cautions and Cockpit Indications

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Figure 2-28. FCS Status Display

2.10.16 FCS Status Display. When an FCS failure has occurred, the FCS status display (figure 2-28)

can be used to determine the location and type of failure. The FCS status display is selected by the FCS

option on the SUPT MENU. For FCS components/functions which are displayed in the matrix, an ″X″

is displayed in the failed channel(s) along with a corresponding BLIN code. For FCS components/

functions which are not displayed in the matrix, BLIN codes are the only indication of failure location.

A. Surface Position: In degrees from streamline for all surfaces. Tolerance is ±1° for all surfaces.

B. Surface Arrow: Direction of surface deflection relative to the surface hinge point except for

stabilators. For stabilators, surface arrow indicates trailing edge position.

C. Column of Xs: An entire FCC channel is failed due to a processor fault or loss of power.

D. Bold X across surface position: Surface failed; FCCs are no longer commanding movement of

that surface in any channel.

E. G-LIMX.XG: ″X.X″ is the current Nz REF value as calculated by the MC. This value is

decremented if in the transonic g-bucket. INVALID is displayed in this position if the interface

between FCC CH 1, FCC CH 3, and the MC is invalid. In this case, all data on the display is invalid.

During FCS IBIT, G-LIM0.0G is displayed in this position.

F. Bold X over G-LIMX.XG: Nz REF data from the MC is invalid or out of range, and Nz REF has

defaulted to +7.5g. This X also appears when gross weight is above 57,405 lb, indicating that Nz REF

has been set to +5.5g even though this may result in an overstress.

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G. BLIN Codes. Up to eight FCC BLIN codes are displayed in octal format for each channel. The codes are

displayed in the order of occurrence. If the list exceeds eight, additional codes may be viewed with a memory

inspect of unit 14 or 15 with address 2253.

In-flight Memory Inspect (MI) of FCC (UNIT 14 or 15) addresses (ADDR)

greater than six digits long is prohibited since it may cause all four FCC channels

to shut down which will result in loss of aircraft control.

H. BLIN Code Channel: The FCC channel corresponding to the list of BLIN codes. The channel is

incremented from 1 thru 4 and back to 1 by selecting the BLIN option.

I. L/R XX.X: FCC air data function corrected true AOA for the left and right AOA probes (see AOA Select in

paragraph K below).

J. AOA XX.X: True AOA based on INS data. INS true AOA is displayed for reference only to aid the pilot in

determining which AOA probe is valid when one is damaged. INS true AOA is normally boxed, indicating that

an average of the left and right probes has been selected for display in the HUD and for use by the AOA indexer

lights and approach lights.

K. AOA Select: The AOA option is displayed only when GAIN ORIDE is selected. If one AOA probe is

damaged, the AOA option can be used to select output from the good probe for display in the HUD and for use

by the AOA indexer lights and approach lights. AOA probe selection does not affect the fixed gains used by the

FCCs in GAIN ORIDE.

NOTE

If a single probe is declared invalid (a two channel AOA failure), and that

probe is selected, the AOA indexer lights and the HUD AOA are blanked

immediately.

L. DEGD Xs: An FCC failure has occurred in the Xd channel that is not covered by other matrix Xs. BLIN

codes should be used to determine the degraded FCC channel function.

M. PTS Xs: The static or total pressure data is failed in the Xd channel. If a three channel PTS failure occurs

(three Xs), the FCC control laws use data from the remaining PTS channel. If a total PTS failure occurs (four

Xs), the FCCs use fixed PTS values. If a PTS failure clears, PTS Xs are removed automatically with or without

an FCS RESET attempt.

NOTE

With a four channel PTS failure, HUD airspeed and altitude are blanked.

N. AOA Xs: AOA data failed in the Xd channel. A three or four channel AOA failure sets four Xs (AOA Four

Channel failure). In UA, P CAS uses the AOA estimator for control law scheduling. If a UA failure clears, AOA

Xs are removed automatically with or without an FCS RESET attempt. In PA, P CAS uses a fixed 8.1° AOA

value, and R CAS uses the AOA estimator for control law scheduling. If a PA failure clears, AOA Xs are not

removed until an FCS RESET is attempted.

O. Sensor Xs (CAS P, R, or Y; N ACC, L ACC, STICK, or PEDAL): The corresponding sensor (rate gyros,

normal or lateral accelerometers, stick or pedal position) is failed in the Xd channel. A three or four channel

sensor failure sets four Xs (total sensor failure). However, for a three channel failure, the FCCs average the

remaining channel with the last channel that failed. For N ACC and L ACC only, the FCCs use a single channel

if the signal from the third failed channel exceeds 90% of full range. A single X for CAS P, CAS R, or CAS Y is

not possible. Any single gyro failure in P, R, or Y sets Xs in CAS P, CAS R, and CAS Y for that channel. Similarly,

an AHRS acceleration failure sets Xs in the N ACC and L ACC for that channel. A failed AHRS sets Xs in all

five rows for that channel.

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ORIGINAL W/IC 34

A1-F18EA-NFM-000

P. 1 2 3 4: Column legends for each FCC channel. FCC A contains channels 1 and 2 while FCC B

contains channels 3 and 4.

Q. Actuator Xs: The actuator is no longer commanded by the FCC in the Xd channel due to a

detected failure (for all actuators except spoilers). The actuator is still commanded by the other

operating channel(s).

R. SPOIL Xs: A single X is caused by a difference between commanded and actual position or by a

SOV over-current. If a two channel failure (two Xs) was caused by a difference between commanded

and actual position, the FCCs will continue to command the spoilers in both channels. This condition

is indicated by two Xs, a blanked surface position, and no bold surface position X. The vent surface

position value is always blanked.

S. Blank Surface Position: FCCs and/or MCs unable to report actuator position.

T. ROLL TRIM: With MC OFP H3E AND UP, roll trim is displayed in the right column with

WoffW. Arrows indicate trim direction. The roll trim value is dimensionless. Roll trim effects for the

same trim value are different at different airspeeds. The roll trim value provides a quantitative

measure of how much roll trim has been commanded.

U. PITCH TRIM: With MC OFP H3E AND UP, pitch trim is displayed in the left column with

WoffW. Arrows indicate trim direction. The pitch trim value is degrees AOA with flaps in HALF or

FULL and g-level with flaps in AUTO.

2.11 AFCS - AUTOMATIC FLIGHT CONTROL SYSTEM

The AFCS or autopilot provides three basic functions: pilot relief, coupled steering, and data link

control.

Different pilot relief modes are provided for the pitch and roll axes. Pitch-axis pilot relief modes

include barometric altitude hold (BALT), radar altitude hold (RALT), and flight path attitude hold

(FPAH). Roll-axis pilot relief modes include roll attitude hold (ROLL), ground track hold (GTRK),

ground track select (GSEL), heading hold (HDG), and heading select (HSEL).

Coupled steering modes allow the roll-axis to be coupled to a TACAN station (CPL TCN), to a

waypoint (CPL WYPT), to the azimuth steering line (CPL ASL), or to bank angle (CPL BNK).

Data link control modes include automatic carrier landing (ACL) and vector (VEC).

