Pilot Training Manual (FS)

8/22/2019 Pilot Training Manual (FS)

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NOTICE

The material contained in this training manual is based on information obtained from

the aircraft manufacturer’s Airplane Flight Manual , Pilot Manual, and Maintenance

Manual. It is to be used for familiarization and training purposes only.

FOR TRAINING PURPOSES ONLY



C f h L j 30 S i h h f ll i Fli h S f l i



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CONTENTS

Chapter 1 AIRCRAFT GENERAL

Chapter 2 ELECTRICAL POWER SYSTEMS

Chapter 3 LIGHTING

Chapter 4 MASTER WARNING SYSTEM

Chapter 5 FUEL SYSTEMChapter 6 AUXILIARY POWER UNIT

Chapter 7 POWERPLANT

Chapter 8 FIRE PROTECTION

Chapter 9 PNEUMATICS

Chapter 10 ICE AND RAIN PROTECTION

Chapter 11 AIR CONDITIONING

Chapter 12 PRESSURIZATION

Chapter 13 HYDRAULIC POWER SYSTEM



CHAPTER 1 AIRCRAFT GENERAL

CONTENTS

Page

INTRODUCTION................................................................................................................... 1-1

GENERAL .............................................................................................................................. 1-1

STRUCTURES........................................................................................................................ 1-2

General............................................................................................................................. 1-2

Fuselage ........................................................................................................................... 1-4

Wing................................................................................................................................. 1-9

Empennage .................................................................................................................... 1-10

AIRCRAFT SYSTEMS........................................................................................................ 1-10

LEARJET 30 SERIES PILOT TRAINING MANUAL



ILLUSTRATIONSFigure Title Page

1-1 Learjet 35/36 ............................................................................................................ 1-2

1-2 General Dimensions................................................................................................. 1-2

1-3 Turning Radius......................................................................................................... 1-3

1-4 Danger Areas............................................................................................................ 1-3

1-5 Fuselage Sections..................................................................................................... 1-4

1-6 Radome .................................................................................................................... 1-5

1-7 Nose Compartment .................................................................................................. 1-5

1-8 Passenger-Crew Door............................................................................................... 1-51-9 Door Latch Inspection Port ...................................................................................... 1-6

1-10 Emergency Exit........................................................................................................ 1-7

1-11 Windshield ............................................................................................................... 1-8




CHAPTER 1 AIRCRAFT GENERAL




STRUCTURESGENERAL

Figure 1-1 shows the Learjet 35/36. The structure consists of the fuselage, the wing, theempennage, and flight controls. The discussionof the fuselage includes all doors and windows.Figure 1-2 shows the general dimensions of

the aircraft.

Figure 1-3 displays the aircraft turning ra-dius.

Figure 1-4 displays the danger areas around the

Learjet 35/36 presented by the weather radar emission cone, engine intakes, and engine ex-haust cones.


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14 FT447.0 cm

Figure 1-1. Learjet 35/36




42 FT



FUSELAGE

General

The fuselage is constructed of stressed all-metal skin with stringers. It employs the arearule design to reduce aerodynamic drag, and has four basic sections (Figure 1-5). They are:

• The nose section that extends from the

radome aft to the forward pressure bulk-head.

• The pressurized section, which includesthe cockpit and passenger areas, extendsaft to the rear pressure bulkhead. On 36models this bulkhead is further forward than on 35 models to provide space for the larger fuselage tank.

• In both models, the fuselage fuel sectionstarts just aft of the rear pressure bulk-head and extends to the tailcone.

• The tailcone section extends aft of thef el section

Nose Section

The nose of the fuselage (Figure 1-6) is formed by the radome. Aft of the radome is the nosecompartment.

The nose compartment access panels are on topof the fuselage (Figure 1-7), forward of thewindshield. The panels must be removed for access to various electronic components,

oxygen bottle (when installed in the nose),emergency air bottle, and the alcohol anti-icing reservoir.

Pressurized Section

The pressurized cabin lies between the forward p ressu re bu lkhead and the a f t p ressu re bulkhead, and includes the cockpi t and

passenger compartment. Within the passenger compartment is a 500-pound-capacity baggagearea at the back of the cabin, a lavatory, a cabinet for storage of provisions, and galleyequipment (depending on the aircraft).

Th d i l d h l f


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Passenger-Crew Door

The primary entrance and exit for passengersand crewmembers is through the clamshelldoor, located on the left side of the forward fuselage (Figure 1-8). The standard entrancedoor is 24 inches wide, but there is an op-tional 36-inch door. The upper door serves asan emergency exit, and the lower door has in-tegral entrance steps.

The upper portion of the door has both outsideand inside locking handles connected to acommon shaft through the door. Rotating either of these handles to the closed position drivessix locking pins into holes in the fuselageframe (three pins forward and three aft) and two pins through interlocking arms that securethe two door halves together.

The lower door has a single locking handle on

the inside. Rotating the lower door handle tothe closed (forward) position drives two pinsinto holes in the fuselage frame (one forward and one aft). There are a total of 10 locking pins on the two door sections.

To facilitate alignment of the upper door locking p ins during c losing , an e lec t r ic actuator motor, torque tube assembly, and one

or two hooks are installed in the lower door.The hooks engage rollers installed on the upper door and draw the two halves together. The actuator motor is operated from inside the aircraft by a toggle switch on the lower door,and from the outside by a key switch. Should the motor fail, the hooks can still be operated manually from inside. Access is provided tothe torque-tube mechanism through a panel in




the lower door, and a ratchet handle provided

in the aircraft tool kit can be used to operatethe torque-tube manually.

NOTE

One hook and roller is used on 24-inch doors, while two hooks and rollers are used on 36-inch doors.

When the door handles are in the closed posi tion, the pins al l contact microswi tches.If any one of the switches is not actuated, a red DOOR light illuminates on the annunciator panel. (See Annunciator Panel sect ion.) If thelight illuminates while the door is closed,

eight inspection ports enable the crew to

confirm the position of the door frame latching pins by observing the posi tion of two whitealignment marks (Figure 1-9). The two latch pins that connect the upper and lower doors arevisible through the upholstery gap at the in-terface and do not have white lines.

When closing the doors from the inside, closeand latch the lower door f irst. Then, close the

upper door and actuate the door motor switchto the closed position. This engages the hooksover rollers in the upper door, and cinches theupper door down tight while allowing the locking pins to line up properly and meet themicroswitches as the upper door handle is rotated to the closed position. The DOOR light will remain illuminated until the hooksare backed away from the upper door rollers

by reverse operation of the door motor switch.

A secondary safety latch is installed on thelower door and is separate from the door- locking system. It consists of a notched pawlattached to the door. The pawl engages a striker

l h d h f h h d i


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NOTE

Anytime the aircraft is occupied withthe entry doors locked, the hooksmust be released. This permits open-ing the upper door for emergencyegress.

If the red DOOR light illuminates, it means:

• Any one of the 10 latch pins is not engaged with its respective microswitch

• The hook drive mechanism is not completely ret racted

• The door is unsafe for takeoff

A hollow neoprene seal surrounds the door-

frame; the seal has holes to allow the entry of

pressurized cabin air, forming a posit ive seal

around the door.

Emergency Exit

A hatch near the right rear of the cabin (Figure1-10) serves as an emergency exit for all occupants. A latching mechanism is accessiblefrom inside and outside the cabin.

The inside latch handle, located at the topcenter of the window, is pulled inward to unlock. To open from the outside, depressinga PUSH button above the window releases ahandle that must then be turned in the directionof the arrow stamped on the handle; then thehatch may be pushed inward.




Windows

Windshield

The windshield (Figure 1-11) is divided intotwo sections, the pilot and copilot halves, and is made up of three laminated layers of acrylic plas tic. The windshield is approximate ly oneinch thick. It is impact-resistant, heated or not, and was tested against 4-pound bird strikes

at 350 knots.

Passenger Windows

The cabin windows (Figure 1-12), includingthe emergency exit window, are made up of two panes of stre tched acryl ic plastic with an air space between them. They are held apart and sealed air tight by a spacer.

Fuel Section

The fuel section, located aft of the rear pres-sure bulkhead, contains the fuselage fuel cells.

The fuel section on 35 models is differentfrom that on 36 models (see Figure 1-5). On36 models, the rear pressure bulkhead has been moved forward, al lowing for four blad-der cells rather than two, almost doublingfuselage fuel capacity.


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Figure 1-12. Windows Locations (Typical)



WING

The Learjet 35/36 has a swept back, can-tilevered, all metal wing (Figure 1-14) that ismounted to the lower fuselage and joined to-gether at the fuselage. Most of the wing issealed to form an integral fuel tank.

Eight fi t t ings at taching the wings to the fu s e l a g e a r e d e s i g n e d t o p r e v e n t w i n g

deflections from inducing secondary loads inthe pressurized fuselage. Ailerons are attached to the rear spar at three hinge points. The single-slotted Fowler flaps are attached to theinboard rear spar by tracks, rollers, and hinges.The spoilers are attached to the top of thewing surface by two hinges just forward of theflaps. The tip tanks are secured to the wing attwo attach points.

The Learjet 35/36 wing is fitted with either vortex generators or boundary layer energizers.Whichever is used, they function to delay airflow separation over the ailerons at highMach numbers.

Subsequent serial-numbered aircraft and those

modified with AAK 79-10 incorporate a soft-flight wing modification, which includes:

• Three rows of boundary layer energiz-ers (BLEs) on each wing that perform thesame function as vortex generators, butare more efficient. If any are missing,MMO is reduced to 0.78 M1 (FC200) or .77 M1 (FC530)

• A full-chord stall fence on each wing, inboard of the aileron, which delays disruption of the airflow over the aileronat high angles of attack




• A stall strip, affixed to the inboard sec-

tion of each wing leading edge, whichgenerates a buffet at high angle of attack to warn of an impending stall

• An aileron gap seal along the leadingedge of each aileron

EMPENNAGE

The high-T-tail empennage (Figure 1-15) includes a vertical stabilizer with an attached rudder and a horizontal stabilizer with attached elevators.

The swept back vertical stabilizer is formed by f ive spars secure ly connected in the tail -cone. It is the mounting point for the rudder and horizonta l stabilizer. At the lower leading

edge of the stabilizer is a dorsal fin that housesa ram-air scoop. Later model airpcraft have theoxygen bottle located within the dorsal fin.

The horizontal stabilizer is a swept back, fullspan unit, constructed around f ive spars. It is

h d h i l bili i

• The center leading edge attaches to an

e lec t r ica l ly opera ted screwjack to provide pi tch axis t rim.

AIRCRAFT SYSTEMS

ELECTRICAL POWERSYSTEMS

Primary DC electrical power is provided bytwo engine-driven generators. Secondary power is suppl ied by two 24-volt ba tt er ie s.The aircraft may be equipped with a single or dual emergency battery system. The aircraftalso has the capability of accepting DC power from a ground power unit.

DC power is used by either two or three solid-state static inverters that, in turn, supply AC power for equipment and ins truments .

LIGHTING

Interior lighting is supplied for general cockpit


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FUEL SYSTEM

Fuel is contained in integral wing tanks, tiptanks, and in a bladder cell fuselage tank justaft of the rear pressure bulkhead. The 36 modelhas a larger fuselage tank than the 35 model.

Fueling is accomplished through f iller caps inthe top of each tip tank.

POWERPLANT

The Learjet 35/36 is powered by two GarrettTFE731 turbofan engines. The TFE731 is alightweight, two-spool, front fan-jet engine.It has a reverse-flow annular combustion chamber that reduces the overal l length and resul tsin more eff icient combustion and cooler external surfaces of the turbine section.

The low-pressure rotor consists of a four-stage, axial compressor and a three-stage,axial turbine rotating on a common shaft. The axial-flow fan assembly is located at the for-ward end of the engine and is gear-driven by

control, and the Aeronca thrust reversers,

if installed.

ICE AND RAIN PROTECTION

The anti-icing systems use engine bleed air,electric heating, and alcohol.

Bleed air is used to heat the wing leading edge,the horizontal stabilizer leading edge, wind-shields, nacelle lips, and on some aircraft, theengine fan spinners. Bleed air is also used toremove rain from the windshield.

Electrically heated systems include pitottubes, static ports, P2T2 sensors, and the stallwarning vanes.

An alcohol system is used for radome anti-icing and to back up the pilot windshield bleed-air anti-icing.

AIR CONDITIONING AND PRESSURIZATION




HYDRAULIC POWER SYSTEMS

The hydraulic system supplies pressure for the operation of the landing gear, gear doors, brakes, flaps, spoilers, and Dee Howard thrustreversers, if installed. A single reservoir supplies fluid to the two engine-driven pumpsthrough fire shutoff valves.

An electric auxiliary pump can pressurize all

systems except the spoilers. It draws fluid from the same reservoir. The auxiliary supplyline is not affected by the fire shutoff valves.

LANDING GEAR AND BRAKES

The Learjet 35/36 has a retractable tricyclelanding gear that is electrically controlled and hydraulically operated.

An emergency air bottle, located in the rightside of the nose compartment, can be used toextend the landing gear or for emergency braking, or both, in case of hy draulic or electrical failure.

Secondary flight controls (spoiler/spoileron

and flaps) are electrically controlled and hydraulically operated.

AUTOMATIC FLIGHT CONTROLSYSTEM

The automatic flight control system (AFCS) includes a fl ight director, autopilot , and

yaw dampers.

The flight director system generates roll and pit ch co mmands by me ans of a si ng le -cueV-bar display in the pilot attitude director indicator. Programming and annunciation of selected modes is accomplished on the AFCScontrol panel in the center glareshield.

The two-axis autopilot provides control of therol l and p i tch axes . When engaged, the autopilot responds to the flight director as programmed, or the pi lot may elect to operatethe autopilot in a basic attitude-hold mode bycanceling all flight director modes, in whichcase the command bars are biased out of view


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sensor uses the copilot pitot line for pitot pres-

sure, while its static pressure is provided bytwo additional heated static ports installed onthe nose, forward of the windshield. An alternate unheated static port inside the nosecompartment is provided for the pilot staticsystem.

FC 530 models use a Rosemount-designed pi tot-stat ic system that physically integra tes

two static ports into each of two pitot tubes,one mounted on each side of the nose section.The air data sensor uses the copilot pitot and static lines.

An unheated static port is located on the right

side of the nose compartment to provide astatic source for the pressurization controlmodule.

OXYGEN SYSTEM

The oxygen system consists of the crew and passenger distribution systems connected to ahigh-pressure oxygen storage cylinder located in the nose compartment on early 35 and 36models. On SNs 35-492 and 36-051 and sub-sequent, the cylinder is located in the verticalstabilizer.




CHAPTER 2ELECTRICAL POWER SYSTEMS

CONTENTS

Page

INTRODUCTION................................................................................................................... 2-1

GENERAL ............................................................................................................................. 2-1

DC POWER............................................................................................................................. 2-2

Batteries ........................................................................................................................... 2-2

Generators........................................................................................................................ 2-5

Ground Power .................................................................................................................. 2-6

Circuit Components ......................................................................................................... 2-6

Distribution.................................................................................................................... 2-12




ILLUSTRATIONS

Figure Title Page

2-1 Component Locations .............................................................................................. 2-2

2-2 Battery Location....................................................................................................... 2-2

2-3 Battery Switches ................................................................................................. 2-3

2-4 Equipment Powered by Battery Charging Bus and Generator Buses ......................... 2-3

2-5 Generator Indicators ............................................................................................ 2-4

2-6 Battery Temperature Indicator .............................................................................. 2-5

2-7 Generator Location ............................................................................................. 2-5

2-8 Generator Switches ............................................................................................. 2-5

2-9 Ground Power Connector ..................................................................................... 2-6

2-10 Basic DC Distribution ......................................................................................... 2-7

2-11 Current Limiter Panel .......................................................................................... 2-7




2-22 AC Distribution ................................................................................................ 2-19

2-23 Emergency Battery Controlsand Indicators.......................................................... 2-20

2-24 Electrical System—SNs 35-002 through 35-205 and 36-002 through

36-040 (Not Incorporating AMK 78-13) ............................................................. 2-22

2-25 Electrical System—SNs 35-202 through 35-204, 35-206 through 35-508,

36-041 through 36-053, and Prior Aircraft Incorporating AMK 78-13 ................... 2-23

2-26 Electrical System—SNs 35-509 and Subsequent, 36-054 and Subsequent, and Prior Aircraft Incorporating AMK 85-1 ...................................... 2-24


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CHAPTER 2ELECTRICAL POWER SYSTEMS




The batteries are capable of operating the min-

imum equipment for night instrument flight for approximately 30 minutes if both generators become inopera tive. An emergency bat tery is provided to operate an emergency at ti tudegyro and the gear and flap systems if a totalaircraft electrical system failure occurs.

