EA-18G. FLIGHT MANUAL (2008)

A1-E18GA-NFM-000

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.

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 2-25.

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

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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. With flaps HALF or FULL, 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. With flaps HALF or FULL, this

interconnect is stabilator to aileron. With flaps AUTO, it is stabilator to aileron and differential TEF.

At high airspeed with flaps AUTO, 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

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),

above 200 KCAS (flaps HALF), or above 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

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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

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.

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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.

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.

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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

FCS HOT

Note

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 +5.5 g 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.

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). With flaps AUTO, P CAS uses the AOA estimator for control law

scheduling. If a flaps AUTO failure clears, AOA Xs are removed automatically with or without an FCS

RESET attempt. With flaps HALF or FULL, P CAS uses a fixed 8.1° AOA value, and R CAS uses the

AOA estimator for control law scheduling. If a flaps HALF or FULL 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

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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.

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.

S. Blank Surface Position: FCCs and/or MCs unable to report actuator position.

T. ROLL TRIM: 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 qualitative measure of how much roll trim has

been commanded.

U. PITCH TRIM: 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 AUTOMATIC FLIGHT CONTROL SYSTEM (AFCS)

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 angle 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.

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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).

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. This is true for all modes except when

disengaging the ACL mode by unboxing the CPL P/R option. In that case, the AUTO PILOT caution

is displayed and all autopilot modes disengage. 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 and the AUTO PILOT caution is displayed. If the stick is moved either longitudinally or laterally

with CPL P/R engaged, the AUTO PILOT caution is displayed and all autopilot modes disengage. The

basic autopilot mode 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 Barometric Altitude Hold (BALT). 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 Radar Altitude Hold (RALT). 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 Flight Path Angle Hold (FPAH). 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 angle 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 Attitude Hold (ROLL). 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.

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2.11.5.2 Ground Track Hold (GRTK). 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°).

2.11.5.3 Ground Track Select (GSEL). 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 Heading Hold (HDG). 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 Heading Select (HSEL). 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 Coupled Steering Modes (CPL). 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 NTRP 3-22.2-EA-18G (EA-18G Classified 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

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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, or the rear cockpit hand controllers. 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.

Detailed descriptions of the functionality of the weapon systems controls are contained in the NTRP

3-22.2-EA-18G (EA-18G Classified Manual).

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 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.

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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.

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,

etc.) perform a specific action (e.g., track, break track, etc.) when the castle switch is subsequently

bumped toward that display.

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2.12.1.3 Trigger. The 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 selected A/A missile. Activates AGI protection. In

Detent

A/G master mode, fires the laser, if either is selected. Commands the HUD event

marker.

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 to break lock, if it is tracking. 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.

2.12.2 Throttle Grip Switches/Controls. 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.

2.12.2.1 Chaff/Flare/ALE-50 Dispense Switch. The chaff/flare/ALE-50 dispense switch is located

on the top inboard side of the right throttle.

2.12.2.2 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) toggle between

weapon modes, or (3) reset sequenced HARM targets.

2.12.2.3 Throttle Designator Controller (TDC). 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. 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. If in ACM mode, pressing

TDC will exit the radar from ACM mode.

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Figure 2-31. Throttle Grip Switches/Controls

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.

When display priority assigned to TSD, the elevation control switch changes the range scale. When

display priority is assigned to other AEA formats with a selection box, the elevation switch provides a

page up and down feature.

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, or (2) toggle between narrow and wide field of view (FOV).

2.12.4 Hand Controllers. The functions of the left and right hand controllers are not identical. The

left hand controller contains the following switches: Countermeasures, 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, Undesignate, and Right Trigger. See

figure 2-32. Refer to NTRP 3-22.2-EA-18G (EA-18G Classified Manual) for complete Hand controller

Switch description.

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2.12.4.1 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.

Forward

Not functional.

2.12.4.2 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.

2.12.4.3 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.

2.12.4.4 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.

2.12.4.5 Chaff/Flare Dispense Switch. The chaff/flare dispense switch is located on the outboard

side of each hand controller. Dispense switch functionality differs between the left and right hand

controllers.

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2.12.5 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).

2.12.6 Grab Handle Chaff/Flare Dispense Switches. 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)

2.13 ENVIRONMENTAL CONTROL SYSTEM (ECS)

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 external fuel tank pressurization,

canopy seal inflation, g-suit operation, radar waveguide pressurization, windshield anti-ice and rain

removal, and on-board oxygen generating system (OBOGS) operation. Cold, dry conditioned air is

provided for avionics cooling, and supplied to the Airborne Electronic Attack (AEA) pallet for

ALQ-218(V)2 system precooling or extended maintenance. 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 for component identification.

2.13.1

Airborne Electronic Attack (AEA) pallet Cooling.

2.13.1.1 Nose wheelwell AEA PRECOOL TEST switch. A two position switch, TEMP CHECK and

LAMP TEST, and a light emitting diode PRECOOL are located in the nose wheelwell and are used

during preflight inspection to determine if AEA pallet pre cooling is required.

2.13.1.2 Cockpit AEA Precool switch. This two position switch is located on the forward cockpit

right hand console. With weight on wheels, operating the switch to ON diverts cool air from the cockpit

to the AEA pallet to precool the ALQ-218 components. This switch is magnetically held ON. When

Precooling is no longer required, returning the switch to OFF terminates AEA precooling, and cockpit

cooling resumes. The switch automatically reverts to OFF at weight off wheels. It may take as long as

20 minutes for Precool to end. Cockpit Cooling is unavailable during Precool if the precool switch was

used.

