hawker 800 fctm 2

1-1Pilot’s Operating ManualOriginal Issue: Feb, 2002

Pro Line 21

Section - IIISYSTEMS DESCRIPTION

Sub-section 11ICE PROTECTION SYSTEM

Table of Contents

Page

GENERAL ................................................................................................11-3SYSTEM OPERATION.............................................................................11-3

ROTARY-CUTTER ICE DETECTOR ....................................................11-4WING SPOTLIGHTS .............................................................................11-4

Figure 1 - Ice Detection Components................................................11-5FLUID STORAGE..................................................................................11-6

Warning Annunciators.....................................................................11-6POWER SUPPLIES...............................................................................11-6

Figure 2 - Airframe Ice Protection System........................................ 11-7ENGINE BLEED AIR ANTI-ICING.......................................................... 11-8

Warning Annunciators.....................................................................11-8Figure 3 - Engine Anti-icing System ................................................. 11-9

ICE PROTECTION - WINDSCREENS.................................................. 11-10ELECTRICAL HEATING SYSTEM......................................................11-10

Figure 4 - Windscreen Electrical Heating ....................................... 11-11POWER SUPPLIES.............................................................................11-12

PITOT, STATIC, RUDDER BIAS andAIRFLOW ANGLE SENSOR HEATING............................................... 11-13

OPERATION........................................................................................11-13POWER SUPPLIES.............................................................................11-14

2 Pilot’s Operating ManualOriginal Issue: Feb, 2002

Sub-section 11ICE PROTECTION

Hawker 800XP Pro Line 21Section III - SYSTEMS DESCRIPTION

Intentionally left blank

3Pilot’s Operating ManualRevision A1: Nov, 2002



GENERAL

An airframe fluid ice protection system is provided for the leading edges of the wings and the horizontal stabilizers. The system is controlled by a single timer switch. An audio warning is given when the system switches off. Location of components in the vestibule area is shown in Figure 2.

Hot air is used to keep the main engine air intake and starter/generator cooling air intake free of ice with electric heating provided for windscreens, pitot heads, forward static plates and stall vanes, rudder bias struts and engine inlet temperature and pressure sensors Pt2 and Tt2.

SYSTEM OPERATION

A WING/TAIL ANTICE timer switch controls an electrically-operated pump for up to 10 minutes. When initially selected, the first minute of operation is at a high flow rate, after which, the system reverts to normal flow. If icing conditions still prevail or are expected, and therefore a further period of operation is required, this should be selected before the timer switch reaches zero.

Using this procedure the system will remain on the normal flow rate, without first delivering a high rate flow and therefore fluid will be conserved. When the timer switch returns to zero, the pump is de-energized and a warning chime sounds via the airplane audio system.

NOTE: At very low temperatures (-28° C or less) ice crystals can exist in the atmosphere, but do not present a hazard. If the airframe ice protection system is used at these low temperatures, the water/alcohol content of the fluid will evaporate, leaving solidified glycol which together with the impinging ice crystals can give the appearance of ice. Use of the airframe ice protection system, under these conditions, is not advisable.

Therefore, operation of the WING/TAIL ANTICE timer switch should be limited to the priming procedures, and additional use in flight only when weather conditions warrant.

MINS10 0

ANTICELO PRESS

ANTICELO QTY

ICEPROT

MWS Panel

Overhead Roof Panel

4 Pilot’s Operating ManualRevision A1: Nov, 2002



De-icing fluid is drawn from the tank through a suction filter to the pump, and then via a pressure filter and check valve to the head compensating valve. The head compensating valve ensures equal fluid pressure at the wings and horizontal stabilizers proportioning units.

Fluid is fed from the head compensating valve to the three proportioning units, one located in each wing, the other between the horizontal stabilizers. A check valve is incorporated in each proportioning unit outlet to prevent back-flow when the system is inoperative. Each proportioning unit splits the main flow down to the requirements of the individual distributor panels. This arrangement makes sure the fluid supply is maintained to the remaining outlets should a pipe become disconnected.

At each distributor panel, fluid is fed through a metering tube into a cavity. From the cavity the fluid passes through a micro-porous plastic sheet and through a titanium outer skin of greater porosity to escape into the atmosphere. Airflow then causes the fluid to spread rearward over the wings and horizontal stabilizer surfaces.

ROTARY-CUTTER ICE DETECTOR (Figure 1)

Formation of ice is detected automatically after takeoff and manual selection of the detector is available for operation on the ground.

Power supplies to the ice detector are fed through the weight-on-wheels switch relay system and controlled by an ICE DET AUTO-OVRD switch. With the switch selected to AUTO, the detector operates when the airplane becomes airborne. Selecting the switch to OVRD by-passes the weight switch relay so that the detector runs on the ground and in flight.

NOTE: The ICE DET switch should be selected to OVRD before taxiing in icing conditions.

The ice detector unit consists of an AC powered motor driving a serrated rotor which rotates in close proximity to a fixed knife-edge cutter.

When ice forms on the rotor, the gap between the rotor and adjacent cutter is filled. The skimming action of the cutter against the ice causes a rise in motor torque which rotates the motor slightly within its mounting. Rotation of the motor actuates a microswitch which connects a DC power supply, via a time delay relay, to illuminate an ICE DETECTED annunciator located on the overhead roof panel. The ice warning is also indicated on the MWS by the illumination of the ICE PROT repeater annunciator.

Pushing an ICE DET TEST button illuminates both annunciators.

The time delay relay maintains the ice warning signal during intermittent rises in motor torque. When ice ceases to form, a spring returns the motor to the normal position, the microswitch opens and after a delay (60 seconds) the warning is cancelled.

WING SPOTLIGHTS (Figure 1)

Two spotlights, one on each wing fairing and controlled by a ICE ON-OFF switch, illuminates the left and right wing leading edges for night visual inspection.

5Pilot’s Operating ManualOriginal Issue: Feb, 2002



ICE DETECTIONSPOTLIGHT

(Same Right Side)

ROTARY-CUTTERICE DETECTOR

ICEDETECTED

ICEPROT

Figure 1Ice Detection Components

ICE DETAUTO

OVRD

MWS Panel

Overhead Roof Panel

LOGO/ICE

LOGO




FLUID STORAGE

De-icing fluid for the system is stored in a tank with an approximate capacity of 10.04 gallons (38 liters). For a tank indicating full, priming and protection for at least 85 minutes is provided. The fluid contents indicator on the right side console reads FULL above 8.19 gallons (31 liters), and reads EMPTY when there are approximately 12 minutes protection still available.

A tank filler cap is accessible from inside the airplane forward of the main entry door. After filling a completely empty system, the vent valve, located below the tank filler, should be pushed for 10 seconds to bleed the pump.

NOTE: The vent valve must not be operated while the pump is running.

Warning Annunciators

With the pump running, system low pressure is indicated by the illumination of an amber ANTICE LO PRESS annunciator on the overhead roof panel and the MWS ICE PROT flashing repeater annunciator.

Fluid low quantity is indicated by the illumination of an amber ANTICE LO QTY annunciator on the overhead roof panel and the MWS ICE PROT repeater annunciator flashing. When these warnings occur, 30 minutes of fluid usage remains.

POWER SUPPLIES

Electrical power distribution to the equipment is as follows:

Rotary-Ice Detector .......................................Busbar XS 2

Ice Warning Annunciators .............................Busbar PS2

Left Wing Inspection Spotlight.......................Busbar PS1

Right Wing Inspection Spotlight ....................Busbar PS2

ICEPROT

ANTICELO PRESS

ANTICELO QTY

Overhead Roof Panel

MWS Panel




Figure 2Airframe Ice Protection System




ENGINE BLEED AIR ANTI-ICING

Air is bled from two stages of the engine compressor to provide supplies for:

• Nacelle inlet cowl anti-icing

• Airplane services

An ENG ANTICE ON-OFF switch, located on the roof panel ice protection section, is provided for each engine. With either or both switches selected to ON, an ICE PROT SELECTED annunciator on the MWS panel is illuminated.

Each switch controls a servo-operated anti-icing on-off valve. When ON is selected, the following events occur:

• The anti-icing valve opens and high pressure air is bled from the HP compressor and ducted forward to anti-ice the nacelle inlet cowl.

• Electrical power is supplied, via the fuel computer switch when set to AUTO, to the Pt2 and Tt2sensor probe heaters located in the inlet.

• In flight, the engine digital computers are reset to a schedule that incorporates a raised idle rpm to compensate for the effect on thrust.

• The temperature provided by the A panel windscreen heating film is raised from the normal setting to ensure adequate anti-icing performance.

Warning Annunciators

With the ENG ANTICE switched ON, low pressure flow into the inlet cowl is detected by a pressure switch set at 6 psi and indicated by the illumination of the MWS annunciators ENG A/ICE and ICE PROT repeater. Full details of the bleed air anti-icing system are contained in Sub-section 2 ENGINES.

ICE PROTSELECTED

ENG ANTICE

ON1 2

OFF

ENG 1A/ICE

ENG 2A/ICE

Overhead Roof Panel

MWS Panel




ENG ANTICE

1 2

OFF

PE BUSBAR

PICCOLO TUBE

2 SEC

MWS DIM BUS

ON

ICE PROTSELECTED

ENG 1A/ICE

Figure 3Engine Anti-icing System

TO Pt2 and Tt2HEATING CIRCUITS

TO ENGINE DIGITALCOMPUTER IDLESCHEDULE

PRESSURESWITCH

FROM ENGINE HP BLEED

VALVEANTI-ICING

6 PSI

DELAY




ICE PROTECTION - WINDSCREENS

The two forward facing curved windscreens ('A' screens) and the left and right forward sidescreens ('B' screens) are anti-iced and anti-misted by electrical heating.

ELECTRICAL HEATING SYSTEM

Power for windscreen and sidescreen heating is supplied from two 208V, frequency wild, three phase alternators, one driven from each main engine. Each alternator is controlled by an associated ALTERNATOR 1 (2) ON/OFF switch. The alternator driven from No. 1 engine normally powers the left windscreen and the right sidescreen; the one driven from No. 2 engine, the right windscreen and the left sidescreen. If an alternator fails, the other automatically supplies both windscreens, but both sidescreens are disconnected. Alternator failure is indicated by the illumination of an associated ALTRFAIL 1 (2) annunciator and the MWS ICE PROT repeater annunciator.

The two forward facing panels of the windshield each incorporate a gold film heating element. Powersupplies, from the alternator to the elements, are controlled by SCREEN HEAT ON-OFF switches (L or R). With SCREEN HEAT ON, the panel temperature, detected by integral sensing elements, is regulated by thermal controllers, one for each windscreen.

In the event of overheat occurring in a panel, a related SCREEN OVHT annunciator and the MWS ICE PROT repeater annunciator will illuminate. At the same time, a relay operates to disconnect the power supply to the overheating element.

When the airplane is on the ground or in flight without ENG ANTICE selected, the windshields are heated to a lower temperature setting. In flight, with ENG ANTICE selected, the temperature is controlled at a higher value.

L SCREENOVHT

R SCREENOVHT

ALTR 1FAIL

ALTR 2FAIL

SIDE SCRNOVHT

ICEPROT

OFF

ALTERNATOR1 ON 2

SCREENHEAT1 ON 2

OFF

Overhead Roof Panel

MWS Panel

-11Pilot’s Operating ManualOriginal Issue: Feb, 2002



Figure 4Windscreen Electrical Heating




POWER SUPPLIES

Power distribution to the equipment is as follows:

PE busbar supplies:

• ALTR 1 (2) FAIL annunciators.

• L (R) SCREEN OVHT annunciators.

• SIDE SCRN OVHT annunciator MWS ICE PROT repeater annunciator SCREEN HEAT L ON/OFF control PS2 busbar supplies.

• SCREEN HEAT R ON/OFF control.

No. 1 engine alternator supplies:

• Left windscreen panel heat normal power supply.

• Right sidescreen panel heat supply.

No. 2 engine alternator supplies:

• Right windscreen panel heat normal power supply.

• Left sidescreen panel heat supply.




PITOT, STATIC, RUDDER BIAS STRUT and AIRFLOW ANGLE SENSOR HEATING

Ice protection in the form of electrical heating is provided for the following:

• Two pitot heads, one located each side of the forward fuselage.

• Two forward static plates, one located each side of the nose section.

• Two rudder bias struts, connected to the rudder quadrant.

• Two airflow angle sensors, one located each side of the forward fuselage.

OPERATION

Each pitot head contains an electrical heating element controlled by a PITOT/VANE HEAT L or R ON-OFF switch. Each switch also controls one element of a double element heating muff installed on each of the two rudder bias struts.

L & R PITOT HTR FAIL annunciators illuminate with the MWS ICE PROT repeater annunciator flashing whenever a PITOT/VANE HEAT L or R switch is OFF, or when both switches are ON and the current draw by either pitot head element is insufficient.

Annunciator dimming is via the MWS dimmer.

A single ammeter and a L-R selector switch are provided. Selecting L or R connects the ammeter to the associated pitot head heater circuit. With PITOT/VANE HEAT switched ON for at least 1 minute, readings of between 5 and 10 amps indicate satisfactory operation of the pitot heaters only. Actual power consumption depends on the ambient temperature. The rudder bias heaters are not connected to the ammeter.

The left and right forward static plates are electrically heated. The electrical power supply to the heating element of each static plate is via a relay controlled by the PITOT/VANE HEAT R switch, and the weight switch relay system. Heating is only available when the airplane is in flight.

Ice protection for each airflow angle sensor is provided by a vane heater element, and a case heater element. The case heater element is thermostatically controlled.

The power supply to the heater elements is 115 VAC, and is derived as follows:

(1) Two windscreen alternators on line: left sensor elements from No. 1 alternator - right sensor elements from No. 2 alternator.

(2) One windscreen alternator off line and No. 1 and No. 2 inverters on line: elements of both sensors from No. 2 inverter.

(3) One windscreen alternator off line, and either No. 1 or No. 2 inverter off line: elements of both sensors disconnected.

The heating elements of each airflow angle sensor are controlled by an associated PITOT/VANE HEATL or R switch.

A vane heater failure is indicated by the lighting of an associated L or R VANE HTR FAIL annunciator, and the MWS ICE PROT repeater annunciator.




POWER SUPPLIES

DC power supplies to the pump and chime unit are taken from busbar PS2. Supplies to the ANTICE LO PRESS, ANTICE LO QTY annunciators and the MWS ICE PROT repeater annunciator is taken from busbar PE.

The fluid contents indicator is supplied from busbar PE when the airplane is on the ground, and busbar PS2 when in flight. Switching of power supplies is controlled by a weight-on-wheels microswitch.

L VANEHTR FAIL

R VANEHTR FAIL

L PITOTHTR FAIL

R PITOTHTR FAIL

ICEPROT

PITOT/VANE HEATL ON R

OFF

Overhead Roof Panel

MWS Panel

11113

6 9

12

v

L R

PITOT AMPS

0 15


Pro Line 21


Sub-section 8LANDING GEAR

Table of Contents

Page

GENERAL ......................................................................................................8-3CONTROLS - ANNUNCIATORS - INDICATORS..........................................8-3

GEAR POSITIONS......................................................................................8-3Gear Locked Down ...............................................................................8-3Gear Unlocked ......................................................................................8-3Gear Locked Up....................................................................................8-3

STANDBY INDICATIONS............................................................................8-4Main Gear .............................................................................................8-4Nose Gear.............................................................................................8-4

WARNING HORN........................................................................................8-5LANDING GEAR SELECTOR and BAULK OVERRIDE...............................8-6

Figure 1 - Landing Gear Selector and Baulk Override ............................8-6RETRACTION and EXTENSION ...................................................................8-7

Figure 2 - Landing Gear Hydraulic System .............................................8-8MAIN GEAR ...................................................................................................8-9

WEIGHT-ON-WHEELS SWITCHES ...........................................................8-9NOSE GEAR ..................................................................................................8-9

NOSE GEAR BAY DOORS.........................................................................8-9Figure 3 - Main Landing Gear................................................................8-10Figure 4 - Nose Landing Gear ...............................................................8-11Figure 5 - Nose Gear Doors Release Strut ...........................................8-12

POWER SUPPLIES...................................................................................8-12AUXILIARY HYDRAULIC SYSTEM ............................................................8-13

LOWERING SEQUENCE ..........................................................................8-13Figure 6 - Auxiliary Hydraulic System....................................................8-14

WHEELS and BRAKES...............................................................................8-15LOCATION of CONTROLS and INDICATORS .........................................8-15MAIN WHEELS..........................................................................................8-15NOSE WHEELS ........................................................................................8-15WHEEL BRAKES ......................................................................................8-15

Normal System Operation...................................................................8-15Wheel Brake Lever .............................................................................8-15

Figure 7 - Wheel Brake Lever................................................................8-16




Page

Figure 8 - Combined Hydraulic Pressure Indicator and Emergency Wheel Brake Annunciators................................ 8-17

Emergency System Operation with Main System Pressure Exhausted ..................................................... 8-17

POWER SUPPLIES .................................................................................. 8-18NOSE WHEEL STEERING.......................................................................... 8-18

OPERATION ............................................................................................. 8-18Figure 9 - Nose Gear Steering System................................................. 8-19




GENERAL

The airplane has a hydraulically-powered retractable landing gear incorporating nitrogen charged shock absorber struts and nose wheel steering. Each main gear has two wheels and retracts inboard into wheel wells in the fuselage. Each main wheel well has a fairing attached to the landing gear and a hydraulically-operated door.

The nose gear has two wheels which retract forward into a bay with hinged doors and a fairing on the landing gear.

Hydraulic pressure is supplied from the main hydraulic system for the normal lowering and retraction of the gear. An auxiliary hydraulic system is provided for lowering the gear should the main hydraulic system or landing gear selection controls fail.

Details of the hydraulic system are provided in Sub-section 5 - HYDRAULICS.

CONTROLS - ANNUNCIATORS - INDICATORS

The gear position annunciators are located in a pyramid cluster on the center instrument panel to the right of the master warning system panel.

GEAR POSITIONS

Gear Locked Down

• Green annunciators illuminated

NOTE: The red GEAR annunciators are also illuminated when the gear selector lever is not in the down position with landing gear locked down.

Gear Unlocked

• Red annunciators illuminated

Gear Locked Up

• Both green and red annunciators extinguished

N GEAR

N GEAR

L GEAR

L GEAR

R GEAR

R GEAR

Gear Position Annunciators




STANDBY INDICATIONS

Main Gear

Standby main gear downlock indication is provided by an independent circuit connected to greenL GEAR and R GEAR annunciators located on the right side console.

Nose Gear

As the nose gear locks down, a mechanical post indicator extends from the top left of the center control pedestal.