It is possible to power the entire DC and ACelectrical systems from the aircraft batteries,

an engine-driven generator, or a GPU.

Figure 2-1 shows major electrical power systemcomponent locations.

DC POWER

BATTERIES

Two batteries in the tail cone (Figure 2-2) pro-vide the secondary source of DC power.

Each battery has a removable cover and a case

Lead-Acid vs. Nicad

Lead-acid batteries are enclosed in a plastic

case. Nickel-cadmium (nicad) batteries areenclosed in a stainless steel case . On SNs 35-341 and 36-050 and subsequent equipped withlead-acid batteries, a sump jar has been added to contain any electrolyte spillover. A spongesaturated in a baking soda and water solutionneutralizes the acid AMK 81 5A makes this


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Figure 2-2. Battery Location



Charging nicad batteries with a GPU is not

recommended.

Charging lead-acid batteries in the aircraft isnot recommended because of poor GPU outputregulation.

Controls

Two battery switches (Figure 2-3) connect the

batter ies in para llel to the battery-charging bus when the switches are on. The switches ,BAT 1 and BAT 2, correspond to the respec-tive battery. Each switch has two-positions,ON–OFF, that complete a ground circuit toclose its respective battery relay in the on posit ion (Figure 2-4) .

The battery relays require approximately 16

V (minimum) from the respective battery. If either battery voltage is less than 16 V, the re-spective battery relay will not close; the bat-tery cannot be connected to the aircraftelectrical system for the purpose of operating

electrical equipment (except equipment hot-

wired to its battery bus), nor can it be charged by a GPU or the generators.


Figure 2-3. Battery Switches



The aircraft batter ies are always connected in

parallel (including during engine star ts) when both battery switches are on.

Indicators

Electrical system indicators (Figure 2-5) aregrouped in a cluster on the upper portion of the center instrument panel. A single DCVOLTS meter, connected to the battery charg-

ing bus through a 5 A cur rent limiter, indicatesthe highest voltage input to the bus by batter-ies, generators, or GPU. To read individual battery voltage, only one battery at a time may be connected to the battery charging bus withthe generators off and a GPU not connected.Aircraft generators and GPUs normally put out

a higher voltage than the batteries; therefore,

when either of these is powering the batterycharging bus, generator, or GPU voltage will be indicated.

Aircraft with nicad batteries are equipped with battery temperature indicators and overheatwarning light systems. These are attached through two electrical connectors on the faceof each battery case to temperature sensors and

thermal switches on each batter y.

A dual-indicating temperature gage is on thelower portion of the copi lot instrument panel(Figure 2-6). Two red warning annunciators labeled BAT 140 and BAT 160 in the annun-ciator panel illuminate if either or both bat-teries reach 140 to 160°F, respectively.


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GENERATORS

The generators supply DC power to all DC

powered equipment on the a ircraf t.

Generator voltage is regulated to 28.8 VDC for lead-acid batteries and to 28.5 VDC for nicad batter ies. On aircraft SNs 35-148 and subse-quent and 36-036 and subsequent, single-generator voltage is reduced as load increasesduring ground operation and any time a starter is engaged to limit amperage. This design

feature protects the 275 A current limitersduring engine start. The generator control panel in the tail cone contains re lays for the batter ies, star ters , GPU overvoltage contro l,and an equalizer circuit for load sharing.

Controls

Two starter-generator switches on the center

switch panel (Figure 2-8) are three- posi tionswitches labeled GEN, OFF, and START. InGEN, current is provided to the generator f ield through the IGN & START circuit breaker,which automatically connects the generator bus; the amber GEN caution annunciator ex-


Figure 2-6. Battery Temperature Indicator



switch panel (see Figure 2-8) provide for re-

setting the generator in case of failure. If theGEN-OFF-START switch is in GEN, mo-mentarily depressing the reset button resets theovervoltage relay, completes a power circuitto the voltage regulator, and restores the gen-erator to normal operation.

Indicators

Two AMPS meters—one for each generator— indicate the load in amps being carried byeach generator (see Figure 2-5). The load indication is measured at the voltage regulator.

Generator voltage is displayed on the DCVOLTS meter.

An amber L or R GEN caution annunciator on

the glareshield panel illuminates if the asso-ciated generator switch is turned off, if the generator fails, or if the generator is tripped off by the overvoltage cutout relay.

GROUND POWER

The receptacle connects GPU power to the

batter y charging bus through a power re laycontrolled by an overvoltage circuit. The over-voltage circuit samples GPU voltage provided through a control relay (Figure 2-10). At leastone battery switch must be turned on to closethe control relay, allowing the overvoltagecircuit to sample GPU voltage , and, if below33 V, the power relay closes to complete theGPU-to-battery charging bus connection.

The GPU should be regulated to 28 V. Due totower shaft torque limits, it must be limited to1,100 A for engine starts. It should be capable of producing at leas t 500 A. If GPU volt-age exceeds 33 V, the overvoltage circuitcauses the power relay to open, thereby dis-connecting the GPU from the electrical sys-tem to prevent damage to voltage-sensitive

equipment.

CIRCUIT COMPONENTS

Current Limiters


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with a small window that allows visual in-

spection of current- limiter integrity.

There are two 275 A current limiters in themain current-limiter panel; these connect thegenerator buses with the battery charging bus.

• On SNs 35-002 through 35-147 and 36-002 through 36-035, testing of these current limiters is accomplished manu-

ally.

• On SNs 35-148 through 35-389, except35-370, and 36-036 through 36-047, test-ing of the current limiters is accomplished using the rotary systems test switch.

For all of the above aircraft, AMK 80-17 pro-vides two amber annunciators, one for each

current limiter, which allow continuous mon-itoring.


LEFT ENGINESTART FUNCTIONS

RIGHT ENGINESTART FUNCTIONS**

Figure 2-11. Current Limiter Panel



On SNs 35-370, 35-390, and 36-048 and sub-

sequent, a single red CUR LIM annunciator onthe glareshield panel allows continuous mon-itoring of the 275 A current limiters.

The 275-amp current-limiter annunciator(s)are illuminated by 1 A overload sensors wired across the current- limiter terminals. Failure of a current limiter results in a surge of currentthrough the overload sensor, causing it to tr ip

and thereby illuminate the light.

In flight, it is important to know if the currentlimiters have blown. On all aircraft with or without current limiter annunciator(s), currentlimiter status may be determined by close ob-servation of voltmeter and ammeter indica-tions. If only one fails, no difference will benoted on either indicator since power from

each generator still flows to the battery charg-ing bus through the opposite current limiter.

Failure of both current limiters, however, could be recognized since the voltmeter will read battery voltage (i .e ., < 25V). On aircraft prior

SN 35 509 d 36 054 difi d b

equipment that draw power from either bat-

tery bus or the battery charging bus. Batterycondition should be monitored using the DCvoltmeter.

On aircraft with the single CUR LIM annun-ciator, if one limiter blows in flight, DC voltsand amps should be monitored closely sincethe CUR LIM annunciator remains illumi-nated and will not alert the pilot to subsequent

failure of the othe r limiter.

Relays

Relays are used at numerous places throughoutthe electrical distribution system, particularlyin circuits with heavy electrical loads. The relaysfunction as remote switches to make or break

power circuits. This ar rangement al lows thecontrol circuit wiring to be a lighter gage sinceless current is required to operate the relay.Relays control the power circuits for the bat-teries, GPU, starters, generators, inverters, and left and right main buses. Instrument panelswitches or CBs complete the control circuits


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Circuit Breakers

Circuit breakers are on two CB panels in thecockpit, one left of the pilot seat and oneright of the copilot seat. On FC 200 AFCSaircraft, three additional CBs under the pilotseat on the autopilot electric box provide power for the autopilot flight director and the yawdamper annunciator lights.

The DC circuit breakers are the thermal type,and the AC circuit breakers are the magnetictype. Amperage ratings are stamped on thetop of each CB.

The CBs are arranged in rows according to the buses that serve them to simplify the isolationof individual buses or circuits. Basically, allCBs in the top row (both sides) are on the 115

VAC and 26 VAC buses; in the second rowthey are on the main DC buses (except threethat are power bus CBs). Additionally, thrustreversers (if installed) are controlled by main bus CBs that are physically installed on the leftand right panels, third and fourth rows.

The third and fourth rows on SNs 35-002

through 35-201 and 35-205, and 36-002through 36-040 are on the DC essential bus.On SNs 35-202 and subsequent, except 35-205,and 36-041 and subsequent, and earlier aircraftincorporating AMK 78-13, the third and fourthrows are on the essential A and B DC buses.

Circui t breakers on the th i rd andfourth rows, but not powered by the essential

buses, are:• L STALL WARN, DOOR ACTR and

ENTRY LTS (left battery bus items)

• R STALL WRN (right battery bus item)

• T/R EMER STOW and T/R POS IND(left main bus item—Aeronca)

• T / R CO N T ( r i gh t ma i n b u s it e m— Aeronca)

• T/R POWER and T/R CONT (left and right main bus items—Dee Howard)





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S Y S T E M S

PILOT

PANELCOPILOT

PANEL

L ACBUS

AUXBUS

FLOODLT

PRI VM

PRIDIR GY

NOSESTEER

MACHTRIM

AIRDATA

AFCSPITCH

AFCSROLL

PRI

VERT GY

FLTDIR

AT TD

PRI YAWDAMP

ANTISKID

CAB HTAUTO

CABBLD

FREONCONT

RDNGLTS

NOSESTEER

STROBELTS

NAVLTS

HFCOMM

DME 1

ADF 1

PRIINV

L IGN& START

L MAINBUS

L ESSBUS

AUD 1

COMM1

NAV 1

ATC 1

L FANRPM

WRNLTS

INSTRLTS

L PITHT

PRI FLTDIR

R NAVCONT

PRIAFCS

AFCSPITCH

FUSEVAL

OXYVAL

AIRBL

YAW

ROLL

PITCH

RAM AIRTEMP

L ICEDET

R ICEDET

OILTEMP

FUELQTY

FLAPS

GEAR

AUXCOM

SPOILER

L STBYPMP

R STBYPMP

L JETPMP VAL

R JET

PMP VAL

L FWSOV

R FWSOV

L FIREEXT

R FIREEXT

L FIREDET

R FIREDET

L AIRIGN

R AIRIGN

ESSBUS TIE

R ESSBUS

AUD 2

COMM 2

NAV 2

ATC 2

UHF

WRNLTS

INSTRLTS

R PITHT

CABPRESS

VLFRCVR

SECAFCS

TAB FLAP

POSN

SEC

MAINBUS TIE

R MAINBUS

R IGNSTART

SECINV

EMER

BAT 1

EMERBAT 2

ADF 2

R LDG &TAXI LT

BCNLTS

RADLTDM

RADAR

ALCPMP

STEREO

CMPTR

SECVERT GY

SECDIR GY

RADAR

SECVM

E.L.LTS

SPOILERON

R AUXBUS

R ACBUS

AC BUSTIE

SEC YAWCAMP

AIRDATASEN




PILOT

PANELCOPILOT

PANEL

L ACBUS

FLOODLTS

PRI VM

NOSESTEER

MACHTRIM

AIR

DATA

AFCSPITCH

AFCSROLL

FLTDIR

ATTD

L VACHT

CABBLOW

FREONCONT

RDNGLTS

NOSESTEER

STROBE

LTS

NAVLTS

MODEPWR

DME 1

ADF 1

PRIINV

L IGN& ST

L MAINBUS

L ESSA BUS

AUD 1

COMM1

NAV 1

ATC 1

L FANRPM

WRNLTS

AIRBLEED

VLFNAV

CLOCK

AFCSYAW

AFCSROLL

PRIAFCS

YAW

ROLL

PITCH

L PITOTHT

L ICEDET

R ICEDET

OILTEMP

FUELQTY

FLAPS

GEAR

AUXCOM

SPOILER

L STBYPMP

R STBYPMP

L JETPMP VAL

L FWSOV

R FWSOV

L FIREEXT

AUXINV

L FIREDET

R ESSB BUS

L ESSB BUS R ESS

A BUS

AUD 2

COMM 2

NAV 2

ATC 2

R FANRPM

WARNLTS

R FIREDET

R FIRE

EXT

CABPRESS

SECFLT DIR

SECAFCS

MAINBUS TIE

R MAINBUS

R IGN& ST

SECINV

EMER

BAT 1

EMERBAT 2

ADF 2

RECOGLT

R LDG &TAXI LT

BCNLTS

CABINTEMP

RADAR

ALCSYS

STEREOSAT/

SEC

FLT DIR

SEC

VM

E.L.

LTS

SPOIL -

ERON

R AUX

AC BUS

R AC

BUS

AC BUS

TIE

AIR

DATASEN

L AUXAC BUS

PRIDIR

GY

PRIVERT

GY

PRIYAW

DAMP

PRI

LHMODVAL

L ITT

RAMAIR

TEMP

PRI

FLTDIR

FUEL

COMPTR

ESS BBUSTIE

R JETPMPVAL

TAB &FLAPPOSN

ESS ABUSTIE

R ITT

CLOCK

SEC

DIR

GY

SEC

VERT

GY

SECRATE

GYRO

SEC

YAW

DAMP



DISTRIBUTION

The aircraft uses a multiple bus, multiple con-ductor, electrical distribution system. Busesand major circuits are protected by relays,current limiters, overload sensors, and CBs to preclude total fai lure . This arrangement alsoallows isolation of malfunctioning buses. AllCBs are accessible to the crew during flight.

Battery BusesThe left and right battery buses are connected to the left and right batteries, respectively,through 20 A current limiters (see Figure 2-10).The battery buses are always hot, provided the battery quick-disconnects are connected.

The battery buses supply power to the fol-

lowing items:

• Left bat tery bus

° Left stall warning system

° Entry lights (step lights, baggaget t li ht d t il i

Battery bus items must be turned off before

leaving the aircraft to prevent battery discharge.

Battery Charging Bus

The battery-charging bus enables the gener-ators or GPU to charge the batteries and is thecentral distribution point for the DC electricalsystem. It is powered by the batteries and GPUthrough their associated power relays by either

generator through the respective left and right275 A current limiters (see Figure 2-10).

One or both batteries can power the entireelectrical system for a limited period of time,with the exception of the Freon air conditioner and auxil iary heater. Because their high amperage requirement would quickly depletethe batteries, these items are isolated by an

open relay that does not close until a GPU or generator is on and operating.