2.13.1.3 UFCD/CNI-AEA power control format. ALQ218 PRECOOL REQD is displayed on the

ALQ-218 pushtile when AEA precooling is required. When precool is initiated by placing the

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PRECOOL switch ON, the ALQ-218 pushtile changes to PRCOOL ON. When precool is no longer

required, PRCOOL ON clears.

2.13.1.4 AEA Pallet Ground Cooling Mode. The AEA PRECOOL switch energizes relays that: opens

the cabin air pressure emergency relief valve, opens the cabin/defog ram air valve, and closes the cabin

flow valve. This enables maximum airflow through the avionics cooling system to the EAU on the AEA

pallet as well as into the pallet bay. The AEA Ground cooling mode continues until either the AEA

PRECOOL switch is set to OFF, or it is deenergized by the right weight-off-wheels relay. ECS Related

Warnings, Cautions, and Advisories.

2.13.2 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

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.

2.13.2.1 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.

2.13.2.2 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:

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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).

2.13.3 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

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.

2.13.4 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.

2.13.4.1 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 0.33 Mach and are always closed above 0.40 Mach. Between 0.35 and 0.40 Mach,

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.

2.13.4.2 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.

2.13.5 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: external

fuel tank pressurization, canopy seal, g-suit, radar waveguide pressurization, and OBOGS.

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The ECS controller modulates the warm air temperature control valve (see FO-29 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, ECS

turbine anti-ice, and cabin heating.

2.13.6 Air Conditioning System (ACS) Pack. The ACS pack provides cold, dry conditioned air for

cabin, and avionics cooling. The primary components of the ACS pack are a compressor, turbine,

condenser, reheater and water extractor.

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.

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.

2.13.7 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.

2.13.8 Avionics Cooling Fans. Two avionic cooling fans augment ECS cooling of avionics. The

avionics ground cooling fan is located in the nose wheelwell. 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.

2.13.8.1 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.

2.13.9 ECS Operating Modes. The ECS has three operating modes: AUTO, MAN (manual), 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).

2.13.9.1 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 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. If AEA precool is required, advance one throttle

to 76% until precool indication is no longer present. 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 cooling system.

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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2.13.9.2 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.

The ECS MAN mode is a degraded operating mode. At high power settings, ECS MAN mode can

significantly increase the amount of bleed air drawn from the engines, resulting in higher turbine

temperatures and reduced engine life. Therefore, the ECS MAN mode should only be used when the

ECS AUTO mode is degraded and temperatures are out of limits (e.g., cabin temperature high or AV

AIR HOT caution inflight).

ECS AUTO mode should be selected unless ECS performance is

degraded, cabin temperature control is lost, or an AV AIR HOT caution

occurs. Extended ECS MAN mode operation, particularly at high power

settings, significantly reduces engine life.

ECS valves are configured to produce maximum ECS flow. The division of airflow between the cabin

and the avionics is dependent on WonW status and throttle setting.

Avionics cooling fan operation is identical to ECS AUTO mode.

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 MAN 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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2.13.9.3 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 control is identical to ECS MAN mode.

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.

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2.13.9.4 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. Directs APU air to the ECS for cabin and avionics cooling.

2.13.9.5 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

Selects ECS MAN mode, a degraded backup ECS operating mode. Provides full,

fixed cabin and avionics airflow, manual temperature control, and full cabin pres-

surization.

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.

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.

2.13.9.6 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 MAN or OFF/RAM

modes, the CABIN TEMP knob controls the cabin add heat valve directly, producing a nonlinear

temperature response.

NOTE

D In ECS AUTO mode, cabin temperature may take 1 to 2 minutes to

stabilize following a large movement of the CABIN TEMP knob.

D In ECS MAN mode, changes in the position of the CABIN TEMP

knob should be held for 30 seconds due to a slower temperature

response.

2.13.9.7 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.

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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.

2.13.9.8 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.

2.13.9.9 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.

2.13.10 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.

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2.13.10.1 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 and MAN modes).

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.

2.13.10.2 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.

2.13.10.3 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.

2.13.10.3.1 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.

• 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.

2.13.10.3.2 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.

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2.13.11 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 airflow 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.

2.13.11.1 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.

2.13.12 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.

2.13.13 ECS RESET and AV COOL Switch.

2.13.13.1 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.

2.13.13.2 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.

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2.13.14 Liquid Cooling System (LCS). 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 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 30°C, the liquid coolant pump

and LCS ground cooling fan may also be commanded on if the RADAR knob is in OFF in order to

provide LCS cooling of the fuel system. T ECS controller will only allow liquid coolant to the liquid

coolant/fuel heat exchanger if the RADAR knob is in OFF.

NOTE

The RADAR knob must be in OFF in order to provide any postflight

LCS fuel cooling.

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 30°C, ram air is cooler than fuel, and AOA is below 15°C. 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.

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 LOFLOW indication, and the radar stops transmitting.

NOTE

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.

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2.13.15 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 EAU OVRHT caution

D AV AIR HOT caution

D ECS ICING caution

D L BLEED and R BLEED warning lights

D ECSDR advisory

D BLEED AIR LEFT (RIGHT) voice alert

D EXT TANK caution

D L or R BLD OFF cautions

D FCS HOT caution

D CABIN caution light

D WDSHLD HOT caution

D CK ECS caution light

2.14 OXYGEN SYSTEMS

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

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

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

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

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

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

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

services panel oxygen disconnect to the mask will prevent either normal

OBOGS breathing gas or emergency oxygen from being delivered to the

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

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

system degrade occurs, immediate descent below

10,000

feet cabin

altitude is required to prevent severe or incapacitating hypoxia.

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

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

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

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

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

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

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

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Figure 2-33. OBOGS Monitor

Loss of electrical power to the OBOGS monitor prevents reporting of

OBOGS DEGD conditions.