Right Side Console

L GEAR R GEAR

Standby Main Gear Downlock Annunciators

L GEAR R GEAR

A RUDDER BIAS B

Upper Left AreaCenter Control Pedestal

Nose GearMechanicalPost Indicator

OAT

PUSH

Fuel Temp Switch




WARNING HORN

A warning horn sounds if the gear is not locked down when the flaps are extended beyond the 15°position. In this instance, the warning horn cannot be cancelled.

The warning horn will also sound if the gear is not locked down and either thrust lever is closed to obtain between 60% and 70% N1 RPM (nominal) with IAS below 150 kts. In this case, the warning horn can be cancelled by a switch on the forward side of the LH thrust lever. The audible warning is repeated if the second thrust lever is closed after a previous warning has been cancelled and not reinstated. The warning horn system resets when IAS is greater than 160 kts.

LEFT THRUST LEVER < 70%N1, > 60% N1

RIGHT THRUST LEVER > 70%N1, < 60% N1

PE BUSBAR

FLAPS 25° (SELECTED)

FLAPS 45° (NOT SELECTED)

GEAR POSITION SWITCH(MADE WHEN GEAR IS NOT LOCKED DOWN)

HORN

GEARWARNCTL

IAS < 150 KTS (RESETS WHEN IAS >160 KTS)

Gear Warning Horn Diagram




LANDING GEAR SELECTOR and BAULK OVERRIDE

A LANDING GEAR selector lever is provided to control the normal raising and lowering of the landing gear. The lever operates in a two-gated slot in the center instrument panel and is connected to a hydraulic selector valve located in the nose gear bay.

When the airplane is on the ground, a solenoid-operated pawl engages to lock the landing gear selector in the down position. At take-off, when the airplane weight comes off the wheels, the solenoid operates to withdraw the pawl allowing the selector lever to be moved away from the LANDING GEAR DOWN position. The pawl re-engages when the airplane lands.

The airplane has a LANDING GEAR BAULK OVRD PUSH button located next to the LANDING GEAR selector lever. If the Baulk fails to disengage when the airplane is airborne, pushing the button disengages the pawl. While the push button is pushed the landing gear selector can be moved to the up position.

CAUTION: THE LANDING GEAR BAULK OVRD CONTROL COULD BE OPERATED WHEN THEAIRPLANE IS ON THE GROUND. A WARNING THAT THE LANDING GEAR SELECTOR ISNOT IN THE DOWN POSITION IS PROVIDED BY ALL RED AND GREEN GEAR POSITIONANNUNCIATORS BEING ILLUMINATED AT THE SAME TIME.

SHOULD THIS WARNING BE IGNORED, A STEADY HORN AUDIBLE WARNING WILLSOUND IF THE ENGINE START PWR IS SELECTED.

Figure 1Landing Gear Selector and Baulk Override

CENTERINSTRUMENTPANEL




RETRACTION and EXTENSION (Figure 2)

With the weight of the airplane off the wheels, selecting LANDING GEAR up allows main hydraulic system pressure into the gear up pipelines. This pressure operates hydraulic jacks, one on each main gear, and one on the nose gear. Additionally, main pressure is routed via sequence valves and reversing valves to operate the two main gear wheel well door actuators.

Retraction sequence is as follows:

• Main gear wheel well doors open.

• Nosewheel steering isolated (nose wheel self-centers).

• Nose gear doors open.

• Nose gear locks up and doors shut.

• Main gear locks up.

• Wheel well doors shut and lock.

Subsequent LANDING GEAR down selection diverts main hydraulic system pressure to the gear down pipelines.

Extension sequence is as follows:

• Main and nose gears unlock and doors open.

• Main and nose gears lock down.

• Main gear wheel well doors shut but do not lock up.

• Nose gear doors shut.

• Nosewheel steering reconnects.

Main system pressure is dumped when the auxiliary hydraulic system is used to lower the landing gear.

In this case, the wheel well doors are pushed open by the extending main gear and remain open with the gear locked down.




SHUTTLE

AUXILIARY

VALVE

SYSTEMHYDRAULIC

AUXILIARY

SYSTEMHYDRAULIC

AUXILIARY

SYSTEMHYDRAULIC

SHUTTLEVALVE

NOSEGEARACTUATOR

NOTE: For Auxiliary Hydraulic System,See Figure 6.

LANDINGGEARSELECTOR

DUMPVALVE

MAIN GEARACTUATOR

MAIN GEARACTUATOR

SEQUENCEVALVE

SEQUENCEVALVE

DOORACTUATOR

DOORACTUATOR

DIAGRAM SHOWS:LANDING GEAR SELECTED ‘UP’GEAR ‘UP’ PIPELINES

GEAR ‘DOWN’ PIPELINESVENTED TO RETURNDOOR ACTUATORS RETRACTED(DOOR CLOSED)GEAR ACTUATORS EXTENDED(GEAR UP)

FILTER

RESTRICTOR AND IN-LINE FILTERS

NON RETURN VALVE

MAIN SYSTEM PRESSURE TO RETURN

PRESSURIZED

Figure 2Landing Gear Hydraulic System




MAIN GEAR (Figure 3)

Each main oleo-pneumatic leg retracts inwards into a wheel well in the wing. The well is covered with the gear up by a fairing hinged to the wing, and linked to the gear. Additional fairing is provided by a wheel well door, hydraulically-actuated to shut when the gear locks up.

When the gear is selected down, the door opens until the gear locks down, then closes to cover the well. Each main gear is stabilized in the down position by a side stay located between the leg and the main wing structure. The side stay also forms the main locking component in both the extended and retracted positions.

WEIGHT-ON-WHEELS SWITCHES

Weight-On-Wheels (WOW) microswitch clusters are installed on the left and right gear. These switches provide control function to various circuits when the airplane is airborne or on the ground.

NOSE GEAR (Figure 4)

The self-centering nose landing gear retracts forward, and is faired in the up and down positions by two doors hinged to the nose structure and attached to the retracting mechanism. A small fairing attached to the rear of the landing gear completes the closure of the bay when the gear is retracted.

The oleo-pneumatic leg is attached to a fitting on each side of the nose gear bay, and is stabilized in the down position by a drag stay which also forms the main locking component in both the extended and retracted positions. A spring strut maintains the drag strut in the locked position.

The nose gear leg incorporates an attachment for towing purposes. When this attachment is used, the steering must be disconnected. A steering disconnect pin is located immediately under the towing pin hole.

NOSE GEAR BAY DOORS (Figure 5)

The door operating mechanism incorporates a release strut which allows the doors to be opened on the ground for access.

Access to the release strut when the doors are closed is via an aperture between the rear of the doors and the gear leg. The strut assembly consists of a lower and upper strut. To latch the doors closed, the upper strut telescopes into the lower and is retained by a hook engaging a pin. The hook is pivoted on and off the pin by a lever which is retained in the closed position by a spring-loaded latch.

When the lever is open, or not latched closed, a microswitch illuminates the N GEAR red annunciator to indicate the doors are either open or not correctly latched.

M6951_0 HA00C 017046AA.AI

N GEAR

N GEAR

R GEARL GEAR

CENTER INSTRUMENT PANEL

INDICATION OF NOSE GEAR DOORS OPEN OR UNLATCHED



Section III - SYSTEMS DESCRIPTIONHawker 800XP Pro Line 21

HYDRAULICACTUATOR

RETRACTION

DOORACTUATOR

Figure 3Main Landing Gear




Figure 4Nose Landing Gear

HYDRAULICRETRACTIONACTUATOR

STEERINGACTUATOR

STEERING DISCONNECT PIN(REMOVED WHEN TOWING AIRPLANE)




POWER SUPPLIES

DC power distribution is as follows:

PE Busbar

Normal landing gear position annunciators (6)

Warning horn control and warning horn

PS1 Busbar

Landing gear lever lock solenoid

PS2 Busbar

Standby landing gear downlock annunciators (2)

Figure 5Nose Gear Doors Release Strut




AUXILIARY HYDRAULIC SYSTEM (Figure 6)

The auxiliary hydraulic system is selected by pulling the AUX HYD SYSTEM handle. This action dumps any pressure in the normal landing gear system lines to the reservoir and isolates the auxiliary system from the main system return line.

It is not necessary to select the gear down in order to lower it. Operating the auxiliary system hand pump directs hydraulic fluid to the gear actuators via lines independent of the normal extension circuit.

LOWERING SEQUENCE

• The nose gear uplock is released, the nose bay doors open and the gear extends.

• The main gear uplocks are released, the wheel well doors are unlocked and the main gear extendspushing the doors open.

• As the nose gear locks down, the mechanical linkage closes the nose bay doors.

• When the main gear has locked down, the wheel well doors are left open.

Although it is not necessary to select the gear down when using the auxiliary system, it is recommended to do so after the gear has locked down, to avoid the possibility of a subsequent retraction of the gear when the airplane is on the ground. Selecting the gear down will also cancel the gear red annunciations.

The annunciations presented below are illuminated when the gear has been lowered via the auxiliary system, but the selector lever has not been moved to the down position.

M6951_0 HA00C 017046AA.AI

N GEAR

N GEAR

R GEARL GEAR

CENTER INSTRUMENT PANEL

INDICATION OF NOSE GEAR DOORS OPEN OR UNLATCHED




ON/OFF VALVE

LANDINGGEARSELECTOR

DUMPVALVE

RESERVOIR

LEVEL INDICATOR

FILTER

MAINGEARACTUATORS

SHUTTLEVALVE

KEY

AUXILIARYSYSTEM

MAINSYSTEM

SUCTION

RETURN

NOSEGEARACTUATOR

CIRCUIT SHOWN WITH SELECTOR HANDLE PULLED

MAINSYSTEM

FLAP CONTROLUNIT

AUX HYDLO LEVEL

Figure 6Auxiliary Hydraulic System

SHUTTLEVALVES

MWS PANEL




WHEELS and BRAKES

LOCATION of CONTROLS and INDICATORS

MAIN WHEELS

The main landing gears each have two identical wheels with tubeless tires. Each wheel contains a fusible plug which releases air from the tire in the event of excessive wheel heat.

Each pair of wheels is mounted on a staggered stub axle so that the outer wheel is slightly forward of the inner wheel. During retraction, the gear twists to bring the inner wheel directly over the outer wheel. This action permits the wheels to be stowed in a wheel well of the smallest possible size.

NOSE WHEELS

The nose landing gear has two identical wheels with tubeless tires rotating on a common axle.

WHEEL BRAKES

Normal System Operation

Main hydraulic system power is used to operate calliper-type disc brakes via Maxaret anti-skid units.

Master cylinders, operated by toe brake pedals through spring-struts, provide straight line and differential braking during normal and emergency operation. With the WHEEL BRAKE lever fully forward, main hydraulic system pressure, backed by the main accumulator, passes to a brake control valve via a main reducing valve. The control valve, in response to movement of the master cylinders, directs the related pressure through modulator units and Maxaret units to each brake unit.

Wheel Brake Lever (Figure 7)

Selection of braking for normal, emergency and parking use is controlled by a WHEEL BRAKE lever moving over a notched rack between a NORMAL mark (fully forward) and a rearward PARK BRAKE section.

An EMERGY position (first rearward notch) is marked in red. The WHEEL BRAKE lever is held in the emergency position and in progressive parking positions by a pawl engaging into the notched rack. A push button on the WHEEL BRAKE lever releases the pawl to permit movement of the lever from the EMERGY or PARK BRAKE positions. The lever can be moved rearwards from NORMAL to EMERGY without releasing the pawl.

Item Location

WHEEL BRAKE lever (Figure 7) Right side of center control pedestal

BRAKES (2) and SUPPLY combined hydraulic pressure indicator (Figure 8)

Center instrument panel below the MWS panel

WHEEL BRAKE EMERG annunciators L and R(Figure 8)

Right of combined hydraulic pres-sure indicator




Further rearward movement of the WHEEL BRAKE lever into the PARK BRAKE range progressively applies the brakes and at 1000 psi, the white WHEEL BRAKE EMERG annunciators illuminate. Full parking pressure is applied when the WHEEL BRAKE lever is in the PARK BRAKE position.

Pressure, supplied by the emergency brake accumulator, is maintained at the brake units by a lever mechanism which operates the brake control valve. A spring strut, initially loaded by the rearward movement of the WHEEL BRAKE lever, operates to reset the control valve when the WHEEL BRAKE lever is released to the NORMAL or EMERGY position.

Figure 7Wheel Brake Lever




A combined hydraulic pressure indicator shows the BRAKE pressures (upper indication) applied to the left and right brake units during normal operation.

The BRAKE indications are electrically operated from pressure transmitters connected into the normal supply lines to the brake units. The indicator also provides the main hydraulic system SUPPLY pressure (lower indication).

Emergency System Operation with Main System Pressure Exhausted

Movement of the WHEEL BRAKE lever rearward into the EMERGY notch effects the change from normal to emergency operation. An emergency reducing valve opens to allow emergency brakes accumulator pressure to enter the brake control valve. Brake pedal movement now results in pressure being applied directly to the brake units. Shuttle valves operate to isolate the normal supply so that the modulator units and Maxaret units are inoperative.

With the emergency system selected, the BRAKE indications continue to show normal system pressure to the brakes (provided there is still pressure in this system) but do not show emergency system braking pressure. WHEEL BRAKE EMERG annunciators with the legends L and R are provided to the right of the combined indicator and provide an indication of applied pressure to each brake unit.

Each annunciator is controlled by a pressure switch connected into the emergency supply line to the related brake unit. During brake pedal operation, the white annunciators are illuminated when the applied pressure reaches 1000 psi and remain on until the pressure is released to below this value.

The WHEEL BRAKE EMERG annunciators also act as a reminder that the anti-skid facility is isolated and that braking is direct.

OFF

221 1

00

BRAKE

23 1

4 OFF

SUPPLY 0

PSI x 1000

L REMERGWHEELBRAKE

Figure 8

Combined Hydraulic Pressure Indicator and Emergency Wheel Brake Annunciators

Left Gear Brake Supply Pressure

Right Gear Brake Supply Pressure

Main Hydraulic System Supply Pressure




POWER SUPPLIES

DC power distribution is as follows:

PE Busbar

WHEEL BRAKE EMERG annunciators

PS1 Busbar

BRAKES applied pressure indicators

NOSEWHEEL STEERING (Figure 9)

CAUTION: A STEERING DISCONNECT PIN MUST BE REMOVED BEFORE TOWING THE AIRPLANE. (Refer to Section 6, Sub-section 1 - GROUND HANDLING)

OPERATION

Hydraulic pressure is provided from the main hydraulic system for operation of the nose wheel steering.

CAUTION: NOSE WHEEL STEERING IS NOT AVAILABLE WHEN THE MAIN HYDRAULIC PRESSURE IS LESS THAN 2300 PSI. THE AUXILIARY HYDRAULIC SYSTEM CANNOT POWER THE STEERING SYSTEM.

When the nose gear is locked down, nose wheel steering is available through a range of 45° left and right of the center line. Steering is controlled from a handwheel located on the left-side console.

Rotation of the handwheel operates the selector valve input via cables and linkage. Movement of the selector valve directs hydraulic pressure to either extend or retract the steering actuator. The nose gear is turned in the required direction by the steering actuator.

When the required degree of turn has been reached, feedback through linkage connected to the landing gear moves the selector valve input to a neutral position. The nose gear stops turning and the selected angle is maintained. A steering on/off valve is operated by the nose gear mechanical indication linkage. The on/off valve is only selected on when the nose gear is locked down. After lift-off the nose gear is centered by the action of cams in the oleo strut.

WARNING: THE STEERING HANDWHEEL MUST BE FREE FROM OBSTRUCTION DURINGGEAR LOWERING. FAILURE TO MAKE SURE THE HANDWHEEL IS FREE TOMOVE MAY RESULT IN THE AIRPLANE VEERING OFF THE RUNWAYIMMEDIATELY ON TOUCHDOWN.

The reason the handwheel must be free from obstruction during gear lowering is that the geometry of the linkage from the handwheel to the selector valve causes the handwheel to rotate while the gear is lowering or retracting. The handwheel must be free to turn and to find its neutral position again prior to the gear locking down.

When the steering on/off valve is selected off, the steering actuator is isolated from the main hydraulic pressure for the purpose of preventing the gear being turned and striking the wheel bay.




Figure 9Nose Gear Steering System

1Pilot’s Operating ManualInitial Issue: Feb, 2002

Pro Line 21


Sub-section 13LIGHTING SYSTEMS

Table of Contents

Page

GENERAL ................................................................................................. 13-3ENTRANCE LIGHTING............................................................................. 13-3

Figure 1 - Main Entry Doorway Lighting and Controls........................ 13-3FLIGHT COMPARTMENT LIGHTING ...................................................... 13-4

INSTRUMENT and PANEL LIGHTING .................................................. 13-4Figure 2 - Flight Compartment Lighting .............................................. 13-5Figure 3 - Flight Compartment Bulkhead Lighting .............................. 13-6

WANDER LIGHT .................................................................................... 13-6SPOTLIGHT ........................................................................................... 13-6CHART LIGHTS ..................................................................................... 13-6STORM LIGHTS..................................................................................... 13-6PAD LIGHTS .......................................................................................... 13-6STANDBY LIGHTING............................................................................. 13-7EMERGENCY LIGHTING....................................................................... 13-7AUXILIARY POWER UNIT PANEL LIGHTING (if installed)................... 13-7FLIGHT COMPARTMENT LIGHTING POWER SOURCES Table 1: ... 13-7PRINCIPAL DIMMER CONTROLS Table 2: ......................................... 13-8ANNUNCIATORS................................................................................... 13-9

GALLEY LIGHTING................................................................................ 13-10Figure 4 - Typical Galley................................................................... 13-10

PASSENGER CABIN LIGHTING............................................................ 13-11GENERAL LIGHTING........................................................................... 13-11READING LIGHTS ............................................................................... 13-11AISLE FLOOD LIGHTING .................................................................... 13-11WARDROBE LIGHTING....................................................................... 13-11VESTIBULE LIGHTING........................................................................ 13-11PASSENGER NOTICES ...................................................................... 13-11STANDBY LIGHTING........................................................................... 13-11CABIN LIGHTING POWER SOURCES Table 3: ................................. 13-12

EMERGENCY LIGHTING (if installed) .................................................. 13-13TOILET COMPARTMENT LIGHTING .................................................... 13-14

Figure 5 - Toilet Compartment.......................................................... 13-14AFT BAGGAGE COMPARTMENT LIGHTING....................................... 13-15REAR EQUIPMENT BAY LIGHTING ..................................................... 13-15MISCELLANEOUS INTERIOR LIGHTING and CONTROLS................. 13-15




Page

Figure 6 - Interior Lighting Control Panel .......................................... 13-15EXTERIOR LIGHTING............................................................................. 13-16

NAVIGATION LIGHTS .......................................................................... 13-16ANTI-COLLISION BEACONS ............................................................... 13-16LANDING and TAXI LIGHTS ................................................................ 13-16WING ICE-INSPECTION SPOTLIGHTS .............................................. 13-16

Figure 7 - Exterior Lighting Control Panel ......................................... 13-16Figure 8 - Exterior Lighting................................................................ 13-17

STROBE LIGHTS ................................................................................. 13-18LOGO LIGHTS (if installed) .................................................................. 13-18BOARDING LIGHT ............................................................................... 13-18PULSE LIGHT....................................................................................... 13-18EXTERIOR LIGHTING POWER SOURCES Table 4: .......................... 13-19




GENERAL

Particular lighting installations may vary from the standard installation. This section provides a typicalinstallation of interior and exterior lighting.