On SNs 35-002 through 35-508 and 36-002through 36-053, when not incorporat ingAMK 85-1 the essential buses are connected


2

E L E

C T R I C A L P O WE R

S Y S T E M S




BATTERY POWER

GENERATOR POWER

GROUND POWER

LEGEND

DC VOLTS

010

20

30

R ESSL ESS

R

E

G

R

E

G

L

GEN

R

GEN

L

GEN BUS275 A

BAT CHG BUS

L

BAT

GPU

R

BAT

275 A

R

GEN BUS

40 A

50 A

20A

40 A

50 A

Figure 2-14. Essential DC Bus Power—SNs 35-002 through 35-201 and 35-205, and 36-002through 36-040 (Not Incorporating AMK 78-13)



On all aircraft, the following equipment is di-

rectly connected to the battery charging bus(Figure 2-16):

• DC VOLTS meter

• Freon air conditioner and auxiliary heater

• Recognition light(s)

• Auxiliary hydraulic pump

• Fuel flow indicating system

• Auxiliary inverter (if installed)

• Utility light (if installed)

• Primary pitch trim motor (FC-530 AFCSonly)

• Left and right engine starters

Generator Buses

The left and right generator buses distribute

mary and secondary inverters, and the left and

right power buses. On SNs 35-509 and sub-sequent, 36-054 and subsequent, and prior aircraft incorporating AMK 85-1, the gener-ator buses also power the respective essentialA and B buses. On all aircraf t, the landing/taxilights are connected to the respective gener-ator bus (Figure 2-16). The generator buses can be powered by the batter ies, a GPU, or ei ther generator.

Power Buses

The left and right power buses are powered from the respective generator bus through a 10A current limiter. Each power bus provides power to three CBs that control the respect iveengine starting and generator functions, main bus power, and inver ter power, as fo llows:

• The L or R MAIN BUS circuit breaker controls the respective main bus power re l ay tha t connec t s t he r e spec t ive generator bus to the main bus when DC power is available (Figure 2-17).


2

E L E


S Y S T E M S




DC VOLTS

010

20

30

LMAIN

RMAIN

50 A

BUS

TIEPOWER

RELAY

70 A

OVERLOAD

SENSOR

L MAIN BUS

CB

POWER

RELAY

70 A

OVERLOAD

SENSOR

R MAIN BUS

CB

L

PWR BUS

R

PWR BUS

L

GEN BUS

L

GEN

R

GEN

10 A

R

GEN BUS

10 A

BAT CHG BUS

275 A

CL

275 A

CL

RBAT

LBAT

GPU

BATTERY POWER

GENERATOR POWER

GROUND POWER

LEGEND

C



overload and cannot; this results in automatic

isolation of the faulty bus.

Essential DC Buses

One of three different bus configurations willapply to a g iven a i rcraf t , depending on p roduc t ion s e r i a l number a nd AMK applicability (see Figures 2-14, 2-15, and 2-18).

The left and right essential buses are powered from the battery charging bus, or from the respective left or right generator buses (as applicable) through a 50-amp current limiter and a 40-amp ESS BUS circuit breaker, and areconnected to each other by a 20-amp ESS BUSTIE circuit breaker which is normally closed for load equalization.

In the event of an overload on one of the

essential buses, the respective ESS BUS cir-cuit breaker opens, followed by the ESS BUSTIE circuit breaker which is forced to acceptthe overload and cannot, resulting in auto-matic isolation of the faulty bus. The currentlimiters provide backup for their respectiveESS BUS circuit breakers.

AC POWERINVERTERS

Alternat ing current to the AC electrical instruments and electronic equipment is provided by two or three 1,000 VAC, solid-state static inverters in the tail cone (Figure 2-19). The third (auxiliary) inverter is optional.During normal operation both, or all three,


2

E L E


S Y S T E M S

BATTERY POWER

LEGEND

L ESS A R ESS A

20 A



inverters are on and operate in parallel. It is

recommended that the auxiliary inverter, if installed, be operated in conjunction with the primary and secondary inverters to extend in-verter life.

The primary and secondary inverters are powered from the respect ive le ft and rightgenerator buses through a 60 A overload sen-sor and a power relay. The power relay is en-ergized closed whenever there is power on therespective power bus, the associated PRI or S C i i b k i l d d h i

C B s , a n d i n v e r t e r a n n u n c i a t o r s o n t h e

glareshield for pr imary, secondary, and aux-iliary inverters.

The paralleling box is the central control unitfor the AC electrical system. It incorporatesload equalizer and frequency synchronizer/ phas er ci rcu it s th rough whic h it ma in ta in s inverter load balance and frequency/phasesynchronization. It also causes illumination of

the associated annunciators for certain mal-functions.

CONTROLS

A switch for each inverter (PRI, SEC, and op-tional AUX) is on the pilot lower instrument pane l (Figure 2-20). The pr imary an d sec-ondary inverter switches have two positions:

PRI-OFF and SEC-OFF, respectively. The aux-iliary inverter switch, if installed, is labeled ON-OFF.

If the auxiliary inverter is installed, an addi-tional switch labeled L BUS-R BUS is also in-


Figure 2-19. Inverter



INDICATORS

Two red inverter warning annunciators labeled PRI INV and SEC INV are on the glareshie ld.If the optional auxiliary inverter is installed,there is also an amber AUX INV annunciator on the glareshield.

The corresponding inverter annunciator illu-minates when inverter output is below 90 VAC

or if bus load is less than 10 volt-amps.

The primary and secondary inverter annunci-ators also illuminate when the respective in-verter switch is off; the AUX INV light,however, illuminates only when the auxiliaryinverter fails with the switch on.

A single AC voltmeter (Figure 2-21) indicates

the voltage on the left or right AC bus, depending on the position of the AC BUS switch.The two-position switch— PRI and SEC—se-lects the bus from which AC voltage is mea-sured. To check individual inverter voltage,only the inverter to be checked should be on.

DISTRIBUTION

115 VAC Buses (L and R)

Alternating current from the inverters is distributed through the paralleling box tothe respective left and right AC buses (Figure2-22). Primary inverter output goes to the left bus ; seconda ry to the right bus . Auxil iar y inverter output (if installed) may be selected to either the left or the right bus.

All CBs on the lef t 115 VAC bus are on the toprow of the left CB panel. The right 115 VAC bus CBs are on the top row of the right CB panel. The f irst CB on the top row of the right panel is the 7.5 amp AC bus-t ie CB. The sec-ond CB on the top row of the right panel and the f irst CB on the top row of the left panel arethe L and R AC BUS 10 A bus feeder CBs .

26 VAC Buses (L and R)

Two step-down transformers draw 115 VAC power from the left and right 115 VAC buses ,


2

E L E


S Y S T E M S




26 V

AC

BUS

26 V

AC

BUSTRANS

TRANSL AC

BUS

R AC

BUS 2 A2 A SEC VM7.5 APRI VM

P S

010

50

AC VOLTS

30

L AUX

AC BUS CB10 A 10 A10

A

10

A

L AC BUS

CB

R AUX

AC BUS CB

L BUS

AUX INVL BUS/R BUS SW

R AC BUS

CB

PARALLELING BOX

SECONDARY

INVERTER

POWER

RELAY

SEC INV

SW

AUX INV

CB

AUX INV

ON-OFF SW

60 A

POWER

RELAYPOWER

RELAY

PRI INV

SW

PRIMARY

INVERTERAUX

INVERTER*

R BUS



35-462 and 36-052. The battery packs containa built-in inverter; these aircraft also have AC powered s tandby att itude indicators. On later aircraft, lead-acid batteries and a DC pow-ered standby attitude indicator are standard.Lead-acid batteries may be retrofitted to ear-lier aircraft.

The nicad battery provides 25 VDC at 3.8ampere-hours and contains an inverter and

transformer that provide 115 VAC and 4.6VAC. The lead-acid battery provides 24 VDCat 5.0 amp-hours.

Both emergency batteries receive a trickle-charge from the normal aircraft electrical sys-tem through the EMER BAT 1 and EMER BAT 2 circuit breakers on the r ight main bus,respectively, when power is on the bus. The

trickle-charge is provided even when theswitches are off, but at a reduced rate. Con-trols and indicator location are illustrated inFigure 2-23.

SINGLE EMERGENCY

POWER SYSTEMIf an aircraft is equipped with a sing le emer-gency battery, the cockpit switch is labeled EMER PWR. An amber EMR PWR annunci-ator on the pilot instrument panel illuminateswhen power from the emergency battery is being used but the trickle-charge from the air-craft electrical system is lost.

The EMER PWR switch has three positions:ON, STBY, and OFF. The emergency battery powers the following equipment with theswitch in ON or STBY:

• ON

° Standby attitude indicator, indicator lighting, and annunciator light

° Landing gear control circuits and gear posi tion lights

° Flap control circuits


2

E L E


S Y S T E M S



dicator. On RVSM equipped aircraft the air data computer and the pilot altimeter are also powered. If power is available from the aircraftelectrical system, the emergency battery isreplenished as it provides power for the standbyattitude indicator. Other equipment tied to theemergency battery and normally powered bythe aircraft electrical system is powered by theemergency battery only when normal electri-cal power is off or failed.

Normally, the EMER PWR switch is in ON. If the electrical system fails, the EMR PWR an-nunciator illuminates when power from the as-sociated emergency battery is in use and the battery is not receiving a t rickle-charge.

In the event of a total aircraft electrical sys-tem failure, the approved AFM recommends

that the EMER PWR switch be placed in STBYuntil gear or flap operation is required to con-serve battery life. Since only the standby at-titude indicator is powered in STBY, batterylife is approximately 3 hours and 45 minutesversus 30 minutes in the ON position.

power supply is on the pilot instrument panel.The applicable annunciator illuminates when power from the associated emergency batteryis being used and is not receiving a trickle-charge.

The BAT 1 switch operates the same systems asdescribed under Single Emergency Power System. The BAT 2 switch has two positions:OFF and BAT 2. When turned on, power fromthe No. 2 emergency power supply is availableto illuminate the EMR PWR 2 annunciator and operate predetermined electrical equipmentshould the normal electrical system fail. Theauxiliary communication radio is the most com-mon equipment powered by BAT 2; however, itsinstallation and use is optional.

The pilot must turn off the emergency battery

switch(es) before leaving the aircraft. If air-craft power is turned off with the emergency bat-tery switch(es) in ON or STBY, the emergency bat teries continue to power the emergency bat-tery equipment and lose their charge.





2


S Y S T E M S

26 V

AC

BUS

26 V

AC

BUS

TRANS TRANSL AC

BUS

R AC

BUS 2 A7.5 A

AC VOLTS

010

3050

P S

2 A

10 A 10 A 10 A 10 A

PARALLELING BOX

PRIMARY

INVERTERAUX

INVERTERSECONDARY

INVERTER

L

MAIN

R

MAIN50 A

70 A 60 A

R I G

N / S T A R T / G E N

R ESS BL ESS B20 A

60 A

60 A 70 A

L I G N / S T A R T / G E N

*




AC VOLTS

010

3050

26 V

AC

BUS

26 V

AC

BUS

TRANS TRANSL AC

BUS

R AC

BUS2 A 7.5 A 2 A

50 A

P S

10 A10 A 10 A10 A

PARALLELING BOX

PRIMARY

INVERTER

AUX

INVERTER

SECONDARY

INVERTER

R I

G N I S T A R T / G E N

R PWR

L

MAINR

MAIN

60 A

70 A60 A


L PWR

R ESS A

20 A

40 A 40 A

70AR ESS B

L ESS A

L ESS B

20 A

60 A

*




2


S Y S T E M S

AC VOLTS

010

3050

26 V

ACBUS

26 V

ACBUS

TRANS TRANS

L AC

BUS

R AC

BUS2 A 7.5 A 2 A

50 A

P S

10A

10A

10 A 10 A

PARALLELING BOX

PRIMARY

INVERTER

AUX

INVERTERSECONDARY

INVERTER

R I

G N I S T A R T / G E N

60 A

R PWR

L

MAIN

R

MAIN

60 A

60 A


L PWR

L ESS A

L ESS B R ESS B

R ESS A20 A

20 A

40 A 40 A



1. The DC voltmeter indicates:

A. Battery voltage only

B. Generator voltage only

C. Voltage on the battery buses

D. Voltage on the battery charging bus

2. When a GPU is used for engine start, the

output value should be:A. Regulated to 24 V

B. Regulated to 28 V and limited to 1,100A

C. Regulated to 33 ±2 V

D. Regulated to 28 V and l imited to500 A

3. The buses that the aircraft batteries power are:

A. Battery buses only

B. Battery and battery charging busesonly

5. If aircraft electrical power fails and theEMER PWR BAT 1 switch is ON, the sys-tems powered by the emergency batteryare:

A. Standby attitude gyro only

B. Flaps and gear only

C. Flaps, gear, and spoiler

D. Standby attitude indicator, gear, and flaps

6. If both 275 A current limiters fail in flight:

A. The essential buses will remain pow-ered by the aircraft batteries

B. The essential buses will remain pow-ered by the generators

C. The battery charging bus will fail im-mediately

D. Both inverters will fail

7. Illumination of a PRI or SEC inverter light indicates:


QUESTIONS



9. If an overload sensor shuts off power toa main bus, power may be restored by:

A. Resetting the control CB after theoverload sensor resets

B. Changing the overload sensor

C. Automatic action after the currentlimiter cools

D. Automatic action after the overload sensor cools

10. To unlock the entrance door when the batter ies are dead:

A. Plug in a GPU and use a key

B. Plug in a GPU with 33 ±2 VDC or less on the small pin and use a key

C. Remove both batteries for chargingand reinstall

D. Enter aircraft through the emergency

hatch, place the emergency batteryswitch to ON, and activate the interior door switch

11. With a dual-generator failure in flight,h i f b i ill h i

12. Inverter output is:

A. 115 VAC, 400 HzB. 115 VAC and 26 VAC, 400 Hz

C. 26 VAC, 400 Hz

D. 115 VAC and 26 VAC, 1,000 Hz

13 . The approved AFM recommends that aGPU be used for engine start when the ambient temperature is:

A. 10°C or belowB. 0°F or below

C. 15°F or below

D. 32°F or below

14. W h e n e i t h e r p r i m a ry o r s e c o n d a ry inverter light illuminates, the first step of corrective action is:

A. Pull the AC bus-tie CBB. Turn the respective inverter switch

off

C. Check for open INV or AC BUS cir-cuit breaker(s)

D R d th l d th f il d AC b


2


S Y S T E M S



CHAPTER 3

LIGHTING

CONTENTS

Page

INTRODUCTION................................................................................................................... 3-1

GENERAL .............................................................................................................................. 3-1

INTERIOR LIGHTING .......................................................................................................... 3-2

Cockpit Lighting.............................................................................................................. 3-2

Cabin Lighting ................................................................................................................. 3-4

Emergency Lighting ........................................................................................................ 3-6

EXTERIOR LIGHTING ......................................................................................................... 3-8

General............................................................................................................................. 3-8




ILLUSTRATIONS

Figure Title Page

3-1 Interior Lighting Controls ........................................................................................ 3-3

3-2 Cockpit Map Lights ................................................................................................. 3-4

3-3 Reading Lights (Typical).......................................................................................... 3-4

3-4 Overhead Lights Control (Typical) .......................................................................... 3-53-5 Advisory Lights and Controls.................................................................................. 3-6

3-6 Emergency Cabin Door Light, Emergency Exit Light,and Wing Inspection/Egress Light ........................................................................... 3-7

3-7 Emergency Lights Control ....................................................................................... 3-7

3-8 Exterior Lighting Locations ..................................................................................... 3-83-9 Exterior Lighting Controls ....................................................................................... 3-9

3-10 Landing/Taxi Lights ................................................................................................. 3-9




CHAPTER 3LIGHTING




Optional map lights may be ins talled, whichconsist of a flexible-neck light located on each

pi lot’s sidewall panel or one of two overhead light installations, depending on aircraft serialization.

Cabin lighting consists of eight fluorescentupper center-panel lights (four on 36 mod-els), two door entry lights, baggage compart-ment lights, individual reading lights, and theno smoking/fasten seat belts sign. The op-tional emergency lighting systems illuminatethe fluorescent upper center-panel lights, and other lights at the exits.