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

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

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

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

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

to create a low oxygen concentration condition. Momentarily pressing and releasing the electronic BIT

pushbutton tests the monitor electronically. Successful completion of either test activates the OBOGS

DEGD caution, which clears automatically after BIT is complete and OBOGS resets to normal

operation.

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

indicates that there is no protection against inadequate oxygen concen-

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

gas flow alone does not ensure adequate oxygen concentration.

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

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

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

NOTE

Oxygen concentration may drop below the predetermined OBOGS

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

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

aircrew hose connections can overwhelm system capacity and may

result in an OBOGS DEGD caution.

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

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

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

pressure and an abrupt inability to breathe.

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

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

physiological requirements.

When cabin altitude is above 10,000

feet, OBOGS DEGD caution

procedures shall be executed for low mask flow, increased breathing

resistance, or OBOGS DEGD cautions of any duration.

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

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

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

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

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

Repeated unexplained MASTER CAUTION lights/tones may be an

indication of OBOGS system degradation.

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

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

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

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

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

Pure oxygen accelerates recovery from hypoxia. Emergency oxygen shall

be used whenever hypoxic symptoms are recognized.

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

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

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

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

the seat survival kit oxygen disconnect.

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

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

valve.

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

OFF OBOGS system off.

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

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

ON OBOGS flow supplied to the mask.

OFF OBOGS flow secured.

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

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

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

rotated to the ON or OFF detent position.

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

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

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

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

services panel oxygen disconnect to the mask will prevent either normal

OBOGS breathing gas or emergency oxygen from being delivered to the

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

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

system degrade occurs, immediate descent below

10,000

feet cabin

altitude is required to prevent severe or incapacitating hypoxia.

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

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

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

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

available.

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

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

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

release tab immediately forward of the green ring.

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

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

emergency oxygen from reaching the breathing regulator. Additionally,

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

FLOW knob.

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

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

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

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

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

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

extinguishing capability.

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

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

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

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

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

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

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

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

AMAD bay.

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

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

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

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

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

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

directed by NATOPS (following throttle OFF for actual emergencies or

as specifically delineated for an FCF A profile).

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

stripes is visible around the outer edges of the light.

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

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

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

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

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

visible around the outer edges of the light.

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

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

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

RIGHT, then APU.

2.15.3.1 FIRE Warning Lights (Rear Cockpit). The rear cockpit left, right, and APU FIRE warning

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

bottle or shut down the engines or APU.

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

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

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

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

illuminated.

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

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

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

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

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

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

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

within 5 seconds.

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

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

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

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

until the DISCH light does come on.

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

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

both fires.

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

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

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

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

APU time to spool down before extinguishing agent is introduced.

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

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

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

should be performed to backup the automatic system.

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

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

turned off during the 10 second delay time.

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

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

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

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

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

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

Fire testing indicates that the probability of extinguishing a fire and

preventing relights is greatly increased by immediately discharging the

fire extinguisher.

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

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

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

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

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

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

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

FIRE lights during the loop test.

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

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

FIRE voice alerts.

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

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

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

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

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

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

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

both primaries).

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

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

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

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

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

by the aircrew.

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

BLEED warning lights prior to aircrew recognition and may not trigger

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

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

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

automatic shutdown and isolation capability may be lost. Extensive

damage and/or fire may result.

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

appropriate sensing element.

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

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

leak codes.

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

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

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

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

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

successfully passed.

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

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

NORM with ac power applied.

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

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

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

TEST A

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

NORM

Provides normal fire and bleed air leak detection.

TEST B

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

2.16 ENTRANCE/EGRESS SYSTEMS

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

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

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

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

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

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

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

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

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

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

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

and be stored on the windshield and canopy. If possible, ensure that

ground crew discharge the static electricity prior to egress. Otherwise,

avoid direct contact with the outside of the windshield and canopy to

prevent electrical shock.

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

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

aircraft in motion.

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

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

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

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

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

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

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

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

canopy at its present position.

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

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

handles or the external canopy jettison handle.

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

canopy sill in the front cockpit. The rear cockpit CANOPY switch is located on the lower right portion

of the instrument panel. The CANOPY switch is spring loaded to the HOLD position and is

solenoid-held in the OPEN position only with WonW. The solenoid can be overridden at any time by

placing the switch to HOLD. With WoffW, the switch must be held in the OPEN position to raise the

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

result in a fail-safe OPEN command.

OPEN

Unlocks and/or raises the canopy.

HOLD

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

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

closed and locked).

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

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

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

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

solenoid held. The switch is guarded to prevent inadvertant actuation.

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

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

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

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

provided to prevent loss of the handle if dropped.

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

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

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

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

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

unlatch the canopy by moving it 1½ inches rearward, after which two canopy frame mounted rocket

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

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

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

jettisoned closed, open, or in any intermediate position.

2.16.1.2.1 CANOPY JETT Handle (Front Cockpit). The CANOPY JETT handle is black and yellow

striped and is located on the left inboard canopy sill just aft of the instrument panel in the front

cockpit. Pressing the button on the tip of the handle unlocks the handle. Pulling the handle aft initiates

the canopy jettison sequence. A REMOVE BEFORE FLIGHT pin is used to manually secure the

CANOPY JETT handle between flights.

2.16.1.2.2 CANOPY JETT Handle (Rear Cockpit). The rear cockpit CANOPY JETT handle is black

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

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

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

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

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

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

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

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

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

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

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

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

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

footing.

NOTE

The ladder is not visible from the cockpit.

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Figure 2-35. Boarding Ladder

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

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

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

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

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

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

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

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

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

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

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

LADDER caution goes out.