ENTRANCE LIGHTING

Entrance lighting is provided by a step light mounted on the rear face of the forward cabinet and a twin light assembly mounted in the vestibule roof. Lights are installed in the main entry doorway top and bottom step risers. Switching is by an ENTRY LIGHTS switch on an interior lighting control panel located straight across from the entry door, left of the refuel panel.

LIGHTINGFIXTURES

Figure 1Main Entry Doorway Lighting and Controls

ENTRYLIGHTS

AISLELIGHTS

VESTLIGHTS

BRTUP

LIGHTSDIM

GALLEYMASTER

CABINLIGHTSO’RIDE

Main Entry DoorwayInterior Lighting Control Panel

Refuel Panel(Shown Open)




FLIGHT COMPARTMENT LIGHTING

INSTRUMENT and PANEL LIGHTING (Figure 2)

Circuit control for the electroluminescent display panels, instruments and the pedestal lighting is provided by a “successive action” LTS MASTER push button switch on the overhead roof panel and the dimming controls on the overhead roof panel and the left and right glareshields. The LTS MASTER switch incorporates an ON caption, which illuminates when pushed.

The main instrument panel and glareshield panel lighting is provided by electroluminescent display panels individually attached to the instrument panels and the glareshield.

The instrument electroluminescent display panels are controlled by the two rotary dimmer switches located on the left and right glareshield side control panels and labelled INSTRUMENT PANEL - PANELS. The glareshield side and upper electroluminescent display panels are controlled by the two rotary dimmer switches located on the left and right glareshield side control panels and labelled GLARESHIELD PANEL.

Floodlights and emergency lights are mounted on brackets in two places on each lower canopy rail and in four places under the main instrument panel glareshield. The flood lights are controlled by a rotary dimmer switch located on the left and right glareshield side control panels and labelled PANEL LTS - FLOOD. The emergency lights are controlled by a three position switch on the left and right glareshield side control panels and labelled PANEL LTS - EMERG/OFF/STORM.

The panel-mounted instruments have integral lights controlled by the two rotary dimmer switches located on the left and right glareshield side control panels and labelled INSTRUMENT PANEL - INST.

The primary flight displays and the multi-function displays have integral lighting and are controlled by the two rotary dimmer switches located on the left and right glareshield side control panels and labelled INSTRUMENT PANEL - DISPLAYS.

Where a requirement for panel or instrument lighting is desirable and no integral lighting exists, panel-mounted pillar lights are utilized. These are controlled by a rotary dimmer switch located on the left glareshield side control panel and labelled PEDESTAL.

The pilot side console lighting is controlled by a rotary dimmer switch located on the left glareshield side control panel and labelled CONSOLE - PANEL. The copilot side console lighting is controlled by rotary dimmer switches located on the right glareshield side control panel and labelled CONSOLE - PANELand DIGIT.

The overhead roof panel lighting is provided by two electroluminescent display panels individually attached to the rear and forward panels. The roof panel instruments have integral lights. The overhead roof panels and instrument lighting is controlled by two rotary dimmer switches located on the right side of the overhead roof panel labelled PANEL DIM and INSTR DIM.




M5532HA00B995356AA

NOTE:FOR INDIVIDUAL INSTRUMENTPANEL ARRANGEMENTSINCLUDING ELECTROLUMINESCENTPANELS REFER TO CHAPTER 31

INST PANELEMERG LAMPS

INST PANELFLOOD LAMPS

INST PANELFLOOD LAMPS

SIDE CONSOLEFLOOD LAMP

SIDE CONSOLEEMERGENCY LIGHT

(RH SIDESIMILAR)

(LH SIDESIMILAR)

CHART LAMPCHART LAMP

FLIGHT ANNUNTEST

PAD LAMP PAD LAMP

MASTER LIGHT SW

STANDBY COMPASSINTEGRAL LAMP

PANEL DIM

INSTR DIM

STORM LAMPSTORM LAMP

S

F

LANDINGGEAR

DIGIT

FD BARS

GLARESHIELDPANEL

PAD

PANELS INST DISPLAYS

PANEL

MICKEY

FLOOD EMERG

STORM

OFF

PANEL LTS

ATCIDENT

CONSOLE

INSTRUMENT PANEL

+

<

CVRMIKE

C

AB

A(LH SIDESIMILAR)BC

DIM

FD BARSANNUN

BRTGLARESHIELD

PANEL

PAD

PANELS INST DISPLAYS

PEDESTALMICKEY

FLOODEMERG

STORMOFF

PANEL LTS

ATCIDENT

CONSOLEPANEL

INSTRUMENT PANEL

+

<

CVRMIKE

Figure 2Flight Compartment Lighting

LIGHT

CHART LIGHT CHART LIGHT

PAD LIGHT PAD LIGHT

STORM LIGHT STORM LIGHT

STORM LIGHT

FLOOD LIGHT FLOOD LIGHTS

EMERGENCY LIGHT




WANDER LIGHT (Figure 3)

A CHART/WANDER LIGHT three position switch plus a wander light socket is located on the upper inward face of panel DA. The wander light is stowed in the right glove compartment and power is supplied from the PE busbar.

SPOTLIGHT (Figure 3)

A bull’s-eye type spotlight, with universal mounting, is installed on the forward face of the flight compartment left bulkhead with a control switch provided directly above. The primary use of the spotlight is to light the inward face of panel DA. Electrical power is supplied from the PE busbar.

CHART LIGHTS (Figures 2 & 3)

Two chart lights, one for each crew member, are mounted on the crew services panel. The lights are controlled by either the associated dimmer controls, also located on the crew services panel, or the CHART/WANDER LIGHT switch, located on the inward face of the right bulkhead, panel DA. The CHART/WANDER LIGHT switch will override the dimmer controls and turn both lights to full intensity.Electrical power is supplied to the chart lights from the PE busbar.

STORM LIGHTS (Figures 2 & 3)

Storm lighting is provided to give high intensity white light to prevent crew flash blindness during lightning conditions. The lights are installed above the pilot and copilot instrument panels, on the flight compartment right bulkhead and one above each side console. The EMERG-OFF-STORM switches located on the left and right glareshield side panels control all storm lights when STORM is selected. Electrical power for the lights is supplied from the PS2 busbar.

PAD LIGHTS (Figure 2)

Two pad lights are mounted, one on each upper canopy rail, and are controlled by rotary dimmer switches located on the left and right glareshield panels labelled PAD. The left pad light is supplied from the PS1 busbar and the right from the PS2 busbar.

STORM LIGHT

CHART/WANDERLIGHT SWITCH WANDER LIGHT

SOCKET

SPOTLIGHTSWITCH

SPOTLIGHT(DA PANEL LIGHTING)

PANEL DA

Looking Aft

Figure 3Flight Compartment Bulkhead Lighting




STANDBY LIGHTING

Electrical power for standby lighting is available when the PE busbar is energized. The following lights are connected to the PE busbar, and are switched by their normal operating switches to provide standby lighting.

• Standby compass (E2B) light

• Chart lights

• Spotlight

• Wander light

EMERGENCY LIGHTING

Electrical power for the flight compartment emergency lights, positioned on the lower canopy rail and under the instrument panel glareshield, is supplied from the essential busbar PE. The emergency lights provide the vital instrument panel fascia lighting and are controlled by the STORM-OFF-EMERG switches mounted on the left and right glareshields, when selected to EMERG.

AUXILIARY POWER UNIT PANEL LIGHTING (if APU installed)

The translucent APU panel is illuminated from behind by parallel wired bulbs. These bulbs illuminate when the APU master switch is ON.

FLIGHT COMPARTMENT LIGHTING POWER SOURCES

Table 1

PS1 Busbar

Flood lights Left pad lightLeft and right glareshield side and upper display electroluminescent panelsLeft, center and right main panel instrument lightingE2B compass (NORM) Center console pillar and panel lightsCenter console instrument lightingOverhead Roof panel lighting

PS2 Busbar

Storm lightsRight pad light Overhead Roof panel instrument lighting Left, center and right main instrument panel electroluminescent display panels

PE Busbar

Emergency lights Chart lights SpotlightDA panel wander lightAnnunciator Brt/Dim supply 1Annunciator Brt/Dim supply 2E2B compass (EMERG)




PRINCIPAL DIMMER CONTROLS

Table 2

PanelDimmer(voltage)

Lighting ServiceSupplyBusbar

Overhead Roof Panel *INSTR DIM(5 VDC)

Overhead roof panel instrument integral lighting

PS2

*PANEL DIM(5 VDC)

E2B compass lighting (NORM)Overhead roof forward panel lightingOverhead roof main panel lighting

PS1

Left Glareshield PAD(28 VDC)

Left pad light PS1

*GLARESHIELD PANEL(5 VDC)

Left glareshield side and upper electroluminescent panels

PS1

*INSTRUMENT PANELPANELS(5 VDC)

Left electroluminescent panelsCenter electroluminescent panels

PS2

*INSTRUMENT PANELINST

(5 VDC)

Secondary flight displayAngle of attackBrake pressure indicatorOutside air/Fuel temperature indicator

PS1

INSTRUMENT PANELDISPLAYS(28 VDC)

Left PFDLeft MFD

-

PANEL LTS FLOOD(28 VDC)

Left and center flood lights PS1

PANEL LTS EMERG/OFF/STORM

(28 VDC)

Emergency lights or storm lights

PE (emerg)PS2 (storm)

*PEDESTAL(5 VDC)

Control pedestal pillar lightsDisplay panel lighting

PS1

CONSOLE - PANEL Pilot console lighting PS1

* Circuits controlled by LTS MASTER switch through relays.




ANNUNCIATORS

The lighting intensity control for the MWS annunciators is provided by a dimmer and a switch located on the center instrument panel. A MWS DIM FAIL annunciator is positioned at the top of the left instrument panel. The MWS ANNUN test push switch is located on the test panel at the top of the overhead roof panel and will illuminate most of the annunciators at full intensity. For details of the operation of annunciators associated with the master warning system, refer to Sub-section 1 - MASTER WARNING SYSTEM.

The lighting control of the MWS annunciators is by variable dimming effected by a rotary MWS DIM control in conjunction with a NORM - DIM OVRD switch. Should the dimming circuit fail, resulting in loss of light to the annunciator, the NORM - DIM OVRD switch may be selected to the OVRD position; this effectively bypasses the dimming circuit and connects a 28 VDC supply direct to the annunciators. Both power supplies are taken from the PE busbar.

The flight annunciator illumination intensity is controlled by the ANNUN BRT/DIM switch located on the left side glareshield panel. This switch also controls the landing gear and wheel brake emergency annunciator illumination levels.

Right Glareshield PAD(28 VDC)

Right pad light PS2

*GLARESHIELD PANEL(5 VDC)

Right glareshield side and upper electroluminescent panel

PS1

*INSTRUMENT PANELPANELS(5 VDC)

Right electroluminescent panelsCenter electroluminescent panels

PS2

*INSTRUMENT PANELINST

(5 VDC)

Cabin pressure controllerTriple indicator gauge

PS1

INSTRUMENT PANELDISPLAYS(28 VDC)

Right PFDRight MFD

-

PANEL LTS FLOOD(28 VDC)

Right and center flood lights PS1

PANEL LTS EMERG/OFF/STORM

(28 VDC)

Emergency lights or storm lights

PE (emerg)PS2 (storm)

CONSOLE - PANEL - DIGIT Copilot console lighting PS2

PanelDimmer(voltage)

Lighting ServiceSupplyBusbar

* Circuits controlled by LTS MASTER switch through relays.




GALLEY LIGHTING (Figure 4)

An incandescent light strip is installed on the back of the upper shelf of the upper galley section and controlled by the GALLEY UP LIGHTS switch on the galley switch panel. The galley work lighting is provided by a twin fluorescent light assembly to illuminate the working surface and controlled by the GALLEY WORK LIGHTS switch on the galley switch panel.

The circuit breaker panel is located in the galley circuit breaker compartment.

CUSTOM WATER CONTAINER

CUPDISPENSERS

WASTE

COFFEEBREWER

SWITCH PANEL

MICROWAVE

COUNTERSTORAGE

SANDWICHTRAY

DRAWER

UTENSILS

CONDIMENTS

PULL OUT WORK

SURFACE

BOWLS

PLATES

MINIATURES

WINE / SPIRITS STORAGE

MISC.STORAGE

NAPKINS

SODA CANS

ICE / COLD STORAGE

Figure 4Typical Galley

GALLEYUP

LIGHTS

GALLEY SWITCH PANEL

MISCELLANEOUSSTORAGE

GALLEYWORK

LIGHTS

COFFEEHEAT

WATERHEAT

M6356_0HA03C014655AA.AI




PASSENGER CABIN LIGHTING

GENERAL LIGHTING

NOTE: For the position of lights and switches refer to Figure 1 or Figure 3 and for details of powersupplies refer to Table 1.

Passenger cabin general lighting is provided by left and right roof-mounted fluorescent tubes.

Two inverter/ ballast units provide momentary high voltage outputs to four power units to enable the lights to become fully illuminated. The roof-mounted light installations each contain two fluorescent tubes, one tube of each installation being supplied from the left-hand inverter ballast unit and the other from the right-hand inverter ballast unit. This arrangement ensures continuity of lighting in the event of failure of a lighting supply.

READING LIGHTS

An individual reading light and push switch is provided above each passenger position.

AISLE FLOOD LIGHTING

Cabin aisle flood lighting is provided by six flood light units each containing 4 bulbs, which are installed in the cabin forward and rear ankle ducts on the left-side. The aisle flood lighting is controlled by the interior lighting control panel AISLE LIGHTS switch located left of the refuel panel.

WARDROBE LIGHTING

The wardrobe interior lighting is provided by a twin light assembly positioned in the wardrobe upper section. The lights are connected in parallel, supplied from the PS2 busbar, and controlled by an illuminated push-switch mounted on the upper right hand section of the wardrobe.

VESTIBULE LIGHTING

Vestibule lighting is provided by a step light mounted on the rear face of the forward cabinet, and a twin light assembly mounted in the vestibule roof and are controlled by the interior lighting control panel VEST LIGHTS switch located left of the refuel panel.

PASSENGER NOTICES

FASTEN BELTS/NO SMOKING notices are installed in the passenger cabin forward and rear bulkheads. Notice illumination is provided by integrally installed lights, controlled by a SEAT BELTS ON-OFF switch and a three-position NO SMKG ON-OFF-AUTO switch. Both switches are located on the flight compartment overhead roof panel and selection of either switch to ON, will illuminate the appropriate part of the safety notices.

Notice illumination is accompanied by activation of an audio system chime unit. When selected, the NO SMKG switch AUTO position will automatically illuminate the safety notices NO SMOKING characters and operate the chime unit when the nose landing gear is locked down and the relevant busbars are energized.

STANDBY LIGHTING

Electrical power for the passenger cabin and toilet standby lights is supplied from the essential busbar PE. Two standby lights are installed in the passenger cabin roof and one in the toilet compartment.




CABIN LIGHTING POWER SOURCES

Table 3

PE Busbar

Cabin standby lightsToilet standby lights

PS1 Busbar

Cabin floor lights (left-hand inverter)

Reading lights (left-hand inverter)

PS2 Busbar

Cabin flood lights (right-hand inverter)

Toilet lights

FWD and rear luggage bay lights

Reading lights (right-hand PSU's)

Battery 1 (in flight)

Vestibule roof lightStep lightsRefuel panel (DB) lightsAisle lights

NOTE: With the landing gear microswitch in the GND position the following services aresupplied from Battery 1:

• Vestibule roof light.

• Cabin flood lighting. (right-hand inverter)

• Step lights.

• Toilet lights.

• Refuel panel lights.

• Aisle lights.




EMERGENCY LIGHTING (if installed)

The emergency lighting system is powered by a forward and aft battery supply and provides emergency cabin and exterior lighting automatically whenever PE power is interrupted or lost with the capability to be manually overridden.

A three position switch, MAN-ARM-OFF, located on the flight compartment overhead roof panel, controls the DC power from a forward and aft battery power supply.

The forward power supply provides lighting for the forward entrance door EXIT sign, floor lighting for the entry way (in the aft base of the crew cabinet) and four cabin aisle lights. The aft power supply provides lighting for the escape hatch area, over/under right wing, three cabin aisle lights, the L/H forward and aft overhead cabin reading lights and the R/H two middle overhead reading lights.

An amber EMERG LTS OFF annunciator, adjacent to the emergency light switch, will illuminate anytime the switch is in the OFF position and PE is powered.

A sonalert system warns the flight crew whenever the emergency lights switch is in the ARM or MANposition and PE power has been interrupted or lost. The sonalert may be cancelled by pushing the HORIZON WARN CANCEL button located on the pilot instrument panel.

"ARM" - ARMING Emergency LightsPE Power (Battery Master) - ONEmergency Lights Switch - ARMEMERG LTS OFF annunciator - Extinguished

"MAN" - Manual Emergency LightsPE Power (Battery Master) - ONEmergency Lights Switch - MANEMERG LTS OFF annunciator - Extinguished

"OFF" - ARMING Emergency LightsEmergency Lights Switch - OFFEMERG LTS OFF annunciator - IlluminatedPE Power (Battery Master) - OFF

NOTE: Turning PE power OFF prior to turning the Emergency Lights switch OFF soundsa sonalert and illuminates the emergency lights. Select the Emergency Lightsswitch to OFF and turn the PE Power (Battery Master) ON then OFF. This actiondisables the Emergency Lighting System.




TOILET COMPARTMENT LIGHTING (Figure 5)

The toilet compartment is illuminated by six fluorescent tubes mounted in the roof trim to provide a concealed lighting effect when the main airplane power is turned on. Additional lighting is controlled by switching on the toilet lighting control panel located above the mirror. Supply is from the PS2 busbar.

Figure 5Toilet Compartment

LAVLIGHTS

TOILET LIGHTING CONTROL PANEL

READLIGHT

MIRRORLIGHTS

BAGGAGELIGHTS




AFT BAGGAGE COMPARTMENT LIGHTING

A roof-mounted twin-bulb tube is installed in the aft baggage compartment and is controlled by the BAGGAGE LIGHTS switch on the toilet lighting control panel. One bulb of the aft baggage compartment light is illuminated whenever the main toilet lights are on, the other is illuminated together with the toilet bulb whenever the main cabin lights are on.