2.11.1 AFCS Mode Selection. Selection of the various autopilot (A/P) modes is accomplished from

the A/P sublevel of the CNI format on the UFCD. Before any autopilot mode can be selected, bank

angle must be less than 70°, pitch attitude must be less than 45°, and the A/P sublevel must be

displayed on the UFCD. The left column of the A/P sublevel displays the couple (CPL) option and the

pitch-axis pilot relief mode options: BALT, RALT, and FPAH. The right column displays the roll-axis

pilot relief mode options: ROLL, GTRK, and HDG. See figure 2-29.

An autopilot mode is enabled by selecting the corresponding option on the UFCD. Once selected, a

highlighted box appears around the option and the corresponding autopilot advisory appears on the

LDDI. If an option is not available, it is not displayed.

Once in the GTRK mode, subsequent selection of the GTRK option enables the GSEL mode (GSEL

replaces GTRK on the UFCD). Once in the HDG mode, subsequent selection of the HDG option

enables the HSEL mode (HSEL replaces HDG on the UFCD).

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2.11.2 Basic Autopilot. The basic or default autopilot mode is FPAH/HDG. When any autopilot

mode is requested from the UFCD, the AFCS first engages FPAH/HDG and then engages the

requested mode. This makes sure that the AFCS is controlling both the pitch and roll axis whenever

an autopilot mode is engaged.

2.11.3 AFCS Mode Deselection. Autopilot modes can be disengaged by either reselecting (unboxing)

the UFCD option or by actuating the paddle switch. If an autopilot mode is unboxed on the UFCD, the

AFCS reverts to the basic autopilot mode in that axis and the AUTO PILOT caution illuminates. If the

stick is moved longitudinally with BALT or RALT engaged or laterally with CPL engaged, the AFCS

reverts to the basic autopilot mode in that axis as the AUTO PILOT caution illuminates. The basic

autopilot mode, however, cannot be disengaged (unboxed) from the UFCD and must, therefore, be

disengaged with the paddle switch. Paddle switch actuation is the only means to make sure that all

autopilot modes have been completely disengaged.

2.11.4 Pitch-Axis Pilot Relief Modes. Any of the pitch-axis pilot relief modes can be engaged in

conjunction with any coupled steering mode (except ACL) or with any roll-axis pilot relief mode.

However, only one pitch-axis mode can be selected at a time. If one pitch-axis mode is requested while

another is engaged, the AFCS switches to the requested mode.

2.11.4.1 BALT - Barometric Altitude Hold. When BALT is engaged, the baro-inertial altitude at the

time of engagement is captured and maintained. While the mode is engaged, this reference altitude

cannot be changed. Longitudinal stick inputs or trim changes disengage BALT, and the AFCS reverts

to FPAH and the selected roll-axis mode.

2.11.4.2 RALT - Radar Altitude Hold. RALT is not available above 5,000 feet AGL. When RALT is

engaged, the radar altitude at the time of engagement is captured and maintained. While the mode is

engaged, this reference altitude cannot be changed. Longitudinal stick inputs or trim changes

disengage RALT, and the AFCS reverts to FPAH and the selected roll-axis mode.

2.11.4.3 FPAH - Flight Path Angle Hold. When FPAH is engaged, the flight path angle at the time

of engagement is captured and maintained. This reference flight path angle can be changed by

longitudinal stick inputs (stick sensitivity similar to CAS) or by pitch trim changes (2°/sec in flaps

AUTO or 0.5°/second in flaps HALF or FULL). When pilot inputs cease, the flight path attitude at

release is captured and maintained.

2.11.5 Roll-Axis Pilot Relief Modes. Any of the roll-axis pilot relief modes can be engaged in

conjunction with any pitch-axis pilot relief mode. However, only one roll-axis mode can be engaged at

a time. If one roll-axis mode is requested while another is engaged, the AFCS switches to the requested

mode.

2.11.5.1 ROLL - Roll Attitude Hold. When ROLL is engaged, the roll attitude at the time of

engagement is captured and maintained. This reference roll attitude can be changed by lateral stick

inputs or roll trim changes (stick and trim sensitivity similar to CAS). When pilot inputs cease, the roll

attitude at release is captured and maintained.

2.11.5.2 GTRK - Ground Track Hold. When GTRK is engaged, aircraft response depends on the roll

attitude at the time of engagement. If roll attitude is less than ±5°, ground track is captured and

maintained. If roll attitude is greater than or equal to ±5°, roll attitude is captured and maintained.

While in GTRK, the aircraft responds to lateral stick or roll trim inputs (stick and trim sensitivity

similar to CAS). When pilot inputs cease, GTRK holds roll attitude (if greater than or equal to ±5°)

or ground track (if less than ±5°).

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ORIGINAL

A1-F18EA-NFM-000

2.11.5.3 GSEL - Ground Track Select. The desired ground track angle is selected by slewing the

command heading marker with the HDG/TK switch, located to the left of the MPCD. When GSEL is

engaged, the aircraft turns from the existing ground track through the smallest angle to the selected

ground track. While in GSEL, the aircraft responds to lateral stick inputs. However, the selected

ground track angle is not changed by stick inputs, so the aircraft returns to the selected ground track

angle upon stick release.

2.11.5.4 HDG - Heading Hold. When HDG is engaged, aircraft response depends on the roll attitude

at the time of engagement. If roll attitude is less than ±5°, magnetic heading is captured and

maintained. If roll attitude is greater than or equal to ±5°, roll attitude is captured and maintained.

While in HDG, the aircraft responds to lateral stick or roll trim inputs (stick and trim sensitivity

similar to CAS). When pilot inputs cease, HDG holds roll attitude (if greater than or equal to ±5°) or

magnetic heading (if less than ±5°).

2.11.5.5 HSEL - Heading Select. The desired heading is selected by slewing the command heading

marker with the HDG/TK switch, located to the left of the MPCD. When HSEL is engaged, the

aircraft turns from the existing heading through the smallest angle to the selected heading. While in

HSEL, the aircraft responds to lateral stick inputs. However, the selected heading is not changed by

stick inputs, so the aircraft returns to the selected heading upon stick release.

2.11.6 CPL - Coupled Steering Modes. The coupled steering modes couple the aircraft in the

roll-axis only. If the CPL option is selected, the AFCS disengages any currently engaged roll-axis pilot

relief mode. The AFCS has the ability to couple to the following sources: a waypoint, waypoint

courseline, or offset aimpoint (CPL WYPT); to a TACAN station or TACAN courseline (CPL TCN);

or to an auto sequence (CPL SEQ#). Refer to chapter 24 for detailed navigation steering information

on waypoint/OAP, auto sequential, and TACAN steering.

2.11.7 Coupled Data Link Modes. The AFCS can couple to data link commands in one of two modes:

ACL and VEC. With ACL boxed on the HSI format, the CPL P/R option appears on the A/P sublevel

when pitch/roll couple capability is available. Selecting the CPL P/R option couples the aircraft to

pitch and roll commands for a Mode 1 carrier approach. With VEC boxed on the HSI format, selecting

the CPL option couples the aircraft (roll-axis only) to data link steering commands. Refer to chapter

24 and the Tactical Manual for detailed information on the ACL and VEC modes.