Exterior lights include landing/taxi lights,wing and tail navigation lights, anticollision beacons, one or two (optional ) recognit ionlights, and high-intensity strobe lights. A winginspection and egress light, which may be partof the emergency lighting option, illuminatesthe right wing area to check for ice accumu-lation and for emergency egress. An optionalwing ice inspection light available on latemodels is not par t of the emergency lighting

A i l li h i id h il

Instrument Lights

Incandescent lighting is installed for pilot,copilot, and center instrument panels, pedestalindicators, and the magnetic compass. Thelights are controlled with the INSTR rheostaton the pilot side panel and both INSTR and PEDESTAL rheostats on the copilot side panel .DC power for the lights is supplied through therespective INSTR LTS circuit breakers on therespective left and right essential buses.

Pilot INSTR Lights

The pilot INSTR rheostat provides lightingcontrol for the pilot flight instruments, en-gine instruments, clock, electrical indicators,oil temperature indicators, altitude indicator,and the radar edge lighting.

Copilot INSTR Lights

The copilot INSTR rheostat provides lightingcontrol for the copilot flight instruments, themagnetic compass, cabin temperature indica-


3

L I G H T I



EL PANEL Rheostat

(Pilot Sidewall)The EL rheostat controls all edge lighting onthe switch panels to the left of a line runningvertically between the radar and radio panels.This control includes dimming for the audiocontrol panel, the left CB panel and the pilotmicrophone jack panel.

EL PANEL Rheostat

(Copilot Sidewall)The EL PANEL rheostat controls all edgelighting on switch panels to the right of the ver-tical line established in the preceding para-graph. It also controls lighting for the copilotmicrophone jack panel, audio panel, and theright CB panel.




Map Lights

When installed, the aircraft may have one or more of three different map light options:

• Flexible neck light on each pilot sidewall panel, with an ON-OFF rheostat for intensity control (see Figure 3-1)

• A reading light and gasper assembly, in-stalled in the cockpit headliner for each pi lot, incorporating a rheosta t for light

intensity adjustment and a light patternadjustment lever (Figure 3-2)

• A dome light assembly, mounted on eachside of the headliner just forward of theupper air outlets incorporating a rocker-operated switch (labeled ON-REMOTE)with an unlabeled center off position(Figure 3-2) and a swivel-mounted light.

All installations are powered through theINSTR LTS circuit breakers on the left and right essential buses. In the REMOTE position,the dome lights are powered from the ENTRYLT circuit breaker on the lef t battery bus.

CABIN LIGHTING

GeneralPassenger compartment lighting consists of reading lights, overhead lights, entry lights,no smoking / fas t en sea t be l t s igns , and refreshment cabinet lights.

Reading Lights

The reading lights are mounted in the upper center panel above the seats on each side of the cabin. There are individual switches for each light. The lights are adjustable for posi-tion and use DC power supplied through theRDNG LTS circuit breaker on the left main bus(Figure 3-3).

Overhead Lights

The cabin overhead light system consists of four (three on 36 models) fluorescent lights recessed in each side of the upper panel, a


3

L I G H T I



cabin lights power supply, a three positionswitch, a cabin lights relay assembly, and a

CAB LTS circuit breaker on the left main bus.

Normally, the lights are controlled with thethree position switch located on the left servicecabinet forward of the entry door (Figure 3-4).

In the event of cabin depressurization, thelights automatically illuminate full brightwhen the cabin altitude reaches 14,000 ft. On

aircraft with the optional emergency lightingsystem, three overhead lights illuminate au-tomatically in the event of aircraft electrical power fa ilure.

Entry Lights

The entry light system consists of a STEPLIGHT switch and light on the left service cab-inet forward of the entry door (Figure 3-4), and another direct ly over the door opening. Power from the left battery bus is supplied through theENTRY LT circuit breaker on the left battery bus; therefore, the lights are operational whenthe aircraft BAT switch is in OFF.

Baggage Compartment Lights

Two lights are installed in the aft baggage compartment; on 36 model aircraft , one light isinstalled in the forward baggage compartment.Aft baggage compartment lights are controlled by a switch on the left service cabinet forward of the entry door (Figure 3-4) and are powered through the ENTRY LT circuit breaker on the

left battery bus. The forward baggage compartment light is control led by a switch on theforward end of the upper center panel.




Passenger Advisory Lights

The no smoking/fasten seat belts advisorylight system consists of two fixtures (one on36 models) (Figure 3-5), a switch on the center switch panel, and the RDNG LT circuit breaker on the left main bus. The switch has three positions: NO SMOKING/FASTEN SEATBELT–OFF –FASTEN SEAT BELT. When theswitch is moved from OFF, an audible tonesounds and the appropriate symbols illumi-

nate. A RETURN TO SEAT light (if installed)in the lavatory is a part of the advisory lightsystem. Location of the fixtures varies withcabin configuration.

Cabinet Lights

The cabinet light system varies with cabinconf iguration and consists of various lightswithin the refreshment cabinet, microswitchesactuated by doors or drawers, power sup - pl ie s, an d a CB on th e right es sen ti al bu s.

EMERGENCY LIGHTING

Cabin Interior and WingInspection and Egress Lights

If these lights are installed, the aircraft isequipped with two nickel-cadmium (nicad) batter y power supplies and a control module.The lights illuminate selected areas auto-matically in the event of aircraft DC power failure.

An emergency light in the upper cabin door (Figure 3-6) illuminates the lower cabin door and the immediate door area. A second lightilluminates the emergency exit window area.An exterior wing inspection/egress light op-tionally installed on the right side of the air-craft is adjacent to the emergency exit windowand illuminates the exterior egress area.

The fluorescent cabin upper-center panel lightsilluminate the cab in interior. When activated,one of the power supplies turns on the cabinupper-center panel lights, while the other


3

L I G H T I




The nicad battery packs charge through theEMER LTS circuit breaker on the right es-

sential bus.

Control Module

The EMER LIGHT TEST switch on the pilot(or center) switch panel (Figure 3-7) providesthe test function for the system and automati-cally illuminates the emergency lights in theevent of an interruption of normal DC electrical power. The switch has three posi tions: TEST,ARM, and DISARM. Setting the switch toTEST simulates a failure of normal DC elec-trical power and illuminates the upper cabinentry door light, the emergency exit light, and the cabin overhead fluorescent lights. Settingthe switch to ARM arms the system to illumi-nate the emergency lights in the event of a fail-

ure of normal DC electrical power. Setting theswitch to DISARM isolates the emergency

lights from the emergency batteries.

The switch should be set to ARM prior to take-off. If the switch is in the DISARM positionand at least one BAT switch is on, the amber light adjacent to the switch illuminates to re-mind the pilot that the switch should be set toARM. The switch should be set to DISARM prior to set ting the BAT switches to OFF.

The WING INSPECTION light switch (in-cluded as part of the emergency lighting system), located adjacent to the EMERGLIGHT TEST–ARM–DISARM switch, may be used independently of the rest of the emer-gency lighting system to visually check for ice accumulation on the wing leading edge.





Turning the switch on illuminates the exterior wing inspection/egress light.

The EMERGENCY LTS. switch on the leftservice cabinet near the entry door (see Figure3-7) provides a means for manual illuminationof the interior emergency lights. When theswitch is set to EMERGENCY LTS., the upper cabin entry door light, the emergency exitlight, the cabin overhead fluorescent lights, and the wing inspection/egress light (if installed)

illluminate. For normal operation, the switchshould be set to OFF, allowing automatic illumination of the emergency lights in theevent of a failure of the normal electricalsystem.

EXTERIOR LIGHTING

GENERAL

The exterior lighting systems consist of thelanding/taxi lights, navigation lights, anti-collision lights, recognition light(s), strobelights, and an optional wing ice inspectionlight (Figure 3-8). The exterior lighting con-trols are shown in Figure 3-9.

LANDING/TAXI LIGHTS

The landing light system consists of one 450W lamp mounted on each main landing gear strut (Figure 3-10), one 20 A current limiter for each side in the current-limiter panel, relays, dimming resistors, and the L and R LDG LT switches on the center switch panel.


3

L I G H T I





Figure 3-9. Exterior Lighting Controls




Regardless of switch position, the lights willnot illuminate unless the respective landing

gear down-and-locked switches are closed and provide a ground. It is recommended that thelights be operated in the L and R LDG LTmodes as sparingly as possible. Lamp servicelife is shortened in the LDG LT mode becauseof the higher current flow.

RECOGNITION LIGHT

A 250 W recognition light is installed in thenose of the right tip tank (Figure 3-11). Thelight is controlled with the RECOG LT switchon the copilot lighting control panel. Whenturned on, DC power from the RECOG LT cir-cuit breaker on the right essential bus closesa control relay and connects power througha 30 A current limiter to the light.

NAVIGATION LIGHTS

The navigation light system consists of onelamp in the outboard side of each tip tank,two lamps in the upper a ft tail fair ing, a NAVLT switch on the copilot lighting control panel,and a NAV LTS circuit breaker on the leftmain bus.

All three navigation lights are controlled bythe NAV LT switch. Addi tional ly, setting the

NAV LT switc h to ON auto ma tically di msmost instrument panel and pedestal “peanut”lights and activates the landing gear positionlight dimmer rheostat.


3

L I G H T I




ANTICOLLISION LIGHTS

Anticollision lights are installed on top of thevertical stabilizer and on the bottom of thefuselage (Figure 3-13). The lights are con-trolled by a BCN LT switch on the copilotlighting control panel. Each light is a dual- bulb light; each bulb oscillates 180° at 45 cy-cles per minute. The beam is concentrated byan integral lens; an illusion of 90 flashes per minute occurs due to the oscillation.

The lights operate on DC power through theBCN LT circuit breaker on the right main bus.





WING INSPECTION LIGHTS

Two separate installations are designed to illuminate the wing area for signs of ice (Figure3-14). Both are optional. One light is installed on the right side of the fuselage adjacent to thelower forward corner of the emergency exitwindow. This light is designed to illuminatethe leading edge of the right wing and addi-tionally serves as an illumination source for emergency egress over the wing. The light is

designated the wing inspection and egresslight, and may be installed as an integral par tof the earlier emergency lighting system or asan option not involving the emergency light-ing system. In either case, a second optionmay include a second light installed on the leftside of the fuselage directly opposite the right-hand light, which serves as an inspection lightfor the left wing.

The WING INSPECTION control switch islocated on the emergency lighting panel or on

the instrument panel (see Figure 3-7).

On SNs 35-416 and 36-048 and subsequent,another option provides a light installed inthe fuselage below the copilot window. It is de-signed to illuminate a black spot on the rightwing leading edge. A covering of ice obscuresthe spot,which enables ice detection at nightwhen the light is turned on. This light is des-

ignated the WING INSP light (Figure 3-14)and is operated by a push-button switch located forward of the rheostats on the copilot rightside panel (Figure 3-15).


3

L I G H T I




TAIL CONE AREA INSPECTION

LIGHTWhen installed, this light is located in the tailcone, direct ly above the entry door. AnON–OFF switch is positioned inside the door at the forward left side of the opening. A microswitch installed on the forward rightside of the opening breaks power to the lightwhen the door is closed (Figure 3-16). Power for operating the tail cone area inspection

light is provided by the left battery bus throughthe ENTRY LT circuit breaker (pilot CB panel) , which permi ts operat ion of the lightwithout turning aircra ft power on. However,on some aircraf t, the light is powered by the battery charging bus through a 5 A current lim-iter; in such a case, an aircraft battery must be turned on to operate the light.


Figure 3-15. Wing Ice Inspection Light

Control




3

L I G H T I

INTENTIONALLY LEFT BLANK




1. The instrument panel flood light control islocated:

A. On the light

B. Just forward of the warning panel

C. On the pilot side panel

D. On the copilot side panel

2. The cockpit map lights are controlled:

A. With an ON-OFF switch on the copilotside panel

B. With the overhead map light rheostaton the copilot side panel

C. With an integral rheostat and a patternlever

D. Automatically, relative to ambientlight

3. The cabin overhead light control switchesare located on the:

6. The emergency lighting switch positionused during normal operation is:

A. DISARM

B. ARM

C. TEST

D. EMER LT

7. The lights that come on when cabin alti-tude reaches 14,000 ft or higher are the:

A. Passenger advisory lights

B. Lavatory lights

C. Cabin overhead panel lights

D. Reading lights

8. The wing ice inspection light switch (if

installed) is located on the:A. Pilot switch panel

B. Light assembly

C. Overhead panel

D C il t i ht id ll

QUESTIONS




CHAPTER 4

MASTER WARNING SYSTEM

CONTENTS

Page

INTRODUCTION................................................................................................................... 4-1

GENERAL .............................................................................................................................. 4-1

GLARESHIELD ANNUNCIATOR LIGHTS ........................................................................ 4-2

MASTER WARNING LIGHTS.............................................................................................. 4-2

TEST ....................................................................................................................................... 4-2

INTENSITY CONTROL ........................................................................................................ 4-3

BULB CHANGE..................................................................................................................... 4-3

ILLUMINATION CAUSES .................................................................................................... 4-3




ILLUSTRATION

Figure Title Page

4-1 Test Switch ............................................................................................................... 4-2

TABLE

Table Title Page

4-1 Annunciator Illumination Causes............................................................................. 4-3




CHAPTER 4

MASTER WARNING SYSTEM




GLARESHIELD

ANNUNCIATOR LIGHTSThe red, amber, and green glareshield lightsreceive power from the left and/or right essential DC buses through the respectiveWRN LTS circuit breakers. The red lights areused to indicate the more critical malfunctions.Amber lights denote cautionary items, and green lights indicate conditions that may be

normal but need to be announced.

If a glareshield annunciator light illuminatesand the condit ion is corrected, the l ight extinguishes; should the condition recur, thelight again illuminates.

Five of the glareshield annunciator lightsgive a flashing indicat ion under the f ol-

lowing conditions:

1. SPOILER—If spoilers and flaps are both extended (f laps more than 13°)

2. STALL (L or R)—If the angle-of-at-t k i di t h h k li it

glareshield annunciator light remains illuminated as long as the causative condition exists.

TEST

Depressing either of the two test switchesunder the glareshield (Figure 4-1) causes thefollowing lights to illuminate:

• All glareshield annunciator lights and

both MSTR WARN lights• FIRE warning lights

• Marker beacon lights (if installed)

• Thrust reverser panel annunciator lights(if installed)

• AFCS/control panel annunciator lights(FC 530 AFCS)

• ANTISKID l ights

• AIR IGN l ights

• Fuel panel lights

C il t fli ht di t i t li ht




INTENSITY CONTROL

A photoelectric cell outboard of each FIREhandle (Figure 4-1) automatically adjusts theglareshield annunciator light intensity for existing cockpit light conditions. The other instrument panel and pedestal annunciator lights dim when the navigation light (NAV LT)switch is turned on.

BULB CHANGEGlareshield annunciator light lenses can be removed for bulb replacement.

ILLUMINATION CAUSES

Table 4-1 shows each annunciator light label,color, and cause for illumination.

NOTE

S o m e l i g h t s a r e o p t i o n a l , a n d arrangements may vary between aircraft.

ANNUNCIATOR CAUSE FOR ILLUMINATION ANNUNCIATOR CAUSE FOR ILLUMINATION

DH

LOW

FUEL

FILTER

At or below altitude set on radioaltimeter.

Fuel is below 400–500 lb in either

Differential pressure is 1.25 psiacross one or both airframe fuelfilters. Fuel is bypassing the filter.