REAR EQUIPMENT BAY LIGHTING

A light and socket are installed in the roof of the rear equipment bay to give general area lighting and power supply for a plug-in wander light. Control of the switch is by manual selection to ON or OFF and automatic selection to OFF when the door is closed. Two pillar lights supply light for the hydraulic tank level indicators. Electrical 28 VDC supply to both the light and power point socket is taken from busbar PE through an ON/OFF switch on the structure adjacent to the bay door hinge.

MISCELLANEOUS INTERIOR LIGHTING and CONTROLS

The dome/exit light is a dual bulb unit (one general purpose, one emergency) above the MED header panel. The control is from VEST LIGHTS switch on the interior lighting control panel left of the refuel panel.

Figure 6Interior Lighting Control Panel

ENTRYLIGHTS

AISLELIGHTS

VESTLIGHTS

BRTUP

LIGHTSDIM

GALLEYMASTER

CABINLIGHTSO’RIDE

Refuel Panel(Shown Open)




EXTERIOR LIGHTING

NAVIGATION LIGHTS (Figures 7 and 8)

These lights give a high-intensity red (left wingtip), green (right wingtip), and white (tail unit) light. All three lights are controlled by a single switch labelled NAV-OFF on flight compartment overhead roof panel forward extension.

ANTI-COLLISION BEACONS (Figures 7 and 8)

These provide two rotating beams of red light. Each beacon contains two reflector bulbs mounted in tandem on oscillating platforms. Both beacons are controlled by a single switch labelled BEACON - OFF on the overhead roof panel forward extension. A moisture drain is provided by the hollow lens cover retaining bolt when the beacon is installed in the lower (inverted) position. The drain is closed with a grubscrew when the beacon is installed in the upper position. Additional drainage for the lower-mounted beacon includes drain holes in the lens cover forward and aft of the attachment bosses.

LANDING and TAXI LIGHTS (Figures 7 and 8)

Mounted together in each wing leading edge, these lights are sealed-beam units controlled by switches labelled L LANDING R - TAXI - OFF, on the overhead roof panel forward extension. Selecting the switches to the LANDING position will operate the landing lights and switching to the TAXI position will operate the taxi lights.

CAUTION: THE 450 WATT LANDING LIGHTS MUST BE USED ONLY IN FLIGHT CONDITIONS, OR ONLY BRIEFLY ON THE GROUND TO CHECK FUNCTIONING. THE TRANSPARENT PANEL WILL SUSTAIN HEAT DAMAGE.

WING ICE-INSPECTION SPOTLIGHTS (Figures 7 and 8)

A wing ice inspection light is installed in both the left and right wing fairing, and consists of a sealed beam unit assembly, mounted on a bracket with provision for light beam angle adjustment. The lights illuminate the wing leading edges when the LOGO-ICE-OFF control switch, on the flight compartment overhead roof panel forward extension, is selected to ICE.

Figure 7 Exterior Lighting Control Panel

Overhead Roof Panel

Forward ExtensionEXTERIOR LIGHTS

OFF OFF OFF OFF

LOGO/ICE STROBE NAV BEACON LANDINGL R

OFF

TAXI

LOGO




Figure 8Exterior Lighting

NAVIGATION LIGHT (GREEN)

STROBELIGHT

LANDING LIGHTS

TAXI LIGHT

LANDING LIGHT

LANDING LIGHT

TAXI LIGHT

LEFT LOGOFLOOD LIGHT(if installed)

RIGHT LOGOFLOOD LIGHT(if installed)

NAVIGATION LIGHT (RED)STROBELIGHT

NAVIGATION LIGHT(WHITE)

STROBE LIGHT

WING ICE INSPECTION SPOTLIGHT (Same Right Side)

BOARDING LIGHT

LEFT LOGOFLOOD LIGHT(if installed)

1 2




STROBE LIGHTS (Figures 7 & 8)

Strobe lighting provides additional airplane identification to the standard airplane position light presentations.

The strobe lighting system consists of three flashing white condenser discharge strobe lights, three power supply units (PSU's 1, 2 and 3) and a STROBE light control switch. The control switch is located on the flight compartment overhead roof panel forward extension labelled STROBE - OFF.

NOTE: Turn off strobe light when taxiing near other airplanes or when flying in fog or clouds.Standard position lights must be used for all night operations.

A strobe light is installed in each wing tip, and one on the rear fuselage extremity. Each strobe light is connected to its own power supply unit.

All three PSU's are installed in the passenger cabin left side below the passenger seating. In operation the lights flash simultaneously at 60 flashes per minute (±5 FPM at the rated voltage).

LOGO LIGHTS (if installed Figures 7 & 8)

The vertical stabilizer logos are illuminated by lights mounted on the underside of the left and right horizontal stabilizers. Both lights are controlled by a single LOGO-ICE-OFF switch located on the EXTERIOR LIGHTS section of the flight compartment overhead roof panel forward extension.

BOARDING LIGHT (Figure 8)

The ground area and bottom steps of the main entry doorway are illuminated by the boarding light which is also located with the wing ice inspection light in the left wing to fuselage front fairing assembly.

The boarding light is controlled by a switch on panel DA and a second switch on the forward vestibule cabinet, labelled ROOF/STEP. The supply is taken from battery No. 1 so that it may be switched ON without selection of the BATTERY master switch.

PULSE LIGHT

The pulse light is a four channel electrical switching device which connects to the external lighting system of the airplane. The system operates by flashing the landing and nose taxi lights 45 times per minute in a variety of patterns. Thus, creating an illusion of exaggerated motion that other pilots can immediately recognize and avoid.

The pulse light may be utilized any time the pilot desires, although it is recommended that the landing lights are switched to steady rate (full time) when the airplane is within 200 feet AGL at night. The pulse light should not be operated in clouds at night or in close proximity of other airplanes on the ground, due to possible pilot disorientation.

The pulse light is powered by 24/28 VDC with an amp load no greater than 25 AMPS per channel or 600 WATTS per channel.




EXTERIOR LIGHTING POWER SOURCES

Table 4

PE Busbar

Navigation lightsLeft wing landing lightLeft wing taxi lightVestibule/ground lights

PS1 Busbar

Left and right wing inspection spotlights

Strobe lights

Right nose taxi light

PS2 Busbar

Right wing landing light

Right wing taxi light

Anti-collision beacons

Tailplane flood lights

Left nose taxi light

Battery 1

Boarding light

FAA ApprovedRevision A3: Mar 22, 2005

Hawker 800XP Pro Line 21 Airplane Flight Manual

SECTION 2

LIMITATIONS

Table of Contents

Page

KINDS OF OPERATION.................................................................................... 3WEIGHT LIMITATIONS ..................................................................................... 3PERFORMANCE LIMITATIONS ....................................................................... 3

Take-off Weight........................................................................................... 3 Landing Weight .......................................................................................... 3Take-off Field Length .................................................................................. 4

OPERATIONAL LIMITATIONS.......................................................................... 4Altitude ........................................................................................................ 4Maximum Permissible Altitude .................................................................... 4Air Temperature .......................................................................................... 4Wind Component ........................................................................................ 4Runway Slope ............................................................................................. 5Airplane Configurations............................................................................... 5

COMPARTMENT LOADING LIMITATIONS ...................................................... 5LOAD LIMITATIONS.......................................................................................... 5

Center of Gravity Limitations....................................................................... 5Figure 2.1 - LOADING and FLIGHT ENVELOPE - POUNDS/FEET ................................................................................. 6Figure 2.2 - LOADING and FLIGHT ENVELOPE - KILOGRAMS/METERS ..................................................................... 7

ICE PROTECTION LIMITATIONS..................................................................... 9Icing General............................................................................................... 9Airframe Icing .............................................................................................. 9Wing/Tail Antice System ............................................................................. 9Engine Icing ................................................................................................ 9

SEVERE ICING CONDITIONS LIMITATIONS ................................................ 10ENGINE LIMITATIONS.................................................................................... 11

Engine Type .............................................................................................. 11Engine Limitations..................................................................................... 11Approved Engine Oils ............................................................................... 12Engine Oil Consumption ........................................................................... 12Engine Fuel Computer .............................................................................. 12Engine Synchronizer ................................................................................. 12Automatic Performance Reserve (APR) ................................................... 12Engine Instrument Markings ..................................................................... 12Thrust Reversers....................................................................................... 12Engine Ice Protection System................................................................... 13

FUEL LIMITATIONS ........................................................................................ 13Fuel Specifications .................................................................................... 13Fuel Additives............................................................................................ 14Maximum Fuel Temperature ..................................................................... 15

Section - 2LIMITATIONS

FAA Approved Revision A3: Mar 22, 2005


PageMinimum Fuel Temperature...................................................................... 15Fuel Quantity............................................................................................. 15Fuel Loading ............................................................................................. 16Pressure Refueling ................................................................................... 16Fuel System Management ........................................................................ 16

ELECTRICAL LIMITATIONS ........................................................................... 16Battery Limitations .................................................................................... 16Generator Limitations................................................................................ 16Main Engine Starter Duty Cycle................................................................ 16Operation of Electrical Circuit Breakers .................................................... 16

AVIONICS LIMITATIONS ................................................................................ 17General ..................................................................................................... 17HF Radio................................................................................................... 17Electronic Standby Instrument System (ESIS) ......................................... 17Flight Management System ...................................................................... 17Autopilot .................................................................................................... 19VNAV ........................................................................................................ 20EGPWS..................................................................................................... 21TCAS II ..................................................................................................... 21

AIRSPEED LIMITATIONS ............................................................................... 22Maximum Operating Speed ...................................................................... 22Maximum Operating Mach Number .......................................................... 22Maneuvering Speed.................................................................................. 22Wing Flaps Extended/Operating Speed.................................................... 22Procedural Use of Flaps 15° for Descent and Holding ............................. 22Air Brakes ................................................................................................. 23Landing Gear Extended/Operating Speed................................................ 23Bird Strike Speed ...................................................................................... 23

MISCELLANEOUS LIMITATIONS................................................................... 25Air Brakes ................................................................................................. 25Cabin Emergency Overwing Exit .............................................................. 25Cabin High Datum..................................................................................... 25Crew Seats ............................................................................................... 25Inter-compartment Door............................................................................ 25Lift Dump................................................................................................... 25Maneuvering Load Factor Limitations....................................................... 25Minimum Flight Crew ................................................................................ 25Nosewheel Tires ....................................................................................... 25Number of Occupants ............................................................................... 25Pressure Cabin ......................................................................................... 25Rudder Bias .............................................................................................. 25Smoking .................................................................................................... 26System Gage Markings............................................................................. 26Weather Radar.......................................................................................... 26Wheel Brakes............................................................................................ 26

Table 1 - Take-off Weights for Wheel Brakes Waiting Periods .................................. 27




KINDS OF OPERATION

This airplane is eligible for certification in the Transport Category and is approved for thefollowing kinds of operation when the appropriate instruments and equipment required by theairworthiness and/or operating certificate are installed and approved and are in operablecondition.

• Atmospheric icing conditions

• Day and night VFR

• IFR

• RVSM

However, the certificate of airworthiness may restrict this airplane to some other category andto a particular use.

WEIGHT LIMITATIONS

Maximum Taxiing (Ramp) Weight.......................... 28,120 lb (12,755 kg)

Maximum Take-off Weight ..................................... 28,000 lb (12,701 kg)

Maximum Landing Weight ..................................... 23,350 lb (10,591 kg)

Maximum Zero Fuel Weight................................... 18,450 lb (8369 kg)

Minimum Zero Fuel Weight.................................... 14,120 lb (6405 kg)

PERFORMANCE LIMITATIONS

Take-off Weight

Maximum Take-off Weight is limited by the most restrictive of the following:

• 28,000 lb (12,701 kg).

• As shown on the MAXIMUM TAKE-OFF WEIGHT FOR ALTITUDE AND TEMPERATUREgraphs (see Sub-section 5.15).

• The maximum permitted by field length considerations (see Sub-section 5.20).

• The maximum permitted by maximum brake energy considerations (see Sub-section5.20).

• The maximum permitted by obstacle clearance considerations (see Sub-section 5.25).

Landing Weight

Maximum Landing Weight is limited by the most restrictive of the following:

• 23,350 lb (10,591 kg).

• As shown on the MAXIMUM LANDING WEIGHT FOR ALTITUDE AND TEMPERATUREgraph (see Sub-section 5.45).

• The maximum permitted by field length considerations (see Sub-section 5.50).




Take-off Field Length

The take-off weight shall not exceed the maximum permitted by field length considerations asdescribed in Sub-section 5.20 for the restricted range of conditions listed.

When complying with the above, the following conditions shall be met:

• The take-off distance used in the graphs shall not be greater than the length of runway plusthe length of clearway if present, except that the length of clearway shall not be greaterthan one-half of the length of the runway.

• The take-off run used in the graphs shall not be greater than the length of the runway.

OPERATIONAL LIMITATIONS

Altitude

Maximum/Minimum field pressure altitudesfor takeoff or landing............................................13,000 ft and -2000 ft

NOTE: Performance appropriate to the lowest published elevation shall be used when thefield pressure altitude is below the lowest published elevation.

Maximum Permissible Altitude

Maximum permissible operating altitude is 41,000 ft.

Maximum permissible altitude with flaps lowered or landing gear extended is 20,000 ft.

Air Temperature

Maximum

All Flight Regimes .............................................. ISA +35° C

Ground (prior to engine start):

Flight compartment exposed to direct sunlight (sky with less than 10% cloud cover)

• With flight compartment sunshields.......... ISA +35° C

• Without flight compartment sunshields..... ISA +31° C

Flight compartment exposed to indirect sunlight (sky with more than 10% cloud cover)

• ISA +35° C

NOTE: If sunshields are utilized, a sufficient quantity shall be installed to protect the entire flight compartment.

Minimum

• Takeoff/Landing........................................ -40° C

• Enroute..................................................... -75° C

NOTE: When the temperature is below the lowest scheduled, performance appropriate tothe lowest scheduled temperature shall be used.

Wind Component

The maximum tailwind component for takeoff and landing appropriate to a height of 33 ft (10.1 m) is 10 knots.




Runway Slope

The maximum effective runway slopes for takeoff are 2% uphill and 2% downhill.

Airplane Configurations

The airplane configurations as stated in Sub-section 5.05 must be observed.

COMPARTMENT LOADING LIMITATIONS

The airplane must be loaded in accordance with Section 6 - WEIGHT & BALANCE of thisAirplane Flight Manual and as provided on placards in the Baggage/Stowage Compartments.

LOAD LIMITATIONS

Center Of Gravity Limitations

The center of gravity must always lie between the forward and aft limits as defined in theenvelope shown in Figures 2.1 and 2.2.

The limits apply with the landing gear up. The effect of the landing gear in the down position isnegligible.

The center of gravity datum is 11 ft (3.4 m) forward of the reference point on the fuselage. Thereference point is defined by a screw on the fuselage skin located beneath the right enginepod.




LOADING and FLIGHT ENVELOPE - POUNDS/FEET

(Landing Gear Retraction Moment Change is Negligible)

Percent Standard Mean Chord

Figure 2.1




LOADING and FLIGHT ENVELOPE - KILOGRAMS/METERS

Percent Standard Mean Chord

(Landing Gear Retraction Moment Change is Negligible)Figure 2.2

of 3 P/N 140-590032-0005TC16 (Issue 2)March 2, 2006

SECTION 2 - LIMITATIONS

ICE PROTECTION LIMITATIONS

Read the following information in place of the existing information under Icing General on Page 9:

Icing General

Icing conditions exist when Outside Air Temperature (OAT), on the ground and during takeoff,is 10° C SAT or below and visible moisture in any form is present (e.g. clouds, fog with visibilityof 1 mile (1600 meters) or less, rain, snow, sleet and ice crystals).

Icing conditions also exist when the OAT, on the ground and for takeoff, is 10 C or below whenoperating on ramps, taxiways, or runways where surface snow, ice, standing water or slushmay be ingested by the engines or freeze on the engines, nacelles or engine sensor probes.

Read the following information in place of the existing information under Airframe Icingon Page 9:

Airframe Icing

Takeoff is prohibited with frost, ice, snow or slush adhering to the wings, control surfaces,engine inlets or other critical surfaces, with the exception of the following areas:

• Frost is allowable on the underside of the wings over the general area of the fuel tanks provided that the depth does not exceed 1/8 inch (3 mm).

If frost is present in this region, the WAT limited take-off weight must be reduced by 1000 lb (454 kg) and the net flight path reference and fourth segment climb gradients must be obtained using a weight 1000 lb (454 kg) higher than the actual weight.

• Frost is allowable on the fuselage provided the layer is thin enough to distinguish the surface features such as paint lines or markings underneath, but all vents, probes and ports must be clear of frost.

A visual and tactile (hand on surface) check of the wing leading edges and the wing upper surface must be performed to ensure the wing is free from frost, ice, snow or slush when the outside air temperature is less than 50° F (10° C) or if it cannot be ascertained that the wing fuel temperature is above 32° F (0° C) and;

• There is visible moisture (rain, drizzle, sleet, snow, fog, etc.) present;or

• Water is present on the wing;or

• The difference between the dew point and the outside air temperature is 5° F (3° C) or less;or

• The atmospheric conditions have been conducive to frost formation.




ICE PROTECTION LIMITATIONS

Icing General

Icing conditions exist when Outside Air Temperature (OAT) on the ground and during takeoff,or Total Air Temperature (TAT) in flight is 10° C or below, and visible moisture in any form ispresent (e.g. clouds, fog with visibility of 1 mile (1600 meters) or less, rain, snow, sleet and icecrystals).

Icing conditions also exist when the OAT on the ground and for takeoff is 10° C or below whenoperating on ramps, taxiways, or runways where surface snow, ice, standing water or slush maybe ingested by the engines or freeze on the engines, nacelles or engine sensor probes.

Airframe Icing

The airplane must be clear of snow, ice and frost before takeoff with the exception of thefollowing areas:

• Frost is allowable on the underside of the wings over the general area of the fuel tanksprovided that the depth does not exceed 0.125 inch (3.175 mm).

If frost is present in this region, the WAT limited take-off weight must be reduced by 1000lb (454 kg) and the net flight path reference and fourth segment climb gradients must beobtained using a weight 1000 lb (454 kg) higher than the actual weight.

• Frost is allowable on the fuselage provided the layer is thin enough to distinguish thesurface features such as paint lines or markings underneath, but all vents, probes andports must be clear of frost.

Wing/Tail Antice System

Only de-ice fluids TKS80, R328 or fluid to specification DTD 406B must be used.

NOTE: A tank indicating FULL provides priming and protection for a period of at least85 minutes.

Engine Icing

Refer to ENGINE LIMITATIONS - this section.




SEVERE ICING CONDITIONS LIMITATIONS

WARNING: SEVERE ICING MAY RESULT FROM ENVIRONMENTAL CONDITIONS OUT-SIDE OF THOSE FOR WHICH THE AIRPLANE IS CERTIFICATED.

