NATOPS FLIGHT MANUAL NAVY MODEL AV--8B/TAV--8B 161573 AND UP AIRCRAFT (2008)

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2.17.7 Antiskid System

The antiskid system is an electro--hydraulic system that controls hydraulic pressure to the brakes providing full skid

protection above 16 knots. The system also provides partial skid protection from 16 knots down to 8 knots. An

impending skid is detected by measuring wheel deceleration. Wheel speed information is provided by a wheel speed

sensor, located on the right brake unit, and an exciter ring mounted on the wheels. As the exciter ring rotates, the

sensor develops an electrical signal at the frequency of the wheel speed. This signal is routed to a control unit which

develops and transmits a signal to operate the antiskid valve. The antiskid valve, when operating in conjunction with

the control unit, relieves brake pressure to arrest tire skid.

Theantiskid system is selected by theANTISKID switchlocated onthelandinggear/flaps controlpanel (fwdcockpit

only). The switch is labeled TEST, ON, and NWS. Power is not supplied to the antiskid system when NWS is

selected. A SKID light, on the caution light panel, comes on when the antiskid system is OFF or failed, providing

essential 28 volt dc power is available.

2.17.7.1 Skid Caution Light

On AV--8B 161573 through 165312, TAV--8B, the SKID light also comes on when a nosewheel castering failure is

detected. On the TAV--8B, the SKID light comes on in both cockpits. See nosewheel steering this chapter.

The TEST position on the ANTISKID switch, allows the pilot to check the antiskid system. Antiskid is inoperative

when the parking brake is engaged or essential 28 volt dc power is lost.

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 Pitot Static System

The pitot--static system employs dual pitot and static sources, one on each side of the forward fuselage near the leading

edge of the windshield. Each tube contains one pitot source and two static sources. See Figure 2-18.

2.18.1.1 Probe Heat Switch

The probe heat switch on the miscellaneous switch panel on the lower main instrument panel has positions PRB HT

(PROBE HEAT on some aircraft) and AUTO. The switch controls power to the left and right pitot--static probes, the

totaltemperatureprobe,thecaseandprobeheateroftheangleofattackprobe,andtheDECStotaltemperatureprobes.

AUTO - With weight on wheels, power is removed from all heaters except AOA case heater. With aircraft

airborne, all probe heaters receive power.

PRB HT - With weight on wheels, all heaters are energized but the left and right pitot static probes are

energized at reduced power. With aircraft airborne, all probe heaters receive power same as AUTO. The switch

is magnetically held in the PRB HT position. When power is removed, the switch drops into the AUTO

position.

2.18.1.2 Pitot Pressure

Pitotpressurefromtheleftpitot--staticprobeissuppliedtotheairdatacomputerandtheQ--feelsystem. Pitotpressure

from the right pitot--static probe is supplied to the standby airspeed indicator and the ejection seat airspeed/altitude

sensor.

2.18.1.3 Static Pressure

One static source from each pitot--static tube are tied together and the pressure is routed to the air data computer and

the Q--feel system. The other static source from each pitot--static tube is tied together and the pressure is routed to

the pressure--operated standby indicators and the ejection seat airspeed/altitude sensor.

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Figure 2-18. Pitot--Static System

2.18.2 Angle of Attack Probe

The angle of attack probe is an airstream direction sensing unit. The probe is located on the right forward fuselage

below the windshield except on radar aircraft where it is located on the left forward fuselage below and forward of

the windshield. It contains a case heater and a probe heater, operation of which is covered under Probe Heat Switch

in Instruments procedures. The standby angle of attack indicator and the air data computer utilize signals from the

angle of attack probe.

2.18.3 Standby Angle of Attack Indicator

The standby angle of attack (AOA) indicator (see cockpit, foldout section) is on the main instrument panel. The

indicator is calibrated from --5° to +25° An adjustable reference index two degrees wide is centered on the 10° mark

to indicate optimum speed approach angle of attack. When electrical power is interrupted, the word OFF appears in

a window in the face of the indicator.

2.18.4 Turn and Slip Indicator

The turn and slip indicator contains a scale, turn pointer, power warning flag and inclinometer ball. A 2--minute turn

is indicated with the needle over the index to the left and right of center. A 4--minute turn is indicated with the needle

half way between the center and the right or left index. The gyro is driven by an inverter, which is powered from the

emergency 28 volt dc bus. An OFF flag is provided to indicate loss of power.

2.18.5 Clock

Astandard8dayclockisinstalledinthecockpitandontheTAV--8BintherearcockpitpedestaladjacenttotheBUNO

PLACARD.

2.18.6 Stopwatch

A mechanical stopwatch is located to the left of the glareshield near the canopy rail (not in rear cockpit of TAV--8B).

The stopwatch contains one pushbutton for winding, starting, and stopping and one pushbutton for resetting.

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2.18.7 Standby Magnetic Compass

A conventional aircraft magnetic compass is installed to the left of the main instrument panel (not installed in the

rear cockpit). The standby magnetic compass is installed on the left archway.

2.18.8 Standby Vertical Velocity Indicator

The standby vertical velocity indicator displays rate of ascent or descent on a scale from 0 to 6,000 feet per minute.

2.18.9 Standby Attitude Indicator

Thestandbyattitudeindicatorisaself--containedelectricallydrivengyro--horizontypeinstrument.Thegyroisdriven

by an inverter which is powered by the emergency 28 volt dc bus. An OFF flag appears whenever power is lost or

the unit is caged. The gyro cages to 0° pitch and roll regardless of aircraft attitude. 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 92° climb and 78° dive. The caging knob on the lower right hand corner, besides being pulled for

caging, is used to adjust the pitch of the miniature aircraft. A pitch--trim scale measures displacement of the miniature

aircraft. Pulling the caging knob and rotating fully clockwise to a detent locks the inner gimbal of the gyro. This

position is for storage and transport, and should never be used during flight. A minimum of 9 minutes of reliable

attitude information (error less than 6°) is available after power loss, even though the OFF flag is in view.

2.18.10 Standby Altimeter

The standby altimeter displays altitude from --1,000 feet to 50,000 feet. The 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

barometer setting. This setting is also used by the air data computer. An electrical altimeter vibrator is provided to

insure smooth travel of the internal mechanism.

Note

The standby altimeter may indicate in excess of 400 feet low at high

airspeeds.

2.18.11 Standby Airspeed Indicator

The standby airspeed indicator displays airspeed from 20 to 600 knots. The indicator contains two pointers and a

singlescalegraduatedfrom1 through10. Thepointers appearonly oneat atime. Atlow airspeedsthescalerepresents

0 to 100 knots and the thin pointer indicates the airspeed. At higher airspeeds the scale represents 100 to 1,000 knots

and the thick pointer indicates the airspeed. However, the thick pointer will not proceed beyond the 600 knot

indication.

2.18.12 Horizontal Situation Indicator

On TAV--8B and Day Attack aircraft the horizontal situation indicator (HSI) (see Figure 2-19) is on the main

instrument panel. The HSI provides horizontal or plan view of the aircraft with respect to the navigation situation.

The knobs, pointers, windows, and flags which are on the HSI are described in the following paragraphs.

2.18.12.1 Aircraft Symbol

The aircraft symbol in the center of the HSI is the aircraft superimposed on a compass rose.

2.18.12.2 Aircraft Heading

The aircraft magnetic heading is read under the lubber line.

2.18.12.3 Heading Marker

The heading marker is manually set to the desired heading.

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Figure 2-19. Horizontal Situation Indicator

2.18.12.4 Course Arrow

The course arrow and the course selector window are set manually with the course set knob to the desired tacan course.

2.18.12.5 Course Deviation Indicator

Any displacement of the aircraft from the selected course causes the course deviation bar to move to the right or left

of the course arrow.

2.18.12.6 Bearing Pointer

The bearing pointer displays the bearing to a selected tacan station. The bearing tail indicates reciprocal course.

2.18.12.7 Range Indicator

Distance is displayed in nautical miles and has a range of 00.0 to 399. A shutter will cover the window if distance

information is invalid.

2.18.12.8 To--From Indicator

The to--from indicator indicates whether the course selected, if intercepted and flown, takes the aircraft to or from

the selected tacan station.

2.18.12.9 Deviation Warning Flag

The flag is in view when the bearing data is invalid.

2.18.12.10 Compass Flag

The flag is in view when the compass data is invalid.

2.18.12.11 Course Set Knob

The knob, labeled CRS, is used to set a predetermined course to be steered to a tacan station, all weather landing

system (AWLS) station, or steer--to--point (waypoint, markpoint, or targetpoint).

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CAUTION

The course line displayed on the EHSI/EHSD and the course in the Course

Selector Window of the HSI is not necessarily the same. The bearing

displayed in the Course Line Data block in the lower right corner of the

EHSI/EHSD and the course line displayed on the EHSI/EHSD should be

used instead of the bearing in the Course Selector Window of the HSI.

Note

With OMNI 7.1 and C1+, a predetermined course cannot be used to steer

to an AWLS station or a targetpoint.

2.18.12.12 Heading Set Knob

The heading set knob, labeled HDG, is used to set the heading marker to a desired heading. If the knob is pulled out,

the compass is disabled and the compass flag comes into view. Ensure the knob is pushed in fully.

2.18.13 Sideslip Vane

A sideslip vane is mounted externally forward of the windshield for use during slow or hover flight. It always points

intotherelativewind.Therearcockpitsideslipvaneismountedonthecentertopofwindshieldarchframe,externally.

2.18.14 Radar Altimeter

The radar altimeter indicates surface clearance directly under the aircraft from 0 to 5,000 feet up to pitch and roll

attitude of ±45°. Operation is based on precise measurement of time required for an electro--magnetic energy pulse

to travel from the aircraft to the ground terrain and return. Audio and visual warnings are activated when the aircraft

is at or below a selected low altitude limit. When target of opportunity (TOO) function is engaged on the upfront

control, the radar altimeter is turned on momentarily and the radar altitude is used by the mission computer to

determine target elevation. The radar altimeter also has an update function available to adjust barometric altimeter

error for mission computer navigation and air--to--ground calculations.

2.18.14.1 Controls and Indicators

The controls and indicators for operation of the radar altimeter are on the HUD, warning/threat lights panel, upfront

control, option display unit and digital display indicator (DDI).

2.18.14.2 Altitude Switch

The ALT switch, on the HUD control panel, has positions of barometric (BARO) and radar (RDR). When the switch

is set to RDR, radar altimeter altitude preceded by an R is displayed in a box in the upper right hand part of the HUD

display. In BARO, barometric altitude is displayed without theR in the box. If radaraltimeter altitudeis invalid with

switch in RDR, the R disappears and a flashing B appears to the right of the box indicating barometric altitude is being

displayed. The flashing B remains until either radar altitude becomes valid again or the altitude switch is placed to

BARO. In the backup mode, barometric altitude replaces radar altitude but will not flash. This switch is

non--functioning in the rear cockpit. With H4.0, the box around the altitude is no longer displayed.

2.18.14.3 Low Altitude Warning Light

The Low Altitude Warning (LAW) light is located on the warning/threat light panel on the upper right of the instrument

panel. If the system is operational and the aircraft descends below the preset low altitude threshold (0 to 5,000 feet above

surface level) the LAW light will come on. The mission computer requires at least three radar altimeter returns above the

threshold altitude to subsequently trigger the LAW during descent. Due to aircraft maneuvers during ascent or descent,

variances in transmitter and receiver capabilities, or mountainous terrain, LAW thresholds above 4,500 feet may be

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unreliable. The LAW light will stay on until the aircraft ascends above the low altitude threshold, the pilot changes

the low altitude warning threshold, the radar altimeter is turned off, or the MASTER CAUTION (or MASTER

WARNING light button on Radar or Night Attack aircraft) light button is pressed. The MASTER CAUTION shutoff

function resets when the aircraft ascends above the threshold altitude. With loss of the mission computer, the LAW

light automatically operates at 200 feet above ground level. To eliminate false LAW indications with speed brake

extended, radar altitude is invalid with gear handle up and radar altitude less than 20 feet. On TAV--8B 163856 and

up, AV--8B 163519 and up, an ALTITUDE, ALTITUDE voice warning is provided in conjunction with the LAW

light.

CAUTION

LAW thresholds above 4,500 feet greatly reduce the probability of the

LAW being triggered during descent.

2.18.14.4 Upfront Control

The pushbuttons and indicators on this control are used for radar altimeter operation and display, to select altitudes

for LAW light operation, and also to enable the BOMB option.

2.18.14.4.1 Altitude Function Selector Pushbutton

Pressing the altitude (ALT) function selector pushbutton enables the status window on the scratch pad to display ON

if radar altimeter is turned on, and enables display of the LAW threshold altitude on the scratch pad when set in by

the keyboard. The display on the scratch pad indicates the altitude that the LAW light will illuminate and the LAW

warning tone or ALTITUDE, ALTITUDE voice warning will activate. Pressing the ALT pushbutton also enables

display of BOMB on the option number 1 display window on the option display unit.

2.18.14.4.2 On/Off Selector Pushbutton

Pressing this pushbutton turns the radar altimeter system on after first pressing the ALT function selector pushbutton.

This causes ON to appear on the scratchpad of the upfront control. To turn the system off the ON/OFF selector

pushbutton is pressed after first enabling the altitude function with the ALT function selector pushbutton.

2.18.14.4.3 Scratchpad and Keyboard

With the ALT function enabled, the LAW altitude is displayed in the scratchpad window. A new LAW altitude is

inserted by typing out the new altitude and then pressing the ENT pushbutton on the keyboard.

2.18.14.4.4 EMCON Pushbutton

Pressing the emission control (EMCON) pushbutton enables the EMCON functionality and EMCN is displayed in

the Options Display Unit (ODU) number 1 display window. Pressing the EMCON pushbutton again disables the

EMCON functionality and EMCN is removed from the ODU. With H4.0 EMCON can also be enabled or disabled

by pressing down and holding the Sensor Select switch on the control stick for greater than 0.8 seconds. When

EMCON is enabled, the RADAR, RADALT, IFF, and TACAN are inhibited. With H4.0, EMCON is overridden by

the RADALT if the pilot connects a TOO or WOF.

2.18.14.4.4.1 Head--Up Display

EMCONshallbedisplayedin themiddleoftheHUDas anindication thatEMCON isON. Itwill displayin allmaster

modes and all of the reject levels.

2.18.14.5 Option Display Unit

The pushbuttons and displays on this panel which affect operation of the radar altimeter are option select pushbutton

number 1 and option number 1 display window. They are used to enable the BOMB function.

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2.18.14.5.1 Option Number 1 Pushbutton and Display Window

With the ALT function selector switch pressed, BOMB is displayed in the option number 1 display window. With

the radar altimeter operating, pressing the option number 1 pushbutton causes a colon to be displayed to the left of

BOMB on the option number 1 display window. The colon indicates that radar altitude is being used by the mission

computer for ballistic computations. The last selected LAW altitude remains in the scratchpad window after the

BOMB option is selected. With weight on wheels the system will initialize to BOMB option enabled so that radar

altitude will be used for ballistic computations unless radar altitude becomes invalid. Pressing the option number 1

pushbutton again removes the colon from BOMB on the option number 1 display window and enables GPS altitude

for ballistic computations if GPS is cued. Otherwise, barometric altitude is used for ballistic computations.

2.18.14.6 DDI Display

Radar altitude is displayed on the DDI when dual mode tracker air--to--ground (A/G) video, forward looking infrared

(FLIR), air--to--air (A/A) radar or air--to--surface (A/S) radar program video is displayed, RDR is selected on the HUD

control panel, and the radar altimeter is within operational parameters.

2.18.14.6.1 Radar Altimeter BIT Checks

To perform a radar altimeter BIT check, press the BIT pushbutton on the DDI menu display to initiate a BIT display.

Press the CNI pushbutton and the word TEST appears next to RALT. After 5.5 seconds the word TEST disappears.

Check LAW light off, and LAW and MASTER CAUTION tones in headset for one second. If a number 1 appears

next to RALT the radar altimeter has failed the BIT check. If the space next to RALT remains blank the radar altimeter

has checked good.

2.18.15 Upfront Control

The pushbuttons and indicators on this control are used for entering (ENT) or clearing (CLR) data (i.e. 0--9, .,--) for

theselectedfunction.(InTAV--8Baircraft),therearetwoUFCSs,oneintheforwardcockpitandoneintheaftcockpit.

If only one UFCS (forward or aft) is being used to enter data, the system shall accept data entry from that UFCS and

cause that data to be displayed on both the entry UFCS and the non--entry UFCS as the entry is taking place. In the

event of entries occurring on both UFCSs simultaneously, the system shall cause each UFCS to display its own entry

as long as neither entry is completed. As soon as the first entry is completed, the data shall be displayed on both

UFCSs.

Note

D If the first key is to be an alpha instead of a numeric, the first pilot to enter

an alpha key will get the result as an alpha. The second pilot will get a

number, which will result in an invalid entry.

D If the data format contains only one decimal point (a Comm frequency is

an example), only the first pilot to enter a decimal point will get the decimal

point.

2.18.16 Digital Display Indicator and/or Multipurpose Color Display

The digital display indicator (DDI), on the left main instrument panel on Day attack (TAV--8B and AV--8B Day

Attack) aircraft or the multipurpose color display (MPCD), on either side of the main instrument panel (Radar and

Night Attack aircraft), are the primary aircraft head down displays. They consist of a 5 by 5--inch CRT display

surrounded by 20 multi--function pushbutton switches. DDI/MPCD mode selection is accomplished either

automatically, as determined by the mission computer, or manually, as selected by the pilot on the DDI/MPCD or

by the hands on throttle and stick (HOTAS). The display computer converts information received from the mission

computer to symbology for display on the DDI/MPCD. Some of the displays options are: MENU, stores status

(STRS), head--up display (HUD), engine parameters (ENG), electronic horizontal situation indicator/display

(EHSI/EHSD), dual mode tracker (DMT) (trainer, day attack, and night attack only), built--in--test (BIT),

VSTOL--REST (VRST), and electronic countermeasures/warfare (ECM/EW). Additionally, in the Radar and Night

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Attack Aircraft (and Trainer aircraft with H4.0), four other pushbuttons are available, FLIR-- when boxed the

NAV/FLIR image is displayed on the MPCD; EMER-- when boxed the emergency checklist menu page is displayed;

CARD-- when boxed the pre--programmed kneeboard cards are displayed; and CAS-- when boxed the close air

support page is displayed. With H4.0 four other pushbuttons are available; CONF-- when boxed the software

configuration page is displayed; TPOD-- when boxed the Litening Pod video page is displayed; SDAT-- when boxed

the system data page is displayed; and COMM-- when boxed the COMM data page is displayed, see Figure 2-20,

Menu Display. The display options are selected by pressing the MENU pushbutton (center bottom pushbutton). The

word MENU is displayed above the center bottom pushbutton for all displays except the MENU display itself and

multipurposedisplay (MPD)test pattern. Useofthevarious displays aredescribed in otherparts ofthe manual where

the affected system(s) is covered.

Note

The display computer in TAV--8B aircraft with OMNI 7.1 does not display

the same page on the front and rear cockpit DDIs at the same time (other

than the MENU page). TAV--8B aircraft with H4.0 display the same page

on both the front and rear cockpit DDIs.