1. Switch ON–Insufficient pressure to

Table 4-1. ANNUNCIATOR ILLUMINATION CAUSES




CAUSE FOR ILLUMINATIONANNUNCIATORCAUSE FOR ILLUMINATIONANNUNCIATOR

MACH

TRIM

PRI

INV

SEC

INV

AUX

INV

LO OIL

PRESS

STAB

OV HT

WING

OV HT

WSHLD

HT

ALC

AI

BAT

140

BAT

160

ENG

SYNC

TO

System is inoperative with speed above0.69 Mach and autopilot disengaged. Ifabove 0.74 Mach, the overspeedwarning horn sounds.

1. Inverter is off.

2. Inverter switch is on and output is

less than 90 V, or less than 10volt-amperes

Inverter has failed with the switch on.

Oil pressure on one or both enginesis below 23 ±1 psi.

Stabilizer structural temperature isabove 215°F.

Wing structural temperature is above215°F.

The windshield anti-ice valve is open.

1. Late ECS–The alcohol tank isempty.

2. Early ECS–Alcohol systempressure is low.

One or both batteries' temperature is140°F or more.

One or both batteries' temperature is160°F or more.

The engine sync switch is on with thenose gear down and locked.

Aircraft is on the ground and the

Table 4-1. ANNUNCIATOR ILLUMINATION CAUSES (Cont)




ANNUNCIATOR CAUSE FOR ILLUMINATION ANNUNCIATOR CAUSE FOR ILLUMINATION

LH ENG

CHIP

RH ENG

CHIP

EMERPWR 1

EMER

PWR 2

START L

START R

COMPTR

WARN

AIR IGN L AIR IGN R

ANTI-SKID

GEN

L CUR

LIMITER

R CUR

LIMITER

R PITOT

HEAT

Ferrous metal particles are detectedin indicated engine’s oil.

Indicated emergency battery ispowering the connected systems.

Indicated starter is engaged.

HSI headings are not within 7°.

Ignition system is activated.

Indicated antiskid generator isinoperative.

Indicated 275 A current limiter hasfailed.

1. Indicated pitot heat switch is off.

2. Switch is on and indicated pitotheat has failed.

(AMK 80-17)

L R

START L

L PITOT

HEAT

Table 4-1. ANNUNCIATOR ILLUMINATION CAUSES (Cont)








1. All glareshield annunciator lights and system advisory annunciator lights can betested by:

A. The rotary test switch

B. Depressing each individual light

C. Depressing either glareshield TESTswitch

D. Shutting the represented system off

2. When a red glareshield annunciator lightilluminates, another annunciation thatoccurs is:

A. Only the pilot MSTR WARN lightflashes

B. Both MSTR WARN lights illuminatesteady

C. Only the copilot MSTR WARN lightilluminates

D. Both MSTR WARN lights flash

4. The glareshield annunciator light inten-sity is adjusted:

A. Automatically by photoelectric cells

B. By depressing the TEST button

C. By depressing each individual capsule

D. By depressing the DIM button

5. The flashing MSTR WARN lights can bereset by depress ing either MSTR WARNlight:

A. Unless a red glareshield annunciator is flashing

B. Anytime

C. Unless a red glareshield annunciator is illuminated steady

D. Unless an engine FIRE PULL lightilluminated steady

QUESTIONS




CHAPTER 5

FUEL SYSTEM

CONTENTS

Page

INTRODUCTION................................................................................................................... 5-1

GENERAL .............................................................................................................................. 5-1

FUEL TANKS AND TANK VENTING SYSTEM ................................................................ 5-3

General............................................................................................................................. 5-3

Tip Tanks ......................................................................................................................... 5-3

Wing Tanks ...................................................................................................................... 5-3

Fuselage Tank .................................................................................................................. 5-3

Ram-Air Vent System...................................................................................................... 5-4




Crossflow System .......................................................................................................... 5-10

Normal Transfer System................................................................................................ 5-11

Gravity-Flow Transfer System ...................................................................................... 5-11

Float and Pressure Switches .......................................................................................... 5-12

Pressure-Relief Valves................................................................................................... 5-12

Fuselage Fuel Fill Transfer Operations ......................................................................... 5-12

TIP-TANK FUEL JETTISON SYSTEM.............................................................................. 5-13

FUEL SERVICING............................................................................................................... 5-13

General .......................................................................................................................... 5-13

Safety Precautions ......................................................................................................... 5-14

Aviation Gasoline .......................................................................................................... 5-14

Anti-icing Additive........................................................................................................ 5-14

Refueling ....................................................................................................................... 5-14




ILLUSTRATIONS

Figure Title Page

5-1 Fuel System.............................................................................................................. 5-2

5-2 Ram-Air Scoop and Overboard Drain...................................................................... 5-4

5-3 Fuel Vent System ..................................................................................................... 5-5

5-4 Fuel Control Panels.................................................................................................. 5-6

5-5 Fuel Flow Indicator .................................................................................................. 5-7

5-6 Jet Pump Schematic ................................................................................................. 5-8

5-7 Fuel Drain Locations ............................................................................................. 5-10

5-8 Aircraft Grounding Points...................................................................................... 5-14

5-9 Prist Blending Apparatus ....................................................................................... 5-15

5-10 Refueling Filler Cap............................................................................................... 5-15




CHAPTER 5

FUEL SYSTEM




TOSUMP

FUEL JETTISONSHUTOFF VALVE

LOW FUELPRESSURESWITCH

DIFFERENTIALPRESSURESWITCHFUEL

SHUTOFFVALVE

75-PSIRELIEFVALVE

CROSSFLOWVALVE

WINGPRESS SW

TRANSFER

VALVE

FUSELAGETANK

TRANSFERLINE

EMPTY LIGHTPRESSURESWITCH

MOTIVEFLOWVALVE

MOTIVEFLOW

FUEL

MODEL 35

WITHOUT

GRAVITY-FLOW LINE

F

P

PP

P

PP

TOSUMP

CROSSFLOWVALVE

MODEL 36 AND MODEL 35

WITH

MODEL 36 AND MODEL 35

WITH




FUEL TANKS AND TANK

VENTING SYSTEMGENERAL

The total usable fuel capacity is approximately6,238 lb for the 35 model and approximately7,440 lb for the 36 model. Unusable (i.e.,trapped) fuel is included in the aircraft basicweight and is not reflected in any fuel quan-

tity indications.

TIP TANKS

Each tip tank capacity is 1,215 lb of usablefuel; capacity is reduced to 1,175 lb with in-stallation of a recognition light. The tanks are permanently attached to the wings and are po-sitioned at 2° nosedown relative to the air-craft centerl ine. Baffles are instal led tominimize slosh and prevent adverse effectson the aircraft center of gravity during extreme pi tch a tt itudes .

WING TANKS

Each wing tank extends from the aircraft cen-terline to the tip tank and holds 1,254 lb of us-able fuel. Areas that are not part of the wingfuel cell are the main landing gear wheel well,the leading edge forward of spar 1 (i.e., wingheat area), and the trailing edge between spars7 and 8 (i.e., flap, spoiler, and aileron areas).

The 2.5° wing dihedral makes the inboard por-

tions of the wing tanks the lowest areas. In eachwing tank, a jet pump and an electric standby pump are located within these areas and wil lremain submerged in fuel until the tanks arenearly empty.

Wing tank ribs and spars act as baffles to mini-mize fuel shifting. Flapper valves located inthe wing ribs allow unrestricted inboard flow

of fuel and limit outboard flow. Two pressure-relief valves at the centerline rib equalize in-ternal pressures between the two wing tanks.The wing tanks begin to f ill through the twotip tank flapper valves as tip tank fuel in-

b d h lf f ll




RAM-AIR VENT SYSTEM

A ram-air scoop located on the underside of each wing (Figure 5-2) supplies positive air pressure in fl ight to a manifold that directlyvents the fuselage tank and both tip tanks.Each wing tank is indirectly vented to its owntip tank through a length of tubing, the ends of which extend to the uppermost area of eachtank (Figure 5-3). The ram-air scoops, by design,do not require heating to remain ice free.

Two vent float valves are located in each tiptank, and one in the fuselage tank on 35 mod-els. The float valves close when the fuel levelreaches the vent ports, preventing fuel from en-tering the vent lines. A vacuum relief valve ineach tip tank and the fuselage tank opens toallow air to enter the tanks should vacuumconditions occur.

Each tip tank has two pressure relief valves that protect the tanks from excessive pressure. The pressure re lief valves are set at 1.0 and 1.5 psi; the second valve provides a backup in case

h fi l f il

of the main landing gear, collects any fuelthat might enter the vent lines. A vent drain

valve permits draining of the sump to ensurethat the vent line to the fuselage tank isunobstructed.

FUEL INDICATINGSYSTEMS

FUEL QUANTITY INDICATINGSYSTEM/LOW FUEL WARNING

The fuel quantity indicating system includesan indicator and tank selector switch located on the fuel control panel (see Figure 5-4). Ared LOW FUEL warning light (Annunciator Section) illuminates when either wing tank

fuel level is low.

The fuel quantity indicating system uses DC power from the right essential bus through theFUEL QTY circuit breaker. The six-position




OPENVENTTUBE**

FLOATVALVE*

VACUUMRELIEF

FUELVENT

DRAIN

OVERBOARDDRAIN

FLAME

TO AMBIENT

VACUUMRELIEFVALVE

1.5-PSIRELIEFVALVE

1.0-PSIRELIEFVALVE




*OPTIONAL ON SNs 35-299 THROUGH 35-596.

STANDARD ON SNs 35-597 AND SUBSEQUENT.

L ON R

L ON R

FUEL JTSN

OPEN

CLOSECROSS FLOW

EMPTY

XFER

OFF

FILL

FULL

}OPEN

CLOSE

FUS VALVE

FUS

TANK

JET PUMP

L WING

1254L TIP

1215

TOTAL

6238

LBS

R WING

1254R TIP

1175

FUS

1340

STANDBY PUMPS

L ON R

MODEL 35

0 0 0 0

OPEN

0

1

2

LBS x 1000

3 FUELQUANTITY

4

5

6

7




There are nine capacitance fuel probes. Onefuel probe is located in each tip tank and in the

fuselage tank. Each wing tank has three probeswired in parallel. The inboard probe in the leftwing contains a temperature compensator thatadjusts quantity readings for all switch selec-tions for fuel density change due to temperature.

If the compensator probe i s uncovered , erroneous fuel quantity indications could beencountered at all switch positions.

Each probe uses an electrical capacitance measuring system to sense the fuel level. It thentransmits an electrical signal to the cockpit in-dicator where it is read as pounds x 1,000 onthe dial.

Each wing tank has a fuel low-level float switch.When either wing tank fuel level reaches 400

to 500 lb remaining, the respective float switchactuates the red LOW FUEL light on the an-nunciator panel to indicate low wing fuel quan-tity (Annunciator Section). When flying withthe LOW FUEL light on, limit pitch attitude and h h i i i d




The standby pumps are used:

• For engine start (automatically ener-

gized with starter switch activation)

• As a backup for the wing jet pumps

• For wing-to-wing crossflow

• For filling the fuselage tank (automati-cally energized with the XFER–FILLswitch in the FILL position)

Both standby pumps are deactivated when theXFER–FILL switch is in XFER.

The transfer pump is used to transfer fuse-lage tank fuel to the wing tanks.

The s tandby pumps are powered by the respective L or R STBY PMP circuit breakerson left and right essential buses; the fuselage pump receives power from the FUSLG PMPcircuit breaker on the right main bus.

MOTIVE-FLOW FUEL

through the motive-flow valves to the jet pumps, where it passes through a small orif iceinto a venturi. The low pressure created in theventuri draws fuel from the respective tank, re-sulting in a low-pressure, high-volume outputfrom the jet pump (Figure 5-6).

Motive-flow pressure varies with engine rpmand is regulated to 300 psi maximum. Conse-quently, jet pump discharge pressure also varieswith engine rpm. At idle, discharge pressure is

approximately 10 psi, while at full-power set-tings, discharge pressure is approximately 12 psi.

There are four jet pumps: one in each wing tank adjacent to the standby pump, and one in eachtip tank. The wing tank jet pumps draw fuelfrom the wing tanks and supply low pressurefuel to the engine-driven, high-pressure fuel

p ump s . Wing j e t p ump ou tp u t c a n b e supplemented by the wing standby pump to ensure positive pressure to an engine. The tiptank jet pumps draw fuel from the tip tanks and deliver it directly to the cavities where the

db d j l d




electrically open and close the motive-flowvalves. Power is provided by the respect ive Lor R JET PUMP VAL circuit breaker on the leftand right essential buses. The amber indicator lights next to the switches illuminate when themotive-flow valves are in transit or are not inthe position selected on the switch. Each jet pump switch (and motive-flow valve) controls both jet pumps (wing and tip) on that side.

FILTERS

A fuel filter is installed in each engine feed lineto filter the fuel before it enters the engine-driven fuel pump. Should the filters becomeclogged, the fuel is allowed to bypass them. Adifferential pressure switch installed in eachfilter assembly illuminates the one amber FUEL FILTER annunciator light if either or both f il ters are bypassing fuel (Annunciator

Panel section).

MAIN FUEL SHUTOFF VALVES(FIREWALL)

The engine-driven pump is capable of suctionfeeding enough fuel to sustain engine opera-tion without either the wing standby pump or jet pump. However, 25,000 ft is the highest al-titude at which continuous operation should be at tempted in this event .

PRESSURE-RELIEF VALVES

A 75-psi relief valve is installed in each mainfuel line on the engine side of the main shut-

off valve. The valves vent fuel overboard to re-lieve pressure buildup caused by thermalexpansion of trapped fuel when the engines areshut down.

FUEL DRAIN VALVES

Drain valves (Figure 5-7) are located at low points throughout the fuel system for drain-

ing condensation or sediment. A small amountof fuel should be drained from each valve dur-ing the exterior preflight inspection. Thevalves, spring-loaded to the closed position,are located as follows:




RH ENGINE FUEL

LINE DRAIN

FLUSH SUMP

WING SUMP

FUEL VENT DRAIN

FUEL FILTER DRAIN

FUEL COMPUTER DRAINS

CROSSFLOW DRAIN

LH ENGINE FUEL

LINE DRAIN

FLUSH SUMP DRAIN

WING SUMP DRAIN

FUSELAGE LINE DRAIN*

FUSELAGE TANK SUMP DRAIN*

FUEL FILTER DRAIN

* THE 35 MODELS WITH OPTIONAL

FUEL LINE AND THE 36 MODELS

HAVE TWO FUSELAGE LINE DRAINS

AND TWO FUSELAGE TANK SUMP DRAINS.

RIGHT WING LEFT WING




settings may be used to balance fuel, if nec -essary. The above considerations do not applyto single-engine operations, and normal cross-flow operations may be performed as usual.

NORMAL TRANSFER SYSTEM

The Learjet models 35/36 each have a fueltransfer line connecting the fuselage tank transfer pump with the crossflow manifold (see Figure 5-1). A DC motor-driven transfer

valve installed in the line controls fuel move-ment between the fuselage and wing tanks.The valve is controlled by the XFER-FILLswitch located on the fuel control panel.

When the switch is positioned from OFF toXFER, the transfer and crossflow valves aresequenced open and the transfer pump is en-ergized automatically while both standby

pumps are deactivated. When the switch is positioned from OFF to FILL, the transfer and crossflow valves are sequenced open, and bothstandby pumps are energized automatically.When the switch is positioned from either

subsequent and on all 36 models, a DC motor-driven fuselage valve is installed in a second fuel line; it connects the fuselage tank with thecrossflow manifold on the right side of thecrossflow valve (see Figure 5-1). The valve iscontrolled by the FUS VALVE switch on thefuel control panel.

When the FUS VALVE switch is positioned toOPEN, both the fuselage valve and the cross-flow valve simultaneously open, a llowing fuelto gravity-flow from the fuselage tank to bothwings. When fuselage fuel is transferred in thismanner, 162 lb of fuel remain in the fuselagetank. The fuselage valve is also controlled bythe XFER–FILL switch.