2.11.8 AFCS Related Caution and Advisories. The AUTO PILOT caution and the following AFCS

related advisories are described in the Warning/Caution/Advisory Displays in Part V:

• BALT

• GSEL

• HSEL

• CPLD

• GTRK

• RALT

• FPAH

• HDG

• ROLL

2.12 WEAPON SYSTEMS CONTROLS

All of the primary controls for the aircraft’s weapon systems (weapons, sensors, and displays) are

located on the front cockpit throttles and stick, the rear cockpit throttles and stick (trainer configured

F/A-18F), or the rear cockpit hand controllers (missionized F/A-18F). This concept, hands on throttles

and stick (HOTAS), allows the aircrew to manipulate the weapon systems without removing the hands

from the aircraft’s primary flight controls. Additionally, the canopy sill DISP switch(es) and the rear

cockpit grab handle switches provide secondary controls for dispensing expendables from the ALE-47

self-protect system.

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Detailed descriptions of the functionality of the weapon systems controls are contained in the

Tactical Manual, F/A-18EA-TAC Series.

2.12.1 Stick Grip Switches/Controls (Front Cockpit). The weapon systems controls located on the

front cockpit stick grip are the A/A weapon select switch, the sensor control switch, the gun/missile

trigger, the A/G weapon release button, and the undesignate/NWS button. See figure 2-30.

Figure 2-30. Stick Grip Switches/Controls

2.12.1.1 A/A Weapon Select Switch. The A/A weapon select switch, located on the left side of the

stick grip, is spring loaded to the center/up position. The weapon select switch is used to select the

desired A/A weapon and can be used to enter A/A master mode. When the weapon select switch is

actuated while in NAV or A/G master mode, the A/A master mode is automatically entered, the RDR

ATTK format is automatically displayed on the RDDI, and the appropriate A/A weapon/radar format

is selected. Once in the A/A master mode, actuating the weapon select switch merely changes the

selected A/A weapon/radar format.

Center

Neutral

Forward

Selects AIM-7 and the corresponding radar format.

Down

Selects AIM-9 and the corresponding radar format.

Aft

Selects A/A GUN and the gun acquisition mode (GACG).

Inward

Selects AIM-120 and the corresponding radar format.

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Selection of an A/A missile initiates launch preparation of the priority missile, if more than one is

carried and brings up the corresponding radar format (e.g., scan volume presets). Subsequent selection

steps to the next available missile in the priority sequence.

2.12.1.2 Sensor Control Switch. The sensor control switch, commonly called the ″castle″ switch, is

located on the top center of the stick grip and is spring loaded to the center/up position. The castle

switch is used to assign throttle designator controller (TDC) priority to a particular cockpit display or,

once air combat maneuvering (ACM) mode functionality is enabled, to select a particular ACM mode.

Center

Neutral

Forward

Assigns TDC priority to the HUD in NAV or A/G master mode. In A/A

master mode, selects the boresight acquisition mode (BST) and enables

ACM mode functionality.

Forward (twice

Selects/deselects EMCON.

within 0.5 seconds)

Depress/release then

Assigns TDC priority to the UFCD in all master modes. If the top level

forward (within 1

CNI, a CNI sublevel, or a data entry format is displayed, selects the last

second)

displayed DDI format.

Left

Assigns the TDC to the LDDI in all master modes. With ACM mode

functionality enabled in A/A master mode, selects the wide acquisition

mode (WACQ).

Right

Assigns the TDC to the RDDI in all master modes. With ACM mode

functionality enabled in A/A master mode, selects the automatic acquisi-

tion mode (AACQ).

Aft

Assigns the TDC to the MPCD in all master modes. With ACM mode

functionality enabled in A/A master mode, selects the vertical acquisition

mode (VACQ).

When the TDC is assigned to a display which cannot accept TDC priority, automatic format

initialization occurs, typically selecting the format which is most commonly used on that particular

display (e.g., RDR ATTK on the RDDI). With TDC assignment, certain displays (e.g., RDR ATTK,

FLIR, etc.) perform a specific action (e.g., track, break track, etc.) when the castle switch is

subsequently bumped toward that display.

2.12.1.3 Trigger. The gun/missile trigger, located on the front of the stick grip, has two detented

positions.

First

Initiates strike camera automatic mode operation (based on the selected A/G

Detent

weapon) and if CVRS is running, commands HUD recording.

Second

In A/A master mode, fires the gun or selected A/A missile. Activates AGI protec-

Detent

tion. In A/G master mode, fires the gun or the laser, if either is selected. Com-

mands the HUD event marker.

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2.12.1.4 A/G Weapon Release Button. The A/G weapon release button, commonly called the

″pickle″, is located on the top center of the stick grip to the left of the castle switch. The pickle is used

to command weapon release while in A/G master mode. The pickle also initiates AGI protection, strike

camera automatic mode operation, and, if CVRS is running, commands HUD recording and the HUD

event marker.

2.12.1.5 Undesignate/NWS Button. The undesignate/NWS button is located on the lower front of

the stick grip. In NAV or A/G master mode, the undesignate button undesignates all A/G designated

targets and commands the radar and FLIR to break lock, if either is tracking. If TDC priority is

assigned to the FLIR format, pressing the button twice in 1 second selects/deselects FLIR velocity

vector slave (VVSLV) pointing mode. In A/A master mode, the undesignate button creates a launch

and steering (L&S) target designation. Subsequent actuation steps target designation to the next

priority trackfile or to a second designated trackfile (DT2), if one exists. If ACM mode functionality is

enabled, the undesignate button exits the ACM mode and returns the radar to the previous search

mode.

2.12.2 Stick Grip Switches/Controls

(Trainer Configured F/A-18F). In the F/A-18F

(trainer

configuration), the front and rear cockpit stick grips are identical. However, the rear cockpit trigger

and A/G weapon release button are not functional. The rear cockpit A/A weapon select switch does not

automatically select A/A master mode. From the rear cockpit, A/A master mode must be entered by

actuation of the A/A master mode light. The front and rear cockpit control sensor switches are

functionally identical, including ACM mode selection. However, the rear cockpit TDC can be assigned

to a sensor different from the front cockpit TDC. The systems controlled by the stick grip

switches/controls respond to the last crewmember action taken from either cockpit.

2.12.3 Throttle Grip Switches/Controls (Front Cockpit). The weapon systems controls located on

the front cockpit throttle grips are the chaff/flare/ALE-50 dispense switch, the cage/uncage button, the

throttle designator controller (TDC), the radar elevation control, and the raid button. See figure 2-31.

Figure 2-31. Throttle Grip Switches/Controls (Front Cockpit)

2.12.3.1 Chaff/Flare/ALE-50 Dispense Switch. The chaff/flare/ALE-50 dispense switch is located

on the top inboard side of the right throttle. The switch is used to dispense expendables from the

ALE-47 and ALE-50 self-protect systems.

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A1-F18EA-NFM-000

Forward

Provides semi-automatic consent. Dispenses chaff singles (C/F mode).

Aft

Initiates the selected manual program. Dispenses flare singles (C/F mode).

Dispenses a single ALE-50 decoy and provides transmit enable.