The ladder is extended remotely by opening the external power receptacle door and pressing and

holding the guarded ladder remote release button. The three latches are opened by a battery powered

actuator. The ladder drops to the open position while being restrained from free falling by a dampening

strut (drop time approximately 3 to 4 seconds). All other procedures for securing the ladder are the

same as manual opening.

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2.16.2.1 Ladder Remote Release Button. The guarded external ladder remote release button is

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

and LEX. Pressing and holding the button applies electrical power to the boarding ladder actuator

which unlocks the three uplock latches so the ladder can free fall.

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

external canopy switch, positive switch identification is required to

prevent inadvertently lowering the canopy and injuring personnel egress-

ing the aircraft.

2.16.3 Ejection Seat. The SJU-17B(V) 2/A and 9/A NACES (Navy Aircrew Common Ejection Seat)

are ballistic catapult/rocket systems that provide the pilot with a quick, safe, and positive means of

escape from the aircraft. See Ejection Seat, foldout section, for ejection seat illustrations. The seat

system includes an initiation system which, after jettisoning the canopy and positioning the occupant

for ejection, fires the telescopic seat catapult. Canopy breakers on the top of the seat give capability of

ejecting through the canopy. As the seat departs the aircraft and the catapult reaches the end of the

stroke, a rocket motor on the bottom of the seat is fired. The thrust of the rocket motor sustains the

thrust of the catapult to eject the seat to a height sufficient for parachute deployment even if ejection

is initiated at zero speed, zero altitude in a substantially level attitude.

NOTE

Safe escape is provided for most combinations of aircraft altitude,

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

airspeed and 0 to 50,000 feet.

Timing of all events after rocket motor initiation is controlled by the electronic sequencer which

utilizes altitude, acceleration, and airspeed information to automatically control drogue and parachute

deployment and seat/man separation throughout the ejection seat’s operational envelope. In the event

of partial or total failure of the electronic sequencer, a 4-second mechanical delay initiates a barostatic

release unit which frees the occupant from the seat and deploys the parachute between 14,000 and

16,000 feet MSL if the ejection occurred in or above this altitude range. The emergency barostatic

release unit operates immediately after the 4-second delay if the ejection occurred below 14,000 feet

MSL. An emergency restraint release (manual override) system provides a backup in the event of

failure of the barostatic release unit. The seat is stabilized and the forward speed retarded by a drogue

chute attached to the top and bottom of the seat. The parachute deployment rocket is automatically

fired to withdraw the parachute from the deployment bag. Full canopy inflation is inhibited until the

g forces are sufficiently reduced to minimize opening shock. There are five modes of operation. See

figure 2-36 for parameters that determine the mode of operation and the corresponding parachute

deployment and drogue chute release times. At high altitude the drogue chute deploys to decelerate

and stabilize the seat. The seat falls drogue retarded to 18,000 feet MSL where the drogue is released,

the main parachute is deployed, and seat/man separation occurs. At medium altitude (between 18,000

and 8,000

feet MSL), and at low altitude

(below

8,000

feet MSL) parachute deployment is

automatically delayed from 0.45 to 2.90 seconds (depending upon airspeed and altitude) after first seat

motion to allow the drogue chute to decelerate and stabilize the seat.

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

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

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Figure 2-36. SJU-17B Ejection Modes

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

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

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

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

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

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

parachute.

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

the dual seat initiators.

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

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

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

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

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

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

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

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

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

the seat occupant for ejection.

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

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

pressure, moved to the desired position, and released.

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

(locked)

slack in the straps back into the reel.

AFT

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

(unlocked)

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

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

on the reel straps.

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

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

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

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

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

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

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

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

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

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

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

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Figure 2-37. Leg Restraint System

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

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

instrument panel during ejection.

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

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

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

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

ejection seat. This survival kit, which fits into the seat bucket, is a contoured rigid platform which

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

the platform provides a seat for the aircrew.

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

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

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

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

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

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

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

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

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

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

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

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

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

connected to the emergency oxygen lanyard.

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

take place: release of survival kit attachment lugs, negative-g strap, leg restraint lines, and inertia reel

straps; firing of the manual override initiator cartridge; firing of the barostatic release unit; and firing

of the parachute deployment rocket, which deploys the parachute. The ejection seat safe/armed handle

automatically rotates to the SAFE position whenever the manual override handle is actuated.

Pulling the manual override handle automatically rotates the ejection

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

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

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

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

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

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

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loaded, stops the seat bucket. The maximum vertical travel of the seat bucket is 6.1 inches. The

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

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

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

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

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

forward backpad position.

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

under the seat bucket may result in an inadvertent ejection.

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

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

seat, leg restraints, and/or Koch fittings.

2.16.3.10 Backpad Adjustment Mechanism. The backpad adjustment mechanism handle is on the

seat bucket adjacent to the top left hand side of the backpad and is connected to the backpad by a

linkage. The backpad has three positions, full-forward, middle, and full-aft, which give a total

forward/aft adjustment of 1.6 inches. When the handle is in the full-up position, the backpad is full-aft,

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

handle is moved within a quadrant until a spring-loaded plunger engages in one of the three detent

positions in the quadrant. Set the backpad for personal comfort and best access to flight controls

during initial strap-in and prior to flight.

2.16.4 Ejection Seat System. The ejection seats are ejected at opposite divergent angles to one

another. The rear seat diverges to the left while the forward seat diverges to the right. The amount of

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

aircrew and the faster the speed, the less the resulting divergent angle. In addition, a sequencing system

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

the rear cockpit. A command selector valve is installed in the rear cockpit to control whether ejection

from the rear cockpit is dual or single.