When placed to FILL, the transfer valve, fuse-lage valve, and crossflow valve are sequenced open, and the standby pumps are energized to

pump wing tank fuel through both fuel linesinto the fuselage tank.

The fuselage valve remains closed when theXFER–FILL switch is positioned to XFER.




FLOAT AND PRESSURESWITCHES

Fuselage Fuel TankFloat Switch

When filling the fuselage tank, a float switchmounted inside the tank actuates when thetank is full. The switch:

• Illuminates the green FULL light on thefuel control panel

• Deenergizes the standby pumps

• Closes the transfer and crossflow valves

• Closes the fuselage valve (all aircraftequipped with the gravity-flow transfer line)

The green FULL light on the fuel control panel

(see Figure 5-4) remains illuminated until theXFER–FILL switch is turned off.

Fuselage Tank Low-Pressure

resets and energizes the pump again when the pressure drops below 2 .5 psi.

PRESSURE-RELIEF VALVES

Two one-way pressure-relief valves are located at wing rib 0.0, which separates the left and right wing fuel tanks. Each valve, relieving inthe opposite direction, opens at 1 PSID toequalize fuel pressure between the wing tankswhen crossflowing or transfer ring fuel.

FUSELAGE FUELFILL-TRANSFER OPERATIONS

Fill Operation

Fuel may be pumped from the wings to thefuselage tank using the FILL position on theXFER–FILL switch. The FILL position may

be used for CG considerations in flight; however,it is normally used only during fuel servicing.

When the XFER–FILL switch is placed to theFILL position:




1. The transfer valve opens

2. The crossflow valve opens

3. The fuselage transfer pump is energized

4. The standby pumps are disabled

5. The white EMPTY light (pressureswitch) is enabled

Gravity-flow is also possible on all aircraftthrough the normal transfer line should thetransfer pump fail. The amount of fuel trapped

(unusable) is approximately 162 lb. The rateof gravity transfer will, however, be slowerthan when using the fuselage valve, if installed.

When the XFER–FILL switch is placed in theOFF position:

• The transfer pump is deenergized, and

• Operation of the standby pumps is once

again possible

• The transfer valve closes, then

• The crossflow valve closes

TIP-TANK FUEL

JETTISON SYSTEMA DC motor-driven valve in the tail cone of each tip tank provides the capability of jetti-soning tip-tank fuel. One FUEL JTSN switchon the fuel control panel (see Figure 5-4) controls both tip-tank jettison valves. Whenthe FUEL JTSN switch is placed to ON, the jet-t i son valves open and two amber l ights illuminate continuously on the fuel control panel to indicate that the valves are fully open.The jettison tubes are scarfed, which createsa low-pressure area that helps pull the fuelout of the tank(s). This, in combination withthe force of gravity, enables the entire contentsof both tanks to be jettisoned.

Fuel jettison is faster while the aircraft is in

a noseup attitude.

It takes approximately five minutes to jettisonfuel from the tip tanks.




of solution and subsequently freeze. The anti-icing additives specif ied for use in the Learjet35/36 are Hi-Flo Prist and QUELL. Either additive prevents the growth of microbiologicalorganisms in the fuel. Fuel containing anti-icing additive conforming to MIL-I-27686 requires no further treatment.

SAFETY PRECAUTIONS

Refueling and defueling should be accomplished

only in areas that permit free movement of fire equipment.

Figure 5-8 shows the aircraft grounding points.

When adding anti-icing additives (Figure 5-9),follow the manufacturer’s instr uctions for blending.

AVIATION GASOLINEAviation gasoline (MIL-D-5572D, Grades80/87, 100/130, and 115/145) may be used asan emergency fuel and mixed in any propor-

ANTI-ICING ADDITIVE

All fuels used must have an approved blended anti-icing additive. Depending upon fueling lo-cation and type of fuel, the additive may or maynot be blended at the ref inery. If not blended at the ref inery, the additive must be blended at the time of fueling. Refer to the AFM for theapproved MIL Specs. Compare the MIL Specof the anti-icing additive to be blended withthe referenced MIL Specs in the AFM to de-termine the correct blending amounts.

REFUELING

Refueling is accomplished through the tiptank f iller caps (Figure 5-10). The fuel beginsto flow by gravity into the wing tanks as thetip tanks reach one-half full. The standby pumps are used to f il l the fuselage tank. (See

Fuel Transfer Systems, this chapter.) A ground power unit should be used, if possible, becauseof the requirement to operate the standby pumps. Refer to the approved AFM for detailed refueling procedures




FUEL ADDITIVE BLENDER HOSE

FUEL NOZZLE

TRIGGER

RING

HANDLE

H I - F L

O

P R I S

T

( O R )

M I L - I -

2 7 6 8 6








1. Trapped fuel weight:

A. Must be added to the weight of fueltaken on board when servicing theaircraft

B. Is included in the aircraft basic weightfor airplanes cer tified in the U.S.

C. Must be accounted for in the fuselagetank for CG purposes

D. May be disregarded since it is lessthan 200 lb

2. With the exception of the FUEL JTSNlights, all other amber lights on the fuelcontrol panel, when illuminated steady,indicate that the respective:

A. Valves are cycling or the pumps are

properly operat ingB. Valves are in the correct position; the

pumps are inoperat ive

C. Switch position agrees with the valveposi tion or pump operat ion

5. The crossflow valve opens:

A. Only when the CROSS FLOW switchis set to OPEN

B. Only when the CROSS FLOW switchis set to OPEN or the XFER–FILLswitch is set to XFER

C. Anytime electrical power is lost

D. Whenever the CROSS FLOW, XFER–

FILL, or FUS VALVE switches aremoved f rom the OFF o r CLOSE posi tion

6. Steady illumination of an amber transfer valve light indicates:

A. The valve failed to close

B. The valve failed open

C. The valve operated correctlyD. Th e v a lv e f a i l e d to m ov e t o t h e

posi tion commanded by the XFER– FILL switch

QUESTIONS




9. Motive-flow fuel for the jet pumps is supplied by the:

A. Engine-driven fuel pumpsB. Wing standby pumps

C. Fuselage transfer pump

D. Motive-flow control unit

10. The amber FUEL FILTER light indicates:

A. Low fuel pressure to the engine-driven pump; the standby pumps should be

turned onB. That both fuel filters are being by-

passed; the light does not il luminateif only one filter is bypassed

C. That one or both fuel filters are being bypassed

D. That only the secondary fuel filters are being bypassed

11. The amount of fuel trapped in the fuselagetank after completion of gravity transfer via the fuselage valve is approximately:

A. 62 lb

12. The wing fuel pressure switch:

A. Turns off the fuselage transfer pump

when wing fuel pressure reaches5 psi

B. Turns on the fuselage transfer pumpwhen wing fuel pressure is below 5 psi

C. Turns off the wing standby pumpswhen wing fuel pressure reaches 5 psi

D. Turns on the wing standby pumpswhen wing fuel pressure is below 5 psi

13. When using any mixture of aviationgasoline:

A. Do not take off with fuel temperaturelower than –54°C (–65°F)

B. Restrict flights to below 15,000 ft

C. Both je t pumps and both s tandby pumps must be on and the pumps must be operat ing

D. All of the above answers are correct

14. The Learjet 35/36 requires anti-icing additive:



The information normally contained in this chapter is

CHAPTER 7




CHAPTER 7

POWERPLANTCONTENTS

Page

INTRODUCTION................................................................................................................... 7-1

GENERAL .............................................................................................................................. 7-1

MAJOR SECTIONS ............................................................................................................... 7-2

Air Inlet Section............................................................................................................... 7-2

Fan Section....................................................................................................................... 7-2

Compressor Section ......................................................................................................... 7-3

Combustor Section........................................................................................................... 7-3

Turbine Section ................................................................................................................ 7-3

St t P R l t 7 12




Start Pressure Regulator ................................................................................................ 7-12

Surge Bleed Valve ......................................................................................................... 7-12

Fuel Flow....................................................................................................................... 7-13

Flow Divider.................................................................................................................. 7-13

Fuel Spray Nozzles........................................................................................................ 7-13

Operation ....................................................................................................................... 7-13

IGNITION SYSTEM............................................................................................................ 7-14

General .......................................................................................................................... 7-14

Automatic Mode............................................................................................................ 7-14

Selective Mode .............................................................................................................. 7-14

Indication....................................................................................................................... 7-14

ENGINE CONTROLS.......................................................................................................... 7-15

STARTERS ........................................................................................................................... 7-16

7

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THRUST REVERSERS (OPTIONAL EQUIPMENT) 7 24




THRUST REVERSERS (OPTIONAL EQUIPMENT)........................................................ 7-24

General .......................................................................................................................... 7-24

Aeronca Thrust Reversers ............................................................................................. 7-24

Dee Howard TR 4000 Thrust Reversers........................................................................ 7-28

QUESTIONS......................................................................................................................... 7-33

ILLUSTRATIONS




ILLUSTRATIONS

Figure Title Page

7-1 Major Sections ......................................................................................................... 7-2

7-2 FAN SPEED Indicator ............................................................................................. 7-3

7-3 TURBINE SPEED Indicator.................................................................................... 7-3

7-4 Airflow Diagram ...................................................................................................... 7-4

7-5 Oil System Schematic .............................................................................................. 7-6

7-6 Oil Servicing Access................................................................................................ 7-7

7-7 ∆P Indicator .............................................................................................................. 7-7

7-8 OIL PRESSURE Indicator....................................................................................... 7-8

7-9 OIL TEMPERATURE Indicator .............................................................................. 7-8

7-10 Engine Fuel System ................................................................................................. 7-9

7-11 Electronic Fuel Computer...................................................................................... 7-10

7-24 ENG SYNC Control Switches 7-22




7-24 ENG SYNC Control Switches............................................................................... 7-22

7-25 Thrust Reverser (Aeronca)..................................................................................... 7-24

7-26 Thrust Reverser Levers .......................................................................................... 7-24

7-27 THRUST REVERSER Control Panel (Aeronca) .................................................. 7-25

7-28 Thrust Reverser (Dee Howard TR 4000) ............................................................... 7-28

7-29 THRUST REVERSER Control Panel (Dee Howard)............................................ 7-297

P O W

E R P L A N T

CHAPTER 7




CHAPTER 7

POWERPLANT

MAJOR SECTIONS and its associated planetary gear drive. The fan




MAJOR SECTIONS

For descriptive purposes, the engine (Figure7-1) is divided into seven major sections as follows:

1. Air inlet

2. Fan

3. Compressor

4. Combustor

5. Turbine6. Exhaust

7. Accessory

AIR INLET SECTION

The air inlet section is a specially designed,sound reducing structure enclosing the fan

and its associated planetary gear drive. The fanshroud is armored for blade containment.

FAN SECTION

The fan section includes the single-stage axialfan, an integral spinner, and the fan planetarygear assembly, which is driven by the low- pressure rotor. The rpm of the LP ro tor is designated N1 and commonly referred to as fanspeed.

The planetary gear provides the required gear reduction for the fan. The rpm of the LProtor (N1) is read on the FAN SPEED indi-cator (Figure 7-2). Engine thrust is set usingthis instrument.

7

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Twelve duplex fuel atomizers (spray nozzles)




The fan performs two functions:

• Its outer diameter accelerates a large air mass at a relatively low velocity intothe full-length bypass duct

Th i di f h f l

p ( p y )and two igniter plugs are located in the combustion chamber.

TURBINE SECTION

The turbine section, consisting of a single-stage axial HP turbine and a three-stage axialLP turbine, is located in the path of the exhausting combustion a ir.

The single-stage HP turbine, rigidly joined with the HP compressor, forms the HP spoolthat rotates independently about the LP rotor shaft. The rpm of the HP spool is designated N2 and commonly referred to as turbine speed.The rpm of the turbine (N2 rpm) is read on theTURBINE SPEED indicator (Figure 7-3). Thisis a supporting engine operation instrument.

11 50 6

TURBINESPEED

% X 10

FAN

SPEED

0

1

2

3

456

7

8

9

10

11 0 0 0 0

Figure 7-2. FAN SPEED Indicator

EXHAUST SECTION In addition to these accessories, a DC starter




EXHAUST SECTION

The exhaust section consists of the primary and

bypass air exhaust ducts. The primary exhaustsection directs the combustion gases to theatmosphere. The bypass air exhaust directsthe fan bypass air to the atmosphere.

ACCESSORY SECTION

The accessory section consists of a transfer gearbox and an accessory drive gearbox located on the lower forward side of the engine.The transfer gearbox is driven by a tower shaftand bevel gear from the HP spool. A horizontaldrive shaft interconnects the transfer gearboxto the accessory drive gearbox to drive thefollowing accessories:

• Oil pump

• Fuel pump and mechanical governor within the fuel control unit (FCU)

• Hydraulic pump

• DC generator

,motor is mounted on the accessory drive gearbox to turn the HP spool for engine starting.

OPERATINGPRINCIPLES

The fan (Figure 7-4) draws air through the engine nacelle air inlet. The outer diameter

of the fan accelerates a moderately largeair mass through the fan bypass duct to provide direct thrust . The inner diameter ofthe fan accelerates a smaller air mass into theLP compressor.

Air is progressively compressed as it passesthrough the LP compressor, then to the HPcompressor where a substantial increase in

pre ssure re su lts. Air leaving the HP com - pressor is forced through a transition duct intoa plenum chamber surrounding the combustor.

7

P O WE R P L A N T

The compressed air enters the combustor The combustion gases continue to expand




pthrough holes and louvers designed to directthe flow of combustion air and to keep the

flame pattern centered within the combustor.Each of the duplex fuel nozzles sprays fuel intwo distinct patterns, resulting in efficient,controlled combustion.

The mixture is initially ignited by the twoigniter plugs. The expanding combustiongases, generating extremely high pressures,are directed to the HP turbine, which extracts

energy to drive the integral HP compressor and the accessory section through the tower shaft.

g pthrough the three-stage LP turbine, which extracts energy to drive the LP compressor

through the LP rotor shaft and the fan throughthe planetary gear.

The combustion gases are then exhausted through the exhaust duct. The resulting thrustcreated by the combustion air adds to the thrustgenerated by the fan through the bypass air duct to produce the total propulsion force. Atsea level, the fan contributes 60% of the total

rated thrust, diminishing as altitude increases.At 40,000 ft, the fan contributes approximately40% of the total thrust. Engine core rotation (looking forward) is clockwise, and fan rotation is counterclockwise.

OIL SYSTEM The engine-driven oil pump incorporates one




OIL SYSTEM

GENERALThe oil system (Figure 7-5) provides coolingand lubrication of the engine main bearings,the planetary gear, and the accessory drivegear.

Oil is contained in a tank on the right side of the engine. Access for servicing and level

checking (Figure 7-6) is located on the outboard s ide of each nacel le.

pressure element, four scavenge elements, and a pressure regulator .

The pressure element draws oil from the t ank and provides pressure lubrication for all bear-ings and gears. The scavenge elements returnoil to the tank.

A bypass oil f ilter removes solids from the oil.A red pop-out ∆P indicator provides visual in-dication of a clogged filte r. It can be checked

through a spring port on the right side of eachengine nacelle (Figure 7-7). The indicator button should be flush with the hous ing; if itis not, maintenance is required before flight.

7

P O WE R P L A N T

VENT

NOS. 4 AND 5

BEARING

NO. 6

BEARINGBREATHER

PRESS

VALVE

Oil cooling is fully automatic and is achieved b b i i f i l i il l

Oil venting is provided and controlled by anl i d i b h i i




by a combination of sect iona l air-o il cooler sin the fan bypass duct and a fuel-oil cooler

mounted on the engine. Temperature and pres sure bypass pro tection is provided fo r the oil coolers.

altitude compensating breather-pressurizingvalve.