2.18.16.1 DDI Switches and Controls

A description of the various switches and controls are discussed in the following paragraph.

2.18.16.1.1 Brightness Selector Knob

This rotary knob is at the top of the DDI. Placing the knob to OFF prevents the indicator from operating. Placing the

knob to NIGHT provides a lower brightness control range and no automatic contrast control. The knob in the AUTO

position allows automatic brightness control circuits to compensate display brightness for changes in ambient

lighting. Turning the knob to DAY provides higher brightness control range with no automatic contrast control.

2.18.16.1.2 Brightness Control

This knob varies the intensity of the presentation.

2.18.16.1.3 Contrast Control

This knob varies the contrast between symbology and the dark background on any level of brightness.

2.18.16.1.4 Pushbuttons

There are 20 pushbuttons on the DDI which are used to select the function and the mode for proper indicator display.

2.18.16.2 MPCD Switches and Controls (Before ECP 306)

There are four two--position rocker switches on the MPCD for display control. A description of the various switches

on the MPCD follows.

2.18.16.2.1 DAY/AUT Switch

The day position of this switch turns the MPCD on and places it in the normal day operating mode. The AUT

(automatic) position turns the MPCD on and selects the auto mode which automatically changes brightness levels

to maintain a fixed contrast ratio based on outside ambience.

2.18.16.2.2 OFF/NGT Switch

The NGT (night) position of the OFF/NGT switch turns the MPCD on and selects the night operating mode. The

OFF position turns the MPCD off.

2.18.16.2.3 BRT Switch

The BRT rocker adjusts the brightness of the display. The switch has a position feedback reference number indication

on the display, which helps match one display to the other, providing consistent image brightness, when the display

alternate toggle (DAT) function is used. The number is automatically removed after a few seconds. The BRT switch

is inoperative with an EHSD display.

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2.18.16.2.4 CONT Switch

The CONT rocker switch adjusts the contrast level of the display. This switch also has the position feedback reference

number indication. Independent contrast settings for color and monochrome displays are stored in the MPCD.

2.18.16.2.5 Pushbuttons

There are twenty pushbuttons on the MPCD which are used to select the function and the mode for proper indicator

display.

2.18.16.3 Digital Display Indicator (TAV--8B)

The displays on the DDIs in both cockpits are always the same. Both DDIs may be switched on from either cockpit,

however, the brightness selector knob on each DDI must be placed to OFF to prevent both DDIs from operating. The

last selection made in either cockpit with the perimeter pushbuttons determines the function selected for both. With

the front cockpit DDI power switch in NIGHT, AUTO, or DAY, changing the position ofthe rearcockpit DDIpower

switch causes a momentary blooming effect on the rear cockpit display.

2.18.17 Multipurpose Color Display (After ECP 306)

The MPCD (Figure 2-21) is an NVG compatible digital display. Four momentary two position rocker switches and

a rotary knob, located on the front of the MPCD, permit control of MPCD off/brightness, night/day viewing modes,

symbology, gain, and contrast.

2.18.17.1 OFF/BRT Control

This rotary switch is located in the upper center of the MPCD and is used to turn the MPCD off (OFF position

selected) or to select the brightness level.

2.18.17.2 NGT/DAY Brightness Selector

This rocker switch is located in the upper left corner of the MPCD and is used to select the lower brightness control

(night) range and (NGT position selected) or to select the higher brightness control (day) range (DAY position

selected). When NGT is selected, the display is NVG compatible. In either NGT or DAY, the display may be manually

adjusted with the CONT, GAIN and SYM controls.

2.18.17.3 SYM Control

This rocker switch is located in the upper right corner of the MPCD. Momentary actuations of the lower half of the

switch incrementally narrows the stroke symbology, making it sharper and dimmer. Momentary actuations of the

upperhalfincrementally widens thestrokesymbology,making itbrighterandless sharp.Iftheswitch isheld ineither

position, the symbology is continuously adjusted to the upper or lower limits. The current level of the SYM control

is displayed near the CONT switch. The range is from 0 to 15. Examples of stroke symbology are the MENU format

or the pushbutton legends on the FLIR format.

2.18.17.4 GAIN Control

This rocker switch is located in the lower left corner of the MPCD. Momentary actuations of the upper half of the

switch incrementally increases the black level of the sensor video and raster symbology. Momentary actuations of

the lower half incrementally decreases the black level of the sensor video. If the switch is held in either position, the

gain is continuously adjusted to the upper or lower limits. The current level of the GAIN control is displayed next

to the GAIN switch when it is depressed. The range is from 0 to 15. The GAIN control is active on the EHSD page

when the Map is selected even though the GAIN setting is displayed as an X when depressed. Examples of raster

symbology aretheEHSD format ortheairspeed and altitude on the FLIR format. Examples of sensorvideo areFLIR

and Map.

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Figure 2-21. Multipurpose Color Display (After ECP 306)

2.18.17.5 CONT Control

This rocker switch is located in the lower right corner of the MPCD. Momentary actuations of the upper half of the

switch incrementally increase the contrast of the sensor video and raster symbology. Momentary actuations of the

lowerhalfincrementally decrease thecontrast ofthe sensorvideo and raster symbology. If theswitch is held in either

position, the contrast is continuously adjusted to the upper or lower limits. The current level of the CONT control

is displayed next to the CONT switch when it is depressed. The range is from 0 to 15. Examples of raster symbology

are the EHSD format or the airspeed and altitude on the FLIR format. Examples of sensor video are FLIR and Map.

2.18.17.6 MPCD Control Setting Retention

There are six different retained settings for the GAIN, CONT and SYM controls. They are:

1. Day Mode selected with all stroke display (examples are MENU and CAS formats).

2. Day Mode selected with monochrome video (examples are FLIR, TPOD, EHSD/EW when MAP is not

selected).

3. Day Mode selected with Map selected (EHSD/EW with Map selected).

4. Night Mode selected with all stroke display (examples are MENU and CAS formats).

5. Night Mode selected with monochrome video (examples are FLIR, TPOD, EHSD/EW when MAP is not

selected).

6. Night Mode selected with Map selected (EHSD/EW with Map selected).

The six settings are remembered and recalled as the operator cycles through different format types.

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The MPCD retains the six settings of the GAIN, CONT and SYM controls plus the position of the NGT/DAY control

when the MPCD is turned off or power is interrupted. Although, if power to the mission systems computer (MSC)

is interrupted, the MSC may command a different type of format to be displayed on the MPCD than what was

previously displayed, which could cause the MPCD to use a different set of retained values. The BRT setting is

dependent on the position of the knob.

2.18.17.7 MPCD Adjustment

2.18.17.7.1 All Stroke (Examples are MENU, CAS or No Sensor Video on MAP)

1. Select NGT or DAY as appropriate.

2. Rotate the BRT knob to a comfortable position.

3. Adjust the SYM for the desired thickness of the symbology.

2.18.17.7.2 Monochrome Video (Examples are FLIR or EHSD without MAP Selected)

1. Select NGT or DAY as appropriate.

2. Rotate the BRT knob to a comfortable position.

3. Adjust the GAIN until the lowest level shade of gray is just visible.

4. Back the GAIN down until the lowest shade of gray just disappears. Do not touch the GAIN switch again.

5. Adjust the CONT to the desired level. Raster symbology should be sameintensity as stroke symbology. FLIR

format contains raster airspeed and altitude and stroke pushbutton legends.

6. Adjust the SYM for the desired thickness of the symbology.

7. Further adjustments primarily use the BRT knob.

2.18.17.7.3 MAP Selected

Note

The map display can be made unreadable if this is adjusted incorrectly.

1. Select NGT or DAY as appropriate.

2. Rotate the BRT knob to a comfortable position.

3. Adjust the GAIN all the way down. The GAIN switch is active even though an X is being displayed instead

of the current level of the GAIN switch.

4. Adjust the GAIN up until the map colors look right. Do not touch the GAIN switch again.

5. Adjust the CONT to the desired level. Raster symbology should be same intensity as stroke symbology. The

pushbuttonlegendsareinraster.RWRsymbolsontheEWpageareinstroke.IfnoRWRsymbols areavailable,

the level of the SYM switch is in stroke.

6. Adjust the SYM for the desired thickness of the symbology. On the EW format, verify that the stroke

symbology (RWR symbols or level of the SYM control) can be easily seen on top of the map.

7. Further adjustments primarily use the BRT knob.

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2.18.18 Head--Up Display

The head--up display (HUD) is on the top of the main instrument panel. The HUD is the primary attitude indicator,

weapon status, and weapon delivery display for the aircraft under all selected conditions. Due to the way the INS

informationistranslatedforpresentationintheHUD,theVSTOLMasterModeprovidesamorereliableIMCattitude

presentation than the other master modes (NAV, AA, and AG). If INS velocity information begins to degrade, the

othermodes may present attitudeinformation that is inaccurate. In IMC conditions thisinaccuratepresentationcould

result in an unrecognized spatial disorientation. This is a particular concern when operating the ASN--130 in a coupled

mode with the GPS. Therefore, the VSTOL Master Mode should be the presentation of choice when flying in IMC

conditions. Use of V/STOL helps to minimize attitude presentation errors when INS velocities are degrading and

should provide a relatively stable attitude reference up to the point of INU failure. The HUD receives attack,

navigation, situation, and steering control information and projects symbology on the combining glass for head--up

viewing. Symbology is unique to the master mode selected. HUD symbology can also be presented head--down on

the DDI/MPCDs by depressing the HUD pushbutton on the DDI/MPCD MENU display. On Radar and Night Attack

aircraft, the HUD can display FLIR video in all master modes provided the HUD symbology brightness selector

switch is in the NIGHT position.

Due partially to new weapons symbology incorporated with H4.0, there are situations where the display computer

tries to write more HUD symbology than it has the time to write. This results in flickering HUD symbology. To

minimizetheoccurrences offlickering HUD symbology, somesymbology is written at alowerintensityso itappears

slightly dimmerin theHUD and somesymbology has been removed. Referto NTRP 3--22.2--AV8B for adescription

of the changes that were made to the HUD symbology in the aircraft NAV and VSTOL master modes. See the

A1--AV8BB--TAC--000 for changes that affect the A/G and A/A master modes.

The HUD displays collimated symbology projected into the pilot’s forward field--of--view (FOV). The HUD has a

22° total field--of--view (TFOV) and an approximately 14° by 14° (16° by 20° on Radar and Night Attack aircraft)

instantaneous field--of--view (IFOV). The optical center of the IFOV is located --6° below the horizontal vision line

from the design eye position. The lower portion of the TFOV coincides with the pilots --17° over the nose vision line.

The HUD is electrically interfaced with the upfront control and the HUD camera. The controls for the HUD are below

the upfront control and are described in the following paragraphs.

2.18.18.1 Head--Up Display (TAV--8B)

The information displayed on both HUDs is identical in both cockpits. On the HUD control panel in the rear cockpit,

the switches for reject level and radar altitude display selection are inoperative. These functions can only be selected

from the front cockpit.

2.18.18.2 HUD Symbology Reject Switch

This three--position toggle switch has positions of NORM, REJ 1, and REJ 2. With the switch placed to NORM, the

normal amount of symbology is provided for all HUD displays. Placing the switch to REJ 1 or REJ 2 changes the

HUD symbology in the different modes. Reject level 1 is automatically selected when altitude alert cue is enabled.

The following paragraphs define the reject levels for the four master modes:

2.18.18.2.1 NAV Mode

Reject level 1 removes AOA legend, FPM legend, airspeed box, altitude box, heading box, and replaces large heading

numerics with nominal sized heading. Also adds AOA and feet per minute (FPM) analog scales. Reject level 2

removes the AOA, FPM, Mach, normal g’s, and ground speed legends, and also removes the heading numerics,

heading scale, altitude box, airspeed box, and heading box. With H4.0, the altitude box, airspeed box, and heading

box are not displayed in any reject level.

2.18.18.2.2 VSTOL Mode

Reject level 1 removes AOA legend, FPM legend, airspeed box, altitude box, heading box, power margin indicator

(digital rpm and JPT indications are displayed) and replaces large heading numerics with nominal sized heading.

Reject level 2 removes AOA legend, FPM legend, airspeed box, altitude box, heading box, vertical flight path

symbol, and nozzle, flaps, rpm, and JPT waterflow legends (power margin indicator if displayed). Also replaces large

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heading numerics with nominal sized heading. With H4.0, the altitude box, airspeed box, and heading box are not

displayed in any reject level.

2.18.18.2.3 A/G Mode

Reject level 1 removes AOA legend, airspeed box, altitude box, heading box, and replaces large heading numerics

with nominal sized heading. Also adds AOA analog scale. Reject level 2 removes AOA legend, airspeed box, altitude

box, heading box, heading numerics, and heading scale. With H4.0, there is no difference between A/G reject level

normal and reject level 1. With H4.0, the altitude box, airspeed box, and heading box are not displayed in any reject

level.

2.18.18.2.4 A/A Mode

Reject level 1 removes AOA legend, airspeed box, altitude box, heading box, and replaces large heading numerics

with nominal sized heading. Also adds AOA analog scale. Reject level 2 removes airspeed box, altitude box, heading

box, heading numerics, heading scale, and pitch ladder. Also removes AOA, Mach, and normal g’s legends. With

H4.0, the altitude box, airspeed box, and heading box are not displayed in any reject level.

2.18.18.3 HUD Symbology Brightness Control

This knob is used to turn on the HUD and then varies the symbology display intensity.

2.18.18.4 HUD Symbology Brightness Selector Switch

This is a three--position toggle switch with positions of DAY, AUTO, and NIGHT. Placing the switch to DAY

provides maximum symbol brightness in conjunction with HUD symbology brightness control. Placing the switch

to AUTO allows automatic control of the contrast by the automatic brightness control circuit. With the switch to

NIGHT, a reduced symbol brightness is provided in conjunction with the HUD symbology brightness control. The

NIGHT position must be selected to have FLIR video on the HUD.

2.18.18.5 Video Brightness Control (Radar and Night Attack Aircraft)

This control is a rotary knob used to adjust the brightness of theHUD rastervideo. It is used to set the black reference

level for FLIR video. Clockwise rotation increases brightness. The brightness control has a pushbutton feature which

isusedforthedisplayalternatetoggle(DAT)function.SelectingtheDATfunction swapsthedisplayson theMPCDs.

2.18.18.6 Video Contrast Control (Radar and Night Attack Aircraft)

This control is a rotary knob that adjusts the contrast of the HUD raster video. Clockwise rotation increases contrast.

2.18.18.7 Standby Reticle Brightness Control (Day Attack Aircraft)

This control turns on the standby reticle and adjusts the symbol’s brightness.

2.18.18.8 Standby Depression Control (Day Attack Aircraft)

This control selects reticle depression angles over the range from 0 to minus 240 milliradians (in 20 milliradian

increments) with respect to the aircraft waterline.

2.18.18.9 Altitude Switch

This is a two position toggle switch with positions of BARO and RDR. This switch is used to select either radar

altitude (RDR) or barometric altitude (BARO) for display on the HUD. Refer to paragraph 2.18.14.2, Altitude

Switch.

2.18.18.10 HUD Camera

The HUD video camera is mounted on the right side of the HUD. It is focused at infinity and records the scene as

viewed by the HUD prism assembly through the HUD combiner assembly. The field of view is 16° vertically and

21° horizontally. This includes all HUD symbology in the camera’s field of view. An exposure control automatically

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adjusts for changing light levels. The HUD camera is part of the Video Recording System (VRS). Two switches on

the miscellaneous switch panel on the center pedestal control the VRS. In most aircraft, the MPCD/HUD switch

determines which displays are recorded. In HUD, only the HUD camera video is recorded. In MPCD, video of the

DDI/MPCD is recorded. The DDI/MPCD symbology is not recorded. The DDI/HUD switches on the miscellaneous

switch panel on the center pedestal controls which displays are recorded. In HUD, only the HUD camera video is

recorded. In DDI, video of the DDI is recorded. The DDI symbology is not recorded. The AUTO/RUN switch

controls the video recorder mode of operation. In AUTO, the mission computer turns on the video recorder when A/A

or A/G master mode is selected. In RUN, the video recorder is turned on for continuous recording, regardless of

master mode. A VRS button is installed on the miscellaneous switch panel on the center pedestal in both trainer

aircraft cockpits. The VRS is controlled by the last button change from either cockpit. Aircrew feedback is provided

by the VRS lights integral to the VRS button. The AUTO light indicates the mission computer will turn on the video

recorder when A/A or A/G master mode is selected. The RUN light indicates the video recorder is turned on for

continuous recording, regardless of master mode.

2.18.18.11 Radar Switch

The RADAR switch is on the miscellaneous switch panel, directly above the INS controls. The switch has four

positions:

OFF -- Removes all radar set power.

STBY -- All radar functions are operational except the radar transmitter and RF transmission circuits. Allows

radar set to warmup before application of high voltage.

OPR -- The radar is placed in the normal mode of operation. Commands radar to full operation if all safety

interlocks have been satisfied and initial warmup and ORT (operational readiness test) is complete.

EMER -- With weight--off--wheels, bypasses temperature and pressure interlocks and allows full radar

operation. The radar is prevented from shutting down due to an overheat condition. If the radar overheats, it

automatically shuts down 30 seconds after the overheat (OVHT) indication appears unless EMER is selected.

Selection of EMER with weight--on--wheels turns the radar off.

CAUTION

Taxi the aircraft with the RADAR switch in STBY or OPR to prevent

damage to the antenna.

2.19

MISSION COMPUTER

The mission computer is a standard general purpose stored program real--time computer with core memory. ECP 285

replaces the mission computer (MC) with a Mission Systems computer (MSC). The MSC is a higher speed computer

with multiple expansion slots. The MSC provides essentially the same functionality as the MC.

2.19.1 Mission Computer Switch

The mission computer switch with positions labeled OVRD, AUTO and OFF is on the miscellaneous switch panel

on thepedestal. Placing theswitch to OVRD (override)inhibits thebackup modeof operation. If in the backup mode

at the time of placing the switch to OVRD, this allows power to be reapplied to the mission computer, enabling it

to reassume control of the MUX BUS and operatenormally. Placing the switch to OFF position turns off themission

computer and enables the display computer for backup mode of operation. When the switch is in AUTO position the

mission computer will be normally utilized but the system automatically reverts to the display computer in case of

MC failure.

2.19.2 DP Switch

This switch controls selection ofmutually redundant display channels in thedisplay computer. Thedisplaycomputer

drives the HUD and DDI, providing display redundancy for attack, navigation, and approach to landing. Selecting

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the PRIM position on the DP switch selects a primary display channel for operation. The ALTER position selects

an alternate display channel for operation. The AUTO position, which is the preferred switch position, randomly

selects the operational channel for operation and provides automatic reselection if there is a display computer channel

failure. If the switch is in PRIM or ALTER and a power interruption occurs, the display may go blank. To regain the

display, cycle the switch from PRIM to ALTER and back to PRIM, or vice versa.

2.20

VREST COMPUTER

To determine the operational capability of the aircraft, the mission computer performs vertical takeoff, vertical

landing, range endurance, speed and time calculations. These calculations are presented on the V/STOL REST

displays. SeeFigure2-22. Thesedisplays areavailableonly inNAV orV/STOL mastermode. Toenablethedisplays,

select VRST on the menu display.