Down

Not functional

2.12.3.2 Cage/Uncage Button. The cage/uncage button is located on the rear inboard side of the

right throttle. In NAV master mode and A/G master mode (AUTO delivery), the cage/uncage button

is used to cage and uncage the velocity vector. Depending on master mode and TDC priority

assignment, the cage/uncage button can be used to (1) cage/uncage weapons, (2) command a radar

STT, (3) toggle between weapon modes, or (4) reset sequenced HARM targets.

(18E OFP) With the velocity vector caged, data such as the flight path/pitch ladder and steering

information is displayed near the center of the HUD despite large yaw rates and/or crosswind angles.

2.12.3.3 TDC - Throttle Designator Controller. The TDC is located on the front right side of the

right throttle. It is used to control the positioning of the acquisition cursor or the slewing of a particular

sensor or weapon (e.g., FLIR or Maverick). The TDC can be pressed for target designation. When held

pressed (action slew) or released (no-action slew), the fore-aft and left-right movement of the TDC

sends X-Y slew commands to the display or to the sensor/weapon to which TDC priority is assigned.

2.12.3.4 Radar Elevation Control. The radar elevation control is located on the front left side of the

right throttle. Momentary actuation changes the elevation of the radar antenna in

1,000

foot

increments at the range where the cursor is positioned on the RDR ATTK format. Press and hold

produces a faster elevation change.

Raises the radar antenna/scan volume.

Center

Neutral

Down

Lowers the radar antenna/scan volume.

2.12.3.5 RAID Button. The RAID button is located on the left side of the left throttle. Depending on

master mode and TDC priority assignment, the RAID button can be used to (1) sequence between

available HARM targets, (2) toggle between narrow and wide field of view (FOV) for the FLIR or

Maverick, or (3) select/deselect the STT RAID or SCAN RAID radar modes.

2.12.4 Throttle Grip Switches/Controls (Trainer Configured Rear Cockpit). The rear cockpit

throttle grips contain the same weapon systems controls as those in the front cockpit, except no

chaff/flare/ALE-50 switch is installed. The systems controlled by the throttle grip switches/controls

respond to the last crewmember action taken from either cockpit.

2.12.5 Hand Controllers (Missionized Rear Cockpit Lots 21 thru 25). Two hand controllers are

installed in the rear cockpit, one on the front inboard section of each console. With the exception of

weapons release, the hand controllers are used to provide the same weapon systems control as stick and

throttle switches/controls do in the front cockpit. The right hand controller is a mirror image of the left

hand controller. Each contains a multi-function switch (MFS), a designator controller (DC) assignment

switch, a designator controller (DC), a radar elevation control, a chaff/flare dispense switch, and an

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undesignate button. Except for the chaff/flare dispense switches, the functions of the left and right

hand controllers are identical. See figure 2-32.

2.12.5.1 Multi-Function Switch (MFS). The MFS is located on the lower inboard side of each hand

controller. The MFS provides the same functionality as the front cockpit RAID and cage/uncage

buttons.

Forward

Sequences between available HARM targets.

Aft

Cages/uncages an A/G weapon.

Down

Selects RAID or changes FLIR/Maverick FOV.

Forward

Not functional

2.12.5.2 Designator Controller (DC) Assignment Switch. The DC assignment switch is located on

the top inboard side of each hand controller. It is used to assign left and right DC priority to the

displays in the rear cockpit. ACM modes cannot be selected by the rear cockpit DC assignment

switches.

Forward

Commands FLIR track/break lock if opposite DC is assigned to the FLIR format.

Aft

Assigns DC priority to the AUFCD.

Inboard

Assigns DC priority to the AMPCD

Outboard

Assigns DC priority to the outboard ADDI. Following DC assignment, commands

track/break lock on the outboard format.

Each DC is initially assigned to its corresponding ADDI. One but not both DCs can be assigned to

one of the center displays, AUFCD or AMPCD. If one DC is assigned to the AMPCD, the other is

forced back to its corresponding ADDI.

2.12.5.3 Designator Controller (DC). The DC is located on the top center of each hand controller.

The track and slew functions of the rear cockpit DCs are identical to the front cockpit TDC. While only

one hand controller DC can be used to designate at a time, both DCs may be used simultaneously on

their assigned formats.

If the front cockpit TDC and the rear cockpit DC are both assigned to the same format, control is

captured by the first controller actuated. If one controller is active when another is selected, the second

input is ignored. If both front and rear TDC/DCs are pressed simultaneously, the TDC/DC assignment

diamond, located in the upper right corner of the specific format, flashes. If both front and rear

TDC/DCs are being slewed simultaneously, the SLEW cue is displayed and flashed.

2.12.5.4 Radar Elevation Control. The radar elevation control is located on the top outboard side of

each hand controller. The front and rear cockpit radar elevation controls are functionally identical.

Control of the radar antenna is captured by the first elevation control actuated from either cockpit. If

one control is active when another is selected, the second input is ignored.

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A1-F18EA-NFM-000

2.12.5.6 Undesignate Button. The undesignate button is located on the lower front of each hand

controller. The front and rear undesignate buttons are functionally identical, except the rear cockpit

buttons do not command NWS.

2.12.6 Hand Controllers (Rear Cockpit LOT 26 AND UP). The functions of the left and right hand

controllers are not identical. The left hand controller contains the following switches: Countermea-

sures, Growth, Sensor Weapons Control, Designator Control Assignment, A/G Weapon Release, Left

Trigger, ECM, and Cage. The right hand controller contains the following switches: HARM, Display

Scroll/Toggle, FOV Wheel, Designator Control, A/A Weapon Release, A/A Weapon Select, Undesig-

nate, and Right Trigger. See figure 2-32 Sheet 2. Refer to Tactical Manuals for complete Hand

controller Switch description.

2.12.7 ALE-47 DISP Switch. An ALE-47 DISP switch is located on the left canopy sill in both front

and rear cockpits. Either switch can be used to initiate manual program 6 or to dispense chaff and flare

singles (C/F mode).

2.12.8 Grab Handle Chaff/Flare Dispense Switches (Rear Cockpit LOTs 21 thru 25). The grab

handle dispense switches are located on the left and right ends of the center grab handle in the rear

cockpit. These switches provide a secondary means to dispense expendables from the ALE-47

self-protect system.

Forward

Provides semi-automatic consent. Dispenses flare singles (C/F mode)

Aft

Initiates the selected manual program. Dispenses flare singles (C/F mode)

Inboard

Initiates manual program 6. Dispenses chaff and flare singles (C/F mode)

Outboard

Initiates manual program 6. Dispenses chaff and flare singles (C/F mode)

2.12.9 Grab Handle Chaff/Flare Dispense Switches (Rear Cockpit LOT 26 and up). The grab

handle dispense switches are located on the left end of the left grab handle and the right end of the

right grab handle in the rear cockpit. Switch function is the same as described above.

2.13 ECS - ENVIRONMENTAL CONTROL SYSTEM

The environmental control system (ECS) utilizes engine bleed air to provide pressurization, heating

and cooling air to various aircraft systems. Warm air is provided for internal fuel tank pressurization

(LOT 23 and below), external fuel tank pressurization, canopy seal inflation, g-suit operation, radar

waveguide pressurization, windshield anti-ice and rain removal, gun gas purge, RECCE bay heating,

and on-board oxygen generating system (OBOGS) operation. Cold, dry conditioned air is provided for

avionics cooling. Warm and cold air are mixed to provide temperature controlled air for cabin heating,

cooling, and pressurization and windshield defog.