INDICATION Oil temperature is displayed on individual(Fi 7 9) th i ht id f




Oil pressure is displayed on a single indicator

with dual (L-R) needles (Figure 7-8) on the engine instrument panel that requires 26 VACfrom the L and R OIL PRESS circuit breakerslocated on their respective L and R 26 VAC bus.

gages (Figure 7-9) on the upper right side of the engine instrument panel. Power for these

gages is supplied through the OIL TEMP circuit breaker located on the right essential bus.

7

P O WE R P L A N T

150

100200

50

OILPRESS

P.S.I.

R

L

OIL TEMPC °

30°

60°90° 120°

190°

150°

Figure 7-9. OIL TEMPERATURE Indicator

FUEL SYSTEM aircraft fuel system and directs fuel throughb bl f l filt ith ∆P i di t




GENERALThe engine fuel system (Figure 7-10) pro-vides for fuel scheduling during engine start-ing and accelerat ion to idle, operat ionalacceleration and deceleration, and steady-state operation throughout the entire operat-ing envelope of the aircraft.

FUEL PRESSURE

Engine fuel pressure is generated by a two-stage engine-driven pump. The centrifugal LPstage increases inlet fuel pressure from the

a bypassable fuel filter with a ∆P indicator button to the HP stage. The HP pump increases

fuel pressure to the valve required for efficientoperation of the fuel control unit (FCU). In addit ion, the HP fuel pump supplies the motive-flow fuel for operation of the fuel tank jet pumps (see Chapter 5, Fuel System).

MOTIVE-FLOW LOCKOUT VALVE AND PRESSURE

REGULATORThe lockout valve remains closed initiallyduring engine start to ensure sufficient pres-sure to the FCU. The valve gradually opens

LOW PRESSURE FUEL

ENGINE BLEED AIR

LEGEND

C C

N1N2

PT2TT2

ITT

POWER LEVER

FUEL

COMPUTERHIGH PRESSURE FUEL

fully as fuel pressure increases during thet t O li i ft th ti fl h t

• An ultimate overspeed solenoid valveenergi ed b the f el comp ter at 109%




start. On earlier aircraft, the motive-flow shut-off valve is also closed when the START-GEN

switch is moved to START. A pressure regu-lator maintains motive-flow line pressure for efficient jet pump operation.

FUEL CONTROL UNIT

The fuel control unit (FCU) schedules fuel flowto the fuel nozzles. Its primary mode of oper-ation is the automatic mode (i.e., fuel computer on). In automatic, the FCU responds toelectrical signals from the fuel computer. Thesecondary mode of operation is the manualmode (i.e., fuel computer off or failed). In man-ual, the FCU responds mechanically to thrustlever movement. The FCU includes:

• A mechanical fuel shutoff valve, oper-ated by thrust lever movement between

CUT-OFF and IDLE• A DC potentiometer, mechanically po-

sitioned by thrust lever movement,which electrically transmits this as

energized by the fuel computer at 109% N1 or 110% N2 to shut off fuel

ELECTRONIC FUELCOMPUTER

General

Two electronic fuel computers are located inthe tail cone area (Figure 7-11). They operateon DC power from the L and R FUEL CMPTR circuit breakers on the left and right essential buses, respectively.

7

P O WE

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Figure 7 11 Electronic Fuel Computer




norma l operat io n. Refer to Sect ion IV,Abnormal Procedures of the AFM




Abnormal Procedures, of the AFM .

Manual Mode Operation

When the computer fails or is turned off, thefuel control unit assumes manual control of fuel metering to the engine. The torque motor valve is deenergized and opens fully. The fuelflow is controlled by the mechanical flyweight governor section, functioning as an onspeed governor, utilizing the metering valve. The

surge bleed valve automatically goes to the 1/3open position and remains there.

START PRESSUREREGULATOR

The start pressure regulator (SPR) is a com- pu te r funct io n an d is ava il able on ly in th e

computer-on mode. Manual SPR overrides theautomatic temperature limiting feature of thecomputer. Therefore, ITT monitoring duringSPR operation is extremely important. Itshould be used only during starting and dis

7

P O WE

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Figure 7-13. Fuel Computer andSPR Switches

mode whenever the PLA is 26° or less (42° onearly model computers)

A resettable digital fuel counter (Figure 7-15)is located on the fuel control panel on the




early model computers).

During acceleration, the computer first signalsthe surge bleed valve to assume the 1/3 OPEN posi t ion; i f the surge margin cannot be maintained in this position, the computer willcommand the FULL OPEN position. The opposite is true during deceleration.

In summary, surge bleed valve position is afunction of the fuel computer, relative to N

1,

N2, and thrust lever angle.

FUEL FLOW

Fuel flow is sensed downstream of the FCU and appears on a dual-needle gage on the center in-strument panel (Figure 7-14). The needles arelabeled L and R, and the gage is calibrated in

pounds per hour times 1,000. Electrical power is supplied directly from the battery-charging bus through a 10 A current limiter.

is located on the fuel control panel on the center pedestal. The indica tor is operated by

the fuel flow indicating system and displays pounds of fuel consumed. The indicator should be reset p rior to engine star ting.

FLOW DIVIDER

The flow divider splits fuel flow between the

Figure 7-15. Fuel Counter

IGNITION SYSTEM Ignition is automatically terminated (in a com-puter-on mode) by an electronic speed switch




GENERALA solid-state, high energy ignition systemconsists of a dual ignition exciter on the en-gine and two igniter plugs in the combustionchamber. Two ignition modes are available: au-tomatic and selective.

AUTOMATIC MODE

Automatic ignition occurs during engine starting when the START–GEN switch on thecenter switch panel (Figure 7-16) is positioned to START and the thrust lever i s moved fromCUT–OFF to IDLE.

puter on mode) by an electronic speed switchin the computer at 45% or 50% N2 as deter-

mined by the computer model installed. Power for automatic ignition is provided by the Land R IGN & START circuit breakers on theleft and right power buses, respectively.

If the computer switch is off during a starter-assisted start or if the computer reverts tomanual mode during start, ignition will con-tinue until the START–GEN switch is moved

out of START. Ignition will also terminate(computer on or off) if the thrust lever is moved forward to a position representing approxi-mately 70% N2.

SELECTIVE MODE

Selective ignition is controlled by two-posi-tion switches labeled AIR IGN L and AIR

IGN R located on the center switch panel(Figure 7-16). When the switch is positioned to AIR IGN, the igniters will operate contin-uously. Ignition power is supplied by the L or

7

P O WE

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ENGINE CONTROLS The thrust lever is connected to the FCU by acable. In automatic mode, thrust lever position




Engine control is achieved by thrust levers

mounted on a quadrant on the center pedestal(Figure 7-17). The levers can be moved fromthe fully aft or CUT–OFF position throughthe IDLE position to the fully forward, max-imum power position. A stop is provided at theIDLE position that requires raising a releasetrigger on the outboard side of each lever be-fore the lever can be moved to CUT–OFF.

, pis relayed to the computer as an electrical sig-

nal from a potentiometer inside the FCU thatrepresents thrust lever angle. In manual mode,thrust lever movement changes P3, which op-erates the metering valve. The thrust lever also mechanically operates a rotary fuel shut-off valve.

Optional thrust reverser levers are piggy-back mounted on the thrust levers. (See Thrust Re-

versers, this chapter).

STARTERS The GEN–OFF–START switches are locking-lever switches. They must be pulled out to




GENERALEach engine starter is powered through relayscontrolled by the GEN–OFF–START switchand the fuel computer (during computer-onstarts). A soft start feature incorporates a resistor to minimize the effect of the initialtorque on the mechanical drive components.After a 1.5 -second delay, a relay operates to

allow the starting current to bypass the resistor so that full electrical potential is available tocomplete the start.

Automatic starter disengagement occurs at50% N2 (45% for SNs 35-245 and subsequent,36-045 and subsequent, and earlier aircraftequipped with 1142 fuel computers). On SNs35-370, 35-390, and 36-048 and subsequent,

illumination of a red light under the appropriateGEN–OFF–START switch indicates that thestarter is engaged. On earlier aircraft modified by AMK 80-17, the red lights may be installed

l h h i l

y pmove to the START position. It is not neces-

sary to pull out for movement to any other po-sition.

When either GEN–OFF–START switch is posi tioned to START for a normal computer-on start, the start sequence is initiated for thatengine. The start sequence and circuitry for theleft engine are presented herein; they are iden-tical with those for the right engine.

There are three different designs for the relaycircuits that route power to the starter

• For SNs 35-002 through 35-147 and 36-002 through 36-035, the relays are wired in parallel (Figure 7-18)

One relay is connected to the opposite generator bus and the other to the battery-

charging bus. This arrangement protectsthe 275 A current limiters during initi-ation of each engine start sequence

7

P O WE

R P L A N T

• For SNs 35-148 through 35-389, except35-370, and 36-036 through 36-047, the

For SNs 35-002 through 35-389, except35-370, and 36-002 through 36-047, two




grelays are again wired in parallel, but both

are connected to the battery-charging bus(Figure 7-19)

This design change includes automaticsingle-generator voltage reduction on theground and during airstarts, resulting in275 A current limiter protection when thefirst generator is switched on and duringinitiation of the start sequence on the sec-ond engine

gseparate modifications have been intro-

duced to the starting circuits:

° AMK 80-17 provides a red starter-engaged light for each starter to pro-vide indication of starter engagement(Figures 7-18 and 7-19). Location of the lights is left to customer specifi-cation

GPU

STARTER 275 AMP275

AMP

L

BAT

R

BAT

° AAK 81-1 installs a third starter relayin series between the two existing re-

• For SNs 35-370, 35-390, and subse-quent, and 36-048 and subsequen t, two

l i d i i h




lays and the starter motor; the cir-

cuits that energize the relays areredesigned. AMK 80-17 is a prereq-uisite or concurrent requirement for this modification (Figure 7-20)

starter relays are wired in series to the

ba t t e r y -c harg ing bus , and the r ed s tar ter-engaged l ights are s tandard (Figure 7-21).

OPERATION

SNs 35-002 through 35-389,except 35-370, and 36-002

through 36-047 with or without AMK 80-17

With the aircraft battery switches on, movingthe GEN–OFF–S TART switch t o START connects DC power through the IGN & STARTcircuit breaker to energize the star ter relays.S ta r t e r engagement occur s a long wi th illumination of the starter-engaged light if

AMK 80-17 i s ins t a l l ed . Wi th the fue l computer on, starter disengagement occurs automatically when power is removed fromthe starter relay circuit. At this time the starter-

d li h (if i ll d) i i h

7

P O WE

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STARTER-ENGAGEDLIGHT

L

STARTER

THIRD RELAY:

• ENERGIZED WITH GEN/START SWITCH

IN START


GPU



L STARTER-

ENGAGED

LIGHT

L

BAT

R

BAT

G U

BATTERY-

CHARGING

BUS

275 AMP

CL

275

AMP

CLR

GEN

R

GEN

BUS

NO. 2

RELAYNO. 1

RELAY

LEFT STARTER

CIRCUIT

RIGHT START

CIRCUIT— SAME

AS LEFT

• ENERGIZED WITH START

SWITCH IN OFF OR START

• DEENERGIZED IN GEN

• ENERGIZED WITH START

SWITCH IN START

• DEENERGIZED BY COMPUTER

ABOVE 45% N2

• IF FUEL COMPUTER IS OFF , RELAY REMAINS

ENERGIZED UNTIL START SWITCH IS MOVED

FROM START POSITION

L

STARTER

If the fuel computer is on for the start, it willautomatically deenergize the No. 2 relay whenN h 45% Th t t d li ht

turbine speed reaches 45% or 50% (depend-ing on computer model), the fuel computer

f th t t l ( ) Thi




N2 reaches 45%. The starter-engaged light

extinguishes. Moving the GEN–OFF–STARTswitch to GEN deenergizes the No. 1 relay to backup the release of the No. 2 relay. If bothrelays fail in the energized position, the starter-engaged light remains illuminated, and thestarter remains powered. The only way to dis-engage the starter in this event is to removeelectrical power from the battery-charging bus by turning off both batteries and both gen-

erators.

If the starter-engaged light remains illuminated after start, consult Section IV, Abnormal Pro-cedures, of the approved AFM .

OTHER START FUNCTIONS

In addition to the starter, a number of other circuits are affected when the GEN–OFF–STARTswitch is placed in START. The standby fuel pump in the associated wing is energized, the ig-nition is armed and the Freon air conditioning

removes power from the start relay(s). This

causes the starter to disengage and terminatesignition and standby pump operation. The startsequence can be abor ted at any point by plac-ing the thrust lever to CUT-OFF and theGEN–OFF–START switch to OFF. If enginestart is accomplished with the fuel computer off, the starter is not automatically disengaged after starting. The pilot must position theGEN–OFF–START switch to OFF to terminate

starter engagement and ignition.

Af te r the eng ine r eaches id le rpm, theGEN–OFF–START switch may be placed toGEN. The generator may be turned on whena GPU is connected; however, it is preferableto place the GEN–OFF–START switch to OFFafter starting engines until the GPU is dis-connected.

On SNs 35-002 through 35-147 and 36-002through 36-035 (during battery start), after the first engine is started, one battery switch

b d ff i l i GEN

7

P O WE

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ENGINEINSTRUMENTATION

TURBINE SPEED (N2 )

T rbine speed (N rpm) is remotel sensed




INSTRUMENTATION

GENERAL

The primary engine instruments are in twovertical rows on the center instrument panel(Figure 7-22). From top to bottom these are:

• Turbine speed (N2 rpm)

• Turbine temperature (ITT)

• Fan speed (N1 rpm)

Turbine speed (N2 rpm) is remotely sensed

by a dual monopole transducer installed in thetransfer gearcase. One output signal is sent tothe turbine speed (N2) indicator, and another tothe fuel computer. This indicator includes ananalog scale and pointer calibrated in percent-age of maximum design rpm, and a digitalcounter, calibrated in tenths of percent. A red OFF flag appears on the face of the indicator toindicate loss of DC power to the indicator. The

indicators are powered through the L R TURBRPM circuit breakers located on the left and rightmain buses, respectively.

TURBINE TEMPERATURE (ITT)

Turbine temperature is sensed by ten parallel-wired thermocouples located between the HPand LP turbines. An averager circuit provides

two output signals: one to the turbine temperature indica tor, and the other to the fuelcomputer. The indicator includes an analogscale and pointer , ca l ibrated in degrees

ENGINE SYNCHRONIZERSYSTEM

CONTROL

The system incorporates a single R ENG




SYSTEM

GENERAL

The engine synchronizer system is installed onSNs 35-067 and subsequent, and 36-018 and subsequent as standard equipment. It incorporates a synchronizer control box that uses N1 or N2 inputs from both engine fuel com- pu ter s to enable au tomatic or manua l syn-

chronization of the engines.

The system incorporates a single R ENG

SYNC indicator located on the pilot lower in-strument panel (Figure 7-23), and two ENGSYNC switches located immediately belowthe thrust levers and labeled SYNC–OFF and TURB–FAN, respectively (Figure 7-24). Thesys tem opera tes manua l ly—wi th theSYNC–OFF switch in the OFF position—or automatically—with the SYNC–OFF switch inthe SYNC position—to maintain the right

engine fan or turbine in sync with the left en-gine fan or turbine as determined by the TURB-FAN switch.

7

P O WE R P L A N T

INDICATION

An amber ENG SYNC light (Annunciator Panel

Automatic synchronization is accomplished by selecting SYNC on the SYNC–OFF switch.If the engines are within approximately 2 5%




An amber ENG SYNC light (Annunciator Panel

section) on the glareshield annunciator panel il-luminates anytime the nose gear is down and locked with the SYNC–OFF switch in the SYNC position. The R ENG SYNC (SLOW/FAST) in-dicator indicates right engine rpm deviationfrom that of the left engine.