The V/STOL--REST basic display appears with the last entered values for the basic aircraft weight (BAW), water

weight (H2O) and basic drag index (BDI). Also the UFC and ODU are enabled for data entry. To ensure accuracy

thedisplayed values forBAW, H2O, and BDImust beverified and new values entered if required. The H2O quantity,

as indicated on the EDP, is displayed on the V/STOL--REST basic display in aircraft with H4.0 and cannot be altered.

2.20.1 VREST Displays

There are five VREST displays: vertical landing (VL), vertical takeoff (VTO), short takeoff (STO), cruise (CRUS),

and bingo (BNGO). To enable the desired display select the appropriate pushbutton VL, VTO, STO, CRUS or

BNGO. Selection is indicated by the box around the legend.

2.20.1.1 Vertical Takeoff and Landing Display

The vertical takeoff and vertical landing displays are identical in format. The displays show the maximum weight

of fuel and water (F+W) aboard the aircraft at which the vertical takeoff or vertical landing can be performed. This

data is computed for both WET (water injected in engine) and DRY operation. If the outside air temperature is below

--5 °C/23 °F the WET data is not displayed.

The data displayed at the bottom of the display; outside air temperature Celsius/Fahrenheit (OATC or OATF),

altimeter barometric pressure setting (ALTM), field elevation (FELV), and gross weight (GWT) is normally system

generated and used in calculating maximum F+W. If any of these parameters are not valid then that parameter is not

displayed. If the non--valid parameter is essential to calculations of maximum F+W, then the calculation is not

performed and no data is displayed for maximum F+W.

Some data affecting the calculations may be entered by the pilot using the options on the ODU. The following options

are available: GWT, OATC or OATF, FELV, and engine parameters (ENG). Selecting ENG enables the relative jet

pipetemperature(RJPT),jetpipetemperaturelimit(JPTL)andrelativehover(RHOV)options.Ifdatais pilotentered

(OATC or OATF, GWT, FELV) and asterisk (*) is displayed to the left of the applicable legend. Pilot entered data

overrides system generated data.

The maximum F+W is calculated as follows: first the hover weight is calculated. This is done by calculating the

maximum rpm limited by ambient temperature, then calculating the maximum rpm limited by jet pipe temperature

limiter settings. Since either of these parameters may limit the maximum rpm, the smaller of the two is used in the

equations to calculate the hover weight. The hover weight is then adjusted for either the takeoff or landing

calculations. For a vertical takeoff 97 percent of the hover weight is used. For a vertical landing approximately 95

percent of the hover weight is used. The adjusted hover weight is used along with the aircraft gross weight, fuel

weight, and water weight to calculate the maximum fuel plus water weight at which a vertical operation can be

executed.

2.20.1.2 Short Takeoff Display

The short takeoff display shows the nozzle rotation airspeed (NRAS), nozzle setting in degrees (NOZ), minimum

ground roll distance (GROL), distance required to clear a 50 foot obstacle (DT50), abort speed (ASPD), and stopping

distance (SDST). Setting the nozzles to the displayed NOZ when NRAS is reached results in the displayed GROL,

DT50, ASPD, and SDST. ASPD and SDST are computed and displayed when the abort (ABRT) push button is

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selected on the VRST page. The ASPD and SDST fields display asterisks (* * *) when ABRT is not selected, when

relevant inputs change, and in any weight on wheels condition. The data is computed for both WET (water injected

in engine) and DRY operation having no effect on ASPD or SDST. If outside air temperature is below --5 °C/23 °F

the WET data is not displayed.

The value of OATC or OATF, FELV, and GWT shown at the bottom of the display is system generated but may be

overridden by the pilot. The value for ALTM is equal to the barometric pressure setting set on the standby pressure

altimeter. Runway data (RUNW) and ground wind (GWIND) are pilot entered inputs. If any of these parameters are

not valid then that parameter is not displayed. If the non--valid parameter is essential to the calculations of NRAS,

NOZ, GROL, and DT50 then the calculation is not performed and no data is displayed, with the exception of ASPD

and SDST fields. These two fields display solid asterisks if they become invalid, and during initial fire up.

Data affecting the calculations may be entered by the pilot using the options on the ODU. The following options are

available: GWT, OATC or OATF, FELV, field data (FDAT), ENG. Selecting the FDAT option enables runway

distance (RDIS), runway heading (RDHG), runway wet/dry (RWET/RDRY), and ground wind (GWND) options;

selecting the ENG option enables relative jet pipe temperature, jet pipe temperature limit, and relative hover options.

The RDRY selection toggles between RDRY and RWET, with the default being RDRY. RDIS allows entries from

1,000 to 13,000 feet, with the default being 1,000 feet. If system generated data is overridden by pilot entered data

(OATC or OATF, GWT, and FELV) an asterisk (*) is displayed to the left of the applicable legend.

Note

(ABRT USES 6,000) is displayed when field elevation is greater than 6,000

feet and (CALC 1,400) is displayed when runway condition is wet and

runway length is less than 1,400 feet.

The nozzle rotation airspeed, nozzle setting in degrees, ground roll distance, and the distance required to clear a 50

foot obstacle are calculated as follows: first the hover weight is calculated based on the limiting rpm; maximum rpm

limited by ambient temperature or maximum rpm limited by the jet pipe temperature limiter setting. Next, the gross

weight to hover weight ratio is calculated. If the ratio is less than 1.35 the nozzle rotation angle is set to 55°. If ratio

is greater than 1.35 the nozzle rotation angle is set to 50°. The airspeed for nozzle rotation is based on aircraft gross

weight and the gross weight to hover weight ratio. If the aircraft gross weight is greater than 27,000 pounds and

ambient temperature is greater than 35 °C/95 °F the nozzle rotation airspeed and ground wind are used to compute

the ground roll distance. The ground roll distance and outside air temperature are used to compute the distance to

clear a 50 foot obstacle. Refer to A1--AV8BB--NFM--400, see Short Takeoff Rotation Speed charts for additional

details.

2.20.1.3 Cruise Display

The cruise display presents the best flight profile for altitude cruise (ACR) and optimum cruise (OPCR) performance.

The ACR column displays the profile necessary to obtain the maximum cruise performance at the existing altitude.

The OPCR column displays the flight profile at which maximum cruise performance can be obtained. The data in

the ACR column is system generated as is most data in the OPCR column. In the OPCR column the exceptions are:

system generated values for calibrated airspeed (CAS) and cruise altitude(CALT) can be overridden by pilot entries.

The WIND entry must be pilot entered.

The ODU options available for data entry are gross weight (GWT), drag index (DI), optimum cruise wind (OWND),

cruise altitude (ALT), and calibrated airspeed (CAS).

When the display is selected CALCULATIONS IN PROGRESS appears while the parameters are being generated.

This may take as long as 12 seconds. The aircraft gross weight, drag index, current altitude, ambient temperature,

wind direction and speed, and range to selected waypoint are computed for use in determining the parameters in the

ACR and OPCR columns.

Altitude cruise calculations are performed first. The system tests for and calculates the Mach number which results

in the best fuel efficiency at the current altitude. The average gross weight, total drag, dynamic pressure, and wind

effects during the cruise are some of the factors considered when calculating the Mach number. The Mach number

and equivalent calibrated airspeed are displayed in the ACR column. Fuel consumption is calculated by taking the

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cruise range and dividing it by the fuel efficiency. The remaining fuel, displayed in the ACR column, is the total fuel

minus the fuel consumed during the climb, cruise, and descent. The maximum range allows for an 800 pound fuel

reserve and is calculated by multiplying the available fuel by fuel efficiency.

Optimum cruise calculations are performed next. The system tests for and calculates the Mach number and altitude

that results in the best fuel efficiency. First the fuel and distance to climb to the current altitude is calculated. Next,

the optimum cruise altitude is calculated based on the weight and drag index of the aircraft. Altitudes between the

current altitude and the optimum altitude are examined at intervals to see which one provides the greatest fuel

efficiency. Next, fuel consumption calculations are done to determine how much fuel will remain when the waypoint

or markpoint is reached. After all altitudes have been tested the calculated values for the optimum altitude and Mach

numberaredisplayed. Theoptimum Mach number is also converted to theequivalent calibrated airspeed fordisplay.

The pilot can manually enter an altitude and/or airspeed for the optimum cruise calculations if threats or the weather

dictate. The ALT and CAS options are displayed on the ODU when CRUS is selected on the DDI. Any altitude less

thanorequaltotheoptimumcruisealtitudeoranyairspeedbelow600knotscanbeentered byselecting theapplicable

option, ALT or CAS, and keying the desired entry. Altitude entries greater than the optimum altitude and airspeed

entries 600 knots or greater are disallowed and denoted by the flashing entry on the scratch pad. Manual entries of

altitude or airspeed for optimum cruise calculations are denoted by an asterisk (*) preceding the CALT and/or CAS

legend.

2.20.1.4 Bingo Display

The bingo display presents the best flight profile for altitude bingo (ABNG) and optimum bingo (OBNG)

performance. The ABNG column displays the flight profile necessary to obtain maximum bingo performance at the

existing altitude. The data in the ABNG column is system generated as is most data in the OBNG column. In the

OBNG column the exceptions are: system generated values for CAS and CALT can be overridden by pilot entries.

The WIND entry must be pilot entered.

The ODU options available for data entry are gross weight, drag index, optimum cruise wind, cruise altitude, and

calibrated airspeed.

When the display is selected CALCULATIONS IN PROGRESS appears while the parameters are being generated.

This may take as long as 12 seconds. The aircraft gross weight, drag index, current altitude, ambient temperature,

wind direction and speed, and range to selected waypoint are computed for use in determining the parameters in the

ABNG and OBNG columns. Constant altitude computations are done first. A minimum drag index is set, a fuel

reserve (800 pounds) is subtracted from the total fuel on board and gross weight is adjusted based on fuel to be used.

Maximum range is computed based on total fuel and fuel flow. Constant altitude data is stored for display.

Optimum altitude computations are done next. The fuel required to climb from sea level to the existing altitude is

computed. The optimum altitude is computed based on total fuel. If the optimum altitude is greater than the existing

altitude, the amount of fuel and distance to climb to the optimum altitude is computed. The amount of fuel to be used

for cruise leg is the total fuel minus the climb fuel. The constant altitude bingo computations are done and climb range

is added to the cruise range to get maximum range. The amount of fuel and distance to climb from sea level to the

existing altitude is computed. Then the amount of fuel and distance to climb from sea level to the optimum altitude

is computed. The amount of fuel and distance to climb to the optimum altitude is the difference between the two

values.

The pilot can manually enter an altitude and/or airspeed for the optimum bingo calculations if threats or weather

dictate. The manual entries are accomplished in the same manner as previously described for the optimum cruise

calculations.

2.20.2 Engine Data Entry

To receive the proper ODU displays for entering engine data, press the VRST pushbutton on the DDI basic menu

display. The DDI shows the basic V/STOL--REST display (Figure 2-22). The ODU displays BAW, H2O (OMNI 7.1

and C1+ only), and BDI. Pressing the VL, VTO, or STO pushbutton on the DDI display causes OATC or OATF,

FELV, GWT, and ENG to appear on the ODU option display windows. Corresponding displays appear on the DDI.

To enter engine data select the ENG option. A colon appears next to the ENG legend indicating selection. The relative

jet pipe temperature, jet pipe temperature limit, and relative hover options are displayed. Selecting anyone of these

options enables the UFC scratch pad and keyboard for data entry. A colon appears to the left of the selected option.

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Pressing this pushbutton selects short takeoff display and causes STO to be boxed, indicating selection.

Pressing this pushbutton selects cruise display and causes CRUS to be boxed, indicating selection.

Number (3) indicates selected waypoint. Waypoint can be changed by pressing the increment (↑) or

decrement (↑) pushbutton. Changing waypoint does not change waypoint selection on EHSI/EHSD display.

With H4.0, pressing the increment or decrement pushbutton for more than 0.8 seconds enables a quick

access session allowing the pilot to select between all 60 waypoints, 10 markpoints, and 5 targetpoints.

Pressing this pushbutton (BNGO) selects bingo display and causes BINGO to be boxed, indicating selection.

Pressing this pushbutton (MENU) selects menu display.

Basic drag index (BDI) is entered by the pilot for use in drag index computations. Drag index values from 0 to

30 may be entered. The displayed drag index is for aircraft in a clean configuration and does not include

compensation for external stores.

With OMNI 7.1 and C1+, water weight (H2O) is entered by the pilot for use in gross weight computations. The

displayed water weight is the last entered value. With H4.0, the displayed water weight is the same as the

EDP water weight.

Basic aircraft weight (BAW). The operating weight (OWT) is entered here for use in gross weight

computations. The maxi mum allowable entry is 20,000 pounds. The OWT is the BASIC WEIGHT plus those

items which remain constant for the mission. These items include the pilot (180 lbs), ALE 39/47 (111 lbs),

strakes, gun, pylons, probe, etc. This weight does not include expendable stores, empty ITERS, TPOD, water

or fuel. The weight entered here should match the DSU Aircraft Weight at the bottom of section 3 on the

AV--8B UPC generated Form F.

Pressing this pushbutton (VL) selects vertical landing display and causes VL to be boxed, indicating selection.

10.

Pressing this pushbutton (VTO) selects vertical takeoff display and causes VTO to be boxed, indicating

selection.

11.

With H4.0, ODU window 2 has changed to BDI. ODU window 3 and 5 are blank for Radar and Night Attack

aircraft.

12.

Minimum Fuel (MFUL) and Minimum water (MH2O) can be entered by the pilot in TAV--8B aircraft with H4.0.

This will allow the system to display VL performance that is corrected for RJPT shift.

Figure 2-22. VSTOL -- REST Displays (Sheet 1 of 5)

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Wet is displayed above left column. The numerics listed below WET are the values to be used in a water

injected takeoff or landing. WET data are not displayed if outside air temperature is below --5 °C.

Dry is displayed above right column. The numerics listed below DRY are values to be used in a DRY (not

water injected) takeoff or landing.

Aircraft gross weight (GWT) is the total weight of the aircraft including fuel, water, stores including hung stores,

and rounds remaining including spent casings. The GWT is used to compute F+W. If displayed GWT is pilot

entered, an asterisk (*) is displayed to the left of GWT.

Field elevation (FELV) is used to compute the maximum allowable F+W weight. The displayed FELV is either

computed or pilot entered. An asterisk (*) is displayed to the left of FELV if pilot entered field elevation is being

used.

Altimeter barometric pressure setting (ALTM) is used to compute F+W. The displayed ALTM is equal to the

barometric pressure setting that is set on the standby pressure altimeter.

Outside air temperature °C (OATC) or °F (OATF) is used to compute F+W. The displayed OAT is either pilot

entered or is set equal to the last computed ground temperature. Once weight is off wheels, OAT is no longer

automatically updated; however, the pilot may manually change the temperature if desired, at which time an

asterisk (*) is displayed to the left side of OATC or OATF as applicable. The last recorded temperature is

stored.

This pushbutton allows the outside air temperature to be entered and displayed in units Fahrenheit or Celsius.

The pushbutton scrolls between TEMF and TEMC. The outside air temperature on the VL, VTO, STO and

PHOV display is calculated and displayed as OATF or OATC.

Fuel and water weight (F+W) is computed for WET or DRY takeoff or landing. The displayed value is the

maximum weight of fuel and water aboard the aircraft at which the vertical takeoff or landing can be done.

The corrected F+W display line indicates the vertical landing values (corrected for RJPT shift) based on the

minimum landing fuel and water inputs for TAV--8B aircraft with H4.0.

Figure 2-22. VSTOL -- REST Displays (Sheet 2)

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WET is displayed above left column. The data in the WET column are the variables to be used if a water injected short takeoff is to be

done. The data is always computed when STO is selected, but will not be displayed if outside air temperature is below --5 °C.

DRY is displayed above right column. The data in the DRY column are the variables to be used if a dry (not water injected) takeoff is to

be done.

Outside air temperature °C (OATC) or °F (OATF) is used in performance computations to determine the variables required for a STO.

The displayed OAT is either pilot entered or is set equal to the last computed ground temperature. Once weight is off wheels, OAT is

no longer automatically updated; however, the pilot may manually change the temperature if desired, at which time an asterisk (*) is

displayed to the left side of OATC or OATF as applicable. The last recorded temperature is stored.

Altimeter barometric pressure setting (ALTM) is used in performance computations to determine the variables for the STO. The

displayed ALTM value is equal to the barometric pressure setting that is set on the standby pressure altimeter.

Field elevation (FELV) is used in performance computations to determine the variables required for STO. The displayed FELV is either

computed or pilot entered. If pilot entered an asterisk (*) is displayed to the left side of FELV.

Aircraft gross weight (GWT) is the total weight of the aircraft including fuel, water, stores including hung stores, and rounds remaining

including spent casings. The GWT is used in performance computations to determine the variables required for STO. The displayed

GWT is either computed or pilot entered. If pilot entered, an asterisk (*) is displayed to the left of GWT.

Ground wind (GWIND) is displayed indicating wind direction and magnitude. It is pilot entered.

Runway (RUNW) displays runway distance (RDIS), runway heading (RHDG), and runway condition (RDRY/RWET), which are all pilot

entered.

Distance to 50 feet (DT50) is displayed with entries in WET and DRY columns. Number indicates linear distance in feet required for a

50 foot obstacle clearance.

10.

This pushbutton allows the outside air temperature to be entered and displayed in units Fahrenheit or Celsius. The pushbutton scrolls

between TEMF and TEMC. The outside air temperature on the VL, VTO, STO and PHOV display is calculated and displayed as OATC

or OATF.

11.

Ground roll (GROL) is displayed with entries in the WET and DRY columns. Number indicates minimum ground roll distance required

for short takeoff.

12.

Nozzle setting in degrees (NOZ) is displayed with entries in WET and DRY columns. The nozzle setting for a STO is 50° or 55°

depending upon the hover weight ratio. NOZ displays the setting to which the nozzles should be rotated when NRAS is reached.

13.

Nozzle rotation airspeed (NRAS) is displayed with entries in the WET and DRY columns. The NRAS is the airspeed at which nozzles

are rotated to do a STO. Setting the nozzles to displayed NOZ when NRAS is reached results in the displayed GROL and DT50.

Figure 2-22. VSTOL -- REST Displays (Sheet 3)

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Altitude cruise (ACR) is displayed above left column. Numbers in this column are computed to indicate to the pilot the flight data

required to obtain maximum cruise performance at existing altitude. An asterisk (*) next to the ACR legend indicates existing aircraft

altitude is greater than optimum altitude. When this occurs data in the ACR column is extrapolated from the data in the OPCR column.

Optimum cruise (OPCR) is displayed above right column. Numbers in this column are computed to indicate to the pilot the optimum

altitude at which the maximum cruise performance can be obtained.

Drag index (DI) is displayed to indicate total drag of aircraft and stores. The drag index is used in cruise computations. If an asterisk (*)

is displayed to the left of DI, a pilot entered drag index is being used in computations.

Gross weight (GWT) is displayed to indicate total weight of aircraft, including fuel, water, stores including hung stores, and rounds

remaining including spent casings. The gross weight is used in cruise computations. An asterisk (*) to the left of GWT indicates that a

pilot entered value is being used in computations.