FLIGHT IN FREEZING RAIN, FREEZING DRIZZLE, OR MIXED ICINGCONDITIONS (SUPERCOOLED LIQUID WATER AND ICE CRYSTALS) MAYRESULT IN ICE BUILD-UP ON PROTECTED SURFACES EXCEEDING THECAPABILITY OF THE ICE PROTECTION SYSTEM, OR MAY RESULT IN ICEFORMING AFT OF THE PROTECTED SURFACES.

THIS ICE MAY NOT BE SHED USING THE ICE PROTECTION SYSTEMS,AND MAY SERIOUSLY DEGRADE THE PERFORMANCE ANDCONTROLLABILITY OF THE AIRPLANE.

During flight, severe icing conditions that exceed those for which the airplane is certificatedshall be determined by the following visual cues.

If one or more of these visual cues exists, immediately request priority handling from Air TrafficControl to facilitate a route or an altitude change to exit the icing conditions:

• Extensive ice accumulation on the airframe in areas not normally observed to collect ice.

• Accumulation of ice on the wing aft of the protected area.

Since the autopilot may mask tactile cues that indicate adverse changes in handlingcharacteristics, use of the autopilot is prohibited when any of the visual cues specified aboveexist, or when unusual lateral trim requirements or autopilot trim warnings are encounteredwhile the airplane is in icing conditions.




ENGINE LIMITATIONS

Engine Type

Two Allied Signal TFE731-5BR-1H turbofan engines.

Engine Limitations

* These conditions appear on placards.

NOTES:

1. Initial maximum take-off thrust is selected by the pilot on takeoff. When theAutomatic Performance Reserve (APR) System is operative, maximum take-off(APR) thrust will be obtained automatically on one engine if the other engine failsduring takeoff.

The five minute limit of maximum APR thrust must include the duration of operationat initial maximum take-off thrust prior to the operation of APR. Any normal take-off limitations exceeded during APR operation must be recorded in the technicallog.

2. This is not a normal cruise setting.

CONDITION% RPM

MAX ITT °C TIME LIMITN1 N2

Start or Relight * ---- ----978996

over 996

Unrestricted 10 Seconds 5 Seconds

Takeoff * 100 100.8978

10061016

5 Minutes 5 Seconds 2 Seconds

Maximum Take-off Thrust *(APR Operating) (see NOTE 1)

100 100.8996

10061016


Initial Maximum Take-off Thrust(APR Not Operating)

100 100.8978

10061016


Maximum Continuous * 100 100.8 968Unrestricted

(see NOTE 2)

Maximum Overspeed * 103 103 ---- 5 Seconds

Maximum ......................................... 127° C up to 30,000 ft

Maximum ......................................... 140° C above 30,000 ft

Transient Maximum...........................149° C (2 minutes)

Minimum (Starting) ...........................-40° CMinimum (Takeoff) ............................30° CTakeoff and Maximum Continuousand Climb .........................................38 lb/in2 (262 kPa) to 46 lb/in2 (317.2 kPa)

Idle....................................................25 lb/in2 (172.4 kPa) minimumTransient Maximum...........................55 lb/in2 (379.2 kPa) (3 minutes)

TempOil

PressOil




Approved Engine Oils

Engine Oil Consumption

The maximum permissible oil consumption rate of an engine during any flight is 0.01 U.S.gallons (0.036 liters) per hour.

Engine Fuel Computer

Flight must not be initiated with an engine fuel computer inoperative.

Engine Synchronizer

The ENG SYNC switch must be selected OFF for takeoff.

Automatic Performance Reserve (APR)

The Automatic Performance Reserve (APR) system must be armed for takeoff, except for crewtraining only, when it need not be armed.

Engine Instrument Markings

Red Indications .............................................A maximum or minimum limit has beenexceeded.

Yellow Indications ..........................................Cautionary operations permissible for shortduration or in special circumstances - refer to individual limitations.

Green or White Indications............................Normal operations.

Airplanes which have not accomplished Honeywell Service Bulletin

72-3597 or 72-3662

Airplanes which have accomplishedHoneywell Service Bulletin

72-3597 or 72-3662

Aeroshell/Royco Turbine Oil 500 Mobil 254

Aeroshell/Royco Turbine Oil 560 Exxon 2197

BP Turbo Oil 2197 BP Turbo Oil 2197

BP Turbo Oil 2380

Castrol 5000

Exxon/Esso 2380 Turbo Oil

Exxon 2197

Mobil Jet Oil II

Mobil 254

NOTE: Engine oils in this column can bemixed.

NOTE: Engine oils in this column can bemixed. Do not mix these oils with anyother oils.




Thrust Reversers

• Deployment of either thrust reverser is restricted to ground operations only.

• The thrust reverse levers must not be selected until the airplane is on the ground.

• Engine starts with thrust reversers deployed are prohibited.

• Reverse thrust must not be used to taxi backwards.

• Thrust in excess of reverse idle must not be selected below speeds of 50 KIAS, except inan emergency.

• When operating on unpaved surfaces, reverse idle thrust must not be exceeded except inan emergency.

• If the thrust reverser system is known to be inoperative or unserviceable, it must bedisabled and locked in the forward thrust position.

Engine Ice Protection System

The ENG ANTICE switches may be selected ON at any engine speed. If engine anti-icing isrequired during takeoff, it is recommended that they should be turned ON prior to setting take-off power.

Engine inlet anti-icing should be used in flight continuously during expected icing conditions.

When icing conditions do not exist, the inlet anti-icing should not be used above 50° F (10° C)ambient conditions for more than 10 seconds.

FUEL LIMITATIONS

The following fuels and additives are approved for use with this engine installation.

Fuel Specifications

Aviation kerosene to the current approved issue of the following specifications:

British DEF STAN 91-87 (D.E.R.D. 2453)

DEF STAN 91-91 (D.E.R.D. 2494)

American ASTM D1655/JET A

ASTM D1655/JET A-1

MIL-T-83133/JP8

Canadian CAN/CGSB 3.23/JET A

CAN/CGSB 3.23/JET A-1

Russian GOST 10227-86 TS-1

GOST 10227-86 T-1

GOST 10227-86 RT

Chinese GB 6537-94/No. 3 JET FUEL




Aviation wide-cut fuel to the current approved issue of the following specifications:

Fuel Additives

Anti-static

STADIS 450 additive may be used in concentrations not exceeding 3 parts per million (ppm)by volume.

SIGBOL additive TU38-101741-78 may be used in concentrations not exceeding 0.0005% byvolume.

Anti-icing and biocidal additives

For anti-icing and preventative continuous Biocidal treatment DEF STAN 68-252, MIL-I-27686or MIL-I-85470 may be used in concentrations not exceeding 0.15% by volume.

NOTE: The above additives should not be added to fuel to specification DEF STAN 91-87,MIL-T-5624 and MIL-T-83133 as they are already present in these fuels.

TGF to GOST 17477-86; TGF(M) to TU 6-10-1457-79; I to GOST 8313-88; I(M) to TU 6-10-1458-79 may be used in concentrations not exceeding 0.3% by volume.

Biobor JF may be used at concentrations not exceeding 135 parts per million by weight, aspreventative biocidal treatment.

For biocidal shock treatment, Biobor JF may be used at concentrations not exceeding 270 ppmby weight, provided it is subsequently off-loaded prior to engine start (135 ppm is the maximumconcentration for engine operation).

Anti-corrosive Additive

Fuels may contain additives complying with DEF STAN 68-251 or MIL-I-25017 atconcentrations permitted by the fuel specification.

NOTE: Fuel to specification DEF STAN 91-87 already includes HITEC E515.

British DEF STAN 91-88 (D.E.R.D. 2454)

American ASTM D1655/JET B

MIL-T-5624/JP4

MIL-T-5624/JP5

Canadian CAN/CGSB 3.22/JET B

Russian GOST 10227-86 T-2


Page 15Section - 2LIMITATIONS

FAA Approved Revision A1: Nov 13, 2002

Maximum Fuel Temperature

The maximum permissible fuel temperature is 57° C.

Minimum Fuel Temperature

To avoid the possibility of fuel freezing or exceeding engine limitations, the following ambientair and indicated fuel temperature limitations should be observed:

Fuel Quantity

The usable fuel capacity of each tank when gravity filled is as follows:

Credit shall not be taken for any fuel remaining in the tanks when the fuel quantity indicatorsread zero in level flight.

After pressure refueling, the contents of each wing tank will be 2.4 US Gallons (9.1 Liters) lessand the contents of the ventral tank will be 3.6 US Gallons (13.6 Liters) less.

The contents of the ventral tank are reduced by 4.8 US Gallons (18.2 Liters) for airplanes whichhave an external toilet servicing facility.

FUELMINIMUM FUEL OR AMBIENT AIR TEMP

AT TAKEOFF

MINIMUM FUEL TEMP IN FLIGHT

FUEL FREEZING

TEMP

DEF STAN 91-91 (D.E.R.D 2494)DEF STAN 91-87 (D.E.R.D 2453)ASTM D1655/JET A-1MIL-T-83133/JP8CAN/CGSB 3.23/JET A-1

-42° C -45° C -47° C

ASTM D1655/JET ACAN/CGSB 3.23/JET AGB 6537-94/No. 3 JET FUEL

-35° C -38° C -40° C

DEF-STAN 91-88 (D.E.R.D 2454)MIL-T-5624/JP4MIL-T-5624/JP5

-53° C-35° C

-54° C-38° C

-58° C-40° C

ASTM D1655/JET BCAN/CGSB 3.22/JET B

-45° C -48° C -50° C

GOST 10277-86 TS-1GOST 10277-86 T-1GOST 10277-86 T-2

-54° C -54° C -60° C

GOST 10277-86 RT -50° C -53° C -55° C

LOCATION

U.S. GALLONSEQUIVALENT

LITERS

Wing tank (either side) 634 2400

Ventral tank 233 882


Page 16 Section - 2LIMITATIONS


Fuel Loading

Fuel tanks may be replenished in any sequence provided that the appropriate refuelinginstructions are observed and that the following preflight fuel loading conditions are achieved:

1. Fuel contained in the wing tanks shall be equally disposed between the two wing tanks.

2. Fuel must not be carried in the ventral tank unless each main wing tank contains at least3450 lb (1565 kg) of fuel.

3. Before flights on which it is to be utilized, the ventral tank must be filled completely. For other flights it must be empty.

Pressure Refueling

Takeoff must not be initiated if the amber FUEL annunciator on the MWS panel and the amberREFUEL ON annunciator on the roof panel are illuminated.

Flight with the Refuel Power Switch ON is prohibited.

Fuel System Management

1. During flight, including takeoff and landing, the difference in fuel quantity between the twowing tanks must not exceed 500 lb (227 kg).

2. Fuel carried in the ventral fuel tank shall be transferred into the wing tanks when the fuel levelin the wing tanks has fallen to 3300 lb (1497 kg) per side.

3. Overweight landing procedure and inspection is required for any landing with fuel in theventral tank.

ELECTRICAL LIMITATIONS

Battery Limitations

Maximum battery charge on the main airplane batteries (B1 and B2) immediately beforetakeoff shall not be greater than 20 AMPS.

Generator Limitations

Maximum continuous engine generator load: 300 AMPS

NOTE: Transient excursions, up to a maximum of 400 AMPS, are permitted for a maximumof 2 minutes.

Main Engine Starter Duty Cycle

On the ground, the maximum permitted starter operating time is 30 seconds. After an abortedstart, a minimum of 1 minute cooling time must be allowed before making another attempt tostart. A further 1 minute is required before making a third attempt. The cycle may be repeatedafter a further period of 30 minutes.

Operation of Electrical Circuit Breakers

If, during flight, a systems failure is accompanied by a circuit breaker operation, no attemptmust be made to reset the circuit breaker unless specified in the appropriate Emergency orAbnormal procedure or, if deemed necessary for the continuation of safe flight, a circuit breakermay be reset once.




AVIONICS LIMITATIONS

General

1. The following documents must be carried onboard the airplane at all times:

• Collins Pro Line 21 Avionics System for the Hawker 800XP Pilot’s Guide P/N 523-0780409, 1st Edition dated May 31, 2001 or later revision.

• Collins FMS-6000 Flight Management System for the Hawker 800XP Pilot’s Guide, P/N 523-0780705, 1st Edition dated April 17, 2001 or later revision.

• Mk V Enhanced Ground Proximity Warning System Pilot Guide, P/N 060-4241-0000, Rev D, dated March 2000 or later revision.

These publications contain the description and operation of the Collins Pro Line 21 avionics, theFMS-6000, TCAS II and EGPWS installations and must be available for use.

2. The pilot’s and copilot’s Air Data Computers must be operative for takeoff.

3. AHRS 1 and 2 must be operative for takeoff.

HF Radio

1. When the ADF is being used for approaches, the use of the HF radio is prohibited.

2. Fuel quantity indications are not to be used during HF radio transmissions.

Electronic Standby Instrument System (ESIS)

1. The red airspeed warning on the ESIS airspeed tape does not provide an associated auralwarning.

2. During operations solely with references to the ESIS, the standby VMO/MMO indication mustnot be exceeded, as the ESIS altitude and airspeed indications are not corrected for staticerror.

Flight Management System

1. IFR enroute and terminal navigation is prohibited unless the pilot verifies either the currencyof the database or the accuracy of each selected waypoint and navaid by reference to currentapproved data.

2. The FMS position must be checked for accuracy prior to use as a means of navigation andunder the following conditions:

• At or prior to arrival at each enroute waypoint during FMS navigation along approvedRNAV routes.

• Prior to requesting off-airway routing and at hourly intervals thereafter during FMSnavigation off approved RNAV routes.

• Prior to each compulsory reporting point during IFR operation when not under radarsurveillance control.

3. During periods of dead reckoning, the FMS shall not be used for navigation.




4. All FMS navigation operations are approved within the U.S. National Airspace System andlatitudes bounded by 60° North latitude and 60° South latitude at any longitude.

• Operation to 70° North latitude is acceptable East of 75° West longitude and West of 120°West longitude.

• Operation to 80° North latitude is acceptable East of 50° West longitude and West of 70°East longitude.

• Operation to 70° South latitude is acceptable except for the 45° between 120° East and165° East longitude.

• The WGS-84 coordinate reference datum in accordance with the criteria of AC 20-130A,AC 91-49, and AC 120-33 must be used. Satellite navigation data is based upon use ofonly the Global Positioning System (GPS) operated by the United States.

5. FMS-based Instrument approaches must be accomplished in accordance with approvedinstrument approach procedures that are retrieved from the FMS-6000 data base.

• Instrument approaches must be conducted in the approach mode and GPS integritymonitoring (RAIM) must be available at the Final Approach Fix.

• Accomplishment of ILS, LOC, LOC-BC, LDA and SDF approaches are not authorizedutilizing the FMS.

• When an alternate airport is required by the applicable operating rules, it must be servedby an approach based on other than GPS navigation, the airplane must have operationalequipment capable of using that navigation aid, and the required navigation aid must beoperational.

• FMS based approaches that are retrieved from the navigation database with an approachname of RNVxxx may be flown provided the VHF navigation receiver is tuned to thereference facility.

6. Provided the FMS is receiving adequate usable sensor inputs, it has been demonstratedcapable of and has been shown to meet the accuracy specifications of:

• VFR/IFR enroute RNAV operation in accordance with the criteria of AC 20-130A.

• GPS primary means of navigation in oceanic and remote airspace in accordance with AC20-130A, when used in conjunction with the Collins Fault Detection and Exclusionsoftware, dual Collins FMS-6000 Flight Management Systems and dual GPS-4000Areceivers or a single FMS-6000 Flight Management System and/or Collins GPS-4000Areceiver when operating on routes approved for single GPS navigation.

This does not constitute an operational approval.

NOTE: With single Flight Management System operation, cross reference must be madeto the Airplane Flight Manual for operating procedures and performance data.

of 1 P/N 140-590032-0005TC12 (Issue 2)November 8, 2005

TEMPORARY CHANGE P/N 140-590032-0005TC12 (Issue 2)

PUBLICATION AFFECTED:

AIRPLANE SERIAL NUMBERS AFFECTED:

DESCRIPTION OF CHANGE:

FILING INSTRUCTIONS:



The following additional information applies to the Flight Management System on Page 19:

P-RNAV

Provided the FMS is receiving adequate usable sensor inputs, it has been demonstratedcapable of and has been shown to meet the accuracy specifications of Operation in EuropeanP-RNAV airspace in accordance with JAA Temporary Guidance Material, Leaflet No. 10,provided the following equipment is operational:

Quantity Description

2 .......................CDU-6200 Control Display Unit

1 .......................DBU-4100 Data Base Unit

2 .......................VIR-432 / NAV-4000 / NAV-4500 Navigation Receiver (any 2 of the listed)

2 .......................DME-442 / DME-4000 DME Transceiver (any 2 of the listed)

2 .......................GPS-4000A Global Positioning System

This does not constitute an operational approval.

FAA Approved Airplane Flight Manual, P/N 140-590032-0005,Dated Nov 30, 2001, or later revision for Hawker 800XPairplanes equipped with Pro Line 21 avionics.

Hawker 800XP 258541, 258556, 258567 and After.

Revised the required operational equipment installation foroperations in European P-RNAV airspace.

Remove and destroy existing Temporary Change 12 Page 1 of 1 dated January 7, 2005 from Section 2 -LIMITATIONS Page 19.

Insert this Temporary Change 12 (Issue 2) Page 1 of 1to face page 19 in Section 2 - LIMITATIONS.

of 1 P/N 140-590032-0005TC15December 15, 2005

TEMPORARY CHANGE P/N 140-590032-0005TC15







Autopilot

Do not push a vertical mode Flight Director button (FLC, VNAV or VS) while the altitude preselector control is being rotated.

FAA Approved Airplane Flight Manual, P/N 140-590032-0005,Dated Nov 30, 2001, or latest revision for Hawker 800XPairplanes equipped with Pro Line 21 avionics.


Autopilot Flight Director Limitation.

Insert this Temporary Change Page 1 of 1 to face Page 19 in Section 2 - LIMITATIONS,




• Operation in European B-RNAV/RNP-5 airspace in accordance with AC 90-96 and AC 20-130A. This does not constitute an operational approval.

• Minimum Navigation Performance Specification (MNPS) airspace when equipped withdual Collins FMS-6000 Flight Management Systems and dual Collins GPS-4000A GPSreceivers, or a single FMS-6000 Flight Management System and/or single GPS-4000Areceiver on routes approved for single GPS navigation. This does not constitute anoperational approval.

• VFR/IFR enroute, terminal and approach VNAV operation in accordance with AC 20-129.

7. Use of FMS to capture and track a DME arc outside the published end points is prohibited.

8. Fuel management parameters are advisory only and do not replace the primary fuel quantityindications.

Autopilot

1. A satisfactory preflight check of the system must be performed before the first flight of theday and after any power up or maintenance activity.