A liquid cooling system (LCS) is used to cool the radar transmitter. A digital ECS controller is used

to schedule ECS output, regulate system temperatures, monitor system health, and detect and isolate

faults. See foldout section for a schematic of the ECS. Items numbers, listed in ( ) next to ECS valves,

are maintenance and MSP code nomenclature.

2.13.1 Bleed Air Shutoff Valves. Engine bleed air is tapped from the final (seventh) stage of the

engine high pressure compressor. A primary bleed air pressure regulator and shutoff valve (Item 1), one

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on each engine, is used to regulate bleed air output pressure as well as to control flow application/

shutoff.

Both valves are electrically controlled by the BLEED AIR knob and are pneumatically actuated.

The primary bleed air shutoff valves failsafe to the closed position if either electrical power or air

pressure is lost. If an engine is shut down before placing the BLEED AIR knob to OFF, the

corresponding primary bleed air shutoff valve may not fully close, resulting in residual engine fumes

in the cabin on subsequent start of that engine.

Bleed air from each engine passes through a check valve, which prevents reverse flow, and is mixed

through a Y-junction prior to the secondary bleed air pressure regulator and shutoff valve (Item 2).

This valve is electrically controlled by the OFF and AUG PULL positions of the BLEED AIR knob and

is pneumatically actuated. The secondary bleed air shutoff valve fails to the open (safe) position if

either electrical power or air pressure is lost.

All three valves can be automatically commanded to the closed position by the bleed air leak

detection (BALD) system.

2.13.1.1 BLEED AIR Knob. The BLEED AIR knob, located on the ECS panel on the right console, is

used to select the engine bleed air source for the ECS system.

NORM Commands both primary bleed air shutoff valves open, selecting bleed air from both

engines.

L OFF Commands the left primary bleed air shutoff valve closed, selecting bleed air from the

right engine only.

R OFF Commands the right primary bleed air shutoff valve closed, selecting bleed air from the

left engine only.

OFF Commands all three bleed air shutoff valves closed, isolating the ECS. Closes the ECS

auxiliary duct doors.

AUG Commands the secondary bleed air shutoff valve closed, opens the ECS air isolation

PULL valve, and allows APU compressor air to operate the ECS.

2.13.1.2 L or R BLD OFF Cautions. The L or R BLD OFF cautions indicate that the corresponding

primary bleed air shutoff valve(s) are commanded closed. The cautions are not an indication of actual

valve position. The L and/or R BLD OFF cautions are displayed in the following circumstances:

a. BLEED AIR knob in L OFF, R OFF, or OFF (L, R, or both cautions).

b. ENG CRANK switch in L or R (L or R caution, respectively).

c. BALD system detects a leak in one or both bleed air systems (L, R, or both cautions).

d. FIRE switch in TEST A or TEST B (both cautions).

e. Over pressurization in one or both bleed air systems (both cautions).

If a bleed air shutoff valve has been commanded closed other than by ENG CRANK, the BLEED

AIR knob must be cycled to OFF and back to NORM to reopen the valve(s).

2.13.2 Bleed Air Subsystem. Engine bleed air downstream of the secondary bleed air shutoff valve

is routed to the primary heat exchanger and three valves. The primary heat exchanger is used for first

stage cooling of hot engine bleed air. The ECS air isolation valve directs bleed air to the air turbine

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A1-F18EA-NFM-000

starter control valves (ATSCV) for crossbleed start and accepts air from the APU compressor for

alternate (AUG PULL) ECS operation. On the ground and during slow speed flight, the ejector shutoff

valve is opened to direct bleed air to the ram air ejectors to induce cooling airflow through the primary

and secondary heat exchangers. The warm air temperature control valve is used to mix uncooled engine

bleed air with primary heat exchanger output air to regulate the temperature of air in the warm air

manifold.

2.13.3 Primary Heat Exchanger. The primary heat exchanger is located near the base of the right

vertical tail and is used to reject heat from engine bleed air to ram air from one of two inlets.

During medium to high speed flight, ram air cooling is provided by the main ram inlet which draws

air from the engine intake. During slow speed flight and ground operations, ram air cooling is provided

by an auxiliary ram inlet which draws free stream air from the top of the fuselage. A ram air ejector in

the ram air exhaust duct is used to induce more airflow when required, such as on the the ground,

during slow speed flight, and during windshield anti-ice operation with either throttle above IDLE.

2.13.3.1 ECS Auxiliary Duct Doors. Two ECS auxiliary duct doors are located on the upper surface

of the fuselage forward of the base of the vertical tails and immediately in front of the primary and

secondary heat exchangers. The ECS auxiliary duct doors are electrically actuated and open into the

free stream, closing the main ram air inlets and exposing the auxiliary ram air inlets.

On the ground, the ECS auxiliary duct doors are open with the BLEED AIR knob in any position

except OFF. Inflight, the doors are positioned based on Mach number and throttle setting. The doors

are always open below Mach 0.33 and are always closed above Mach 0.40. Between Mach 0.35 and Mach

0.40, the doors are open if either throttle is near MIL (THA greater than 25°).

The ECS auxiliary duct doors should operate symmetrically. If either ECS auxiliary duct door is not

in the commanded position for greater than 8 seconds, an ECS DR advisory is displayed along with an

ECS BIT indication of DEGD. If the door(s) return to the commanded position, the ECS DR advisory

is removed.

2.13.3.2 Ram Air Exhausts. The ECS ram air exhausts are located on the upper fuselage just aft of

the leading edge of the vertical tails. There is one exhaust for the primary heat exchanger (right side)

and one for the secondary heat exchanger (left side). These exhausts discharge the heated ram air

overboard. The exhausts have been redesigned to a five swept stack configuration to prevent

overheating of the aft fuselage structure.

2.13.4 Warm Air Subsystems. Air leaves the primary heat exchanger at a greatly reduced tempera-

ture. Warm air from the primary heat exchanger is routed directly to the following systems: internal

fuel tank pressurization (LOT 23 and below), external fuel tank pressurization, canopy seal, g-suit,

radar waveguide pressurization, and OBOGS.

The ECS controller modulates the warm air temperature control valve (Item 21) to mix air from the

primary heat exchanger with uncooled engine bleed air to regulate the temperature in the warm air

manifold. The warm air manifold supplies air for windshield anti-ice and rain removal, gun gas purge,

RECCE bay heating, ECS turbine anti-ice, and cabin heating.

2.13.5 Air Conditioning System (ACS) Pack. The ACS pack provides cold, dry conditioned air for

cabin, FLIR, and avionics cooling. The primary components of the ACS pack are a compressor, turbine,

condenser, reheater and water extractor.

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The ECS controller modulates airflow through the ECS flow modulator valve (Item 4) to control

ECS flow and the speed of the ECS compressor and turbine. Compressor discharge air is directed to

the secondary heat exchanger for additional cooling. Cooled air from the secondary heat exchanger is

directed to the ECS turbine and is used as another source of OBOGS air inflight.

Expansion through the ECS turbine drives the compressor and greatly reduces the temperature of

the airflow, typically to below freezing. The ECS controller regulates the ECS turbine output

temperature and prevents ECS icing by adding warm air through the anti-ice add heat valve (Item 51).