Wind is displayed in direction and magnitude in both the altitude cruise and optimum cruise columns. Displayed ACR wind is computed

from aircraft sensors. OPCR wind is a pilot entered value.

Maximum range (MRNG) is displayed in altitude cruise and optimum cruise columns. MRNG indicates maximum range that can be

reached if the altitude, airspeed, and Mach of respective columns is followed. The MRNG computations allow a 800 pound fuel

reserve.

Remaining fuel (RFUL) legend is displayed with entries in ACR and OPCR columns. Number indicates remaining fuel in pounds after

arriving at selected waypoint if ACR or OPCR profile is followed.

Range (RANG) legend is displayed with entries in ACR and OPCR columns. Number indicates range in nautical miles to the selected

waypoint. If range is more than 10,000 nautical miles, cruise data is blanked.

Cruise altitude (CALT) legend is displayed with entries in ACR and OPCR columns. Number in ACR column indicates existing aircraft

altitude. Number in OPCR column indicates best altitude for use with other OPCR variables to increase maximum cruising range.

Aircraft can have a pilot entered altitude in the OPCR column, denoted by an asterisk (*) next to CALT legend.

10.

Mach legend is displayed with entries in ACR and OPCR columns. Number in ACR column indicates the best Mach to increase

aircraft range at existing altitude. Number in OPCR column indicates best Mach for use with other OPCR variables to increase

maximum cruising range.

11.

Calibrated airspeed (CAS) legend is displayed with entries in ACR and OPCR columns. Number in ACR column indicates the best

airspeed to increase aircraft range at existing altitude. Number in OPCR column indicates best airspeed for use with other OPCR

variables to increase maximum cruising range. Aircraft can have a pilot entered CAS in the OPCR column, denoted by an asterisk (*)

next to CAS legend.

Figure 2-22. VSTOL -- REST Displays (Sheet 4)

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Altitude bingo (ABNG) legend is displayed above left column. Numbers in this column indicate to the pilot flight data required to obtain

maximum bingo performance at existing aircraft altitude. An asterisk (*) next to the ABNG legend indicates existing aircraft altitude is

greater then optimum altitude. When this occurs data in the ABNG column is extrapolated from the data in the OBNG column.

Optimum bingo (OBNG) legend is displayed above right column. Numbers in this column indicate to the pilot the optimum altitude at

which the greatest bingo performance can be obtained.

Drag index (DI) is displayed to indicate total drag of aircraft. This drag index is used in computations and is set to a clean

configuration for bingo flight. Aircraft can have pilot entered drag index.

Gross weight (GWT) is displayed to indicate bingo weight of aircraft including fuel, water, and spent casings but not including stores

including hung stores, expendables, or rounds remaining. An asterisk (*) to the left of GWT indicates a pilot entered value is being

used in computations.

Descent range (DCRG) is displayed to indicate the range to the waypoint at which descent should begin for altitude bingo or optimum

bingo flight.

Wind is displayed in the altitude bingo and optimum bingo columns. Displayed ABNG wind is computed form aircraft sensors. OBNG

wind is a pilot entered value.

Maximum range (MRNG) indicates maximum aircraft range in nautical miles in the direction of the selected waypoint, including cruise

at the displayed altitude followed by an idle descent to sea level. Fuel remaining after descent is 800 pounds.

Remaining fuel (RFUL) legend is displayed with entries in ABNG and OBNG columns. Number indicates fuel remaining in pounds after

arriving at selected waypoint, if ABNG or OBNG profile is followed.

Range (RANG) legend is displayed with entries in ABNG or OBNG columns. Number indicates range in nautical miles to selected

waypoint. If range is more than 10,000 miles, bingo data is blanked.

10.

Cruise altitude (CALT) legend is displayed with entries in ABNG and OBNG columns. Number in ABNG column indicates existing

aircraft altitude. Number in OBNG column indicates best altitude for maximum cruising range. Aircraft can have pilot entered altitude in

OBNG column, denoted by an asterisk (*) to the left of CALT.

11.

Mach legend is displayed with entries in ABNG and OBNG columns. Number in ABNG column indicates the best Mach to increase

aircraft range at existing altitude. Number in OBNG column indicates best aircraft Mach for use with other OBNG variables to increase

maximum range.

12.

Calibrated airspeed (CAS) legend is displayed with entries in ABNG and OBNG columns. Number in ABNG column indicates the best

airspeed in knots to increase aircraft range at existing altitude. Number in OBNG column indicates best airspeed for use with other

OBNG variables to increase maximum range. Aircraft can have pilot entered CAS in OBNG column, denoted by an asterisk (*) next to

the left of CAS.

Figure 2-22. VSTOL -- REST Displays (Sheet 5)

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2.20.3 VREST Calculation Considerations

2.20.3.1 Lateral Asymmetries

The VREST computer does not account for performance degradations induced by lateral asymmetries. Therefore,

for takeoffs with lateral asymmetries above 32,000 inch--pounds, increase VREST calculated STO NRAS by 10

KCAS. For takeoffs with lateral asymmetries above 80,000 inch--pounds, increase VREST calculated STO NRAS

by 15 KCAS.

2.20.3.2 Gross Weight and Air Temperature

The VREST computer does account for aircraft gross weight in STO NRAS calculation, whereas the charts depicted

in A1--AV8BB--NFM--400 do not. Therefore, there is no requirement to add 5 KCAS to VREST calculated STO

NRAS for takeoffs at gross weights greater than 27,000 pounds when ambient air temperature exceeds 35 °C.

2.21

AIR DATA COMPUTER

The air data computer (see Figure 2-23) is a solid state digital computer which receives inputs from the magnetic

azimuth detector, standby altitude Kohlsman setting, TOT probe, AOA transmitter, mission computer, and

pitot/static pressure. Accurate air data and magnetic heading are computed. The air data provided to the mission

computer accounts for air data compensation, position error calibration, and converts indicated airspeed to calibrated

airspeed. Computed data is supplied to the mission computer system, altitude reporting function of the IFF,

multipurpose display system, SAAHS, aileron high speed stops, flap controller, LIDS, Q--feel, stall warning, landing

gear up warning, APU and DECS.

2.21.1 Total Temperature Probe

The total temperature probe is on the upperleft sideof thevertical stabilizer. Operation ofthe probeheater is covered

under Probe Heat Switch in Instruments, this section. The air data computer uses total temperature to calculate

ambient temperature.

2.21.2 ADC BIT Check

To perform an initiated ADC BIT check, press the BIT pushbutton on the DDI menu display to initiate a BIT display.

Press the ADC pushbutton and the word TEST appears next to ADC. After several seconds the word TEST

disappears. If a failure code then appears next to ADC the system has failed the BIT check. If the space next to ADC

remains blank the system has checked good.

2.22

ENTRANCE/EGRESS SYSTEMS (AV--8B)

2.22.1 Canopy System/Boarding Steps

Thecockpit areais enclosed by asliding typecanopy which consists ofacast acrylictransparency mountedin ametal

frame. The canopy is mounted on rails which slope upward toward the rear of the aircraft. The canopy is

counterbalanced to the open position by a spring and pulley system. Except for the aid provided by the

counterbalance, the canopy is opened and closed manually. A canopy seal, routed around the canopy frame, is

automatically inflated by a solenoid operated pneumatic valve whenever the WOW sensor indicates weight is off

wheels.

Note

The windscreen birdstrike protection capability is analytically estimated at

350 knots for a one pound bird.

Normal entrance/egress is gained by four boarding steps on the right forward fuselage below the canopy. One step

is mechanically linked to the canopy, and moves down when the canopy is opened and up when thecanopy is closed.

The other steps are in the moldline of the fuselage and provide steps/handholds for entrance/egress.

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2.22.2 Normal Canopy System

The canopy operating mechanism is mechanically linked to a boarding step on the right forward fuselage so that as

thecanopy opens thestep extends and as thecanopy closes thestep retracts. When moved to thefully closedposition,

the canopy is automatically locked by two latches at the intersection of the lower leading edge of the canopy bow

and the windshield frame. The controls for normal operation are the external canopy release handle, the internal

canopy unlock handle and the canopy bow handles. The external canopy release handle and internal canopy unlock

handle unlock the canopy and the canopy bow handles are used as grips to open and close the canopy. Close the canopy

using both bow handles to ensure both canopy latches are properly engaged. Once unlocked externally, the canopy

can be opened by applying downward force to the boarding step. Avoid use of the boarding step to close the canopy

due to probability of disengaging the canopy/boarding step interlock described in paragraph

2.22.2.5,

Canopy/Boarding Step Mechanical Link.

2.22.2.1 External Normal Canopy Release Handle

The external normal canopy release handle (Figure 2-24) is on the right side of the fuselage below the windshield.

The handle is labeled NORMAL CANOPY RELEASE HANDLE. Operating a push--type latch causes the handle

to pop out from a slot in the fuselage, and then pulling out and forward on the handle releases the two canopy locks.

After the canopy is unlocked, the counterbalance system causes the canopy to slide back while partially extending

the boarding step. Downward pressure on the step will fully open the canopy.

Figure 2-24. Canopy Controls (AV--8B)

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2.22.2.2 Boarding Steps

In addition to the boarding step which extends when the canopy is opened externally, there are three boarding steps

on theright sideof thefuselage below the canopy. These steps, which arestowed flush with moldlineof thefuselage,

are easily located by three vertical lines extending down from the canopy to the steps. To use the steps as

steps/handholds while entering or egressing the cockpit, the steps are unlocked by pushing PUSH buttons on the top

part of the two upper stowed steps. The lower stowed step is interconnected with the upper right step and operation

is controlled by the upper step. Unlocking the steps causes the steps to spring approximately 15° outboard and down

to form a step/handhold. The steps must be stowed before flight. To stow, the steps are rotated upward and inboard

by hand or foot. The steps lock in the stowed position when they become flush with the moldline of the aircraft.

On AV--8B prior to 162077, there are only two boarding steps in the moldline of the fuselage.

On AV--8B 162077 and up, a third step is added approximately 28 inches forward of the top step, and above and

slightlytotherightofthelowerstep.Thestepoperatesthesameasthestepsintheearlieraircraftandisinterconnected

to the lower step so that the latch of this step must be used to unlock both steps.

CAUTION

Egressing the cockpit without using the steps may cause injury. Pilots

should be familiar with the location of the step release without requiring a

visual so that the steps may be quickly unlocked and used in case of an

emergency egress.

2.22.2.3 Canopy Internal Unlock Handle

The canopy internal unlock handle (Figure 2-24) is on the right side of the cockpit just forward of the lower part of

the windshield arch. Pulling aft on the handle releases the canopy locks to allow the canopy to be moved manually.

The handle is spring--loaded to the locked, or forward, position and there is no requirement to manipulate the handle

when locking the canopy.

2.22.2.4 Canopy Bow Handles

Two canopy handles are on the inside of the canopy bow on either side of the cockpit. After the canopy is unlocked,

the handles afford a means of opening or closing the canopy from within the cockpit.

2.22.2.5 Canopy/Boarding Step Mechanical Link

The mechanical link between the canopy and the drop down boarding step includes a disengagable interlock. Should

theboarding step bind orjam forany reason, such as during acrash landing wheretheground clearanceisinsufficient

for the step to fully drop, the interlock can be separated from the step by applying a sudden aft force to both canopy

handles. The interlock can be recoupled by holding the footsteps in the up position while moving the canopy forward

until resistance is felt, then applying a sudden forward force to both canopy bow handles.

2.22.2.5.1 Canopy Latch Viewports

Canopy latch viewports (Figure 2-25) are located on the right and left sides just forward of the lower edge of the

canopy. These viewports allow a visual check of the canopy latches to ensure that they are properly closed. If the

latches are up (not properly closed), orange alignment lines will not be aligned.

2.22.2.5.2 Canopy Caution Light

A CANOPY caution light on the caution/advisory light panel comes on when the canopy is not closed and locked.

The light operates in conjunction with the MASTER CAUTION light.

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Figure 2-25. Canopy Latch Viewport

2.22.2.6 Emergency Canopy Shattering System

Emergency operation of the canopy consists of detonating a small explosive charge of mild detonating cord (MDC)

which serves to break away or shatter the cast acrylic transparency. After the transparency is removed, the pilot can

depart through the canopy frame during ground egress. The MDC is a small diameter explosive cord attached around

the edge of the transparency near the canopy frame, and also attached in an overhead pattern on the top inside of the

canopy. The explosive is fired by one of the three emergency controls. The emergency canopy control is the internal

emergency canopy shattering handle inside the cockpit (Figure 2-24).

2.22.2.6.1 Internal Canopy Shattering Handle

Theinternalcanopyshatteringhandle(Figure2-24)isretainedbyaspring--loadeddetentinayellowand blackstriped

housing at the left forward corner of the canopy frame. The handle, also striped yellow and black, is attached to a

cable which is looped inside a cover to give approximately 5 inches of slack, before connecting to a mechanically

actuated initiator. After the cable slack is taken up and the handle is pulled, the MDC fires to remove the canopy

transparency. The purpose of the cable slack is to provide clearance for the pilot’s hand when the MDC fires.

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2.23

CANOPY SYSTEM (TAV--8B)

Each cockpit (Figure 2-26) has a sideways opening, manually operated acrylic canopy, hinged on the right side. Each

canopy has a counterbalance torsion bar and a damper strut to assist in opening and closing. The damper strut has

a lock mechanism which locks the canopy when it reaches the full open position. Entry to each cockpit is normally

gained from the left side of the aircraft by means of a boarding ladder. There are no boarding steps or handholds

integral to the aircraft. Each canopy is locked independently by threeinterconnected latches which engageon theleft

side of the cockpit. The locks can be operated by interconnected external and internal controls. The external normal

controls are the external canopy lock handles. The internal normal controls are the canopy internal lock handles. A

CANOPY caution light, on the caution lights panel in both cockpits, provides indication that either or both canopies

are unlocked. Each cockpit has an inflatable canopy seal which operates on air from the anti--g system. The two

cockpit seals are interconnected and are inflated through an inflation control circuit when both canopies are closed

and locked with the engine running. The canopy seal circuit is controlled by operation of weight--on--wheels switches.

A clear polycarbonate blast shield is installed between the two cockpits. The shield serves to protect the rear seat

occupant from wind blast in situations where the front canopy is removed during flight. After AFC--373, a canopy

mounted shielded mild detonating cord (SMDC) actuated thruster is located on the forward canopy aft arch at the

aircraft centerline. The thruster fires a pin into the support located on the aircraft structure during ejection, securing

the aft canopy arch to the aircraft. This prevents the canopy arch from deflecting forward during ejection and

damaging the drogue chute bridle on the main parachute container. Damage to the bridle could result in drogue failure

with subsequent main chute damage and/or injury from high speed parachute deployment.

2.23.1 External Canopy Lock Handles

Each canopy can be unlocked by operation of the external canopy lock handle (Figure 2-26) on the left canopy frame.

The external handle is first extended by pressing the handle lock button. Once extended, rotate the external handle

clockwise till the open detent is engaged. The canopy internal lock handle rotates with the external handle. The

canopy is now unlocked and can be opened as far as possible, about 87°, to engage the damper lock. Before entering

the cockpit the external canopy lock handle should be stowed by pressing it flush with the mold line.

Figure 2-26. Canopy Controls (TAV--8B)

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2.23.2 Damper Lock Handles

Once in the cockpit, the canopy can be unlocked from the fully opened position by pulling forward on the damper

lock handle above the right canopy sill (Figure 2-26). The canopy can then be closed. If the canopy will not move,

the damper lock handle is not fully disengaged. The application of excess force on the canopy with the damper lock

handle engaged creates undue stress on the canopy acrylic and can lead to acrylic cracking and failure.

2.23.3 Canopy Internal Lock Handles

The canopy internal lock handle (Figure 2-26) is used to lock or unlock the canopy from inside the cockpit. Once

the canopy is closed it is locked by rotating thehandle forward till theclosed detent is engaged. The top of thehandle

is hinged and spring--loaded outboard so that it can be stowed behind a guard as it reaches the forward position. This

is to prevent inadvertent unlocking of the canopy. An indicator is provided with lines that are parallel to the canopy

sill when the handle is full forward. To unlock the canopy, pull the handle top inboard away from the guard, and then

rotate the handle back to engage the open detent. After the canopy is unlocked, normal cockpit egress is continued

by fully opening the canopy to engage the damper lock and extending the external canopy lock handle by pressing

the handle lock button. When opening the canopy care must be taken to slowly move the canopy to the full open

position until the damper lock engages. The canopy should not be allowed to freefall to the full open position,

additionally, the canopy should not be left open in windy conditions to prevent undue stress on the canopy acrylic.

Disengage the damper lock by pulling forward on the damper lock handle and close the canopy. Lock the canopy by

rotating the external canopy lock handle counterclockwise until the closed detent is engaged, and then stow the

external handle by pressing it flush with the mold line.

2.23.4 Canopy Caution Lights

A CANOPY caution light on the caution lights panel in both cockpits comes on when either or both canopies are

unlocked with power on the aircraft. The light circuits contain two micro switches which are actuated by the aft

canopy latch in the forward cockpit and the center canopy latch in the rear cockpit.

2.23.4.1 Pilot Assist Handles

Two pilot assist handles are provided in each cockpit. They are on the windshield arch in the front cockpit and the

forward canopy arch in the rear cockpit.

2.23.4.2 Canopy Vent Straps

On TAV--8B a canopy vent strap is provided on the assist handle in each cockpit. In utilization, the vent strap is

unhooked from the stowed position bracket (Figure 2-26) next to the assist handle. The canopy is opened

approximately 30° and the strap is hooked to a similar bracket on the canopy frame. The canopy is now secured in

the ventilation position. As the canopy is opened, the vent strap stops the canopy opening at the ventilation position.

The vent strap should be hooked back to the stowed position bracket when not in use.

2.24

EJECTION SEAT (AV--8B)

Each aircraft is equipped with an SJU--4/A ejection seat (Figure 2-27) which utilizes catapult cartridges and rocket

thrust to propel it from the aircraft. The SJU--4/A provides escape capability during takeoff and landing emergencies

at zero speeds, zero altitude, and throughout the remainder of the flight envelope of the aircraft, except for very

unusual flight conditions. The SJU--4/A was qualified for use for aviators from 136 to 213 pounds. It incorporates

a seat mounted personnel parachute and accommodates a survival package with a pararaft, and is designed for use

with an integrated torso harness. An emergency oxygen supply and an emergency locator beacon are provided. A

non--adjustable headrest with canopy breakers is part of the seat structure and houses the personnel parachute. The

front surface of the seat bucket serves as a buffer for the calves of the legs. The sides of the bucket extend forward

to protect the legs and aleg restraint system is incorporated to prevent flailing during ejection. The ejection sequence

is initiated by pulling theejection control handleto full travel with both hands. This fires theprimary initiators(M99)

which then ignites the catapult cartridges to eject the seat. The left primary M99 fires a thruster which removes an

arming key from an airspeed sensor, and initiates the canopy MDC and IFF switch. A dual port thruster is fuzed by

two SMDC assemblies which in turn arms the airspeed sensor. As the seat and outer catapult tubes travel upward the

emergency oxygen and emergency locator beacon (AN/URT--33 or AN/URT--140 after ACC 689) systems are

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activated, the leg restraint lines are pulled to restrain the legs against the front of the seat bucket. After 31 inches of

seat travel, two seat back rockets are ignited to provide the momentum necessary for man/seat combination to attain

sufficient terrain clearance to permit parachute deployment. Seat stabilization, upon ejection, is controlled by the

directional automatic realignment of trajectory (DART) system by the means of lanyards attached between the seat

and aircraft feeding through tension brake assemblies to counteract excessive pitch and roll conditions. The ejection

seat is capable of four modes of operation, depending on ejection airspeed and altitude. The modes are (1) Low

Airspeed/Low Altitude, (2) High Airspeed/Low Altitude, (3) Intermediate Altitude and (4) High Altitude. See Figure

2-28. If ejection was initiated below 225 ±20 knots (180 ±20 knots with IACC 658, or below 165 ±5 knots depending

on altitude after AFC--449 (Figure 2-29) and 7,000 ±750 feet, 7,000 ±100 feet after AFC--449, the airspeed sensor

striker contacts the low speed selector valve, allowing gas flow to initiate low speed/low altitude mode of operation.