2. During autopilot operations, a pilot must be seated at the controls with seat belt and shoulderharness fastened.

3. The autopilot and yaw damper must not be used for takeoff and landing.

4. Do not manually override the autopilot during normal flight.

WARNING: OVERRIDING THE AUTOPILOT IN PITCH DOES NOT CANCEL THEAUTOPILOT AUTOMATIC TRIM. IF A FORCE IS APPLIED TO THE COLUMNWITH THE AUTOPILOT ENGAGED, THEN AUTOMATIC TRIM WILL RUN TOOPPOSE THE APPLIED FORCE. THIS CAN LEAD TO A SEVERE OUT-OF-TRIM CONDITION DURING ANY PHASE OF FLIGHT.

5. Maximum airspeed for operation of the autopilot system must not exceed the airplaneindicated maximum speed VMO/MMO.

6. Operation of the autopilot system with a pitch trim malfunction is prohibited.

7. Do not use the autopilot or yaw damper below 200 ft above terrain during non-precision orCategory I precision approach operations, or 600 ft above terrain during all other operations.

8. The maximum demonstrated adverse wind conditions for autopilot coupled approaches are17 knots crosswind component and 11 knots tailwind component.

9. Nav and localizer captures must be accomplished with an intercept angle of less than 90°.




10. Category II approaches must be executed while coupled to the autopilot with the followinglimits:

Runway Visual Range (RVR) ........................1200 ft minimum

Decision Height (DH) ....................................100 ft minimum

Headwind ....................................................17 knots

Tailwind ....................................................11 knots

Crosswind ....................................................17 knots

Autopilot must be disengaged at ...................80 ft

Two engine operations only

11. During a Category II approach, if the autopilot malfunctions or disengages below 1000 ftAGL, the Category II approach must be discontinued. Hand flying the approach to CategoryI minimums is allowable.

VNAV

1. When using the VNAV system, the barometric altimeters must be used as the primaryaltitude reference for all operations.

2. Use of VNAV guidance for a V-MDA approach that includes a step-down fix between the finalapproach fix and missed approach point is prohibited.

3. VNAV altitudes must be displayed on the MFD map page or CDU legs page when utilizingVNAV for flight guidance.

4. Use of VNAV while conducting a missed approach procedure is prohibited.

5. Provided the FMS is receiving adequate usable sensor inputs, it has been demonstratedcapable of and has been shown to meet the accuracy specifications of VNAV operation inaccordance with the criteria of AC 20-129. Such VNAV approaches must be flown utilizingeither the flight director or autopilot.

6. VNAV approach guidance to a DA is not authorized if the reported surface temperature isbelow the Baro-VNAV minimum temperature limitation specified on the applicable RNAVapproach procedure chart.

NOTE: Barometric VNAV guidance during approach including the approach transition,final approach segment and the missed approach procedure is not temperaturecompensated. Operating at uncompensated minimum IFR altitudes will notprovide expected terrain and obstacle clearance for temperatures below ISA.




EGPWS

1. Navigation must not be predicated upon the use of the TAD. The Terrain Display is intendedto serve as a situational awareness tool only and may not provide the accuracy and/or fidelityon which to solely base terrain avoidance maneuvering.

2. Pilots are authorized to deviate from their current air traffic control (ATC) clearance to theextent necessary to comply with an EGPWS warning.

3. In order to avoid giving unwanted alerts, the Terrain Awareness alerting must be inhibited byselecting the TERR INHIB switchlight when within 15 nautical miles of takeoff, approach orlanding at an airport not contained in the EGPWS Airport Database. Refer to Honeywelldocument 060-4326-000 for airports contained in the installed EGPWS Terrain Database.

4. When the FMS is operating in the DR mode, the Terrain Awareness alerting must be inhibitedby selecting the TERR INHIB switchlight.

NOTE: The terrain database, displays and alerting algorithms currently account for limitedcataloged human-made obstructions in North America and Europe. If obstacledata is not in the database for a particular obstacle, the Obstacle Awarenessalerting is not available for that obstacle.

TCAS II

Pilots are authorized to deviate from their current ATC clearance to the extent necessary tocomply with a TCAS resolution advisory.

If ATC requires the transponder altitude reporting to be disabled, TCAS II must be turned off.




AIRSPEED LIMITATIONS

Maximum Operating Speed

VMO ................................................................280 KIAS (Ventral fuel tank not empty, Flaps 0°)

VMO ................................................................335 KIAS (Ventral fuel tank empty, Flaps 0°)Sea Level to 12,000 ft, reducing by 1 kt per680 ft to 310 KIAS at 29,000 ft.

Maximum Operating Mach Number

MMO...............................................................0.80 IMN

MMO...............................................................0.73 IMN with Mach Trim System Fail/Inoperative and Autopilot disengaged.

NOTE: The maximum operating speeds and operating Mach numbers as given aboveshall not be deliberately exceeded in any regime of flight (climb, cruise ordescent) except for the purpose of pilot training or routine test flights inaccordance with Section 5, Sub-section 2 of the Pilot’s Operating Manual (POM).

If the limits are inadvertently exceeded, speed shall be reduced to or below thelimiting values as quickly as possible.

Maneuvering Speed

VA ..................................................................196 KIAS

NOTE: Maneuvering speed is the speed below which full application of aerodynamiccontrols will not result in excessive airplane loads. Maneuvers involving anglesof attack near the stall should be confined to speeds below VA.

Avoid rapid and large alternating control inputs, especially in combination withlarge changes in pitch, roll or yaw (e.g. large slip angles) as they may result instructural failures at any speed, including below VA.

Wing Flaps Extended/Operating Speed

VFE/VFO.........................................................220 KIAS (Flaps 15°)175 KIAS (Flaps 25°)165 KIAS (Flaps 45°)

Procedural Use of Flaps 15° for Descent and Holding

The maximum airspeed for procedural use of flaps 15° for descent and holding is 220 KIAS.

The maximum altitude for use is 15,000 ft. Such use of wing flaps is not permitted in icingconditions.




Air Brakes

Air Brakes (Flaps 0° Only) .......................... No Limit

Landing Gear Extended/Operating Speed

VLE/VLO ........................................................ 220 KIAS

Bird Strike Speed

Under normal conditions the maximum permissible airspeed to meet bird strike requirementsis 280 KIAS up to 8000 ft.

Following an airplane ground soak at temperatures below -10° C, the windscreen heat shouldbe operative and selected ON for a minimum of 5 minutes prior to takeoff in ambienttemperatures of below -10° C and for a minimum of 15 minutes prior to takeoff when ambienttemperatures are below -20° C.

If the minimum times for windscreen heat operation have not been achieved or in the case of windscreen heat failure followed by flight in ambient temperature below -10° C, the maximum permissible airspeed is 257 KIAS up to 8000 ft.



FAA Approved Original Issue: Nov 30, 2001


of 2 P/N 140-590032-0005TC13March 14, 2005

TEMPORARY CHANGE P/N 140-590032-0005TC13






MISCELLANEOUS LIMITATIONS

Read the following information as an addition to Section 2:

Ditching

After ditching, do not open the main cabin door.

FAA Approved Airplane Flight Manual, P/N 140-590032-0005,Dated Nov 30, 2001, or later revision for Hawker 800XPairplanes equipped with Pro Line 21 avionics.


Addition of ditching information and procedures.

Insert this Temporary Change Page 1 of 1 to face Page 25 in Section 2 - LIMITATIONS.




MISCELLANEOUS LIMITATIONS

Air Brakes

In flight, the air brakes must not be operated when the flaps are extended to any position.

Cabin Emergency Overwing Exit

The internal cabin emergency overwing exit locking pin, if installed, must be removed and stowed before each flight.

Cabin High Datum

Cabin High Datum shall only be selected when operating into airfields greater than 9000 ftelevation.

Crew Seats

Both crew seats shall be locked in position during takeoff and landing. When installed and in use, the 3rd crew member seat shall be locked in position; when not in use, it shall be folded and stowed or removed from the airplane.

Inter-compartment Door

When a door is provided between the crew and passenger compartments, it shall be securedin the open position during takeoff and landing.

Lift Dump

Lift dump is to be used only when the airplane is on the ground.

Maneuvering Load Factor Limitations

Operation is limited to normal flying maneuvers and aerobatic maneuvers are not permitted.

The maximum accelerations (i.e. load factors) for which the structure is approved are 2.0g withflaps extended and 2.73g with flaps fully retracted.

Maneuvers exceeding these values can cause permanent distortion of the structure and mustbe avoided.

Minimum Flight Crew

The minimum crew is two pilots.

Nosewheel Tires

The airplane must be installed with chined nosewheel tires.

Number of Occupants

The total number of persons carried shall not exceed 17 nor that for which approved seatingaccommodation is provided.

Pressure Cabin

The cabin shall not be pressurized during takeoff and landing. Maximum pressure differential for normal operations is 8.55 lb/in2.

NOTE: The safety valve is set to operate between 8.6 to 8.8 lb/in2 (59.3 to 60.7 kPa).

Rudder Bias

The rudder bias switches must be ON and the systems operative during each takeoff.




Smoking

When the airplane is being operated in the Transport Category (Passenger), smoking in thetoilet compartment is prohibited.

Placards summarizing this limitation shall be provided inside and outside each toilet and shallnot be obscured.

System Gage Markings

Red Arc/Red Radial................................................ Maximum or minimum limit.

Yellow Arc ............................................................... Cautionary range permissible for shortduration or in special circumstances - refer to individual limitations.

Weather Radar

Do not use the weather radar in the vicinity of ground personnel. A hazardous area extends upto 2 feet in front of the radar dish.

Wheel Brakes

NOTE: If any of the wheels’ fusible plugs blow, the brakes must be inspected and certifiedserviceable before the next takeoff.

After the airplane has made a normal landing or a stop from a rejected takeoff, a waiting periodshould be established to make sure the brakes are both sufficiently cool and in a serviceablecondition for a further rejected takeoff (critical case).

After Rejected Takeoff

Required period from completion of taxi-in following a rejected takeoff from a speed of 90 KIASor less, to before start of taxi-out for takeoff.

After a single rejected takeoff ..................................25 minutes

After two or more successive rejected takeoffs .......45 minutes

If the rejected takeoff is made from a speed greater than 90 KIAS, the brakes must be inspectedand certified to be serviceable before the next takeoff.

After Normal Landing

The required waiting period from completion of taxi-in from landing to before start of taxi-outfor takeoff is 5 minutes, except when the take-off weight exceeds the values given in Table 1.

When the weight exceeds these values, a period of 30 minutes must be allowed. The table isbased on still air and a downhill slope not exceeding 1/2%.

Corrections for more adverse conditions are given in the NOTES below Table 1.




Table 1 - Take-off Weights for Wheel Brakes Waiting Period

NOTES:

1. In 1 - 5 knot tailwind subtract 1500 lb (680 kg).

2. In 6 - 10 knots tailwind subtract 3000 lb (1360 kg).

3. If the downhill slope exceeds 1/2%, subtract 250 lb (113 kg).

4. Take-off weight as limited by climb requirements may be more restrictive whenoperating in shaded areas.

5. Performance appropriate to sea level shall be used when the field pressure altitudeis below sea level.

FieldPressureAltitude ft

Temperature °C

-20 -10 0 10 20 30 40 50

Take-off Weight

14,00021,100 lb9570 kg

20,500 lb9298 kg

20,000 lb9071 kg

19,500 lb8845 kg

19,300 lb8754 kg

19,100 lb8663 kg

13,00021,600 lb9797 kg

21,000 lb9525 kg

20,400 lb9253 kg

19,900 lb9026 kg

19,700 lb8935 kg

19,500 lb8845 kg

12,00022,100 lb10,014 kg

21,500 lb9752 kg

20,900 lb9480 kg

20,400 lb9253 kg

20,000 lb9071 kg

19,900 lb9026 kg

11,00022,600 lb10,251 kg

22,000 lb9979 kg

21,400 lb9706 kg

20,900 lb9480 kg

20,400 lb9253 kg

20,300 lb9208 kg

10,00023,100 lb10,478 kg

22,500 lb10,205 kg

21,900 lb9933 kg

21,300 lb9661 kg

20,800 lb9434 kg

20,600 lb9344 kg

900023,800 lb10,795 kg

23,100 lb10,478 kg

22,400 lb10,160 kg

21,800 lb9888 kg

21,200 lb9616 kg

21,000 lb9525 kg

21,000 lb9525 kg

800024,500 lb11,113 kg

23,700 lb10,750 kg

22,900 lb10,387 kg

22,300 lb10,115 kg

21,700 lb9843 kg

21,400 lb9706 kg

21,300 lb9661 kg

700025,100 lb11,385 kg

24,400 lb11,067 kg

23,500 lb10,659 kg

22,900 lb10,387 kg

22,200 lb10,069 kg

21,700 lb9843 kg

21,700 lb9843 kg

600025,600 lb11,612 kg

25,000 lb11,339 kg

24,000 lb10,886 kg

23,500 lb10,659 kg

22,700 lb10,296 kg

22,000 lb9979 kg

22,000 lb9979 kg

500026,300 lb11,929 kg

25,600 lb11,612 kg

24,800 lb11,249 kg

24,200 lb10,977 kg

23,400 lb10,614 kg

22,700 lb10,296 kg

22,400 lb10,160 kg

400027,000 lb12,247 kg

26,200 lb11,884 kg

25,500 lb11,566 kg

24,900 lb11,294 kg

24,100 lb10,931 kg

23,200 lb10,523 kg

22,800 lb10,342 kg

22,800 lb10,342 kg

300027,600 lb12,519 kg

26,800 lb12,156 kg

26,100 lb11,838 kg

25,500 lb11,566 kg

24,700 lb11,203 kg

23,900 lb10,840 kg

23,200 lb10,523 kg

23,200 lb10,523 kg

200028,000 lb12,700 kg

27,500 lb12,473 kg

26,700 lb12,111 kg

26,000 lb11,793 kg

25,400 lb11,521 kg

24,500 lb11,113 kg

23,700 lb10,750 kg

23,600 lb10,704 kg

100028,000 lb12,700 kg

28,000 lb12,700 kg

27,300 lb12,383 kg

26,700 lb12,111 kg

26,000 lb11,793 kg

25,200 lb11,430 kg

24,300 lb11,022 kg

23,900 lb10,840 kg

Sea Level28,000 lb12,700 kg

28,000 lb12,700 kg

28,000 lb12,700 kg

27,300 lb12,383 kg

26,600 lb12,065 kg

25,800 lb11,702 kg

24,900 lb11,294 kg

24,300 lb11,022 kg


Pro Line 21


Sub-section 1MASTER WARNING SYSTEM

Table of Contents

Page

GENERAL .......................................................................................................1-3ANNUNCIATOR TYPES .................................................................................1-3MWS ARRANGEMENT ..................................................................................1-3REPEATER ANNUNCIATORS.......................................................................1-3

Figure 1 - Master Warning Annunciations and Controls ...........................1-4SYSTEM OPERATION ...................................................................................1-5ACKNOWLEDGEMENT .................................................................................1-5ANNUNCIATION SEQUENCE........................................................................1-6

Table 1: Annunciation Sequence .............................................................1-6TESTING .........................................................................................................1-7

DIM CONTROL FAILURE .....................................................................1-7MWS MASTER WARNING LAMP FAILURE .........................................1-7

POWER SUPPLIES ........................................................................................1-7ANNUNCIATORS WITH ASSOCIATED MWS REPEATERS........................1-8

ELECTRICAL.........................................................................................1-8ICE PROTECTION ................................................................................1-9FUEL....................................................................................................1-10ENGINE FIRE WARNING ...................................................................1-11DUCT OVHT REPEATER ANNUNCIATOR ........................................1-11

MWS ANNUNCIATORS - WITHOUT A REPEATER ...................................1-12Table 2: MWS Annunciators - W ithout a Repeater .................................1-12




GENERAL

A master warning system (MWS) consisting of annunciation logic, displays and associated controls isprovided to alert or advise the crew to the status of the airplane systems.

ANNUNCIATOR TYPES

The annunciator captions are illum inated against a black background and are color coded as follows:

MWS ARRANGEMENT (Figure 1)

A main MWS panel with annunciators is located on the center instrument panel with a MWS DIMvariable control and a NORM/DIM OVRD switch located adjacent to the main MWS panel. A MWS DIMFAIL annunciator is located on the copilot instrument panel above the PFD/MFD.

Two red master warning lamps, each with a push-to-cancel switch, are located on the glareshield, onein front of each pilot. Additional annunciators are arranged in groups in the system areas of theoverhead roof panel. An ANNUN test button is located in the test section of the overhead roof panel.

REPEATER ANNUNCIATORS

The MWS panel also provides repeater annunciators which have an upward pointing arrow.

When illuminated, these annunciators indicate to the flight crew that an additional warning annuncia-tor has illuminated on the overhead roof panel.

ENG 1FIRE

EMERGENCY

ABNORMAL

ADVISORY

AIRBRAKE

These indicate a hazardous fault condition which requires immediate flightcrew action and are accompanied by the flashing red MWS warning lamps.

These indicate a fault condition which is not immediately hazardous anddoes not require urgent action by the flight crew. The MWS master warninglamps do not operate with this warning.

These are advisory indications of system status and do not require anyremedial action from the flight crew.

VALVE 2MAIN AIR

Typical Repeater Annunciator

ENG 1FIRE




REVERSER MACH TRIMFAIL

ENG 1FIRE

CABINALTITUDE

ELEV/AILTRIM

HYD 1LO PRESS HYD OVHT

HYD 2LO PRESS

MAIN AIRVALVE 1

MAIN AIRVALVE 2

AUX HYDLO LEVEL

ENG 1CMPTER

ENG 2CMPTER

EMRG BRKLO PRESS

ENG 2A/ICE

ICEPROT

MFDMFDPFD PFD

Copilot Instrument PanelPilot Instrument Panel

Center InstrumentPanel

Glareshield

ENG 1A/ICE

ELECT FUEL DUCTOVHT

ENT DOORUNLOCKED

RUDDERBIAS

APU ON

ICE PROTSELECTED

STALLIDENT

FUELXFD TFR

AIR BRAKE

Figure 1Master Warning Annunciations and Controls

Overhead Roof Panel

ENG 1FIRE

ENG 2FIRE

ENG 1FUEL

ENG 2FUEL

REFUELON

AUX FUELTFR

WING FUELXFD/TFR

REAR BAYDOOR

FUEL 1LO PRESS

FUEL 2LO PRESS IGN ON

BATT 1CNTCTR

BATT 2CNTCTR

GEN 1FAIL

GEN 2FAIL

BUS TIEOPEN

L SCREENOVHT

R SCREENOVHT

ALTR 1FAIL

ALTR 2FAIL

L VANEHTR FAIL

R VANEHTR FAIL

SIDE SCRNOVHT

ICEDETECTED

L PITOTHTR FAIL

R PITOTHTR FAIL

ANTICELO PRESS

ANTICELO QTY

XS 1FAIL

XEFAIL

XS 2

FAIL

INV 1FAIL

INV 2

FAIL

STBY INV

ON

MWS

FAILDIMOIL 1

LO PRESS

HP AIR 1OVHT

HP AIR 2OVHT

REAR BAYOVHT

ENG 2FIRE

OIL 2LO PRESS

APUFIRE

MWS




SYSTEM OPERATION

When a system status change or fault condition occurs, the appropriate annunciator illuminates atmaximum intensity.