2.16.4.1 EJECT MODE Handle. The EJECT MODE handle is located on the right side of the main

instrument panel in the rear cockpit. The EJECT MODE handle is used to select the desired ejection

sequence to be initiated from the rear cockpit, or provide for single ejection for solo flight. Positioning

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

use of a collar to hold the handle in that position. To release from AFT INITIATE, pull then turn

clockwise.

NORM

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

(vertical)

seat first) when initiated from the front cockpit.

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AFT

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

INITIATE

(horizontal)

SOLO

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

(45° CCW)

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

ejection is immediate.

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

SOLO mode is selected when both seats are occupied, simultaneous

ejection initiation may result in a collision between seats.

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

Alternate selection when flying solo results in ejection of unoccupied

seat and possible collision with the front cockpit seat.

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

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

result.

2.16.4.2 SEAT CAUT MODE Switch. The SEAT CAUT MODE switch is located in the rear cockpit

above the EJECT MODE handle. The switch position changes the operation of the CK SEAT caution

for solo or dual flight.

NORM

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

loaded to this position.

SOLO

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

be pinned to remain in this position.

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

repeats the DDI CHECK SEAT caution display. The caution is displayed when the right throttle is at

MIL or above, weight is on wheels, and the front seat is not armed with the SEAT CAUT MODE switch

set to SOLO or either seat is not armed with the SEAT CAUT MODE switch set to NORM.

2.17 EMERGENCY EQUIPMENT

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

selective jettison system.

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

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

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

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

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

is jettisoned and in between each subsequent set.

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

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

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

LDG GEAR handle UP or with WoffW.

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

stores are jettisoned.

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

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

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

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

On).

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

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

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

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

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

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

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

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

tone with the LDG GEAR handle UP.

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

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

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

SELECT JETT knob.

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

pushbutton in the SELECT JETT knob.

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

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

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

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

modes for weapon selection; refer to NTRP 3-22.2-EA-18G (EA-18G Classified Manual).

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

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

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

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

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

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

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

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

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

indicate system malfunctions requiring immediate action. Caution lights and displays indicate

malfunctions requiring attention but not immediate action. After the malfunction has been corrected,

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

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

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

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

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

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

lights panel and the instrument panel, are yellow lights. The advisory lights, scattered throughout the

cockpit(s), are green. Lights that have been lit on the caution lights panel flash when overheated to

prevent light damage.

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

2.17.2.2 MASTER CAUTION Light (Rear Cockpit). A yellow MASTER CAUTION light, on the

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

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

CAUTION is reset.

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

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

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

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

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

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

engine FIRE warning.

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

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

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

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

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

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

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

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

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

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

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

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

repeated at the lowest priority until reset or disabled by the pilot. The BINGO voice alert is repeated

every 30 seconds until the BINGO setting is adjusted.

CAUTION

VOICE ALERT

IFF 4

MODE 4 REPLY

FCS

FLIGHT CONTROLS

FCS HOT

FLIGHT COMPUTER HOT

L (R) OVRSPD

L (R) EGT HIGH

L (R) FLAMEOUT

ENGINE LEFT (RIGHT)

L (R) OIL PR

L (R) STALL

L (R) ENG

L (R) ENG VIB

FUEL LO

FUEL LOW

BINGO

WARNING

VOICE ALERT

ALTITUDE

L (R) BLEED AIR

BLEED AIR LEFT (RIGHT)

L (R) FIRE

ENGINE FIRE LEFT (RIGHT)

APU FIRE

2.17.4

Terrain Awareness Warning System (TAWS).

The terrain awareness warning system alerts

the aircrew of a controlled flight into terrain (CFIT) condition during all mission phases. The system

operates any time that the navigation mission computer (MC1) and TAMMAC digital mapping set

(DMS) are functional. TAWS functions as a safety backup system and not as a performance aid.

TAWS has been designed to eliminate false warnings, minimize nuisance warnings, and generate

consistent aircrew response in all aircraft master modes. Five possible voice warnings are provided to

indicate the correct initial response to an impending CFIT condition, and a visual cue is provided to

indicate the recovery direction of pull, or in some instances, to command an increase in turn rate. All

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TAWS warnings should be treated as though an imminent flight into terrain condition exists. Pilot

response to a TAWS warning should be instinctive and immediate.

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

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

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

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

automatically at start-up.

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

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

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

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

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

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

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

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

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

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

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

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

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

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

against flight into level or descending terrain as GPWS does.

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

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

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

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

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

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

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Figure 2-39. TAWS On/Off Pushbutton

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

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

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

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

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

computed trajectories include the following assumptions:

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

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

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

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

will be used for bank angles ≥ 70°.

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

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

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

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

corner speed.

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

TAWS assumes a terrain clearance of 50 ft.

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

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

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

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

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

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

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

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

locked, a TAWS warning is provided.

2.17.4.3

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

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

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

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

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

″Roll−Left...Roll−Left″,

″Roll−Right

Roll−Right″,

″Pull−Up...Pull−Up″,

″Power...Power″, and

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

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

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

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

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

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

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

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

is above 8.5° with flaps HALF or FULL (or 18° AOA for flaps AUTO) and the throttle is not already

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

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

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

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

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

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

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

has been assessed.

2.17.4.3.2

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

vector with the HUD recovery arrow and perform a dive recovery. Figure 2-42 depicts a situation in

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

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

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

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

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

procedure.

Figure 2-42. TAWS Visual Recovery Cue (Arrow displayed perpendicular to actual horizon)

Figure 2-43 contains the ACI aural cue priorities.