The personnel parachute is deployed by a drogue chute and/or a wind oriented rocket deployment (WORD) rocket

motor, depending on the mode of operation. Besides the ejection control handle, the following controls are

incorporated on the seat: shoulder harness lock lever, the emergency restraint release handle, emergency oxygen

release, ground safety control handle, and a seat positioning switch mounted above and outboard of the left console

fuel panel. After AFC--449, an Electronic Airspeed/Altitude Sensor (EAAS) is used which has two Light Emitting

Diodes (LEDs) to indicate the condition of the internal BATTERY (yellow) or a FAULT (red) condition. When

aircraft power is initially applied to the sensor the LEDs come on for approximately 8 seconds and then go off if no

failures are detected. The EAAS has two independently functioning modules which measure pitot--static pressure

from the aircraft to determine either mode (1) Low Airspeed/Low Altitude or (2) High Airspeed/High Altitude

operation. The FAULT LED indicates that either or both of the modules have failed the start--up self--test. If the

FAULT LED is illuminated, assume the ejection seat mode of operation will be mode (1) Low Airspeed/Low

Altitude.

Note

After AFC--449, any on/off cycle of the aircraft battery shall be of a

sufficientdurationtoallowfortheEAAS tofinish itsapproximate8second

self--test before turning the aircraft battery to the off position. Completion

is indicated by the EAAS Battery and EAAS Fault LEDs turning off.

2.24.1 Front Cockpit Ejection Seat SJU--13/A

The SJU--13/A seat (Figure 2-27) is the same as the SJU--4/A except that it has two divergence rockets installed on

the left side and canopy breakers installed on the top. The divergence rockets provide separation from the rear seat

during dual ejection. The rockets ignite simultaneously during ejection with the two seat back rockets and they will

fire during dual or single ejection. Because of the slope of the canopy transparency in the front cockpit, the canopy

breakers provide break through capability if required during ejection. The SJU--13/A seat has a 0.4 second delay to

provide front and rear seat separation.

2.24.2 Rear Cockpit Ejection Seat SJU--14/A

The SJU--14/A seat is the same as the SJU--4/A except that it has two divergence rockets installed on the right side.

The divergence rocket motors operate the same as on thefront seat except that the rocket action causes the seat’s path

to diverge opposite that of the front seat.

2.24.3 Survival Kit

Thesurvival kit isapostejection lifesupport unitthat alsoacts asastructuralportion oftheejectionseat. Theprimary

structural member of the survival kit is the seat pan (attached to the seat bucket) which serves as a mounting base

for the following post ejection life support equipment: survival package, emergency oxygen supply and emergency

locator beacon. The seat pan also provides secure attaching points for the pilot’s lap belts, and a contour/self--contouring

cushion is fitted on top. The seat is released (separated) automatically during the ejection sequence and the survival

kit is released manually after ejection. The entire survival kit is retained intact as a unit until the survival package

is manually deployed.

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2.24.4 Pilot Harness and Seat Harness

The pilot’s harness is a combined parachute harness and restraint garment that is put on before entering the cockpit.

When seated, the pilot’s upper harness is connected to the parachute risers by two Koch connectors, and the lower

harness is connected to the two lap belt connectors. The parachute risers and lap belts remain in the aircraft between

flights.

2.24.5 Shoulder Harness Inertia Reel and Gas Generator

A shoulder harness inertia reel and gas generator is installed on the ejection seat. The inertia reel provides the pilot

with capability for locking his shoulder harness to prevent forward motion and also capability for unlocking his

shoulder harness so that he is free to move forward or aft. In addition, the reel will lock automatically, although the

shoulder harness lock lever is in the unlocked position, whenever high g conditions are encountered. The shoulder

harness gas generator operates only during ejection to pull the shoulder straps back so the pilot is positioned for

ejection. The seat also contains an inertia reel guillotine which when activated, severs the shoulder straps to free the

pilot’s upper torso automatically following ejection, or manually when the emergency restraint release handle is

operated during manual separation or emergency ground egress.

2.24.5.1 Shoulder Harness Lock Lever

Located on the seat bucket just outboard of the pilot’s lower left thigh, the shoulder harness lock lever has two

positions,forwardandaft. Theforward isthemanuallocked positionand preventsforward movementoftheshoulder

harness. The aft position is the unlocked/auto lock position, and forward movement is provided except for high g

conditions. Once locked automatically while in the aft position, the harness can be unlocked by cycling the handle

forward and then aft to the unlocked position.

2.24.5.2 Seat Adjust Switch

The seat adjust switch, on the left console above and outboard of the fuel panel, operates the actuator. The operating

cycle is 30 seconds on and 1 minute off to permit cooling.

2.24.6 Leg Restrainers

The leg restraining system is designed to prevent leg flailing during ejections. The system consists of leg garters, leg

garter straps, and ratchet/snubbing assemblies. The leg garters are adjustable and worn at the mid--calf on each leg.

The garters are hooked to a key on each leg strap, and the garters can be detached from the keys and thus they can

be worn to and from the aircraft. The leg garter straps are routed through the leg garter keys so that the legs will be

restrained against the front of the seat when the straps are pulled taut during ejection. The garter straps are routed

through the ratchet/snubbing assemblies on the front of the seat, through the garter keys, and the upper end of the

garters are secured to locking devices in the ratchet/snubbing assemblies by releasable pins. The pilot can increase

strap length for adequate leg motion by pulling out a springloaded release pin on the ratchet/snubber assembly and

then pulling the strap forward. To decrease strap length, the pilot reaches behind the ratchet/snubber assembly and

pulls the strap through. During the ejection sequence, the straps are pulled back through the ratchet/snubber

assemblies, drawing the pilot’s legs aft and restraining them against the seat structure. As the seat travels upward,

tension is applied to the rip stitch. The lanyards are sheared at the rip stitch, severing the connection between seat

and aircraft. At seat/man separation, as the pilot is extracted from his seat by his parachute, the pins in the

ratchet/snubber assemblies are released to free the pilot’s legs. Operation of the emergency restraint release handle

also releases the pins during manual separation and emergency ground egress.

2.24.7 Ejection Control Handle

The ejection control handle, on the forward seat pan, is mechanically connected to the ejection initiation system.

When the handle is pulled fully upward, the ejection sequence is initiated. After actuation, the handle remains

attached to the seat and the handle must be released before or during man/seat separation. Initiators will fire before

maximum handle travel of approximately 3.25 inches so that the pilot can retain his grip and prevent arm flailing

when exposed to windblast.

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2.24.8 Ground Safety Control Handle

Theground safety control handleon theright forward sideoftheseat at shoulder level provides ameans forsafetying

or arming the seat. The armed position is down and back against the seat. The safetied position is up and forward

away from the seat. Prior to moving the handle from one position to the other, a handle lock must be released. The

lock is released by pulling on the spring--loaded lower end of the handle.

2.24.9 Post Ejection Sequencing System

The post ejection sequencing system includes all gas operated and cartridge actuated devices required to initiate the

sequencing functions of the four post ejection modes. The system includes a parachute container opener assembly

which is activated by gas pressure from either the 7,000 foot or 14,000 foot aneroid activated initiators or gas pressure

from the man/seat separation initiator. When activated, the opener assembly opens the container so that the personnel

parachute can be deployed. The system also includes the wind oriented rocket deployment (WORD) rocket which

is mounted on the back of the seat and is connected on one end to the WORD bridle assembly (personnel parachute

withdrawal line) and on the other end to the drogue bridle assembly (drogue suspension lines). The WORD rocket

is attached to theback ofthe seat by means of theWORD motor/droguerelease assembly. The WORD motor/drogue

release assembly actuates to release the rocket motor (and with it the personnel parachute withdrawal line) from the

back of the seat during all four modes of operation of the post ejection sequencing system. After release, the WORD

rocket motor is then fired by inertia (I -- WORD deployment sequence) for low speed ejections, or the WORD rocket

motor is fired by the pull of the drogue suspension lines (drogue -- WORD deployment sequence) for higher speed

ejections. See Figure 2-28 for each of the four modes of operation of the post ejection sequencing system.

2.24.10 Parachute

The personnel parachute system includes a WORD bridle assembly (parachute withdrawal line), riser assemblies

with snubbing lanyards which initiate the ballistic spreading gun and initiate man/seat separation, a spring--loaded

internal pilot parachute, a main canopy, a ballistic spreading gun, and an override disconnect assembly. Parachute

deployment begins, propelled via the WORD bridle, by force generated by means of drogue--WORD, I--WORD, or

in the event of WORD bridle failure, by the internal pilot parachute. When the main canopy and suspension lines

are fully deployed, line stretch pulls a lanyard which, in turn, exerts tension on a spring--loaded firing pin in the

ballistic spreading gun assembly. The pin is withdrawn, igniting the spreading gun cartridge. Cartridge energy expels

14 pistons which, in turn, expel 14 slugs, attached to alternate suspension lines, in a 360° pattern, thus spreading the

main canopy. Should the spreading gun cartridge fail to fire, continued pull on the firing lanyard will remove a

piston--retaining band, freeing the pistons, slugs and suspension lines to allow conventional canopy inflation.

Theparachuteisa28foot flatcanopy typeparachutewithalternating coloredpanels (white,olivegreen,international

orange, and sand shade). The parachute is rated for suspension of 100--300 pounds with a descent rate of 13.3 to 23.1

feet per second.

2.24.10.1 Four--Line Release System

With ACC--667 PART 2, the four--line release system is a feature used to reduce canopy oscillation and provide

limited forward motion and directional control of the parachute during descent. Suspension lines 1, 2, 27 and 28 are

rigged so that when the four--line release lanyards are pulled sharply down, the lines are released from the left and

right connector links. Release of the lines allows four gores of the canopy to billow free creating a vent at the rear

of the canopy. Canopy oscillations are reduced or eliminated upon actuation of the four--line release system. The

escape of air through the vent imparts a horizontal motion to the parachute assembly. The two four--line release

lanyards, one attached to each rear riser, are used for directional control.

2.24.11 Sea Water Activated Release System

On TAV--8B, AV--8B 162721 and up, an automatic backup method of releasing the parachute canopy when landing

in sea water after an emergency egress is installed. The sea water activated release system (SEAWARS) system

consists of two releases mounted outboard of the Koch connectors on the parachute risers. Each release contains an

electronics package (sensor), battery, cartridge, and canopy release fitting. Immersion in sea water activates the

sensors which mechanically release the parachute risers from the pilot’s restraint harness. With SEAWARS installed,

the normal procedures for connecting and releasing the Koch fitting are unchanged.

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2.24.12 Man/Seat Separation System

After ejection, as aerodynamic drag is imposed on the personnel parachute, tension on a lanyard sewn to the right

hand and left hand riser assemblies fires a man/seat separation initiator, producing gas pressure which is directed to

an inertia reel guillotine. The guillotine severs two inertia reel shoulder straps, releasing restraint on the pilot’s upper

torso. Simultaneous actuation of the man/seat separation mechanical linkage by the right hand and left hand riser

assemblies lanyard releases the survival kit and the pilot from the ejection seat. The pilot and survival kit are then

withdrawn from the seat assembly by the aerodynamic drag on the personnel parachute.

2.25

MANUAL SEPARATION AND EMERGENCY GROUND EGRESS

2.25.1 Manual Separation

If any component should fail during any of the four automatic ejection modes, the occupant can operate the man/seat

separation system manually. By operating the emergency restraint release handle, the man/seat separation initiator

fires and the following sequence takes place: the inertia reel guillotine actuates to sever the inertia reel shoulder straps,

thetorquereleaserod rotatesto releasethesurvivalkit andleg restraints,theWORDmotoractuatesand theparachute

container is opened. The personnel parachute then deploys, aided by either the drogue--WORD or I--WORD

sequence, depending upon airspeed at the time the emergency restraint release handle is pulled.

2.25.2 Emergency Ground Egress

During emergency ground egress when it is desired to evacuate the cockpit with the survival equipment, as part of

the procedure the pilot must actuate the emergency restraint release handle in order to release the survival kit and leg

restraints.

2.25.3 Emergency Restraint Release Handle

The emergency restraint release handle on the seat bucket just outboard of the pilot’s right lower thigh is actuated

by first squeezing the handle and then pulling up and aft. Once actuated, the handle locks up in the released position

and will remain there until the handle is reset. During ground egress when the handle is pulled the following occurs:

the ejection initiation system is safetied, the inertia reel shoulder straps are severed, the leg restraint straps and

survival kit are released, the main parachute container is opened and the WORD motor/drogue release assembly is

actuated. During manual separation after ejection, the above occurs plus the WORD rocket motor fires to deploy the

main parachute.

2.25.4 Survival Package

The survival package is attached to the survival kit seat pan through a lanyard system which allows the package to

fall free of the seat pan, yet remain in close proximity to the pilot. When the survival package is manually released

by actuating the survival package release, it is allowed to fall approximately 12 feet, where its fall is snubbed by a

lanyard which causes inflation of the life raft. The package then falls another 13 feet below the raft, giving stability

to the raft during parachute descent. The survival package contains: a life raft, signal devices, medical aids and

miscellaneous post ejection survival aids.

The following is a representative list of items contained in the survival package:

Cord, (Nylon), Fibrous Type I 50 feet.

Signal, (Flare), Smoke and Illumination, MK--124 MOD 0.

Sea (Dye) Marker, Fluorescent 2.

Sponge, (Bailing), Cellulose Type II, Class 2.

SRU--31/P Survival Kit, Packet (#1) (Medical).

SRU--31/P Survival Kit, Packet (#2) (General).

SRU--31A/P Optional.

Bag, Drinking Water (50 ml).

Opener, Can, Hand.

Ground/Air Emergency Code Card.

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Blanket, Combat Casualty, (3 oz).

Clear Vinyl Envelope 2 00--334--4120 or Equivalent.

Beacon Set, Radio AN/URT--140.

Liferaft, Inflatable.

2.25.4.1 Survival Package Release

The survival package release is a loop on the back, right side of the survival kit. Pulling the loop releases the survival

package from the seat pan.

2.25.5 Emergency Oxygen

The emergency oxygen supply is a completely self--contained unit, attached to the bottom of the seat pan, that

provides 100 cubic inches of breathing oxygen. It can be operated either automatically (during ejection) or manually

if a failure occurs in the aircraft main oxygen system. Automatic emergency oxygen control is provided by a lanyard

assembly located on the underside of the seat panel left thigh support and is connected to the seat catapult cartridge

manifold. Upon upward movement of the seat panel, as in an ejection, automatic actuation is provided. A pressure

gauge, visible through the cutout on the forward left hand side of the seat cushion, should register 1,800 psi (needle

in the green area) with full bottle. Duration of emergency oxygen supply is approximately 15 minutes, depending

upon altitude (the higher the altitude, the longer the duration), since oxygen is delivered by the mask regulator only

upon demand. Oxygen from the bottle is supplied by release of a valve in the pressure regulator at the forward left

edge of the survival kit. Two keeper yokes on the valve shaft keep the valve in the closed position until emergency

oxygen is required. One of these yokes is attached by a cable to the manual release (emergency oxygen actuator),

which is stowed on the forward left--hand inboard edge of the survival kit. The other is attached by cable, through

a quick disconnect fitting to a lanyard attached to the ejection seat catapult cartridge manifold. Either cable will

dislodge a yoke and actuate the emergency oxygen supply valve to provide oxygen from the bottle and shut off the

main aircraft supply. When the seat is ejected or the pilot leaves the aircraft still attached to his survival gear, the cable

attached to the catapult cartridge manifold is pulled and emergency oxygen is supplied automatically. Manually

pulling the emergency oxygen actuator provides emergency oxygen at any time.

2.25.5.1 Emergency Oxygen Release

Manual emergency oxygen release (emergency oxygen actuator) is provided by a handle/pull ring located on the

inboard side of the left thigh support. An upward pull on the handle provides emergency oxygen to the pilot in the

event of a failure in the aircraft main oxygen system or a failure of the emergency oxygen actuation system during

ejection. Once activated, the emergency oxygen supply can not be turned off.

2.26

EMERGENCY LOCATOR BEACON (AN/URT--33 OR AN/URT--140 AFTER ACC--689)

An emergency locator beacon on top of the seat pan, is automatically actuated during emergency egress via the

auto--actuation lanyard that is connected to the radio, emergency oxygen supply, and cockpit deck fitting.

2.26.1 Canopy/Interseat Sequencing System

A canopy/interseat sequencing system is provided to allow dual or single ejection initiated from either cockpit

depending on the position of the ejection sequence selector.

2.26.1.1 Ejection Sequence Selector

Through use of the ejection sequence selector on the left side of the rear main instrument panel (foldout FO--2), six

modes of ejection sequencing can be selected. The selector has three positions: DUAL, FWD and AFT/SOLO. The

DUAL position is with the selector handle aligned horizontally. The FWD position is with the handle in the 45°

counterclockwise position from the horizontal. The FWD position can only be maintained by use of a collar placed

around the shaft of the handle. The AFT/SOLO position is with the handle aligned vertically. If the handle is released

at any point other than the AFT/SOLO position, the handle will return to the DUAL position.

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A1-AV8BB--NFM--000

DUAL

Dual ejection results from ejection initiation in either cockpit. Rear seat ejects

(horizontal)

first, followed by front seat after a 0.4 second delay.

FWD

Dual ejection results from ejection initiation in front cockpit. Rear seat ejects

(45° ccw)

first, followed by front seat after 0.4 second delay.

Single ejection results from ejection initiation in rear cockpit. Front cockpit can

then eject solo, with a 0.4 second delay.

AFT/SOLO

Dual ejection results from ejection initiation in rear cockpit. Rear seat ejects first,

(vertical)

followed by front seat after a 0.4 second delay.

Single ejection results from ejection initiation in front cockpit, with a 0.4 second

delay. Rear cockpit can then eject solo.

2.27

ENVIRONMENTAL CONTROL SYSTEM

The environmental control system (ECS) provides conditioned air and pressurization for the cockpit and avionics

equipment. The ECS also provides conditioned air to the windshield defog, anti--g and canopy seal systems. See

Environmental Control System, foldout section, for environmental control system schematics.