Output temperature is regulated during low altitude operations where ECS icing would be likely.

The condenser, reheater and water extractor remove water from the cold conditioned air. With the

large heat load of the APG-79 radar, the condenser can run at sub−freezing conditions. During

prolonged operation of the APG-79 in high humidity conditions with warm fuel temperatures, periodic

automatic deicing of the condenser occurs. During de-icing, a short term increase in cabin supply air

temperature may be noticed.

In LOT 26 and up, at altitudes above 40,000 feet the water removal system may be bypassed by

opening the water extractor bypass valve (Item 171) in order to improve ECS operating efficiency. If

this valve fails to close after descending below 37,000 feet, water removal capability is degraded. Water

or ice pellets may be blown into the cabin and erratic system behavior (flow/pressure surges) may

result. An ECS ICING caution may also occur.

2.13.6 Secondary Heat Exchanger. The secondary heat exchanger is located near the base of the left

vertical tail and is used to reject heat from ECS compressor air to ram air. Operation of the secondary

heat exchanger is identical to operation of the primary heat exchanger. However, excess ECS system

moisture is sprayed onto the secondary heat exchanger to increase system cooling. This moisture may

be seen exiting the secondary heat exchanger exhaust duct when the throttles are advanced during

ground operations.

2.13.7 Avionics Cooling Fans. Two avionic cooling fans augment ECS cooling of avionics. The

avionics ground cooling fan is located in the nosewheel well. The aft avionics cooling fan is located in

door 108. These fans normally provide primary avionics cooling on deck, and also provide contingency

avionics cooling in flight. Fan activation is a function of bleed air pressure, which varies with engine N2

rpm (i.e. throttle position). Fan operation is described in the individual ECS mode descriptions.

2.13.7.1 Aft Cooling Fan Shutoff Valve. The aft cooling fan shutoff valve, when closed, secures flow

from the aft avionics cooling fan into the ECS and also prevents ECS backflow to the aft avionics

cooling fan. If this valve fails (indicated by MSP 863) while in the open position then, depending on

system pressure, ECS backflow can overspeed the fan in the wrong direction. This may cause a

structural failure of the fan and result in collateral fragment damage to adjacent systems. ECS

backflow is most likely to occur when ECS MAN mode is selected and ECS flow is commanded by

default to maximum.

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Selection of ECS MAN mode is prohibited. Selecting ECS MAN mode

while the aft cooling fan shutoff valve is open may cause the fan to

overspeed, resulting in a catastrophic fan failure and potential loss of

OBOGS.

2.13.8 ECS Operating Modes. The ECS has three operating modes: AUTO, MAN (manual) (PRO-

HIBITED), and OFF/RAM. Each mode is selected by the corresponding position of the ECS MODE

switch. On the ground only, APU compressor air may be used instead of engine bleed air to run the ECS

and cool the avionics (BLEED AIR knob in AUG PULL).

2.13.8.1 ECS AUTO Mode. ECS AUTO mode is the normal operating mode of the ECS. The ECS

controller modulates ECS output to provide the required airflow to the cabin and the avionics. Cabin

airflow is scheduled as a function of ram air temperature and throttle setting, with the highest flow

delivered at hot and cold temperature extremes. Cabin flow may decrease at IDLE. ECS flow to the

avionics is dependent on WonW status, throttle setting, Lot number, and radar configuration. For all

aircraft configurations, more air is provided to avionics inflight.

With APG−73 radar installed, WonW and both throttles at IDLE, the avionics ground cooling fan

and the aft avionics cooling fan energize. With APG−79 radar installed, WonW and both throttles at

IDLE, only the avionics ground cooling fan energizes. The fan(s) provide the primary source of avionics

cooling on deck. The ECS provides a second source of ground avionics cooling but at a fixed, low flow

rate with the remainder of the flow going to the cabin.

With WonW and at least one throttle advanced to approximately 74% N2 rpm, or with WoffW, the

fan(s) secure. Once secured, the fan(s) will not reenergize until both throttles are retarded below

approximately 70% N2 rpm. With WonW and fans secured, the ECS provides all avionics cooling and

controls to the cabin and avionics airflow schedules. The ECS controller schedules avionics airflow

based on the temperature of the air being delivered (warmer air requires higher flow to maintain

constant cooling).

The ECS controller modulates airflow output to meet scheduled airflow requirements. ECS output

temperature is scheduled by the ECS controller to meet avionics cooling requirements and to prevent

ECS turbine icing. Temperature is regulated by adding warm air as required. The controller divides the

airflow between the cabin and avionics. Inflight, if ECS output is inadequate for demand, cabin airflow

is normally given priority. The avionics typically receive that portion of the ECS output which is not

used by the cabin or FLIR cooling systems.

Cabin temperature is controlled by gradually mixing warm air with cold conditioned air, according

to the position of the CABIN TEMP knob. Cabin temperature is normally selectable in the range of

35 to 135° F between the CABIN TEMP knob positions of COLD and HOT, respectively (60 to 160°

F with the DEFOG handle in the HIGH position).

NOTE

In ECS AUTO mode, cabin temperature may take 1 to 2 minutes to

stabilize following a large movement of the CABIN TEMP knob.

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2.13.8.2 ECS MAN Mode.

Selection of ECS MAN mode is prohibited. Selecting ECS MAN mode

while the aft cooling fan shutoff valve is open may cause the fan to

overspeed, resulting in a catastrophic fan failure and potential loss of

OBOGS.

2.13.8.3 ECS OFF/RAM Mode. ECS OFF/RAM mode is used to terminate normal ECS operation

following a major ECS malfunction. Conditioned ECS air is terminated, the cabin ram air scoop is

deployed, and if inflight the aft avionics cooling fan energizes.

NOTE

If ECS OFF/RAM mode is selected inflight, the AV COOL switch

should be placed in EMERG to deploy the FCS emergency ram air

scoop and maximize the emergency avionics cooling available.

Partial cabin pressurization is provided, but only if the CABIN TEMP knob is above the full COLD

position. Cabin airflow and pressurization is provided by ram air from the cabin ram air scoop. Ram

air supply varies with inflight dynamic conditions (airspeed and altitude), and cabin airflow and

pressurization vary similarly.

In ECS OFF/RAM mode, avoid operations above 25,000 feet MSL in

order to prevent decompression sickness (DCS). Normal cabin pressur-

ization is not provided. Partial cabin pressurization is available under

certain circumstances, but may be lost insidiously without warning.

Inflight, cabin temperature is no longer automatically controlled. The CABIN TEMP knob

manually controls addition of heated air according to the CABIN TEMP knob. Temperature control

in between the full COLD and full HOT positions is difficult and imprecise, and ECS response to

commanded temperature changes is slow and nonlinear.

NOTE

In ECS OFF/RAM mode, changes in the position of the CABIN

TEMP knob should be held for 30 seconds due to a slower temperature

response.

With the CABIN TEMP knob not in full COLD, cabin temperature varies with throttle position and

flight condition. System flow/temperature cycling can be expected with the CABIN TEMP knob in full

HOT since overtemperature protection operates intermittently.