In the event of a red warning, both MWS red master warning lamps will flash and if the red annunciatoris located on the overhead roof panel, the associated repeater annunciator illuminates with a steadyintensity. If the roof panel annunciator is amber, the repeater flashes.

NOTES:

1. A repeater annunciator will illuminate steady should associated red and amberannunciators illuminate together.

2. When any annunciation is initiated, all previously dimmed annunciators will increase inbrightness.

ACKNOWLEDGEMENT

Either pilot can acknowledge the warning by pushing either MWS red master warning lamp on theglareshield with the following results:

• The MWS red master warning lamps are cancelled.

• The brightness of the annunciator (and repeater where applicable) reduces to the level selected bythe MWS dimmer switch.

• The repeater, if flashing, changes to steady.

Should an additional system status change or failure occur, the relevant annunciator illuminates atmaximum intensity and any dimmed annunciator increases in brightness.

Subsequent dimming is achieved by pushing either MWS red master warning lamp. This low lightinglevel is maintained until the system fault clears causing the annunciator to extinguish or until anothersystem status change or failure occurs.

White annunciators are always illuminated at the set dim level and do not cause an increase inbrightness of existing warnings to occur.




ANNUNCIATION SEQUENCE

Table 1: Annunciation Sequence

AnnunciatorLocation

AnnunciatorColor

Annunciator state and intensity level when first

illuminatedAttention Event

OverheadRoof Panel

Illuminates steady at maximum intensity

Flashes at maximum intensity

Illuminates steady at dimmed intensity

(as set on MWS DIM control) None

MWSMain Panel


MWS

Master warning lamps flash (on Glareshield)



None

Illuminates steadyat dimmed intensity

(as set on MWS DIM control) None

AMBER REPEATER

WHITE

RED

REPEATER

AMBER

WHITE




TESTING

Pushing the ANNUN test button on the overhead roof panel results in the following:

• All of the overhead and MWS panel annunciators, including white, illuminate at maximum intensity.

• The amber repeater annunciators will flash and the red repeater annunciators illuminate steady.

• The MWS master warning lamps flash alternatively at an even rate.

DIM CONTROL FAILURE

If a warning occurs and the associated annunciator fails to illuminate due to an open-circuit failure ofthe dimming circuit, the amber MWS DIM FAIL annunciator will illuminate on the MWS panel.

The flight crew should set the NORM/DIM OVRD switch to the DIM OVRD position. This causes thewarning annunciator to illuminate at maximum intensity and the MWS DIM FAIL annunciator willextinguish.

When the warning is acknowledged, the annunciator will remain at the maximum level of intensity.

MWS MASTER WARNING LAMP FAILURE

The MWS red master warning lamp control channel contains dual circuitry which provides a back-up,and a means of indicating a single failure. A failure is indicated by alternate flashing at an uneven rate (one faster than the other) of the two lamps.

POWER SUPPLIES

The MWS is powered from PE busbar via three circuit breakers located on panel DA-D:

• MWS PWR 1 and 2

• MWS TEST

The MWS dimming circuit is powered from a secondary busbar - MWS DIM bus.

NORMANNUN

DIM OVRD DIM

DIMMWS

FAIL

MWS

MWSMWS

FAILED SIDE

SLOW RATE

WORKING SIDE

NORMAL RATEOF FLASHOF FLASH




ANNUNCIATORS WITH ASSOCIATED MWS REPEATERS

ELECTRICAL

MFDMFDPFD PFD


Center Instrument

Panel

GEN 1FAIL

GEN 2FAIL

BATT 1

BATT 2

CNTCTR

CNTCTR

BUS TIEOPEN

XS 1FAIL

XS 2FAIL

XEFAIL

INV 1FAIL

INV 2FAIL

ELECT

The illumination of any of the above electrical annunciators on the Overhead Roof Panel will beaccompanied by the illumination of the associated MWS repeater shown below:

AC Power Section

DC Power Section

Overhead Roof Panel




ICE PROTECTION

The illumination of any of the above ice protection annunciators on the Overhead Roof Panel will beaccompanied by the illumination of the associated MWS repeater shown below:

MFDMFDPFD PFD


Center Instrument

Panel

ICEPROT

L SCREENOVHT

R SCREENOVHT

ALTR 1FAIL

ALTR 2FAIL

HTR FAILL VANE

HTR FAILR VANE SIDE SCRN

OVHT

ICEDETECTED

HTR FAILL PITOT

HTR FAILR PITOT ANTICE

LO PRESSANTICELO QTY

Ice Protection Section

Overhead Roof Panel




FUEL

MFDMFDPFD PFD


Center Instrument

Panel

FUEL

ENG 1FUEL

ENG 2FUEL

FUEL 1LO PRESS

FUEL 2LO PRESS

REFUELON

WING FUELXFD/TFR

AUX FUELTFR

FUELXFD TFR

The illumination of any of the above fuelannunciators on the Overhead Roof Panelwill be accompanied by the illum ination of theassociated MWS repeater shown below:

The illumination of either of the above fuelannunciators on the Overhead Roof Panel willbe accompanied by the illum ination of theassociated MWS repeater shown below:

Overhead Roof Panel Overhead Roof Panel




ENGINE FIRE WARNING

MFDMFDPFD PFD


Center Instrument

Panel

ENG 1FIRE

ENG 2FIRE

ENG 1FIRE

ENG 2FIRE

The illumination of either of the above fire warning annunciators on the Overhead Roof Panel ForwardExtension will be accompanied by the illum ination of the respective MWS repeater shown below:

DUCT OVHT REPEATER ANNUNCIATOR

MFDMFDPFD PFD


Center Instrument

PanelDUCTOVHT

This repeater illum inates on the MWS panel and directs attention to the DUCT TEMP indicator on theOverhead Roof Panel ENVIRONMENTAL section.

Overhead Roof Panel Forward Extension




MWS ANNUNCIATORS - WITHOUT A REPEATER

Table 2: MWS Annunciators - Without a Repeater

Red Annunciators Location Amber Annunciators Location

MWS panel(if installed)

MWS panel

MWS panel MWS panel

MWS panel MWS panel

MWS panel MWS panel

MWS panel MWS panel

MWS panel

MWS panel

MWS panel

MWS panel

MWS panel

APUFIRE

REVERSER

HP AIR 1 or 2OVHT TRIM

ELEV/AIL

OVHTREAR BAY

FAILMACH TRIM

CABINALTITUDE LO PRESS

HYD 1 or 2

LO PRESSOIL 1 or 2 HYD OVHT

MAIN AIRVALVE 1 or 2

AUX HYDLO LEVEL

ENG 1 or 2CMPTER

LO PRESSEMRG BRK

ENG 1 or 2A/ICE




Table 2 continued: MWS Annunciators - Without a Repeater

Red Annunciator Location Amber or White Annunciators Location

MWS panel

MWS panel

MWS panel

Overhead panel

MWS panel(if installed)

MWS panel

Flight CompartmentOverhead Roof Panel

MWS panel

Flight CompartmentOverhead Roof Panel

ENT DOORUNLOCKED

RUDDERBIAS

STALLIDENT

REAR BAY DOOR

APUON

ICE PROT SELECTED

IGN ON

AIR BRAKE

STBY INV ON


Pro Line 21


Sub-section 12OXYGEN SYSTEM

Table of Contents

Page

GENERAL ................................................................................................. 12-3

Figure 1 - Oxygen Cylinder Assembly ................................................ 12-3Figure 2 - Oxygen System.................................................................. 12-4

SERVICING ............................................................................................... 12-5

Figure 3 - Oxygen Box Assembly ....................................................... 12-5OPERATION ............................................................................................. 12-6

Figure 4 - Flight Compartment Oxygen Panel on Left Console .......... 12-6PORTABLE OXYGEN SMOKE SET ...................................................... 12-7

Figure 5 - Portable Oxygen Smoke Set .............................................. 12-7FLIGHT CREW SUPPLY........................................................................ 12-8

Mask - Regulator.............................................................................. 12-8Figure 6 - Oxygen Mask and Regulator (Mod. No. 25A025A) ............ 12-8

Goggles............................................................................................ 12-9Figure 7 - Smoke Goggles and Mask-Regulator ................................ 12-9

THERAPEUTIC SUPPLY ....................................................................... 12-9PASSENGER SUPPLY ........................................................................ 12-10

Figure 8 - Passenger Oxygen Box Locations and Mask Stowage.... 12-10Figure 9 - Oxygen Mask Folding and Stowage................................. 12-11



Hawker 800XP Pro Line 21

Section III - SYSTEMS DESCRIPTION






GENERAL

Two 750 liter oxygen cylinders (with provision for a third 750 liter cylinder) are mounted in the rear equipment bay between frames 24 and 25. The cylinders are charged to 1800 psi and normally supply oxygen to two quick-release sockets in the flight compartment, two therapeutic outlets and eight drop-out mask units in the passenger cabin and one drop-out mask unit in the toilet compartment.

An automatic shut-off valve is located in the oxygen box assembly (Figure 3) which will shut-off the supply of oxygen should there be a rupture of the supply pipeline downstream of the valve. Provision is made to install a regulator and quick-release socket in the forward vestibule cabinet for a third crew member, a drop-out mask in the vestibule, and for additional drop-out mask units in the passenger cabin depending on the number of seats.

Cylinder pressure is reduced to a nominal 70 psi by a pressure regulator incorporating a relief valve operating at 90 psi. The pressure regulator has an integral grounding lug attached to two bonding leads from the adjacent system piping.

The supply for the therapeutic outlets is taken directly from the pressure regulator. The drop-out masks supply is taken from the pressure regulator through a baromatic valve and the passenger supply valve. The baromatic valve automatically causes the masks to fall to the half-hang position at a certain cabin altitude and can be operated manually to release the masks at any altitude.

750 LITEROXYGENCYLINDERS

TAIL

FRAME 24

VENT HOSE

BLANKING PLUG IF

THIRD CYCLINDER

IS NOT

SEALED BOX

SCREENS

OPTIONAL 750 LITEROXYGEN CYLINDER

BLANKING PLUG IF 3rd CYLINDER NOT INSTALLED

Figure 1Oxygen Cylinder Assembly





750 LITER

OXYGEN CYLINDERS

750 LITER

OXYGEN

CYLINDERS

NOTE: A blanking plug is installed if an optional cylinder is not installed.

Figure 2Oxygen System





SERVICING

The oxygen cylinders are charged through a charging valve in the oxygen box assembly which is situated in the right hand rear fuselage between frames 24 and 25.

A contents indicator is mounted next to the charging valve. The charging supply passes through a line filter and bursting disc assembly before joining the pipeline from the cylinders to the automatic shut-off valve.

The automatic shut-off valve is also located in the oxygen box assembly and is provided to shut off the oxygen supply should there be a rupture of the supply pipeline downstream of the valve.

All system piping is made from stainless steel or light alloy, except for hoses which connect the oxygen cylinders, drop-out mask units, therapeutic outlets, and the mask quick-release sockets.

Figure 3

Oxygen Box Assembly





OPERATION

Oxygen from the storage cylinders is fed to the master SUPPLY valve on the flight compartment oxygen panel on the left console.

Opening the master SUPPLY valve allows oxygen to flow to the contents indicator and the pressure regulator, then, at 70 psi to the combined mask-regulators, therapeutic outlets and the baromatic valve.

Figure 4Flight Compartment Oxygen Panel on Left Console

TEST

PASSENGER SUPPLY

EMERGENCY

PULL

PULL TO OPERATE

PUSH FOR OFF

SUPPLY

ON

EMPTY

1/4

1/23/4

FULL

OXYGENCONTENTS

OXYGEN





PORTABLE OXYGEN SMOKE SET

The Puritan-Zep portable oxygen smoke set comprises a single 312 liter capacity oxygen cylinder and a smoke mask. The cylinder rests in a fixture secured to the rear of the flight compartment right console.

Pre-Mod. 252939: The top of the cylinder is secured to the forward face of panel DA by a toggle fastener.

Post-Mod. 252939: The cylinder is secured by a toggle fastener, and hinged bracket which covers electrical power points on panel DA.

2nd FLASHLIGHT

Figure 5Portable Oxygen Smoke Set

OXYGENCYLINDER

HINGED OXYGENCYLINDER MOUNTING PANEL

PANEL DA

FORWARD

MASKSTOWAGE

GOGGLE STOWAGE ONLEFT SIDE SIMILAR

NOTE





FLIGHT CREW SUPPLY

Mask-Regulator

Under normal flight conditions the mask-regulator is selected to the —N“ position. At this setting the ratio of oxygen to air increases with an increase in altitude until at approximately 30,000 feet, when 100% oxygen is supplied.

Between 35,000 and 41,000 feet 100% oxygen at a positive pressure is automatically maintained. However, 100% oxygen is available at any altitude when the mask regulator is selected to the 100% position.

Turning the mask regulator knob to EMERGENCY provides a 100% oxygen supply under positive pressure. The regulator can be functionally tested by setting the selector to the 100% position and pushing the regulator knob to TEST position. The flow of oxygen can be checked by feel; the test can be carried out with the mask in its stowage.

Figure 6

Oxygen Mask and Regulator (Mod. No. 25A025A)




Goggles

Combined smoke goggles and mask-regulators are provided for the flight crew with the goggles stowed in the flight compartment left and right side consoles.

NOTE: Headsets and hats must be removed before donning oxygen masks.

THERAPEUTIC SUPPLY

Oxygen for therapeutic use is available at two self-sealing outlets in the passenger cabin. These outlets incorporate a check valve, spring-loaded against its seating and sealed by two sealing rings.

When the bayonet adapter of the therapeutic mask is inserted, the hollow probe of the adapter unseats the check valve and enters the sealing rings. Oxygen then flows to the mask as shown by an indicator integral with the mask hose.

Figure 7Smoke Goggles and Mask-Regulator




PASSENGER SUPPLY

The emergency drop-out passenger mask unit consists of a mask and hose assembly and an emergency mask stowage. The emergency mask stowage boxes are provided in the airplane ceiling structure above the passengers' heads and supply oxygen to the single face masks. The stowage opens automatically or manually and allows the mask to drop into the half hang position during an emergency.

The mask has a lightweight molded face-piece which can be held against the face with one hand. Metal plates on either side of the base give it support and secure a filter. The mask supply hose assembly consists of two hoses joined by the flow indicator. When the mask is in the stowed position, the flow indicator is held in the carrier clip in the stowage box. In this position the check valve in the flow indicator is held closed and prevents flow of oxygen.

B

A

DETAIL A DETAIL B

Figure 8Passenger Oxygen Box Locations and Mask Stowage





Figure 9

Oxygen Mask Folding and Stowage


Pro Line 21


Sub-section 18PITOT-STATIC SYSTEM

Table of Contents

Page

GENERAL ................................................................................................. 18-3PITOT HEADS........................................................................................ 18-3STATIC VENT PLATES .......................................................................... 18-3STATIC VENTS...................................................................................... 18-3 EQUIPMENT ISOLATION ...................................................................... 18-4STALL VENTS ....................................................................................... 18-4

Figure 1 - Pitot Static System Block Diagram................................... 18-5


Sub-section 18PITOT STATIC SYSTEM






GENERAL

The pitot-static system supplies the source pressure for the No. 1 and No. 2 Air Data Computers, an Electronic Standby Instrument System Air Data Unit, the Stall Warning and Identification System, and theAirplane Cabin Differential Pressure Indicator.

Provisions are made available for the connection of additional equipment.

The pitot-static system pressure lines and components are shown in a block diagram on Figure 1.

PITOT HEADS

Two pitot heads are mounted, one on each side of the fuselage nose and provide independent suppliesof pitot pressure to the following:

Left Pitot Head (P1)

• Air Data Computer No. 1

Right Pitot Head (P2)


• Standby Airspeed Indicator

• Stall Detectors

• Additional equipment connections.

STATIC VENT PLATES

A static vent plate is mounted on each side of the forward fuselage. Each static plate provides two staticports (static ports 5 & 6, reference Figure 1). These sources provide static pressure to the following:

Static 5 (S5)


Static 6 (S6)


• Stall Detectors

STATIC VENTS

Static vents are provided on each side of the aft fuselage (static vents 8 & 9, reference Figure 1). Thesestatic vents are respectively connected and provide static pressure to the following:

Static 8 (S8)

• Electronic Standby Instrument System (ESIS) Air Data Unit

Static 9 (S9)

• Cabin Differential Pressure Indicator

• Additional Equipment Connections




EQUIPMENT ISOLATION

A PITOT ISOLATION valve, located on the copilot’s side console, provides isolation of the stall detectorsand any additional equipment from the Air Data Computer No. 2 and the ESIS Air Data Unit.

With the valve in the NORMAL position, P2 pitot pressure is supplied from the right pitot head to all relevantinstruments and equipment.

Operating the valve from NORMAL to ISOLATE maintains P2 pitot pressure to the Air Data ComputerNo. 2 and the ESIS Air Data Unit but isolates the stall detectors and any additional equipment (referenceFigure 1).

STALL VENTS

A stall vent is mounted under each wing and interconnected by pressure lines to the two stall detectors(reference Figure 1).