2.27.1 ECS Air Sources

2.27.1.1 Bleed Air

Thenormal sourceofECS airis the 6th--stageengine bleed air except for radarequipped aircraft which use8th--stage

engine bleed air. Through a series of manifolds and valves this air is cooled and mixed to reduce temperature and

pressure to usable levels.

2.27.1.2 Ram Air

Asecondary sourceofECS airis ramairwhichcan beused toventilatethecockpit andprovidecoolingairtoavionics

equipment requiring forced air cooling.

2.27.2 Cockpit Air Conditioning

High temperature engine bleed air is routed through the cabin pressure regulator and shutoff valve and venturi to the

primary heat exchanger where it is cooled. The cooled output of this heat exchanger is applied to the compressor

turbine. Thecompressed airis run through thesecondary heat exchanger, and then expanded in the expansion turbine

section, resulting in cold air that is mixed with hot bleed air from the cabin temperature control valve. This valve is

modulated by the temperature controller on the ECS panel. The conditioned air passes through a water separator and

vent/defog changeover valve to the cockpit ECS louvers and windshield defog ducts.

Ram air across the heat exchanger is used for initially cooling the bleed air. The ram air discharge from the heat

exchangers is routed to the engine inlet or vented to the ECS bay on nonradar aircraft or vented overboard on radar

aircraft depending on the flight conditions encountered. At high air speeds the ram air is vented to the ECS bay or

vented overboard as applicable. At lower airspeeds and during ground operation airflow across the heat exchangers

is augmented by the engine turbine fans which pull ambient air across the heat exchangers into the engine inlet.

2.27.2.1 Temperature Management

The pilot can control cockpit temperature by selecting either a manual (MAN) mode or automatic (AUTO) mode with

the temperature controller on the ECS panel.

In the MAN mode, holding the cabin temperature control knob to the COOL or WARM settings applies a control

signal directly to the cabin temperature control valve to open or close the valve as desired. The signal is applied to

thevalveas long as theknob is held in theCOOL orWARM setting. In this mode, thecabin temperaturecontrol knob

is spring loaded to the center position and returns to the center position when released.

In the AUTO mode the temperature is electronically regulated by the cabin temperature control. The control

continuously monitors onboard temperature sensors and compares the sensed temperature to the selected

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temperature. If an imbalance exists, too hot or too cold, a signal is applied to position the cabin temperature control

valve to maintain the selected temperature. In this mode the cabin temperature control knob, when released, remains

in the position selected.

2.27.2.2 Temperature Controller

The temperature controller is on the ECS panel.

AUTO

With knob in automatic section, counterclockwise rotation decreases cockpit and

windshield defog air temperature. Clockwise rotation increases temperature. In

AUTO mode, temperature is electronically regulated. This is the normal mode of

operation. In AUTO mode, knob will remain in position selected when released.

MAN

In MAN mode, knob is springloaded to the center position. With knob in manual

section, counterclockwise rotation increases cockpit air and windshield defog

temperature. Clockwise rotation decreases temperature. In MAN mode

temperature is controlled through direct operation of the cabin temperature

control valve.

Note

Should chunks oficeand/orsnow bedetected discharging from thecockpit

ECS louvers, a higher cockpit air temperature should be selected to restore

the system to normal operation. The ice/snow condition is caused by too

cold a selection of the temperature controller resulting in a freeze--up

condition of the water separator coalescer and operation of the internal

coalescer bypass relief.

2.27.3 Defog System

The defog system uses the same air for defogging as passes to the cockpit ECS louvers. When defog is selected, a

larger portion of airflow is diverted from the pilot to the windshield defog ducts. Temperature of the defog air is

automatically increased when MAX DEFOG is selected and the temperature controller is in the AUTO mode.

Temperature of the defog air may be increased or decreased by the pilot by changing the temperature controller

position in either the AUTO or MAN mode.

2.27.3.1 Defog Switch

The defog switch, labeled CABIN, is on the ECS panel.

NORM

The vent/defog changeover valve is energized to provide the majority of

conditioned airflow to the pilot.

DEFOG

Increased airflow is directed to windshield. Vent/defog changeover valve

deenergized.

MAX

Increased airflow is directed to windshield at an increased temperature. MAX

DEFOG

DEFOG position inoperative with temperature in MAN mode.

Note

For extreme windshield fog conditions, place defog switch to MAX

DEFOG and increase the temperature in the AUTO range on the

temperature controller.

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A1-AV8BB--NFM--000

2.27.3.2 Defog Shutoff Valve

A defog shutoff valve provides for increased pilot cooling during periods of low system pressure (e.g. engine at idle)

by directing all system airflow to the ECS louvers. A pressure switch monitors the primary heat exchanger outlet

pressure and provides a signal to close the defog shutoff valve when pressure is less than 20 psi.

2.27.3.3 Windshield Overheat Caution Light

When the windshield temperature limit of the defog air is exceeded, the WSHLD overheat caution light comes on.

The WSHLD caution light is on the caution/advisory light panel and operates in conjunction with the mastercaution

light. When the WSHLD overheat caution light comes on power is automatically supplied to the cabin temperature

control valve to drive it to the closed (low temperature) position.

2.27.4 Cockpit Pressurization

Cockpit pressure scheduling is maintained by the cabin pressure regulator/discharge valve on the forward cabin

bulkhead. From sealevel to 8,000 feet altitudethecockpit is unpressurized. Between altitudes of8,000 feet to 23,000

feet the system maintains a constant cockpit pressure altitude of 8,000 feet. At altitudes above 23,000 feet, the cockpit

pressure regulator maintains a constant 5 psi pressure differential greater then ambient pressure. The cockpit pressure

can be dumped by setting the cockpit pressure switch, on the ECS panel, to DUMP. On the TAV--8B the air

communication duct connects the two cockpits for pressure equalization and air circulation.

2.27.4.1 Safety Relief Valve

If the cockpit pressure regulator/discharge valve malfunctions the cockpit safety relief valve will open to vent excess

cabin pressure. The valve also provides negative pressure relief. This occurs during rapid aircraft descent. Normal

operation of the system cannot always react to change cabin pressure as fast as the aircraft can descend. When this

occurs, the safety relief valve will open and allow ambient pressure to enter the cabin to equalize pressure.

2.27.4.2 Cockpit Altimeter

A cockpit altimeter is mounted to the right of the caution light panel on the main instrument panel (see cockpit, foldout

section) and indicates cockpit pressure altitude from 0 to 50,000 feet.

2.27.4.3 Cockpit Pressure Switch

The cockpit pressure switch, labeled PRESS, is on the ECS panel on the right console.

NORM

Pressure regulator and shutoff valve open, with normal conditioned air and

pressure provided to system.

DUMP

Cabin safety/dump valve control valve energized to dump cockpit pressure.

Pressure regulator and shutoff valve still open, supplying normal conditioned air

to the crew station.

RAM

Cabin pressure is dumped as in the DUMP position, and pressure regulator and

shutoff valve energized to shut down engine bleed air to system. Nose ground

cooling/ram air valve opens to supply ram air to entire system. Switch position

also energizes the ground cooling fan to facilitate flow of ram air to forward

avionics equipment.

2.27.5 Radar Waveguide Pressurization

On radar aircraft the forward ECS system provides positive pressure to the radar waveguide to prevent arcing at high

altitude. The cooled output of the secondary heat exchanger is split and a portion is provided for waveguide

pressurization. The ECS system provides pressure regulated, clean, dry air to the waveguide.

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2.27.6 Anti--g System

With the cockpit air conditioning and pressurization system operating, the anti--g suit remains deflated up to

approximately 1.5g’s. Above this acceleration, the air pressure applied to the suit increases in proportion to increasing

g’s. While acceleration is constant, the suit remains inflated at a constant pressure and, as acceleration decreases, the

suit will deflate in proportion to the decrease in g’s. A manual inflation button in the anti--g suit valve on the aft left

console allows the pilot to manually inflate the suit for purposes of checking the system or for fatigue relief. The

anti--g system is inoperative with the cockpit pressure switch in the RAM position.

2.27.7 Canopy Seal

The canopy seal inflates automatically on takeoff when the canopy seal control valve deenergizes open with

weight--off--wheels. The valve energizes closed and vents canopy seal pressure with weight--on--wheels. With the

cockpit pressure switch in the RAM position, the canopy seal is inoperative.

2.27.8 Cockpit Equipment Cooling

Cockpit conditioned air is also used to cool cockpit installed avionics equipment. A cooling fan, which operates

continuously, pulls cockpit air through the following cockpit installed equipment: ODU, HUD, DDI(s), and on

non--radar aircraft the DC.

2.27.8.1 Crew Station Cool Caution Light

Failure of the cockpit equipment cooling fan causes the CS COOL caution light to come on. The light is on the

caution/advisory lights panel and operates in conjunction with the master caution light. To protect the avionics

equipment should the CS COOL light come on, the cockpit temperature should be adjusted to a setting as cold as

practical and any unneeded avionics equipment should be turned off.

2.27.9 Forward Equipment Cooling

2.27.9.1 Normal Operation

On non--radar aircraft part of the cold expansion turbine discharge air is combined with cockpit conditioned air and

routed to a plenum for normal INS cooling. Discharge air from the cabin pressure regulator valve is ducted into the

nose cone to cool the angle rate bombing system (ARBS) which contains internal cooling fans.

On radar aircraft the bleed air output from the primary heat exchanger is mixed with the cold air output of the

expansion turbine by way of the avionics cooling valve. This modulating valve is driven by the airflow

temperature/sensor controller and is positioned to maintain a 40 °F supply temperature to the forward avionics

systems (radar, INS, FLIR). The radar system is also cooled by a liquid cooling system. See paragraph 2.27.11.

2.27.9.2 Emergency/Ground Operation

On non--radar aircraft, when the aircraft is on the ground, a ground cooling fan operates through the

weight--on--wheels switches to provide cooling air to forward equipment. The ground cooling fan operates in flight

when the RAM position of the cockpit pressure switch is selected. When RAM is selected, ram air is supplied to the

entire system, except for the canopy seal and anti--g system.

On radar aircraft the ground cooling fan also operates through the weight--on--wheels switch to provide cooling air

to the INS, FLIR, and radar. In flight, when the temperature sensor/controller detects an out of tolerance condition

(FWD BAY caution) or the RAM position of the ECS pressure switch is selected, the ground cooling fan operates

to provide cooling air to the forward equipment, and the emergency ram air valve is opened to provide ram air to the

INS cooling plenum and to the cockpit. The canopy seal, anti--g system, and radar waveguide pressurization system

are not affected.

2.27.9.3 Forward Equipment Bay ECS Switch

On radar aircraft the forward equipment bay ECS switch, labeled FWD EQUIP, is on the ECS panel.

NORM - Switch is spring loaded to this position. System operation is normal.

RESET - Restarts system after temporary malfunction and shutdown.

ORIGINAL

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A1-AV8BB--NFM--000

2.27.9.4

Forward Equipment Bay Caution Light

On radar aircraft the FWD BAY caution light comes on whenever the airflow temperature sensor/controller senses

inadequatecooling, temperatureout oftoleranceand/orlow orno system airflow. Thelight is on thecautionadvisory

light panel and operates in conjunction with the master caution light.

2.27.10 Aft Fuselage Equipment Cooling

2.27.10.1 Normal Operation

High temperature and pressure engine bleed air is routed through the aft equipment pressure regulator and shutoff

valve. This valve functions automatically but can be controlled manually by the EQUIP switch (non--radar aircraft)

ortheAFT EQUIPswitch (Radaraircraft)locatedon theECS panel.With thevalveopen,thebleedairpassesthrough

a heat exchanger cooled by ram air from the intake at the base of the vertical stabilizer. From the heat exchanger, the

air is expanded in a turbine which drives a fan in the ram air exhaust. The fan induces ram air flow through the heat

exchanger, particularly during ground operation. The discharged cool air from the turbine is warmed to the proper

temperature by mixing it with hot bypass bleed air which is regulated by a thermostatically controlled temperature

control valve. The mixed air is then routed to the aft equipment cooling plenum for distribution to avionics

equipment.

2.27.10.2 Emergency/Ground Operation (Radar Aircraft)

On the ground with engine not running, or in flight and aft equipment cooling system pressure is lost or system shut

down, the avionics auxiliary cooling fan will operate to supply ambient airto theaft equipment cooling plenum. The

cooling fan turns on automatically (assuming electrical power is available) whenever any of the following occurs:

main system turned off, turbine inlet pressure too high, bleed air overpressure, incorrect two way valve position,

equipment delivery air temperature outside limits of control system, incorrect auxiliary cooling valve position. Any

of the above actions result in the avionics auxiliary cooling valve opening to allow ambient air from the ram air duct

to go to the aft equipment cooling plenum via the cooling fan.

2.27.10.3 Aft Equipment Bay ECS Switch

The aft equipment bay ECS switch, labeled EQUIP on non--radar aircraft or AFT EQUIP on Radar aircraft, is on the

ECS panel.

OFF - The equipment pressure regulator and shutoff valve is energized close to shutdown the system.

ON - A neutral switch position. This is normal position for switch.

RESET - Restarts system after temporary malfunction and shutdown.

2.27.10.4 Aft Equipment Bay Caution Light

The AFT BAY caution light comes on whenever the cooling fan is operating (except on the ground) and/or in the

emergency/ground cooling mode. The light is on the caution/advisory light panel and operates in conjunction with

the master caution light.

2.27.11 Liquid Cooling System

The aft ECS system is an integral part of the radar liquid cooling system. The liquid/air heat exchanger extracts heat

from the liquid coolant flowing in the closed loop system. A thermostatic temperature control valve set in parallel

with the heat exchanger core, senses coolant temperature and mixes bypass and core fluid flows to maintain delivery

temperatures to the transmitter at 80 ±10 °F. During normal operation ram air is routed through the transmitter

auxiliary cooling valve through the heat exchange, windmills the transmitter auxiliary cooling fan and is then vented

overboard. During ground operation and/or during periods of low ram air pressure, the transmitter auxiliary cooling

fan is used to draw air through the system.

During abnormal operation, ram air temperature too hot or too cold, conditioned bleed air is applied across the heat

exchanger. This occurs when the ram air temperature switch senses an out of tolerance condition. The conditioned

bleed air is routed through the bypass two--way valve, the heat exchanger, and the transmitter two--way valve to the

aft equipment cooling plenum.

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2.28

EMERGENCY EQUIPMENT

2.28.1 Jettison Systems

The jettison systems consist of the emergency jettison system and the selective jettison system.

2.28.1.1 Emergency Jettison Button

The emergency jettison system utilizes the emergency jettison button to jettison all stores and suspension equipment

from BRU--36 bomb racks (bomb rack) on stations 1 through 7. AIM--9 missiles are not jettisoned as normal loading

practice does not install the impulse cartridges in the bomb rack.

The landing gear handle must be up, or the weight must be off the aircraft landing gear, or the armament safety

override switch (inaccessible from cockpit) must be in the override position, to enable the emergency jettison button.

When the landing gear handle is in the UP position, the armament safety override switch will be disengaged if

previously engaged. The emergency jettison button, labeled EMER JETT, is on the landing gear control panel and

is painted with alternating black and yellow stripes. Emergency jettison is performed by pressing the button with the

proper ground interlocks satisfied. All weapons are jettisoned in a safe condition, however, there is no guarantee fuzes

will not arm during release or detonate on impact. Jettison occurs at 50 millisecond intervals starting with stations

1, 4 and 7, then stations 2 and 6, and then stations 3 and 5. Stores are not jettisonable from the outrigger pylons.

On the TAV--8B, an emergency jettison button is provided in the rear cockpit outboard of the landing gear/flaps

control panel. The rear cockpit button parallels operation of the front cockpit button.

2.28.1.2 Selective Jettison

Selective jettison is performed by the selective jettison select knob in conjunction with the selective jettison (JETT)

pushbutton,andinsomecases,inconjunctionwiththestationselectbuttons. Selectivejettison canonly beperformed

with the landing gear handle up and the weight off the aircraft landing gear, or with the armament safety override

switch in the override position. All weapons are jettisoned in a safe condition, however, there is no guarantee fuzes

will not arm during release or detonate on impact. AIM--9 missiles are not jettisonable as normal loading practice

does not install the impulse cartridges in the bomb rack.

On TAV--8B, no selective jettison capability exists from the rear cockpit.

2.28.1.2.1 Selective Jettison Select Knob

The selective jettison select knob on the armament control panel to the left of the station select buttons has rotary

positions STA, STOR, SAFE, CMBT and FUEL, and a center push--to jettison (JETT) pushbutton. The STA and

STOR positions are used in conjunction with the station select buttons. With STA selected, all stores and suspension

equipment hung on BRU--36 bomb racks on selected stations are jettisoned. With STOR selected, jettison is the same

as in the STA position, except that all stores mounted on improved triple ejector racks (ITER) are released while

retaining the ITERs. The CMBT position jettisons all stores, including suspension equipment, suspended from bomb

racks, except that all AIM--9s and suspension equipment are retained. Jettison occurs at 50 millisecond intervals

starting with stations 1, 4 and 7, then stations 2 and 6, and then stations 3 and 5, skipping any AIM--9 station. The

FUEL position jettisons fuel tanks from four stations in pairs, first from stations 2 and 6, then, 50 milliseconds later,

stations 3 and 5. The center JETT pushbutton, when pushed, activates the jettison circuits after the stations and

jettison modes are selected. CMBT/FUEL can be selected before takeoff. If STA/STOR is selected, weapon

programming will be inhibited. The SAFE position prevents any selective jettison.

2.28.1.2.2 Station Select Buttons

The station select buttons are on the armament control panel on the lower left corner of the main instrument panel.

The buttons are numbered 1 through 7 corresponding to the aircraft external stores stations. Pressing a button, or

combination of buttons, selects stations for jettison with the STA and STOR positions of the selective jettison knob.

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The word SEL is displayed in the station window when the station is selected. Nonselected stations display a dash

in their windows.

2.29

MANEUVERING TONE

An aural maneuvering tone based on AOA and Mach is installed to advise the pilot of the existing aircraft flight

characteristics. A 1,600 Hz--10 pps tone (area B, Figure 2-30) is provided when above 0.45 Mach and 225 knots on

AV--8B aircraft and above 0.11 Mach on TAV--8B aircraft. No tone (area A on the chart) denotes a region where full

high speed stop roll inputs will result in positive aircraft response with no concern for departure. The onset of the

1,600 Hz--10 pps tone marks the maximum turn capability of the aircraft and requires near full back stick. Above the

onset of the 1,600 Hz--10 pps tone (area B on the chart), the roll response is sluggish. Large lateral inputs will result

in a correct initial response followed shortly by a roll reversal. If the roll reversal stall warning cue is ignored and

the roll input maintained, a mild rolling departure will occur. Area B is sometimes characterized by a building wing

rock which will end with a departure. This departure can be avoided and the wing rock stopped by releasing back

stick until the tone stops.

Figure 2-30. Maneuvering Tone (with DEP RES)

2.30

WARNING/CAUTION/ADVISORY LIGHTS AND TONES (TAV--8B AND AV--8B DAY ATTACK

AIRCRAFT)

Thewarning/caution/advisory lightsand displayssystem providesvisual indicationsofnormalaircraft operationand

system malfunctions affecting safeoperation ofthe aircraft. The lights are on various system instruments and control

panels in the cockpit (Figure 2-31).