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Pull-Up (TAWS/GPWS)

Roll-Left (TAWS/GPWS)

Roll-Right (TAWS/GPWS)

Roll-Out (Not Used)

Check gear (TAWS/GPWS)

Power (TAWS/GPWS)

Engine Fire Left

Hardware Discrete

Engine Fire Right

Hardware Discrete

APU Fire

Hardware Discrete

Bleed Air Left

Hardware Discrete

Bleed Air Right

Hardware Discrete

Hardware Discrete (primary)

Deedle (Master Caution)

MC (backup)

Missile (Not Used)

Fuel Low

Bingo

Altitude

Whoop (LAW Tone)

Flight Computer Hot

Flight Controls

Mode 4 Reply

Engine Left

Engine Right

HF Comm

Reserved (Climb (TLAS))

Reserved (Sink Rate (TLAS))

Reserved (Guide Slope (TLAS))

Whoop (continuous)

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

All voice alerts play until completed.

Figure 2-43. ACI Aural Cue Priorities

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2.17.4.4 DFIRS Record Code 8. TAWS data is sent to DFIRS and MU maintenance card each time

TAWS/GPWS detects a potential CFIT condition. The data, which is sent as record 8, consists of:

1. TAWS/GPWS pushbutton state.

2. Intended aural warning.

3. Operational state of TAWS.

4. TAWS elevation source (DTED, flat earth).

5. TAWS operational mode.

6. Recovery arrow axis.

7. Ground intercept point latitude, longitude and altitude.

Only data items 1 and 2 are valid when TAWS is not operational.

2.17.4.5 TAWS and MSP code 11A. An MSP code of 11A indicates an incorrect configurable

parameters file for TAWS has been loaded onto the mission card. 11A MSP code indicates a DMC

software degrade. When this code appears, TAWS is no longer operational, and MC logics has switched

to the GPWS algorithm hosted in MC1.

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

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

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

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

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

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

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

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

recovery.

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

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

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

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

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

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

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

nuisance warnings.

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

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

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

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

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

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

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

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

data is restored.

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

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

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

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

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

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

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

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

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

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

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

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

2.17.5.3 CFIT Protection Provided.

Above 150 feet AGL -

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

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

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

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

Below 150 feet AGL -

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

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

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

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

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

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

condition no longer exists for 0.3 seconds.

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

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

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

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

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

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

down.

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

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

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

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Figure 2-44. GPWS HUD Roll Warning Cues

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

second.

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

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

following a warning under extreme flight conditions.

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

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

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

the center of the HUD. See figure 2-44. The HUD recovery arrow is always perpendicular to the horizon

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

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

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

2.17.5.4.2 GPWS Voice Commands. Refer to figure 2-45 for GPWS aural warning cues.

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GPWS Warning

Aural Cue

Repetition Rate

Condition

Excessive bank angle

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

2 seconds

Excessive takeoff

″POWER, POWER″

2 seconds

sink rate

Excessive landing

″POWER, POWER″

2 seconds

sink rate

Gear-up landing

″CHECK GEAR″

8 seconds

ALDR or floor

″POWER, POWER″ for airspeed <210 KCAS and

2 seconds

altitude

AOB ≤45°

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

2 seconds

for AOB >45°

″PULL UP, PULL UP″ for all other flight

2 seconds

conditions

Figure 2-45. GPWS Aural Cues

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

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

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

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

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

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

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

call.)

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

transmitted.

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

as follows:

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

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

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

AOB, 0 to 30° dive, 450 to 560 KCAS).

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

to prevent hard landings.

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

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

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

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

1. Loss of air data or RALT, INS, or either MC1 or MC2 failed or off.

2. Less than 6 seconds after WonW.

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

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

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

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

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

2.18 INSTRUMENTS

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

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

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

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

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

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

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

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

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

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

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

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

KIAS. It operates directly from left pitot and static pressure.

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2.18.3 Standby Altimeter. The standby altimeter is a counter-pointer type. The counter drum

indicates altitude in thousands of feet from 00 to 99. The long pointer indicates altitude in 50-foot

increments with one full revolution each 1,000 feet. A knob and window permit setting the altimeter

to the desired barometric setting. This setting is also used by the flight control computers. The standby

altimeter operates directly from left static pressure.

2.18.4 Standby Rate of Climb Indicator. The standby rate of climb indicator displays vertical speed

on a scale from 0 to ±6,000 fpm and operates directly from left static pressure.

2.18.5 Standby Magnetic Compass. A conventional aircraft magnetic compass is mounted on the

right windshield arch in the front cockpit.

2.18.6 Angle Of Attack Indexer. The angle of attack indexer is mounted to the left of the HUD. It

displays approach angle of attack (AOA) with lighted symbols; corresponding AOA indications are

shown on the HUD (see figure 2-46). The indexer operates with the landing gear down and locked and

weight off the gear. The lighted symbol(s) flash if the arresting hook is up and the hook bypass switch,

on the left vertical panel, is in CARRIER. The symbols will not flash with the arresting hook up and

the hook bypass switch in FIELD. The switch is solenoid held to FIELD and automatically goes to

CARRIER when the arresting hook is lowered or aircraft power is removed. The AOA indexer knob on

the HUD controls dimming of the symbols. All symbols light when the lights test switch on the interior

lights control panel is held to TEST.

2.19 AVIONICS SUBSYSTEM

The avionics subsystem combines the integration and automation needed for operability with the

redundancy required to ensure flight safety and mission success. Key features of the system include

highly integrated controls and displays, inertial navigation set with carrier alignment capability, and

extensive built in test capability. The avionics subsystems operate under the control of two mission

computers with primary data transfer between the mission computers and the other avionics

equipment including the EAU via the mux buses and the high speed data network.

2.19.1 Mission Computer System. The mission computer system consists of two digital computers

(MC1 and MC2) which are high speed, stored program, programmable, general purpose computers

with core memory. Both MCs are capable of providing the same basic navigation and weapon delivery

back up should the other MC fail.