Left ForwardStatic Vent Plate

56

Right ForwardStatic Vent Plate

6S6

S5

Copilot Instrument Panel

Cabin PressureController

No. 1Air Data

Computer

P1 P2

P

S

StallDetector

LeftPitotHead

RightPitotHead

9

8

S9

S9

PitotIsolationValve

Connections forAdditional Equipment

Stall Ident(Autopilot Disconnected)

S6

Left RearStatic Vents

Right RearStatic VentsS8

LEGEND= Pitot Pressure = Pitot Drain= Static Pressure = Static Drain

Figure 1Pitot Static System Block Diagram

5

No. 2Air Data

Computer

ESISAir Data Unit

8

9S9

S8

S8

S6

P2

P2

S9

P

S

StallDetector

V

V

V2V1

P2

Stall Ident3rd Channel

= Static Pressure = Stall Vent Drain

LeftStallVent

RightStallVent


Pro Line 21


Sub-section 7STALL WARNING/IDENTIFICATION

Table of Contents

Page

GENERAL ..................................................................................................... 7-3SYSTEM LOGIC......................................................................................... 7-3CONTROLS and ANNUNCIATIONS .......................................................... 7-3

OPERATION ................................................................................................. 7-4STALL WARNING SYSTEM....................................................................... 7-4STALL IDENTIFICATION SYSTEM ........................................................... 7-4

Figure 1 - Stall Valve Annunciators ........................................................ 7-4Figure 2 - Variation of Stall Identification Angle With

Rate of Increase of Vane Angle............................................. 7-5THIRD STALL IDENTIFICATION CHANNEL ............................................. 7-6

Figure 3 - Stall Warning and Identification.............................................. 7-6Figure 4 - Stall System Pitot Static Block Diagram................................. 7-7

SYSTEM FAULTS and ANNUNCIATIONS .................................................. 7-8Figure 5 - System Fault Annunciators .................................................... 7-8

FLAP ASYMMETRY ................................................................................... 7-9WEIGHT-ON-WHEELS ASYMMETRY....................................................... 7-9SSU SELF TEST ........................................................................................ 7-9

SYSTEM ANNUNCIATORS........................................................................ 7-10Table 1: Summary of Annunciators ...................................................... 7-10

SYSTEM POWER SUPPLIES..................................................................... 7-10Table 2: Power Supplies....................................................................... 7-10

STALL WARNING and IDENTIFICATION LOGIC ..................................... 7-11Figure 6 - Stall System Flow Logic ....................................................... 7-11




GENERAL

NOTE: This section applies to 800XP Pro Line 21 airplanes prior to Serial No. 258675. For airplanes Serial Nos. 258675 and subsequent, refer to Section VIII SUPPLEMENTS for Supplement P/N 140-590032-0041.

A stall warning and identification system is provided to emphasize the airplane’s natural cues available at the point of stall.

The system functions are

• Stick shaker (warning)

• Stick pusher (identification)

The system consists of

• Two sensing channels, each utilizing an airflow angle sensor vane.

• Two Signal Summing Units (SSUs).

• Two stick shaker motors.

• A hydraulic operated actuator with two electro-hydraulic valves.

• A third sensing channel, which utilizes pitot pressure from the right pitot head, static pressure fromthe forward static plates and stall vent pressure from the left and right stall vents.

• A stall identification sensor.

• Annunciators and test switches.

SYSTEM LOGIC1. It is impossible for a stick push to occur before a stall warning (stick shake).

2. No single active fault of an SSU or relay can cause the operation of a stall valve or the associatedred STALL VALVE annunciator.

3. The autopilot is disengaged when a stall warning signal is initiated. This prevents the autopilot fromattempting to counteract the resulting stick shake operation or a subsequent stick push.

CONTROLS and ANNUNCIATIONSSystem faults are indicated on two groups of amber annunciators, one per pilot.

Three STALL switches are located in the TEST section of the overhead roof panel. Anti-icing heating of the airflow sensor vanes is controlled from the PITOT/VANE HEAT switches.

Indication of failure of the vane heaters is provided by the two amber L and R VANE HTR FAIL annunciators also located on the overhead roof panel and by the ICE PROT repeat annunciator on the MWS panel.

There are no control switches, the stall warning part of the system becomes armed on takeoff (no weight-on-wheels), while the stall identification part of the system becomes armed 6 seconds after takeoff.




OPERATION

STALL WARNING SYSTEMThe stall warning system uses an electrically driven stick shaker on each control column to provide a physical warning of an approaching stall to the pilots.

Angle of attack is derived from two electrically heated airflow angle sensor vanes mounted one on each side of the forward fuselage. Electrical signals proportional to vane angle are sensed in the associated Signal Sensor Unit which also receives inputs relating to flap angle. From the flap angle signal, the SSU calculates the point of stall warning. When the vane angle corresponds to that point the SSU provides an output to operate the stick shaker motor on each control column.

The stall warning system is inhibited while the airplane is on the ground with weight-on-wheels to prevent wind gusts triggering false stick shaker operations. At takeoff, the system is armed and begins monitoring the pitch attitude of the airplane.

STALL IDENTIFICATION SYSTEMThe stall identification system uses a hydraulic stick pusher to force the control column forward (pitch down) at the calculated point of stall. The stick pusher is powered by main hydraulic system pressure (backed by the main accumulator), the rate of operation being controlled by a fluid restrictor.

Control of the hydraulic pressure to the stick pusher is via two independent stall valves (A and B), connected in series and mounted integral with the stick pusher. Both stall valves must be open to activate the unit.

The output from one SSU energizes one stall valve, and an associated red STALL VLV (A or B) OPEN annunciator is illuminated to indicate that the valve is being signalled to open (reference Figure 1).

The SSU uses the vane sensor and flap angle inputs to calculate the point of stall. The rate of increase of vane angle may also modify the calculation, i.e. the point of stall being advanced when the rate of increase is high dynamic stall (reference Figure 2).

When the vane angle agrees with the calculated point of stall, the SSU produces an output (stall identification) to energize the associated stall valve and annunciator.

When a SSU produces a stall warning output, which is then followed by an identification signal, the warning signal latches the stall identification relay for the same channel (reference Figure 3). A warning signal from either channel will disconnect the autopilot and operate the stick shaker. The latch is removed when the warning output ceases. This makes sure the stick push is maintained until the airplane has reached a nose-down attitude well below the stall point.

M6950_0 HA00C 017045AA.AI

STALL VLV A OPEN

STALL VLV B OPEN

Figure 1Stall Valve Annunciators




0 1 2 3 4 514

16

18

20

22

VANE RATE - DEGREES/SECOND

24

26

6 7 9

28

FLAP 0° WARN

FLAP 15/25° WARN

FLAP 45° WARN

FLAP > 0° IDENT

FLAP 0° IDENT

VA

NE

AN

GLE

to H

FD

(D

egre

es)

Figure 2Variation of Stall Identification Angle with Rate of Increase of Vane Angle

NOTE: HFD is the Horizontal Fuselage Datum.




THIRD STALL IDENTIFICATION CHANNELA third channel for stall identification is provided by a pitot/static system which uses pitot pressure (P2) from the right pitot head, static pressure (S6) from the forward static vents and a vent pressure (V) from left and right stall vents. The stall vents are located on the under-side of the left and right wings (reference Figure 4).

These pressures are sensed by a capsule operated stall detector, which is set to produce an output at a point between the settings for the stall warning and identification signals from the SSUs.

The output from the third channel sensor energizes a relay which connects the stall identification output from one channel’s SSU to the stall valve of the other channel. Thus, with the third channel output activated, both stall valves A and B will open following a stall identification output from only one SSU, thereby ensuring system integrity should a SSU fail.

VANEANGLESENSOR

INPUT WARN

IDENT

SSU

LEFT and

STICK -

MOTORS

INPUT

WARN

IDENT

SSUVANEANGLESENSOR

3rd CHANNELRIGHT

SHAKER

STALLVALVE A

3rd CHANNELSTALL VENT

3rd CHANNELSTALL VENT

HYDRAULICPRESSURE

STICK PUSHER

FLAPANGLE

STALLVALVE B

Figure 3Stall Warning and Identification

PITOTPRESSURE

STATICPRESSURE




Figure 4Stall System Pitot Static Block Diagram

Left ForwardStatic Vent Plate

6

Right ForwardStatic Vent Plate

6S6

P

S

StallDetector

RightPitotHead

P2PitotIsolationValve

Stall Ident(Autopilot Disconnected)

S6

P

S

P2

StallDetector

V1

V

V

V2

Stall Ident3rd Channel

LEGEND

= Pitot Pressure = Pitot Drain

= Static Pressure = Static Drain

= Stall Vent Pressure = Stall Vent Drain

LeftStallVent

RightStallVent




SYSTEM FAULTS and ANNUNCIATIONSThe duplication of the stall warning system makes sure a single passive fault cannot prevent a stick shake occurring. Should an active fault develop, the faulty system can be isolated via the appropriate circuit breaker.

The remaining good system will still operate both stick shaker motors. The stall identification system is designed that a single active fault cannot give an inadvertent stick push, while making sure a single passive fault would not prevent a push operation occurring, when required.

All annunciator warnings in the stall warning and identification system will also cause a repeater STALL IDENT annunciator on the MWS panel to illuminate (reference Figure 5). The power to energize a stall valve is routed via the identification relay of one channel and the warning relay of the other channel.

A monitoring circuit will cause an IDENT 1 or IDENT 2 annunciator (depending on the channel at fault) to be illuminated after a 4 second time delay, if an identification signal from one SSU has been triggered without a warning signal from the other SSU.

The IDENT 1 or IDENT 2 annunciators are part of the IDENT/INHIB switches provided for both pilots. When an IDENT annunciator illuminates, either pilot can push the associated switch to inhibit the faulty channel. The INHIB annunciator part of the switch will then illuminate. The third channel sensor, together with the remaining SSU would provide a stick push operation when required.

The pilots may attempt to reset the failed channel by operating an INHIBIT RESET switch located on the stall diagnostic panel (forward side of the Pilot’s bulkhead).

Should the channel fail again after reset, it should be inhibited and left in that condition for the remainder of the flight.

M6949_0 HA00C 017044AA.AI

Figure 5System Fault Annunciators

MWS Panel




FLAP ASYMMETRYIf asymmetry between the left and right flap positions exists for more than 2 seconds, then the FLAP annunciator (Pilot’s group only) will illuminate.

WEIGHT-ON-WHEELS ASYMMETRYIf asymmetry between the positions of the left and right weight-on-wheels relays exists for more than 4 seconds, then a SQUAT annunciator (Copilot’s group only) will illuminate.

SSU SELF TESTA built-in test within the SSU detects the following faults:

• Loss of 26 VDC supply

• Loss of internal power or short circuit

• Loss of airflow angle sensor excitation

• Loss of flap position input

• Airflow angle sensor transformer winding open or short circuit

If any of the above occur, an external relay causes a SSU annunciator to illuminate. If the weight-on- wheels switch is in the flight condition, a magnetic indicator associated with the faulty SSU will display white. The magnetic indicators are located on the stall identification diagnostic panel.




SYSTEM ANNUNCIATORS

SYSTEM POWER SUPPLIES

Table 1: Summary of Annunciators

Function Annunciator MWS Panel

Airflow angle sensor left or right vane heating malfunction

VANE HTR FAILL (R)

Flap position asymmetry FLAP

Stall identification channel 1 (2) (3) fault

IDENT 1 (2) (3)

Signal Summing Unit fault SSU

Weight-On-Wheels asymmetry

SQUAT

Stall valve A (B) operating STALL VLV A (B) OPEN

Table 2: Power Supplies

Panel Location Row/Column Circuit Breaker Circuit or Equipment Busbar

DA-D B/1

STALL IDENT 1 26 VAC input to SSU 1 XS 1

DA-D B/4

STALL VLV A 28 VDC to stall valve 1 and annunciators

PS1(a)

DA-D B/6

STALL WARN MOTOR 1 (LH)

28 VDC to stick shaker motor 1 PS1(a)

DA-D B/2

STALL IDENT 2 26 VAC input to SSU 2 XS 2

DA-D B/5

STALL VLV B 28 VDC to stall valve 2 and annunciators

PS2(a)

DA-D B/7

STALL WARN MOTOR 2 (RH)

28 VDC to stick shaker motor 2 PS2(a)

DA-DB/3

STALL IDENT 3 28 VDC to SSU 3 PE

# VANEHTR FAIL

ICEPROT

FLAP

IDENT #

STALLIDENT

SSU

SQUAT

STALLVLV #OPEN


Page 7-11Sub-section 7STALL WARNING/IDENTIFICATION

Pilot’s Operating ManualRevision A2: Nov, 2004

STALL WARNING and IDENTIFICATION LOGIC (Figure 6)

Figure 6Stall System Flow Logic

START

ANGLE OF ATTACK MONITORED BY AIRFLOW ANGLE SENSOR

STALL WARNING TRIGGER POINT &STALL IDENTIFICATION TRIGGERPOINT MODIFIED BY FLAP ANGLE

ANGLEOF ATTACK

INCREASING?

NO

YES

Channel 1SSU 1 CALCULATES APPROACHINGSTALL AND INITIATES A WARNING

SIGNAL

Channel 2SSU 2 CALCULATES APPROACHINGSTALL AND INITIATES A WARNING

SIGNAL

WARNING SIGNAL FROM EITHER CHANNELDISCONNECTS AUTOPILOT AND OPERATES

ANGLEOF ATTACK

INCREASING?

NO

YES

PREVENTIVEACTION TAKEN

BY PILOT

ANGLE OF ATTACK

STALL IDENTIFICATION TRIGGER POINTMODIFIED BY RATE OF INCREASE OF ANGLE OF ATTACK. SSU 1 AND/OR

SSU 2 INITIATE IDENTIFICATION SIGNAL WHEN THE AIRPLANE IS AT POINT OF STALL

APPROACHING STALL SENSED ATUNDER-WING PRESSURE VENTS

3rd CHANNEL SENSOR IS ACTIVATED

Channel 3

BOTH SSU IDENTSIGNALS OPERATED

ONLY ONE SSUIDENT SIGNAL

OPERATEDNEAR

TO STALL?

NO

3rd CHANNEL RELAY ENERGIZESTO CONNECT ACTIVE SSU OUTPUT

TO BOTH STALL VALVES

YES

OR

STALL VALVE AOPEN AND

STICK PUSHER ACTIVATED

STALL VALVE BOPEN

CONTINUES TO INCREASE

THE STICK SHAKER


Page 7-12 Pilot’s Operating ManualRevision A2: Nov, 2004





Pro Line 21


Sub-section 16WATER/WASTE

Table of Contents

Page

GENERAL ................................................................................................. 16-3WATER SYSTEM...................................................................................... 16-3

LAVATORY WATER TANK.................................................................... 16-3LAVATORY WATER PUMP ................................................................... 16-3GALLEY WATER.................................................................................... 16-4

Figure 1 - Galley Master Switch and Typical Galley........................... 16-4WASTE SYSTEM...................................................................................... 16-5

TOILET ................................................................................................... 16-5Figure 2 - Typical Toilet Compartment ............................................... 16-5Figure 3 - Typical Toilet with External Servicing Facility .................... 16-6Figure 4 - Toilet Ground Servicing...................................................... 16-7

2 Pilot’s Operating ManualInitial Issue: Feb, 2002

Sub-section 16WATER and WASTE






GENERAL

The water and waste system consists of a basin with heated running water located in a lavatory at the rear of the passenger cabin and an electrically flushing toilet with external servicing facilities.

WATER SYSTEM

Wash water is stored in a 2.3 gallon heated water tank with a water pump controlled by operation of the faucet.

LAVATORY WATER TANK

The water tank contains a triple element low voltage immersion heater, controlled to 40° C ± 2°, and water level microswitches. Provided the tank is full, the water heater comes into operation immediately when the TOILET WASH WATER push switch, on galley panel FG-A, is selected. The switch light illuminates to show system operation.

The water level microswitches operate as follows:

• The high-level switch contacts are closed when the tank is full and open immediately the float startsto fall.

• The low level switch contacts open when the minimum water level (just above the elements) isreached to isolate water heater power supplies preventing overheating of the elements.

LAVATORY WATER PUMP

A counterclockwise rotation of the faucet will energize the pump for continuous operation until the faucet is released. The pump is a self contained 12 VDC geared pump designed to prevent leakage and contamination of the water supply and requires little maintenance. The pump boosts air pressure into the water lines at 20 psig (approximately) above cabin air pressure.

Electrical power for the water pump is normally provided from busbar PS2 through the GALLEY POWER and TOILET WASH WATER switches on the galley switch panel. For ground operation, the system may be connected to the No. 1 battery through the ROOF/STEP light switch selected to the ON position. Pushing the drain button allows the basin contents to drain away to a heated overboard drain mast.




GALLEY WATER (Figure 1)

The galley water system, depending on options selected, has either a heated 2.23 gallon water tank, an unheated 2.07 gallon water tank or both.

With the heated water tank, the water temperature is thermostatically controlled to approximately 100° by factory setting and cannot be adjusted. The water tank has a spigot and tube assembly located on the front face of the tank to deliver the water to a cup or glass.

When the GALLEY MASTER switch on the interior lighting control panel is pushed, the galley busbar is connected to the PE busbar and indicated by the illumination of a small LED in the top left corner of the switch. This provides a power supply, via circuit breakers, for the galley electrical equipment.

CUSTOM WATER CONTAINER

CUPDISPENSERS

WASTE

COFFEEBREWER

SWITCH PANEL

MICROWAVE

COUNTERSTORAGE

SANDWICHTRAY

DRAWER

UTENSILS

CONDIMENTS

PULL OUT WORK

SURFACE

BOWLS

PLATES

MINIATURES

WINE / SPIRITS STORAGE

MISC.STORAGE

NAPKINS

SODA CANS

ICE / COLD STORAGE

Figure 1Galley Master Switch and Typical Galley

MISCELLANEOUSSTORAGE

GALLEYUP

LIGHTS

GALLEY SWITCH PANEL

GALLEYWORK

LIGHTS

COFFEEHEAT

WATERHEAT

INTERIOR LIGHTING

GALLEY MASTER SWITCH

REFUEL PANEL

CONTROL PANEL

(PANEL IS LOCATED STRAIGHT ACROSS FROM THE ENTRY DOOR, ON THE FORWARD SIDE OF THE REFUEL PANEL)

FWD

REFUEL PANEL DOOR SHOWN OPEN




WASTE SYSTEM

TOILET (Figures 2 & 3)

An electrical flushing toilet (Monogram 4620-17) with external servicing facilities is installed at the rear of the passenger cabin. The toilet is a self-contained removable unit consisting of a tank, motor, pump and a filter.

Operation of the flushing system is by a PRESS TO FLUSH timer button on the vanity unit. Electrical power is normally provided from PS2 busbar and selection of the ROOF LIGHT switch (panel DA, top inboard face) or the ROOF/STEP LIGHT switch (forward vestibule cabinet, rear face) to ON connects a battery No. 1 supply to the coil of the entry lights relay.

Operating the PRESS TO FLUSH timer button will connect the power supply to the motor-driven flushing pump and rotating filter for approximately 8 seconds. Flushing liquid cascades in a thin curtain over the complete inner surface of the toilet bowl from the flushing channel surrounding the upper rim of the bowl. Waste is carried directly to the tank and prevented from re-entry by means of a restrictor in the bottom of the bowl. Flushing liquid is filtered out of the tank through a self-cleaning rotary filter and pumped up to the flush channel (reference Figure 4).

M6359_0HA03C014658AA.AI

VIEW OF TYPICAL LAVATORY

VIEW WITH TOILET COVER REMOVED

Figure 2Typical Toilet Compartment




M6360_0HA03C014659AA.AI

BOWL

GROUND FLUSH INLET

ELECTRICALCONNECTION

MOTOR-PUMP-FILTERCARTRIDGEASSEMBLY

DRAIN VALVE ASSEMBLY

Figure 3Typical Toilet with External Servicing Facility





A

DETAIL A

Figure 4Toilet Ground Servicing

hawker 800 fctm 2

Documents

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icing system

ice protection windscreens

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electrical heating system

ice detection components

engine bleed air