The red warning lights indicate a hazardous condition requiring immediate action. There are two categories of caution

lights, both of which are yellow. The six priority caution lights are located to the left of the upfront control panel and

below the master caution light. They are L FUEL low, R FUEL low, 15 SEC, MFS, BINGO fuel and H₂O. The other

caution lights are on the caution/advisory light panel in addition to the landing gear in--transit lights and the combat

thrust selection light on the water switch panel. All caution lights indicate the existence of an impending dangerous

condition requiring attention but not necessarily immediate action. Illumination of a priority caution light may

require immediate corrective action in certain flight conditions. The green advisory lights indicate safe or normal

configuration, condition of performance, operation of essential equipment, or information for routine purposes. The

advisory lights are on the caution/advisory light panel and on various other panels throughout the cockpit.

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Illumination of most warning lights is accompanied by a warning tone in the headset. The exceptions are the GEAR

and landing gear handle lights. Steady illumination of these indicators will not initiate a warning tone. The (hooter)

tone for most warning lights, except LAW, is a 700 to 1,700 Hz sweep for 0.85 second, with an interruption interval

of 0.12 second. The LAW warning tone is a 1,000 Hz tone with an on and off rate of 2 pulses per second for a duration

of 3 seconds and has a priority over the hooter warning tone for the 3 seconds. The hooter warning tone will be shut

off if the cause for the warning light coming on goes away, or by pressing the MASTER CAUTION light. In addition,

the red threat lights on the warning/threat lights panel have special tones associated with their illumination. The

characteristics of these tones are covered elsewhere. Except the CW NOGO and P NOGO, all caution lights on the

caution/advisory light panel flash when they first come on at a flash rate of three to five flashes per second. Pressing

the MASTER CAUTION light causes these caution lights to go to steady illumination, but they will remain on as

long as the cause for the light coming on exists. Except the L FUEL, R FUEL and 15 SEC lights, the priority caution

lights comeon with asteady illuminationand remainon untilthecauseforthelight comingon goesaway. Allcaution

lights that activate the MASTER CAUTION and all priority caution lights are accompanied by a tone (tweedle dee)

which consists of a 0.3 second, 1,900 Hz signal followed by a 0.15 second, 2,600 Hz signal. The steady (750 pound)

L FUEL and R FUEL lights do not activate the tweedle dee tone. The signals are repeated twice for a total time of

0.9 second. No warning/caution/advisory lights are operational until one of the following occurs: GTS started, engine

started, or compass/lights test switch activated (ground alert). A thermostat will cause the lights to flash if they

overheat on the ground. The thermostat is inoperative in flight and there is no indication if the lights overheat.

Dimming of the warning/caution/advisory lights is covered under Lighting, this chapter.

On TAV--8B 163856 and up, AV--8B 163519 and up, the warning tones (hooter) and caution tones (tweedle dee) for

many of the warning and caution lights are replaced by voice warnings. Refer to Voice Warning in this chapter for

list of lights that have associated voice warnings instead of tones.

2.30.1 Master Caution Light

A yellow MASTER CAUTION light, on the main instrument panel to the left of the upfront control panel, comes

on flashing and is accompanied by the tweedle dee tone when any of the caution lights on the caution/advisory light

panel come on, except for the following: CW NOGO and P NOGO. The MASTER CAUTION light goes off when

it is pressed (reset), will cause all flashing caution lights on the caution/advisory light panel to go from flashing to

steady and will silence the caution/warning tones. For voice warnings, pressing the MASTER CAUTION light will

silence only one warning at a time. The stall warning/maneuvering tones are not silenced.

2.31

WARNING/CAUTION/ADVISORY LIGHTS AND TONES (AV--8B RADAR AND NIGHT

ATTACK AIRCRAFT)

Thewarning/caution/advisory lights (Figure2-31)and displays system provides visual indications ofnormalaircraft

operation and system malfunctions affecting safe operation of the aircraft. The lights are on various system

instruments and control panels in the cockpit. Eleven green warning lights located to the right of the upfront control

panel and below the red MASTER WARNING light indicate a hazardous condition requiring immediate action. Six

green priority caution lights are located to the left of the upfront control panel and below the yellow MASTER

CAUTION light. They are L FUEL low, R FUEL low, 15 SEC, MFS, BINGO fuel, and H₂O. The other green caution

lights are on the caution/advisory light panel in addition to the yellow landing gear in--transit lights and the green

combat thrust selection light on the water switch panel.

All caution lights indicate the existence of an impending dangerous condition requiring attention but not necessarily

immediate action. Illumination of a priority caution light may require immediate corrective action in certain flight

conditions. The green advisory lights indicate safe or normal configuration, condition of performance, operation of

essential equipment, or information for routine purposes. The advisory lights are on the caution/advisory light panel

and on various other panels throughout the cockpit.

Priority caution lights and the MASTER CAUTION light come on flashing when the cause for the light exists (except

for the 750# L/R FUEL lights and the 15 SEC caution light). The 750# L/R FUEL lights come on steady without

illuminationoftheMASTERCAUTIONlight.PressingtheMASTERCAUTIONlightortheMASTERWARNING

light will reset the priority caution lights and MASTER CAUTION light (except 15 SEC caution light). When the

lights are reset the priority caution lights remain on steady as long as the condition continues to exist and the

MASTER CAUTION light goes off. The 15 SEC caution light comes on flashing after 15 seconds if condition still

exists (the MASTER CAUTION light remains off).

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The warning lights come on flashing and are accompanied by a voice warning in the headset. Pressing the MASTER

WARNING or MASTER CAUTION light causes the warning light to go steady and halts the voice warning. The

light will remain on as long as the cause for the light exists.

The green threat lights have special tones associated with their illumination. The characteristics of these tones are

covered elsewhere.

Except the CW NOGO and P NOGO, all caution lights on the caution/advisory light panel flash when they first come

on at a flash rate of three to five flashes per second. Pressing the MASTER WARNING or MASTER CAUTION light

causes these caution lights to go to steady illumination, but they will remain on as long as the cause for the light exists.

A thermostat will cause the lights to flash if they overheat on the ground. The thermostat is inoperative in flight and

there is no indication if the lights overheat. Dimming of the warning/caution/advisory lights is covered under

Lighting. Referto VoiceWarning in this chapterforlist oflights that haveassociated voicewarnings instead oftones.

2.31.1 Master Warning and Master Caution Light

The red MASTER WARNING light on the right side of the upfront control unit comes on flashing and is accompanied

by the appropriate voice warning when a warning light comes on. The yellow MASTER CAUTION light comes on

flashing and is accompanied by the appropriate voice warning/audible tone when any caution light on the right

console caution/advisory light panel comes on, except CW NOGO and P NOGO. Either master light will go out when

it or the other master light is pressed (reset), and will cause all flashing warning or caution lights to go from flashing

to steady. Pressing either master light will also silence the current voice warnings. For voice warnings, pressing the

MASTER CAUTION or MASTER WARNING light will silence only one warning at a time. The stall

warning/maneuvering tones are not silenced. Voice warnings are listed in the following table and are stated twice

(e.g., ALTITUDE, ALTITUDE).

2.31.2 Voice Warnings

TAV--8B 163856 and up, AV--8B 163519 and up. Voice warnings are provided in conjunction with certain

warning/caution lights in place of special tones, Figure 2-32. The voice warnings also replace some other tones not

associated with warning/caution lights. The following table provides a list of voice warnings and the associated

warning/caution light, or in the case of no light, the applicable implication. Refer to Warnings, Cautions, and

Advisories in Chapter 12 for implications of the lights.

All voice warnings are presented twice; e.g., the voice warning associated with the FIRE warning light will be

presented as ENGINE FIRE, ENGINE FIRE. In the case of multiple voice warnings, the highest priority voice

warning is sounded first. Before sounding the next voice warning, the priority list is checked to see if any higher

priority warnings have become active. If so, the appropriate voice warning is sounded. If not, the lower priority

warning is then sounded only if it is still active. The purpose of this mechanization is to keep the pilot informed of

the most important failure as well as keep him from being overloaded with unnecessary voice warnings. Once

CAUTION, CAUTION has been sounded, it cannot be repeated for another sounding for a 5--second period. Once

a FUEL LOW, LEFT, FUEL LOW, RIGHT, or BINGO warning is sounded, it will not be permitted to sound again

for a 60--second period.

2.31.3 Ground Proximity Warning System (Trainer with H4.0, Night Attack and Radar Aircraft)

Ground proximity warning system (GPWS) is a safety backup system that alerts the aircrew of an impending

controlled flight into terrain (CFIT) condition. It operates when the MC is powered on and sensor data is available.

The GPWS option window 4 on the ODU with ALT option selected allows the pilot to disable/enable the system.

A colon in the option window indicates selection. GPWS can be deactivated. Deactivation of GPWS starts a 20

minute timer which automatically activates GPWS when the 20 minutes has expired. GPWS provides warnings of

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

landings, floor altitude violations, limited protection against flight into rising terrain, diving flight depending on

flight stages that include takeoff, cruise, or landing, and Altitude Loss During Recovery (ALDR). ALDR includes

theloss ofaltitudedueto persistency timers,pilot reaction,rolling towings level,g--onset, steadystatediverecovery,

variable safety buffer and clearance altitudes for this warning condition. GPWS also provides for terrain

compensation over downward sloping terrain.

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VOICE WARNING AND ASSOCIATED WARNING/CAUTION LIGHTS

VOICE WARNING

PRIORITY

WARNING

CAUTION

IMPLICATION

NUMBER

LIGHT

ENGINE FIRE

FIRE

—

Same as warning light.

OVERTEMP

—

Same as warning light.

HYDRAULICS

HYD

—

Same as warning light.

FUEL CONTROL

EFC

—

Same as warning light.

FLAP FAILURE

FLAPS

—

Same as warning light.

RIGHT FEED

R FEED

—

Same as warning light.

LANDING GEAR

GEAR

—

Same as flashing GEAR

warning light.

ALTITUDE

LAW

—

Same as warning light.

LEFT TANK

LTANK

—

Same as warning light.

RIGHT TANK

RTANK

—

Same as warning light.

FIFTEEN SECONDS

—

15 SEC

Same as caution light.

BINGO

—

BINGO

Same as caution light.

LIMITER OFF

JPTL

—

Same as warning light.

OBSTACLE

—

Aircraft is at or below the set

obstacle clearance elevation

angle for AWLS.

WATER

—

H₂O

Same as caution light.

FUEL LOW, LEFT

—

L FUEL

Same as flashing caution light.

FUEL LOW, RIGHT

—

R FUEL

Same as flashing caution light.

GENERATOR

GEN

—

Same as warning light.

MANUAL FUEL

—

MFS

Same as caution light.

CAUTION

—

MASTER

A caution light on the

CAUTION

caution/advisory light panel

has illuminated.

ACNIP GO

—

ACNIP BIT passed.

ACNIP FAIL

—

ACNIP BIT passed.

TAV--8B 163856 and up.

Figure 2-32.

Voice Warnings and Associated Warning/Caution Lights

The GPWS is a look--down system with no forward look capability. GPWS uses the radar altimeter as the primary

altitude source with ADC and INS as backup altitude sources when the radar altitude is invalid. GPWS calculates

terrain slope with inputs from the INS and radar altimeter. When radar altimeter information is invalid, the system

switches from operational mode to coast mode for up to 2 minutes. In coast mode, GPWS calculates an estimate of

current aircraft altitude. Coast mode can only be enabled while the aircraft is not transonic and is over a flat surface

(<2° of slope). Warnings can be generated while in coast mode. If there is insufficient valid sensor data, GPWS

transitions to the bypass state. No warnings are generated in the bypass state. Sensor hierarchy defines which

combinations of sensors are required to keep the GPWS valid and providing full protection. When GPWS does not

have sufficient sensors to provide full protection, degraded level of protection is provided (i.e. Coast mode).

All GPWS warnings should be treated as imminent flight into terrain, unless reassessed situational awareness dictates

otherwise. Pilot response to a valid warning should be instinctive and immediate, using the maximum capabilities

of the aircraft to recover until safely clear of terrain.

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2.31.3.1 GPWS Warning Cues

GPWS provides unambiguous directive aural and visual cues to the aircrew for each potential CFIT condition. A

HUD recovery cue, Figure 2-33, indicating the correct direction to recover the aircraft, and voice warnings are

provided. GPWS voice warnings are: ROLL OUT, CHECK GEAR, PULL UP, and POWER. When an excessive

bank angle condition exists, the ROLL OUT aural warning is heard twice every 2 seconds and the visual recovery

cue appears. When an excessive landing or take off sink rate exists, the POWER aural warning is heard twice every

2 seconds and the visual recovery cue appears. When a potential gear up landing condition exists (greater than 60

seconds after takeoff, gear up, altitude less than 150 feet above ground level (AGL), airspeed less than 200 KCAS,

and rate--of--descent greater than 250 FPM), the CHECK GEAR warning is heard once every 8 seconds but the visual

recovery cue does not appear. When the aircraft flies below the floor altitude or an Altitude Loss During Recovery

warning is generated, one of three aural warnings POWER, ROLL OUT, or PULL UP is heard, based on the aircraft

situation. The ALDR alert is heard twice every 2 seconds along with the visual recovery cue. The floor altitude alert

is also heard every two seconds. The aural cue is silenced when corrective action is sensed.

2.31.3.1.1 Recovery Cue

The recovery cue is a steady arrow that is used in conjunction with all GPWS warnings except the CHECK GEAR

warning. The on/off condition of the arrow is warning dependent. The arrow displayed on the HUD shows the

direction of the horizon. The recovery cue is displayed over the existing data on the HUD. When the potential CFIT

condition no longer exists, the recovery cue is removed from the HUD. The aural and visual cues come on together

and normally go off independently.

2.32

ON--BOARD OXYGEN GENERATING SYSTEM

The on--board oxygen generating system (OBOGS) provides a continuously available supply of breathing gas for

the crew while the aircraft’s engine is operating. Engine compressor bleed air is cooled and conditioned through the

use of a heat exchanger and is directed to the concentrator.

The concentrator contains an electrical motor, powered by the emergency dc bus, and molecular sieves which remove

most of the nitrogen from the bleed air and provides an oxygen rich gaseous mixture for the pilot. A temperature

sensor, upstream of the oxygen concentrator, turns on the OXY caution light if the bleed air temperature exceeds the

design operating limits. The oxygen rich gaseous mixture is then routed to a plenum, located in the cockpit, where

the temperature is stabilized and a limited supply is stored for use during peak demands.

2.32.1 OBOGS Monitor

The OBOGS monitor is mounted on the left side of the seat bulkhead in the forward cockpit. The monitor

continuously monitors oxygen concentration and initiates the OXY caution in the event of OBOGS system failure.

The CRU--99 digital monitor has two test buttons; a test plunger and momentary test button. The monitor performs

a power--up BIT during a 2 minute warmup period and periodic BIT on 60 second intervals during normal operation.

No indication of power--up BIT or periodic BIT is provided. The monitor can be tested using the test plunger by

momentarily pressing the test button.

2.32.2 Oxygen Switch

The OBOGS system is turned on and off by an oxygen switch located on the pilot services panel on the left console.

The switch is labeled OXY and OFF.

2.32.3 Rear Oxygen Knob (TAV--8B)

The rear oxygen switch is located on therear pilot services panel, and controls OBOGS airflow to therear pilot when

the OBOGS is turned on by the forward oxygen switch.

2.32.4 Oxygen Breathing Regulator

Thepilot’schestmountedoxygenbreathingregulatordeliverstheoxygenenrichedairtothepilotatpositivepressure,

the limits of which increase with altitude. It is designed to interface with the hose assembly which connects with the

pilot’s survival kit oxygen disconnect.

2.32.5 Oxygen Caution Light

An OXY caution light, on the caution/advisory light panel, comes on if the bleed air temperature is too high or the

oxygen concentration level is too low.

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Figure 2-33. GPWS HUD Warning Symbology

2.32.6 Emergency Oxygen

An emergency supply of gaseous oxygen is contained in an emergency oxygen bottle located in the survival kit.

During emergency operation, emergency oxygen is routed through thepilot’s breathing regulator to the pilot’s mask.

2.32.6.1 Emergency Oxygen Actuator

The emergency oxygen actuator can be actuated by pulling up the green ring on the inboard side of the left thigh

support.

2.32.6.1.1 Emergency Oxygen Supply Gauge

Thepressuregaugeis mounted on thepressurereducing and shutoffvalve on the top of theemergency oxygen bottle.

The gauge has a red refill sector from 0 to 1,800 psi and black full sector from 1,800 to 2,500 psi.

2.33

BUILT--IN--TEST

The built--in--test (BIT) mechanization provided within each avionics subsystem forms the basis for fault isolation.

This provides both the pilot and maintenance personnel with the status of the avionics subsystems. The BIT system

provides the pilot with simple displays of system status without interfering with other essential functions. The

mission computer (MC) displays the subsystem BIT results on the DDI and BIT messages delineate the system

failure/anomaly. Two types of BIT are mechanized, periodic and initiated. Periodic BIT begins functioning upon

equipment power application. It provides a failure detection capability that is somewhat less than that provided by

initiated BIT in that it does not interfere with normal equipment operation. Two forms of BIT derived data are

supplied totheMC.Oneformis validityinformation associatedwith selecteddata. Thesecond formis theequipment

failure information which identifies failed weapon replaceable assembly (WRA). The MC uses these two forms of

BIT data to implement reversion operation and advisories for the pilot, as well as equipment status displays for both

the pilot and maintenance personnel.

ORIGINAL

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2.33.1 Reversion

When the BIT equipment determines, through either periodic or initiated BIT, that a function has exceeded a

predetermined threshold, the data derived from that function is immediately indicated as not valid. The MC, upon

receiving this indication, reverts to the next best available source or to a backup mode of operation. This reversion

is maintained as long as the data remains invalid from the primary source. Automatic tactical reversion mechanized

for flight aids, navigation, air--to--air and air--to--ground weapons delivery are shown in Figure 2-34. When a reversion

from a primary path occurs, the pilot is provided with the appropriate cuing on the HUD and DDI only if the reversion

results in some loss of capability or performance. If no reversion path exists, displayed data is removed.

2.33.2 BIT Display

A MENU selectable BIT display (Figure 2-35) contains the status of all avionics equipment which interface with the

MC. During BIT display, messages are displayed next to each affected equipment legend on the DDI. No indication

(blank) adjacent to equipment legend indicates periodic BIT has detected no faults. Messages displayed as a function

of equipment status are listed as follows:

OFF indicates no multiplex response and no equipment ready signal (equipment not turned on, equipment not

present, power supply failure, etc., and for RWR it could also mean avionics mux failure between the digital

data computer and RWR computer).

Asterisk(*) indicates a single mux bus failure on the ground or double mux bus failure during flight.

WRA fail numeral 1 through 16 indicating which WRA within an equipment group is failed.

TEST indicates initiated BIT is commanded but not completed for the equipment (during AUTO BIT, TEST

next to stores management system (SMS) indicates failures are removed from mission computer because the

SMS does not perform an initiated BIT). When there are SMS failure codes, a post--flight IBIT should not be

performed or SMS failures will be cleared from the MC. Inspect failures through the MAINT mode. With H4.0,

SMS will no longer be a part of the AUTO BIT.