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With WonW and both throttles at IDLE, the avionics ground cooling fan and the aft avionics cooling

fan provide the only source of avionics cooling. Above approximately 74% N2 rpm, both fans secure,

and avionics equipment is deprived of all cooling (AV AIR HOT caution).

NOTE

During ground operations in the ECS OFF/RAM mode, the throttles

should be kept below 70% N2 rpm whenever possible in order to

preserve avionics cooling. If either throttle must be advanced above

approximately 74% N2 rpm more than momentarily or if an AV AIR

HOT caution is present, placing the BLEED AIR knob to OFF

reenergizes both fans and provides avionics cooling.

Inflight, the aft avionics cooling fan and the FCS emergency ram air scoop (with AV COOL in

EMERG) provide emergency avionics cooling.

NOTE

With the FCS emergency ram air scoop extended, avionics cooling

inflight is maximized by maintaining altitude below 25,000 feet MSL

and airspeed between 200 to 300 KCAS.

2.13.8.4 AUG PULL. With both generators online and the APU running, selecting AUG PULL (up

on the BLEED AIR knob):

a. overrides APU automatic shutdown

b. closes the secondary bleed air shutoff valve.

c. opens the ECS air isolation valve.

d. shuts down the aft avionics cooling fans.

e. and directs APU air to the ECS for cabin and avionics cooling.

2.13.8.5 ECS MODE Switch. The ECS MODE switch, located on the ECS panel on the right console,

is used to select the ECS operating mode.

AUTO

Selects ECS AUTO mode, the normal ECS operating mode. Provides automatic

cabin and avionics airflow, automatic temperature scheduling, and full cabin pres-

surization.

MAN

PROHIBITED

Selection of ECS MAN mode is prohibited. Selecting ECS MAN mode

while the aft cooling fan shutoff valve is open may cause the fan to

overspeed, resulting in a catastrophic fan failure and potential loss of

OBOGS.

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OFF/RAM Selects ECS OFF/RAM mode, which terminates conditioned ECS airflow. Provides

ram air for cabin airflow and pressurization depending on flight conditions, and

manual temperature control.

2.13.8.6 CABIN TEMP knob. The CABIN TEMP knob, located on the ECS panel on the right

console, is used to control the temperature of air delivered to the cabin. In ECS AUTO mode, clockwise

rotation of the CABIN TEMP knob linearly increases cabin temperature. In ECS OFF/RAM mode, the

CABIN TEMP knob controls the cabin add heat valve directly, producing a nonlinear temperature

response.

NOTE

• In ECS AUTO mode, cabin temperature may take 1 to 2 minutes to

stabilize following a large movement of the CABIN TEMP knob.

• In ECS OFF/RAM mode, changes in the position of the CABIN

TEMP knob should be held for 30 seconds due to a slower tempera-

ture response.

2.13.8.7 Cabin Louvers/Foot Air Outlet. Three sets of louvers are provided to direct airflow within

the cabin, one on either side of the main instrument panel and one at the base of the center console

behind the control stick. The left and right louvers have controls for elevation and azimuth. The center

louver has a single control for elevation and can be closed completely by moving the lever to the full

aft position. When the center louver is closed, airflow through the left and right louvers is increased.

Maximum aircrew cooling is provided by pulling the DEFOG HANDLE full aft, closing the center

louver, and directing the side louvers towards the body. A fixed foot air outlet directs airflow to the

base of the cabin.

2.13.8.8 Windshield Defog Outlets Fixed windshield defog outlets direct cabin air onto the inner

windshield to remove and inhibit fog. Windshield fogging can occur during rapid environmental

changes, such as high rates of descent and high humidity.

2.13.8.9 DEFOG Handle. The DEFOG handle, located on the right console outboard of the ECS

panel, controls the division of airflow between the windshield defog outlets, the three cockpit louvers,

and the foot air outlet. The DEFOG handle mechanically controls the position of the cabin air/defog

diverter valve.

HIGH Directs all cabin airflow to the windshield defog outlets to maximize defog, and

increases the temperature range controlled by the CABIN TEMP knob in ECS AUTO

mode.

NORM Equally divides cabin airflow between the cabin louvers/foot air outlet and the wind-

shield defog outlets. Provides adequate windshield defog for most conditions.

LOW Directs all cabin airflow to the cabin louvers and the foot air outlet to maximize cabin

cooling.

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2.13.9 Cabin Pressurization. The cabin is pressurized using airflow from the cabin heating and

cooling system. Cabin pressurization is controlled by the CABIN PRESS switch and automatic

operation of the cabin pressure regulator. Cabin pressure altitude is displayed on a cabin pressure

altimeter.

Cabin pressure is controlled by the cabin pressure regulator which regulates exit airflow to maintain

a pressure/altitude schedule. The cabin is unpressurized from sea level to an aircraft altitude of 8,000

feet. Between 8,000 and 24,500 feet aircraft altitude, cabin pressure is maintained at a constant 8,000

feet. Above 24,500 feet, cabin altitude increases slowly to approximately 14,500 feet at 35,000 feet

aircraft altitude and 20,000 feet at 50,000 feet aircraft altitude. A rule of thumb for cabin altitude above

24,500 feet aircraft altitude is aircraft altitude x 0.4.

A cabin safety and dump valve is incorporated to limit cabin pressure if the pressure regulator fails.

When the CABIN PRESS switch is placed to DUMP or RAM/DUMP, the cabin safety and dump valve

opens to release pressure to the cabin pressure regulator, reducing cabin pressure to ambient.

In ECS OFF/RAM mode, avoid operations above 25,000 feet MSL in

order to prevent decompression sickness (DCS). Normal cabin pressur-

ization is not provided. Partial cabin pressurization is available under

certain circumstances, but may be lost insidiously without warning.

2.13.9.1 CABIN PRESS Switch. The CABIN PRESS switch, located on the ECS panel on the right

console, is used to control cabin pressurization. The switch is lever-locked in the NORM position.

NORM Automatically regulates cabin pressurization according to the cabin pressure schedule

using the cabin pressure regulator (ECS AUTO mode).

DUMP Dumps cabin pressurization. Normal ECS airflow to the cabin and the avionics is not

affected.

RAM/ Dumps cabin pressurization. Terminates all ECS airflow to the cabin by closing the

DUMP cabin flow valve (Item 9). Cabin airflow is provided by the cabin ram air scoop. Avion-

ics airflow is provided by a simplified control scheme.

2.13.9.2 Cabin Pressure Altimeter. A cabin pressure altimeter, located on the center console in the

front cockpit and the lower left instrument panel in the rear cockpit, displays the current cabin

pressure altitude.

2.13.9.3 Cabin Pressurization Warning System (CPWS). The CPWS pressure switch monitors cabin

pressure and aircraft relays monitor related controls to warn aircrew of potentially hazardous cabin

pressurization conditions.

2.13.9.3.1 CABIN Caution Light. The yellow CABIN caution light is located on the lower right

caution lights panel. The light illuminates when cabin pressure altitude is above 21,000 +/- 1,100 feet.

The light may not extinguish until cabin pressure altitude is below 16,500 feet.

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• CABIN light may appear with normal cabin pressurization when

aircraft altitude is above 47,000 feet MSL. If altitude is maintained,

aircrew should continuously monitor physiological condition.