There are six avionics mux bus channels with redundant paths (X and Y) for each channel.

The mission computer:

1. Computes and controls the data sent to the cockpit displays,

2. Computes missile launch and weapon release commands,

3. Provides mode control and options for various avionics systems,

4. Generates BIT initiate signals to and equipment operational status from various avionics

systems.

The front and rear DDIs are driven directly by the MC over a high speed interface bus, not by

avionics mux bus commands. The HUD is driven directly by redundant connection to either MC. MC1

drives the front and rear LDDIs and HUD while MC2 drives the front and rear RDDIs and HUD.

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Figure 2-46. Angle of Attack Indexer

When an MC is off or non-functional, the displays driven by that MC show a green square in the center

of the display. Each MC provides the same level of functionality in the single MC backup mode of

operation.

With both MCs inoperative and the left generator operative, the SDC provides a limited HUD

format on the front MPCD/UFCD, prevents the FADEC and ECS controller from going into default

mode operation, and provides left/right ATS cautions when necessary. See Chapter 25 Backup/

Degraded Operations for a description of SDC Backup Mode.

MC2 provides digital video color capability to the AMPCD via the Fiber Channel Network Switch

(FCNS) and High Speed Video Network (HSVN).

The computers receive inputs for navigational data and steering command computations from the

inertial navigation system, electronic flight control system, multipurpose display group, TACAN, and

backup attitude and the navigation system. The computers control display symbology and information

presented to the pilot by the multipurpose display group.

2.19.1.1 MC Switch. The MC switch, located on the aft outboard edge of the left console, is used to

manually turn OFF either of the two mission computers, MC1 or MC2.

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1 OFF Removes power to MC1.

NORM Both MC1 and MC2 are powered with ac electrical power available.

2 OFF Removes power to MC2.

2.19.1.2 Electronic Attack Unit (EAU). The EAU provides the primary interface between the AEA

subsystems and the AMC. Within the AEA suite, the EAU interfaces with the ALQ−218, CCS,

ALQ−99 Pod suite and the MATT. The interface between the EAU and AMC consists of a High Speed

Data Network (HSDN) and MIL−STD−1553 bus interface. The EAU provides jammer management

logic, manages AEA libraries, provides power control, and interfaces Built In Test (BIT) for the AEA

subsystems. The EAU coordinates aircraft navigation data with the AEA systems, provides EW track

files for display and recording, and provides the audio interface between the AEA subsystems and the

avionics suite.

2.19.1.3 Mission Data Entry. Mission data (date and flight number) can be manually loaded into the

mission computer for data recorder documentation. Data is entered by performing the following:

1. On the DDI - Press MENU, MUMI, then ID.

2. On the UFCD - Enter Julian Date (DATE).

3. On the UFCD - Enter Flight Information (FLT).

Mission data can be manually loaded into the mission computer through the Memory Unit Mission

Initialization (MUMI) display or automatically loaded into the mission computer through the Data

Storage Set (DSS). The DSS consists of the Memory Unit (MU) and the Memory Unit Mount (MUM)

and provides memory storage for aircraft parameters, maintenance data, and avionics initialization

data. The DSS receives, stores, retrieves, and transmits data with the mission computer.

2.19.1.3.1 Mission Initialization. The MU provides the capability to load the following mission

initialization files: HARM, RADAR, MU ID, TACAN, WYPT/OAP, Combined Interrogator Tran-

sponder (CIT), Sequential Steering (S/S), data link/ID, Overlay Controlled Stores (OCS), and bomb

wind data. The S/S file can have a 15 point sequence consisting of Geographic Reference Points

(GEOREF), GPS waypoints, and almanac data initialization files. Loading is done at aircraft power up

or when MUX communication is lost for more than 1 second and regained. If MUX communication is

not regained, a MU LOAD caution is displayed and an AV MUX error message is displayed on the

MUMI display. Manual loading may be done using the MUMI display.

2.19.1.3.2 Memory Unit Mission Initialization (MUMI) Format. The MUMI format (see figure

2-47) is accessible from the SUPT MENU and with WonW provides a visual indication of mission

initialization files loaded from the MU. If the MU directory indicates that no user files are present, the

MU ID displays NO IDENT. When the MU directory indicates a user file is present, MC1 displays the

option. When the option is selected and the file is being read by MC1, the option is boxed. If the read

is successful, the file is loaded and the option is unboxed. When a file is present and errors have

resulted from reading the file, the following occurs:

1. The MU ID displays NO IDENT.

2. The applicable load error is displayed (HARM, RDR, TCN, WYPT, S/S, OCS, GPS WYPT, GPS

ALM, ALR 67, WIND, DL13, or CIT).

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Figure 2-47. MUMI Display

3. MC 1 sends the appropriate maintenance code to the SDC.

4. If WonW, an MU LOAD caution is displayed on the DDI.

2.19.1.4 Mission Data Erase. There are three methods to manually and two methods to automati-

cally declassify aircraft mission data.

2.19.1.4.1 Manual Erase - SECURE ERASE Button. The SECURE ERASE button is guarded and

located on the right hand forward vertical console. Pressing the SECURE ERASE button erases the

information stored in all systems and the mission card.

2.19.1.4.2 Manual Erase - ACI CRYPTO Switch. Setting the intercommunications amplifier control

CRYPTO switch to the ZERO position sends an erase signal to the MU. This causes the MU to erase

all data stored between predetermined memory locations.

2.19.1.4.3 Erase and Hold Data. The erase controller (EC) within MC 1 provides the capability to

automatically or manually erase, or inhibit erasing, of classified data contained in the MU, SMS, MC1,

and MC2.