DSEL is unique to the GPS indicating the equipment has been deselected. If GPS is deselected during flight

the BIT page will indicate GPS DSEL, until the next power up.

M is unique to the DSS indicating the DSS memory is full.

D is unique to the DSS indicating the DSS has failed check sums.

0 indicates a software compatibility problem and is unique to the SAAHS and ADC BITs. There are two

instances in which a 0 would be considered normal. One instance is the transfer of the MC between any two

of the following platforms: Radar, Night Attack, and Trainer Aircraft without a reload. The other instance

would be a fresh MC load in a Night Attack or Trainer Aircraft. When either of these two instances occurs,

0 may be displayed, but the altitude box does not flash. If this is the case, an Auto BIT should be performed.

If 0 clears, the mission can be continued. If 0 never clears or is not the result of these instances, the mission

should be aborted. A 0 also causes the Altitude Box Symbol to flash on the HUD, this should never be

considered normal.

The following applies with H4.0.

9. NOT LOADED is displayed next to ALMANAC if the miniaturized airborne gps receiver (MAGR) does not

have the GPS Almanac loaded. It is also displayed next to CRYPTO if the MAGR does not have crypto loaded.

10. LOADED is displayed next to ALMANAC if the GPS almanac is loaded into the MAGR. It is also displayed

next to CRYPTO if the MAGR has crypto keys loaded but has not locked onto a GPS signal yet.

11. INCORRECT is displayed next to CRYPTO if the MAGR determines that the loaded crypto keys are incorrect.

12. OK is displayed next to CRYPTO if the MAGR determines that the loaded crypto keys are correct.

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RADAR AIRCRAFT WITH C1+

Figure 2-35. BIT Displays (Sheet 2)

The following applies to H4.0 with tactical aircraft moving map capability (TAMMAC) installed only.

13. OPEN is unique to the Advanced Memory Unit (AMU) indicating the AMU door is open.

14. MAP is unique to the AMU indicating that a MAP formatted card is installed.

15. F1 is unique to the AMU indicating mandatory files are missing from the installed mission card.

16. F2 is unique to the AMU indicating mandatory files are missing from the installed maintenance card.

17. D1 is unique to the AMU indicating the mission card has failed checksums.

18. D2 is unique to the AMU indicating the maintenance card has failed checksums.

19. M1 is unique to the AMU indicating the mission card memory is full.

20. M2 is unique to the AMU indicating the maintenance card memory is full.

21. U is unique to the AMU indicating the Fatigue Tracking User Program is present on the maintenance card but

the program is not running in the AMU.

Note

If U is displayed the operator should initiate an AMU BIT. If the U persists,

the maintenance card should be reloaded on an AMS system.

Pressing the MENU, then the BIT pushbutton on the DDI selects the BIT display which displays the appropriate

message next to the affected equipment. For the SAAHS, of the 17 possible WRA failures only 10 can be displayed

at a time on TAV--8B (with ONMI 7.1) and Day Attack aircraft or 14 on Radar and Night Attack aircraft. If more

failures are reported than can be displayed only the first 10/14 are displayed (any combination totaling not more than

33 spaces).

ORIGINAL

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Figure 2-36 contains a listing of WRA failure numerals with the associated equipment display nomenclature. In

addition, a brief description of system/aircraft degradation is given for each failed WRA.

All WRA failures which are present when the BIT display is selected or were present for 2 minutes before selection

of the BIT display and all MUX bus failures are displayed on the DDI. Limited duration, non--repeatable WRA

failures which were not present for 2 minutes are not stored in the mission computer.

WRA and mux bus failures are reset (removed from mission computer storage) by performing an AUTO BIT (ground

only) or an initiated BIT (IBIT) on the affected equipment. An exception to this is the SAAHS. The flight control

computer stores SAAHS failures and the mission computer displays their status when the BIT display is selected.

These failures can be reset only by performing IBIT or by placing the AFC switch to RESET. All SMS failures are

cleared from the mission computer by an AUTO or initiated BIT. With H4.0, SMS will no longer be a part of the

AUTO BIT. Failures which are detected by periodic BIT will be displayed again as they are detected; however, failed

systemswhicharetestedonlyonstart--upwillhaveallfailureindicationsclearedandwillnotbere--testedbyperiodic,

initiated, or AUTO BIT. MSC IBIT can be commanded in aircraft with H4.0. Once an MSC BIT is commanded, the

MSC goes offline. Since the MSC is offline, an MSC IBIT cannot be stopped unless power to the MSC is secured.

Backup displays produced by the display computer are displayed if the display computer is functioning correctly.

2.33.3 BIT Reporting

During postflight it is the pilot’s responsibility to record (before engine shutdown) and report to ground maintenance

BIT and AUTO BIT equipment fail indications. This reporting can avoid loss of inflight failure indications and the

need for ground maintenance to apply electrical power to obtain BIT and AUTO BIT fail indications. The velocity

reasonableness BIT failures (ADC 7, INS 2, and GPS 3) are never displayed on the deck even if the failures lasted

more than 2 minutes. They are designed to provide the pilot with additional situational awareness while inflight.

2.33.4 BIT Initiation

In addition to displaying equipment BIT status, the DDI is used to command initiated BIT. Those avionics sets

identified by thelegends on the BIT display periphery have an initiated BIT capability. Thepilot commands initiated

BIT by pressing the adjacent button. The status messages are displayed as required as each equipment set enters,

performs, and completes its BIT routine.

During the AMU BIT, the AMU pauses the Fatigue Tracking User Program. After the AMU completes BIT, the MSC

willwait15secondstoverifythattheFatigueTrackingUserProgramhas restarted.Iftheuserprogramis notrunning,

the MSC will automatically command another BIT (up to 6 times total). Therefore, the AMU BIT process may vary

in length (anywhere from 18 to 108 seconds). If the user program fails to start 15 seconds after the last AMU BIT,

U will be displayed next to the AMU legend on BIT page 2.

To perform a simultaneous initiated BIT on all systems except INS, RWR, SAAHS, SMS, and MSC (H4.0 only),

the AUTO button is pressed. When the AUTO button is pressed, all failure codes are cleared for all systems. The INS,

RWR and SAAHS failure codes may appear as each system performs its periodic BIT. SMS will display TEST for

approximately 45 seconds, allowing the SMS time to report the results of its periodic BIT. With H4.0, SMS will no

longer be a part of the AUTO BIT.

Note

D The SAAHS has 17 WRA failure codes. However, only 10 can be displayed

at any given time on TAV--8B (with OMNI 7.1) and Day Attack aircraft or

14 on Trainer (with H4.0), Radar, and Night Attack aircraft (* and 0--13 or

any combination totaling not more than 33 spaces). If more than 14 WRAs

are reported, only the first 14 (or 33 spaces total) will be displayed.

D When TAMMAC is installed, the AMU has 19 different WRA failure

codes. There is a maximum of 33 spaces available for display of failures

on the AMU line. If more than 33 spaces worth of failures are being

reported, only the first 33 spaces will be displayed.

2-123

ORIGINAL

A1-AV8BB--NFM--000

FAILURE NUMERAL

EQUIPMENT

INDICATION TO PILOT

(WRA)

ACNIP

ACNIP WRA failure

Software compatibility problem

Invalid air data parameter(s) removed from HUD

AOA removed from HUD

TAS not available for display on HUD when A/G selected

ADC

Altitude removed from HUD

Magnetic heading invalid

Velocity Reasonableness Test Failure

OPEN

The AMU door is open

MAP

A MAP formatted card is installed

AMU WRA failure

High Speed Interface Bus (HSIB) failure

Mission Card failure

Mission Card format improper

Mission Card type improper

Maintenance Card failure

Maintenance Card format improper

AMU

Maintenance Card type improper

AMU and PC Card Ambiguity

Mission Card Failed Checksum

Required mission files are missing

Mission Card memory is full

Maintenance Card failed checksum

Required maintenance files are missing

Maintenance Card memory is full

Fatigue Tracking User program is not running but present

Low band receiver failure

High band receiver failure

ASPJ

Processor failure

Low band transmitter failure

High band transmitter failure

A1, Interface Assembly

A2, Switch Assembly

A3, Modem, A Assembly

Reserved for Modem B Assembly

ATHS

A5, Modem C Assembly

Reserved for Modem D Assembly

A10, 1553 and Processor Assembly

A12, 1553 I/O Assembly

A13, Power Supply and Discrete Assembly

AWLS

AWLS receiver failure

BCN

Radar beacon R/T failure

CMDS

CMDS WRA failure

CNIDC

CNI data converter failure

COMM

Radio 1 failure

Radio 1 antenna system failure

COMM

Radio 2 failure

Radio 2 antenna system failure

Figure 2-36.

BIT Failure Indications (Sheet 1 of 3)

ORIGINAL

2-124

A1-AV8BB--NFM--000

FAILURE NUMERAL

INDICATION TO PILOT

EQUIPMENT

(WRA)

DC 1 failure

DC 2 failure

DDI failure (fwd cockpit)

DDI

DDI failure (aft cockpit)

DMC

Digital map computer failure

High speed interface bus (HSIB) failure

DMT

DMT inoperable

Data storage set failed

Incorrect DSU load

DSS

DSS memory is full

DSU failed checksum

DVMS

Digital map computer failure

Digital memory unit failure

EMS

Engine monitoring unit failure

FLIR electronics unit failure

FLIR

FLIR sensor failure

FLIR power supply failure

GPS receiver fail

GPS battery fail

GPS

Velocity reasonableness test failure

GPS antenna fail

Cockpit HUD failure

HUD

Aft cockpit HUD failure

Interference blanker failure

IGV

Inlet Guide Failure

IFF R/T failure

IFF

KIT--1A failure

IFF antenna system failure

Automatic reversion to AHRS mode

INS

Velocity reasonableness test failure

Left MPCD failure (fwd cockpit)

MPCD

Right MPCD failure (fwd cockpit)

RALT

Radar altimeter R/T failure

Radar target data processor

Transmitter

Receiver/Exciter

Computer power supply

Antenna

Antenna electronics

RDR

Transmitter flow low (indicates low liquid coolant)

Waveguide pressure low

Weight--on--wheels/inflight disagree

Radar hardware or aircraft launch discrete to Radar not

AMRAAM compatible

AMRAAM data link RF power test fail

System failure detected, run I--BIT to isolate fault

Figure

2-36.

BIT Failure Indications (Sheet 2)

2-125

ORIGINAL

A1-AV8BB--NFM--000

FAILURE NUMERAL

INDICATION TO PILOT

EQUIPMENT

(WRA)

RWR computer inoperable

Special receiver inoperable

Integrated antenna array inoperable

RWR

Quadrant receiver at 315° inoperable

Quadrant receiver at 225° inoperable

Quadrant receiver at 135° inoperable

Quadrant receiver at 45° inoperable

Software compatibility problem

Invalid mode or function inoperable

Loss of pitch or roll or yaw function

Loss of coordinated turn function

Loss of control stick steering or emergency disengage

Loss of forward pitch stab aug in approach

Loss of pitch stab aug

Loss of forward pitch stab aug in approach

Loss of roll stab aug

SAAHS

Loss of roll stab aug

Loss of rudder trim and yaw stab aug

Loss of roll/yaw interconnect

Loss of particular switch function, or SAAHS switches off

during BIT

Loss of auto pitch trim or manual trim input

Loss or auto roll trim or manual trim input

Forward lateral accelerometer or roll rate gyro failed

Static inverter or contactor failed

SMS computer failure

Armament control panel failure

Station 1 controller failure

Station 2 controller failure

Station 3 controller failure

SMS

Station 4 controller failure

Station 5 controller failure

Station 6 controller failure

Station 7 controller failure

Aircraft wiring

TACTS

AISI failure

TCN

Tacan R/T failure

NOTES:

AV--8B 163853 and up.

TAV--8B, AV--8B Day Attack.

TAV--8B, AV--8B 161573 through

164547.

AV--8B 164549 and up.

AV--8B 165384 and up; also AV--8B 161573 through 165383, TAV--8B 162963 through 164542 after

AFC--354 Rev A/Part 2/Part 3.

AV--8B 165354 and up.

AV--8B 165305 and up; also AV--8B 163853 through 165006 after AFC--326/Part 3.

H4.0 only.

H4.0 with TAMMAC installed.

TAMMAC not installed.

Figure 2-36. BIT Failure Indications (Sheet 3)

ORIGINAL

2-126

A1-AV8BB--NFM--000

During AUTO BIT a mixture of tones is heard in the headset. These tones result from simultaneous BIT checks of

the ACNIP and the two VHF/UHF receiver--transmitters. The ACNIP tone consists of three 2.5 KHz to 3.5 KHz beeps

for approximately 6 seconds, followed for approximately 2 seconds by two cycles of the LAW tone and one cycle

of the hooter tone. Each receiver--transmitter BIT produces a series of 960 Hz on/off tones while BIT is in progress

and a steady 960 Hz tone after BIT completion. The total BIT time for the transmitter--receivers is typically from 3

to 6 seconds. The total time for the mixture of tones heard during the AUTO BIT varies from approximately 3 to 10

seconds. The sound heard will vary in accordance with the volume control settings of the above three pieces of

equipment. AUTO BIT causes the INS and CIP/AUT caution lights to come on for a short period of time. At times

this period will be long enough to trigger the MASTER CAUTION light. This is normal operation.

For subsystems which require operator participation for initiated BIT thoroughness, a maintenance BIT is provided.

The AUTO, ARBS, MSC (H4.0 only) and MAINT BITs are usable only on the ground and must be performed with

the aircraft stationary.

Note

Performing an AUTO BIT during taxi can cause false failure codes.

The STOP button allows the pilot to stop initiated BIT or AUTO BIT at any time. The same effect is also achieved

by pressing MENU. With either method, any test in progress stops and the equipment returns to normal operation.

The COMM 1 and COMM 2 BIT cannot be stopped once initiated, the TEST legend may be removed from the DDI

but tones will be heard as BIT is performed. The GPS test cannot be stopped once initiated. With H4.0, BITs stop

if they complete, or either MAINT or STOP is pressed. BITs continue even if the BIT page is left. With H4.0, if a

system is unable to complete its BIT test, a ? is displayed next to the system’s label. The following legends are not

displayed on the BIT display during flight: MAINT, AUTO, ADC, INS, SAAHS and MSC (H4.0 only).

Systems have to be turned on and warmed up for initiated BIT. Warm up times are: RALT, 3 minutes; AWLS, 15

seconds; IFF, TCN, and beacon (BCN), 30 seconds; and SAAHS has a 2--minute rate gyro run up time. If BIT is

commanded before the system is warmed up the display will indicate TEST until warmup is completed.

Note

If a CNIDC BIT failure occurs, the status of critical IFR flight equipment

(TACAN, RADALT, BCN, IB, IFF, AWLS, and ACNIP) will not be

known.

On AV--8B 164549 and up, an Operational Readiness Test (ORT) is performed on the radar whenever the aircraft is

powered up. The ORT is nearly the same as an initiated BIT, however, it cannot be stopped once started. The ORT

is automatically initiated whenever the radar has been turned off for more than 7 seconds and is powered up. This

interval defines a cold start. When the radar is turned on, it takes 3.5 to 4.0 minutes for warmup and ORT. The ORT

is performed concurrent with the warmup process and begins after 30 seconds.

During ORT and initiated BIT with weight--on--wheels, the radar transmitter is tested with the radiation path into a

dummy load to protect the ground crew from radiation hazard. Transmitter tests take about 30 to 45 seconds to

complete, however the tests are inhibited if STBY or EMCON are selected during warmup. The radar will not

function until the transmitter tests are complete. If the aircraft takes off in EMCON the tests will not be completed

until EMCON is deselected. The pilot has no indication that the radar is not ready.

On AV--8B 161573 through 162973, auxiliary communication, navigation, identification panel (ACNIP) initiated

BIT is also initiated each time the ACNIP mode switch is positioned to MAN. On TAV--8B, AV--8B 163176 and up,

the battery switch must be in the alert mode to enable the ACNIP initiated BIT. A one second warning tone (hooter)

indicates the ACNIP passed BIT. On TAV--8B 163856 and up, AV--8B 163519 and up, ACNIP BIT pass is indicated

by an ACNIP GO, ACNIP GO voice warning and ACNIP BIT fail is indicated by an ACNIP FAIL, ACNIP FAIL

voice warning.

2-127

ORIGINAL

A1-AV8BB--NFM--000

Note

If emission control (EMCON) is selected, initiated BIT for the radar

altimeter and IFF are inhibited, also the radar transmit function on AV--8B

164549 and up.

Initiated BIT procedures are described in the systems coverage throughout this manual where applicable. With a

combined knowledge of reversion operation (Figure 2-34) and WRA failure numerals (Figure 2-36), it is sometimes

possible for the pilot to determine the extent of system/aircraft operational degradation due to avionics equipment

failure/anomaly.

On aircraft afterAFC--392, theEnhanced VariableInlet GuideVane Control System (EVICS)conducts an automatic

BIT check after engine start. The BIT check verifies proper operation of the EVICS, Hydro--Mechanical Unit (HMU)

and T1 temperature sensor. Component failure is indicated by IGV 1 on the BIT 2 Page.

2.34

DDI MISSION COMPUTER OFP DISPLAY (OMNI 7.1 AND C1+)

The ID number of the current operational flight program (OFP) load for the mission computer can be determined by

selecting the maintenance display on the DDI. The maintenance display is selected by the following procedure:

1. Select BIT display from MENU.

2. Select MAINT from BIT display.

With the maintenance display selected, the current mission computer OFP is displayed above the MENU button

legend at the bottom center of the display.

2.35

SOFTWARE CONFIGURATION PAGE (H4.0 ONLY)

The software configurations of certain systems can be viewed by accessing the configuration displays (CONF1 or

CONF2) from the main menu display by pressing CONF. See Figures 2-37 and 2-38. The MSC automatically

performs a configuration check with the peripherals shown on the configuration pages. If a fault such as an incorrect

software load is detected for one of the peripherals during the MSC power up cycle, the appropriate configuration

page with the faulty item shall be displayed on the left MPCD with the offending software load flashing and ACK

in window 1 of theODU. When ACK is selected thesoftware identifierof theoffending item is lined out, as depicted

in Figure 2-37. Items listed on the software configuration pages that are not powered up or are not applicable to that

typeofaircraft, areleft blank. IfTAMMAC is installed in theaircraft, NOT COMPATIBLE may appearflashingnext

to the digital map computer (DMC) legend. This indicates a map theater load or symbol set that is not compatible

with the AV--8B. This can be corrected by loading an AV--8B theater from an AMU MAP card.

Note

Flight with a software configuration fault is prohibited.

2.36

POSTFLIGHT DATA RETRIEVAL

ThemissioncomputerstoresWRAfailuresforpostflightretrieval.Toreportthesefailures,theBITdisplay isselected

before engine shutdown and the failures are recorded for use by maintenance personnel.

ORIGINAL

2-128

NATOPS FLIGHT MANUAL NAVY MODEL AV--8B/TAV--8B 161573 AND UP AIRCRAFT (2008) - page 3