• DCS may be experienced when operating with cabin pressure altitude

above 25,000 feet even with a working oxygen system. Symptoms of

DCS include pain in joints, tingling sensations, dizziness, paralysis,

choking, and/or loss of consciousness.

NOTE

There is no corresponding DDI caution for the CABIN caution light.

2.13.9.3.2 CK ECS Caution Light. The yellow CK ECS caution light is located on the lower right

caution lights panel. The light illuminates when the position of cabin pressurization related controls

will inhibit cabin pressurization.

NOTE

There is no corresponding DDI caution for the CK ECS caution light.

2.13.10 Windshield Anti-ice and Rain Removal. The windshield anti-ice and rain removal systems

use the same air nozzle to direct warm air over the external windshield in order to improve pilot

forward visibility in icing/raining conditions. Warm air is provided by mixing engine bleed air with

output from the primary heat exchanger. Windshield air nozzle orientation is intended to affect (in

flight) an area roughly 20 inches to the left and 9 inches to the right of centerline at design eye level

and below. System operation is controlled by the WINDSHIELD switch.

2.13.10.1 WINDSHIELD Switch. The WINDSHIELD switch, located on the right console outboard of

the ECS panel, is used to select either windshield anti-ice or rain removal. The switch is lever-locked

to the OFF position.

ANTI Delivers a high flow rate of 290 ±20°F air to the external surface of the windshield.

ICE

OFF Terminates anti-ice/rain removal airflow.

RAIN Delivers a low flow rate of 270 ±20°F air to the external surface of the windshield.

2.13.11 Anti-g System. The anti-g system delivers air pressure to the g-suit proportional to sensed

load factor. A button in the anti-g valve allows the aircrew to test system operation by manually

inflating the anti-g suit. The system incorporates a pressure relief valve to prevent over-pressurization.

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2.13.12 ECS RESET and AV COOL Switch.

2.13.12.1 ECS RESET. If the ECS is DEGD, selecting the ECS RESET option from the BIT/

HYDRO-MECH display commands an ECS controller software reset and may restore normal

functionality following a transient fault.

Resetting the ECS controller at mid to high power settings may cause

uncomfortable cabin pressure surges and ear pain.

Resetting the ECS controller results in the loss of system over-

temperature protection for 10 to 70 seconds.

2.13.12.2 AV COOL Switch. The AV COOL switch, located on the lower right instrument panel,

extends the spring−loaded FCS ram air scoop located on the right side of the forward fuselage for

emergency cooling of FCC A, the right TR, and one AHRS unit. The AV COOL switch should be placed

in EMERG if an FCS HOT caution is displayed or during ECS OFF/RAM mode operation as a

preventative measure. Once extended, the FCS emergency ram air scoop cannot be retracted in flight.

EMERG Extends the FCS emergency ram air scoop to provide direct ram air cooling of FCC A,

the right TR, and one AHRS unit, and supplemental cooling to other avionics.

NORM FCS emergency ram air scoop not extended. The switch is spring−loaded to the NORM

position.

2.13.13 LCS - Liquid Cooling System. The LCS is a closed loop system normally used to transfer

heat from the radar transmitter to fuel and/or ambient air. When fuel temperatures are warmer than

the liquid coolant (typically at low fuel levels and high ambient temperatures), the LCS is used to

transfer heat from the fuel to ambient air. The LCS contains a liquid coolant pump, a liquid coolant/air

heat exchanger, and an LCS ground cooling fan (all three located in the left LEX) and two liquid

coolant/fuel heat exchangers, and in LOT 26 and up, two liquid coolant/ECS air heat exchangers.

The LEX liquid coolant/air heat exchanger has one air inlet located on the bottom of the LEX, and

two air exhausts. Only one exhaust path is commanded open at any given time. During LCS ground

cooling fan operation, the exhaust on the bottom of the LEX is commanded open. Inflight, only the air

exhaust on the top of the LEX can be commanded open.

LCS ground cooling fan air is the primary cooling source for liquid coolant on deck. During ground

operations, the liquid coolant pump and LCS ground cooling fan are commanded on when power is

applied to the radar (RADAR knob in STBY, OPR, or EMERG).

During ground operations, when feed tank fuel temperatures exceed 40°C (30°C in LOT 26 and up),

the liquid coolant pump and LCS ground cooling fan may also be commanded on if the RADAR knob

in OFF in order to provide LCS cooling of the fuel system. In LOT 25 and below, the bypass valve,

which allows liquid coolant to the liquid coolant /fuel heat exchangers, is solely controlled by the liquid

coolant temperature, so the amount of fuel cooling provided (if any) depends on the temperature of

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each fluid. In LOT 26 and up, the bypass valve is controlled by the ECS controller and will only allow

liquid coolant to the liquid coolant/fuel heat exchanger with the RADAR knob in OFF.

NOTE

In LOT 26 and up, the RADAR knob must be in OFF in order to

provide any postflight LCS fuel cooling.

In either case, placing the RADAR knob to OFF removes the radar as a heat source and should

extend ground operating time.

Inflight, the LCS ground cooling fan secures, the lower air exhaust path closes, and the upper air

exhaust path is controlled by the mission computer. The upper air exhaust path opens when feed tank

fuel temperatures exceed 40°C (30°C in LOT 26 and up), ram air is cooler than fuel, and AOA is below

15°C. In LOT 25 and below, heat not removed by the LEX liquid coolant/air heat exchanger is removed

by the fuel system through the liquid coolant/fuel heat exchangers. In LOT 26 and up, heat not

removed by the LEX liquid coolant/air heat exchanger is removed by a combination of the liquid

coolant/fuel and liquid coolant/ECS air heat exchangers.

In LOT 26 and up, provisions have been incorporated to limit the risk of fire following a major liquid

coolant leak. When LCS low pressure is detected, the liquid coolant pump secures and check valves

limit coolant leakage. Since the SDC controls the liquid coolant pump, an SDC reset secures the pump,

resulting in a radar OVRHT (APG−73) or LOFLOW (APG−79) indication, and the radar stops

transmitting.

NOTE

In LOT 26 and up, an SDC reset temporarily secures the liquid coolant

pump. Before selecting SDC RESET, the RADAR knob should be set

to STBY until the SDC reset is complete and normal radar cooling

capability is restored.

2.13.14 ECS Related Warnings, Cautions, and Advisories. The following ECS related warnings,

cautions, and advisories are described in the Warning/Caution/Advisory Displays in Part V:

D L BLEED and R BLEED warning lights

D FCS HOT caution

D BLEED AIR LEFT (RIGHT) voice alert

D FLIR OVRHT caution

D AV AIR HOT caution

D GUN GAS caution

D L or R BLD OFF cautions

D TK PRES LO caution

D CABIN caution light

D TK PRES HI caution

D CK ECS caution light

D WDSHLD HOT caution

D ECS ICING caution

D ECSDR advisory

D EXT TANK caution

If ECS troubleshooting has been done while on deck, check for an 8A6

MSP code after takeoff. An 8A6 MSP code is set by a system flow

modulating regulating valve failure. This failure can potentially lead to

an unexpected loss of cabin pressurization.

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F18. FLIGHT MANUAL (2008) - page 18