NOTE

During an ERASE, the DDI controlled by the MC undergoing the

erase will flash STANDBY, then briefly display a green square, then

flash STANDBY until the erase is complete.

When the EC determines classified mission initialization files have been read from the MU, the EC

classified data management system is activated. When activated, MC1:

1. Displays the HOLD and ERASE options on the MUMI display.

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2. Displays the CDATA advisory.

3. Sends applicable maintenance code(s) to the SDC.

2.19.1.4.4 Manual Erase - MUMI ERASE option. Manual erase is a two pushbutton process and is

initiated by pressing the ERASE pushbutton on the MUMI display. When the ERASE pushbutton is

pressed, the option to proceed with the erasure (ERASE) and the option to cancel the erase (CNX)

replaces the HOLD and ERASE options. Selecting the second ERASE option initiates erasure. While

erase is in progress, ERASE is boxed and erasing proceeds the same as automatic erase. When erasing

is complete, the ERASE pushbutton unboxes. While erase is in progress one of the following is

displayed on the MUMI display:

1. ERASING - erasing of unit is in progress.

2. COMPLETE - erasing of unit is complete.

3. FAILED - unit failed to erase.

NOTE

During an ERASE, the DDI controlled by the MC undergoing the

erase will flash STANDBY, then briefly display a green square, then

flash STANDBY until the erase is complete.

When erase fails, the MC 1 retains the MUMI ERASE and HOLD pushbutton options and displays

the ERASE FAIL caution on the DDI. When erasing is complete, MC1 removes the ERASE, HOLD,

and MC SUSPEND pushbutton options from the MUMI display, removes the CDATA advisory from

the display, and resets the applicable maintenance code(s).

2.19.1.4.5 Automatic Erase. The MU, SMS, MC1, and MC2 automatically erase classified data

when all of the following criteria are met.

1. Airspeed is less than 80 KCAS.

2. Left and right engine THA less than 29°.

3. Transition from WoffW to WonW.

4. Pilot does not select erase inhibit (HOLD) or MC SUSPEND options.

NOTE

Automatic erase can be inhibited by selecting the HOLD pushbutton

option.

5. HOLD boxed with MU displayed prevents automatic erase of the MU.

6. HOLD boxed with ALL displayed prevents automatic erase of all units (MU, armament

computer, MC1, and MC2).

The EC commands the MU and the armament computer, then MC1 and MC2 to erase. The MC

ERASE IN XX SEC countdown timer starts (60 seconds). During the countdown, an MC SUSPEND

pushbutton option is displayed. The MC SUSPEND option may be toggled between boxed (selected)

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and unboxed (deselected). When the timer reaches zero and the MU and armament computer have

finished erasing, the decision to continue erasing the remainder of MC1 and MC2 depends on the MC

SUSPEND option being deselected (unboxed); when deselected, the remaining erase of MC2 and MC1

is completed.

NOTE

During an ERASE, the DDI controlled by the MC undergoing the

erase will flash STANDBY, then briefly display a green square, then

flash STANDBY until the erase is complete.

Automatic erase is also initiated by pilot ejection. The state of the HOLD options is ignored during

pilot ejection.

2.19.1.4.6 Backup Erase Controller. If MC1 fails, backup erase capability is provided by MC2 by

providing an ERASE option on the HSI format. When the ERASE pushbutton is depressed, the option

to proceed with the erasure (ERASE) and the option to cancel the erasure (CNX) are provided.

Selecting the second ERASE option initiates erasure. While ERASE is in progress, the ERASE option

is boxed; however, additional cuing is not provided. In backup mode, pilot ejection is the only

automatic erase provided. The MUMI display for the MC being erased disappears while the erase is in

progress.

2.19.2 Master Modes. There are three master modes of operation: navigation (NAV), air-to-air

(A/A), and air-to-ground (A/G). Controls, displays, and the avionics equipment operation are tailored

as a function of the master mode selected. The navigation master mode is entered automatically when

power is applied to the aircraft, when the air-to-air or air-to-ground modes are deselected, when the

landing gear is lowered, when the SPIN mode activates, or when the aircraft has WonW and the THA

is greater than 27°. The A/A master mode is entered either by pressing the A/A master mode button

alongside the left DDI or by selecting an A/A weapon with the A/A weapon select switch on the control

stick. The A/G master mode is selected by pressing the A/G master mode button. The selection is

performed by the stores management set (SMS), and the SMS identifies the selected master mode to

the mission computer.

2.19.2.1 Steering Information. The sources of steering information available in the NAV master

mode are waypoint, TACAN, instrument landing system, and data link. The data link modes available

in the NAV master mode are vector and automatic carrier landing. TACAN and waypoint steering are

mutually exclusive; selecting one automatically deselects the other. Data link, ILS, and TACAN (or

waypoint) steering can be provided simultaneously. The ACL mode is selectable only in the NAV

master mode and the vector mode is available in all master modes. Steering information is used by the

Automatic Flight Control System to provide coupled steering options.

2.19.3 Cockpit Controls and Displays. The cockpit controls and displays which are used for

navigation operation are on the multipurpose display group.

2.19.4 Multipurpose Display Group. The multipurpose display group consists of the right and left

digital display indicators (DDIs), the multipurpose color display (MPCD), the aft multipurpose color

display (AMPCD), the digital map set (DMS), the head-up display (HUD), the CRS (course) set

switch, the up front control display (UFCD) and the HDG/TK (heading/ground track) set switch. The

multipurpose display group presents navigation, attack, and aircraft attitude displays to the pilot. The

multipurpose display group converts information received from the mission computer system to

symbology for display on the DDIs, the MPCD, the UFCD, and the HUD. The HUD camera records

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EA-18G. FLIGHT MANUAL (2008) - page 3