OPERATION

Manual synchronization is accomplished byselecting OFF on the SYNC–OFF switch. TheR ENG SYNC indicator shows SLOW or FASTout-of-sync condition of the right engine (slaveengine) relative to the left engine (master en-gine). The pilot has the option of selecting ei-ther N2 or N1 as the rpm reference by using theTURB–FAN switch.

If the engines are within approximately 2.5%

rpm of each other, the right engine automati-cally synchronizes to the left engine. It is nec-essary, therefore, to manually sync to within2.5% initially. As in manual sync, either N2 or N1 may be selected as the rpm reference.

DC electrical power is supplied to the systemfrom the left essential bus through the leftFUEL CMPTR circuit breaker to the L FUEL

CMPTR switch.

The amber ENG SYNC annunciator lightserves as a reminder that the system should beturned off.

The engine sync system is inoperative if either fuel computer is off or failed.

THRUST REVERSERS(OPTIONAL EQUIPMENT)

microswitches operated by the reverser levers(Figure 7-26) . The sys tem incorporatesautomatic stow and stow prevention features to




(OPTIONAL EQUIPMENT)

GENERAL

The Learjet 35/36 series aircraft may beequipped with either a cascade thrust reverser system, manufactured by Aeronca, Inc., or a tar-get reverser system (TR 4000), manufactured by the Dee Howard Co. Effective with SNs 35-507 and 36-054, either system is available for

retrof it, but only the target system is availableduring production.

AERONCA THRUSTREVERSERS

General

The Aeronca thrust reverser system incorpo-rates a translating structure (Figure 7-25) thatforms the afterbody of the engine nacelle.When deployed, it exposes cascade vanes

hil i lt l ti t bl k

automatic stow and stow prevention features to

minimize the poss ibi l i ty of inadver tent deployment on the ground and in flight as wellas inadvertent stow at high reverse thrust set-tings. The system is self-arming on the ground through control circuits operating through thelanding gear squat switch relay box.

7

P O WE R P L A N T

THRUST REVERSER ControlPanel

DEPLOY Lights

The two white DEPLOY lights illuminate whenh di h i f ll d l d




The reverser levers control the deploy and s tow cycles and engine power when the reversers are deployed. The THRUST RE-VERSER control panel (Figure 7-27) is lo-cated in the center of the glareshield above the annunciator panel. It incorporates a rocker selector switch for normal and emergency op-erations, seven annunciator lights that pro-vide visual evidence of normal sequencing

and certain abnormal conditions, and a testswitch for performing system test functions.

TEST Button

The TEST button provides a means of checkingoperation of the bleed valve and, on some aircraft, also checks the blocker door positionindicating circuits. When depressed, the whiteBLEED VALVE lights should illuminate, and,on aircraft incorporating AMK 81-6 (installationof blocker door position indicator [DPI] switches),the white UNLOCK lights will flash to indicatethat the blocker doors are correctly stowed

the corresponding thrust reverser is fully deployed.

Both DEPLOY lights must be illuminated; otherwise, the reverser lever solenoid interlockswill not release to permit thrust increase.

UNLOCK Lights

In addition to the test function above, the twowhite or amber UNLOCK lights illuminatesteady while the translating assembly is in

transit during the deploy and stow cycles; thatis, the reversers are not fully deployed or locked in the stowed position.

NORM-EMER STOW Switch

In the NORM position, the red rocker switch provides the electrical circuitry for all normaland automat ic functions. In the EMER STOW

po si ti on, all nor mal electr ic al ci rcu its ar e bypassed, and a separate ci rcuit applies stowcommands to the reversers.

System Operation

Arming

The air motor transmits torque to drive the trans-lating structure aft, exposing the cascade vanes.As the assembly approaches its aft limit, the




Arming

The reversers are automatically armed for normal operation when the following con-ditions exist:

• The T/R circuit breakers are closed

NOTE

The T/R POS and T/R EMER STOWcircuit breakers are located on the left

main bus, and the T/R CONT circuit breaker is located on the right main bus.

• The aircraft is on the ground (squatswitch relay box is in the ground mode)

• The NORM-EMER STOW switch is inthe NORM position

• Both th rust l ever s a re a t the IDLE position

Electrical power for deployment will not beil bl l b th th t l t IDLE

As the assembly approaches its aft limit, the

blocker doors c lose , the DEPLOY l igh t illuminates, the UNLOCK light extinguishes,and the reverser lever solenoid-operated interlock releases.

The reverser lever solenoid-operated inter-lock prevents movement of the reverser leversaft of the idle-deploy position until both DEPLOY lights illuminate. If the pilot is ap-

plying excessive af t pressure on the reverser levers when the DEPLOY lights illuminate, thesolenoid-operated interlock will not release,and reverse thrust above approximately55%–60% N1 will not be possible. The inter-lock will release when aft pressure is relaxed.

For single-engine reversing, both thrust leversmust be at IDLE, and both reverser levers must

be raised to the deploy position in order to deploy the reverser on the operating engine.Since the reverser on the inoperative enginewill not deploy, the solenoid interlock will

l h f h h

7

P O WE R P L A N T

At 60 KIAS, the reverser levers should besmoothly returned to the idle-deploy position.

undetected jammed blocker door could resultin inadvertent deployment of the affected thrust reverser. Each blocker door (upper and




When the engines reach the reverse-idle rpm(approximately 55–60% N1), the pilot maystow the reversers by moving the reverser levers to the full forward position.

Normal Stow

When the reverser levers are moved from theidle-deploy position to the full forward and down position (stow), they operate switchesthat send a stow signal to the directional controlsolenoid of the air motor. The bleed valvecloses, admitting bleed air into the air motor,which causes i t to dr ive the t rans la t ing s tructure toward the s tow posit ion. The DEPLOY lights extinguish and, simultane-ously, the UNLOCK lights illuminate. Whenthe thrust reversers are fully stowed and the pneumatic latches engage the translating structure, the UNLOCK lights extinguish. Asin the deploy cycle, bleed air is shut off to thewindshield, nacelle, and wing/stabilizer heatsystems for approximately three seconds when

( pp

lower) actuates a DPI switch when in the prop-erly stowed position. If the stow cycle is complete (i .e ., la tches engaged) and one of theDPI switches is not actuated, the corre-sponding UNLOCK light flashes to indicatethe jammed blocker door. Since damage tothe system has occur red, repairs are required pr ior to the next takeoff .

A flashing UNLOCK light at any other timeindicates a malfunctioning DPI switch, butthe blocker doors are still properly stowed.This does not preclude operating the reverserson landing.

BLEED VALVE Light

When the reversers are stowed, illumination of a BLEED VALVE light means that the bleed valve is open. This isolates the bleed air-sys-tem from the air motor, and deployment of the affected reverser will not be possible.

switch is not functional with the thrust leversset at any power setting above approximately70% N1. It is therefore imperative that if the

DEE HOWARD TR 4000THRUST REVERSERS




1 p

EMER selection is made for any reason due toa reverser malfunction, the amber EMER STOWindicator light be monitored. If the powersetting is sufficiently high to prelude the emer-gency stow circuits from functioning, the amber light will not illuminate, and the appropriatethrust lever must be retarded until the light illuminates. Illumination of the EMER STOWlight gives visual indication that the emergency

stow circuits have, in fact, activated.

In the event of a system malfunction whileintentionally operating in the reversing range,there is nothing to preclude use of the EMER STOW selection at any time, and doing so willimmediately command all components to stowand illuminate the amber EMER STOW light.

All thrust reverser normal, abnormal, and emergency procedures are contained in thesupplement section of the approved AFM .

GeneralThe Dee Howard thrust reversers incorporatea hydraulically operated system consisting of a pair of clamshell doors forming the afterbodyof the engine nacelle (Figure 7-28). When deployed, the doors deflect al l exhaust in a for-ward direction. The reverser hydraulic systemis integral with the aircraft’s hydraulic system

for normal operation. It is equipped with aseparate accumulator and a one-way check valve that enable one deploy and stow cyclein the event of aircraft hydraulic system fail-ure. The accumulator preload pressure is900–1,000 psi.

An automatic emergency stow system, whichincludes an automatic throttle-retard feature,

is incorporated to provide protection againstinadvertent deployments.

Two pairs of spring-loaded latches—one pair

7

P O WE R

P L A N T


that incorporates four separate solenoid valvesthat are electrically sequenced by microswitches.One of the solenoid valves—the isolation

when the aircraft is on the ground and thethrust levers are at IDLE.



valve—blocks hydraulic pressure at the selec-tor valve inlet until the system is fully armed.The other three solenoid valves are for latchrelease, door stow, and door deploy.

THRUST REVERSER ControlPanel

The reverser levers control the deploy and

s tow cycles and engine power when the reversers are deployed. The THRUST RE-VERSER control panel (Figure 7-29) is lo-cated in the center of the glareshield above theannunciator panel. It incorporates two ARM– OFF–TEST switches (one for each reverser)that provide system arming, disarming, and testing. Four annunciator lights—two for eachrever se r—prov ide v i sua l ind ica t ion

of normal sequencing and certain abnormalconditions.

ARM OFF TEST Switches

The TEST posit ion provides a means of checking operation of the hydraulic isolationvalve. When TEST is selected, the isolationvalve is energized open, and hydraulic pressureis applied to a pressure switch that illumi-nates the ARM light.

The ARM position enables all sequencing mi-croswitches and energizes the isolation valve

open. Illumination of the ARM light indicatesthat the isolation valve has opened and hydraulic pressure is available to the other three solenoid valves for normal sequencing.

The OFF position completely disarms the deploycircuits without disarming the automatic emer-gency stow system.

ARM LightsThe green ARM lights illuminate in conjunc-tion with the TEST and ARM functions as de -scribed above However should the ARM light

DEPLOY Lights

The amber DEPLOY lights flash during allstow/overstow cycles and illuminate steady

doors, while pressure is maintained on the latchrelease actuators.




stow/overstow cycles and illuminate steady

when the respect ive reverser is in the fully deployed posit ion during a normal deployment.A flashing DEPLOY light at any other timeindicates that one or more of the door latchesare unlocked (see Automatic Emergency Stow,this chapter).

System Operation

Arming

The reversers are armed for normal operationas follows:

• The T/R CBs—two for each reverser— are closed

• The aircraft is on the ground (i.e., squatswitch relay box is in the ground mode)

• The respective ARM–TEST switch isin ARM

• The respective thrust lever is at the IDLE

When fully deployed, the doors contact aswitch that illuminates the DEPLOY lightsteadily, deenergizes the latch solenoid valveclosed, and energizes the reverser lever solenoid-operated lock, which releases toallow the reverser lever to be pulled further aftto increase reverse thrust.

Reverse Thrust

When the DEPLOY light(s) illuminate and the reverser lever solenoid-operated lock(s) release, the reverser lever(s) can be pulled further aft to increase N1 to achieve the desired results. A second hard stop limits N1 rpm toapproximately 75%, which constitutes maxi-mum reverse thrust.

At 60 KIAS, the reverser levers should besmoothly started toward the idle deploy position.

Use of maximum reverse power below 50KIAS could cause reingestion of exhaust gases

7

P O WE R

P L A N T

Abnormal Indications

ARM Light Fails to Illuminate

latch position switches on the same side— inboard or outboard—indicate an unlatched condition for the doors, the result is as follows:




During TestIf the ARM light fails to illuminate when TESTis selected on the ARM–TEST switch, the isolation valve failed to respond correctly, hy-draulic pressure is not available, or the pressureswitch is faulty; also, the affected reverser will be inoperative.

ARM Light Fails to Illuminateduring Normal Arming (On theGround at Idle)

If the ARM light fails to illuminate when ARMis selected on the ARM–TEST switch (on theground with thrust levers at IDLE), possiblemalfunctions are:

• Isolation valve failure

• No hydraulic pressure available

• Pressure switch failure

• Thrust lever IDLE switch failure

• T h e is o la t i on v al v e op e n s, w hi c h illuminates the ARM light

• The DEPLOY light begins to flash

• The stow solenoid valve is energized open, which applies stow pressure tothe door actuator and throttle retard ac-tuator, which retards the thrust lever tothe idle position.

The steady ARM light and flashing DEPLOYlight remain on until the latches return to thelatched position or until power is removed from the control circuits.

Automatic Throttle Retard

Automatic throttle retard is designed primarily

to minimize severe thrust asymmetry that mayoccur as a result of inadvertent deployment of a reverser during high thrust settings. This isaccomplished by use of the overstow cycle hy-




7

P O WE R

P L A N T


l Th TFE731 2 2B i id 3 500 6 El t i l f i il


QUESTIONS



l. The TFE731-2-2B engine provides 3,500lb of thrust at:

A. Sea level up to 72°F (22°C)

B. All altitudes and temperatures

C. Sea level at any temperature

D. All altitudes up to 72°F (22°C)

2. The engine LP rotor (N1) consists of:

A. A four-stage, axial-flow compressor and a s ing le - s t age cen t r i fuga lcompressor

B. A single-stage fan and a three-stage,axial-flow compressor

C. A single-stage fan, a four-stage, axial-flow compressor, and a three-stage,axial-flow turbine

D. A four-stage, axial-flow compressor and a four-stage, axial-flow turbine

3 During a normal ground start the ignition

6. Electrical power for engine oil pressureindication is provided by the:

A. Left and right essential buses

B. Inverters through the 26 VAC bus

C. Battery charging bus

D. Pilot and copilot 115 VAC buses

7. The primary engine thrust indicating

instrument is the:A. Turbine (N2)

B. ITT

C. Fan (N1)

D. Fuel f low

8. The maximum ITT during engine start is:

A. 832°C

B. 870°C for ten secondsC. 795°C

D. 860°C

CHAPTER 8

FIRE PROTECTION




CONTENTS

Page

INTRODUCTION................................................................................................................... 8-1

GENERAL .............................................................................................................................. 8-1

ENGINE FIRE DETECTION AND INDICATORS............................................................... 8-2

Sensing Elements and Control Units ............................................................................... 8-2

FIRE PULL or ENG FIRE PULL Lights ........................................................................ 8-3

Fire Detection System Test .............................................................................................. 8-3

ENGINE FIRE EXTINGUISHING ........................................................................................ 8-3

Extinguisher Containers................................................................................................... 8-3

ILLUSTRATIONS

Figure Title Page




8-1 Engine Fire Detection System.................................................................................. 8-2

8-2 Engine Fire Warning Lights and Controls (LH) ...................................................... 8-3

8-3 System Test Switch .................................................................................................. 8-3

8-4 Engine Fire Extinguishing System........................................................................... 8-4

8-5 Fire Extinguisher Discharge Indicators.................................................................... 8-5

8-6 Portable Fire Extinguisher........................................................................................ 8-5

CHAPTER 8FIRE PROTECTION




FIRE PROTECTION

ENGINE FIREDETECTION

wire at its center that carries DC power throughthe detection circuit. The electrical resistanceof the ceramic material is relatively high at nor-mal temperatures; consequently there is lit-




AND INDICATORSSENSING ELEMENTS

AND CONTROL UNITS

Within each engine cowling are three heat-sensing elements: one mounted on the engine pylon f irewall, one mounted around the lower

engine accessory section, and one surround-ing the engine combust ion sect ion . The elements are connected to a control unit thatmonitors the electrical resistance of the sensing elements (Figure 8-1). The sensingelements are made of Inconel metal tubingfilled with a pliable, heat-sensitive ceramicmaterial that, in turn, encloses a conductor

mal temperatures; consequently, there is litt le current flow from the conductor wirethrough the ceramic material to ground (i.e.,outer tubing). At high temperatures, however,the electrical resistance decreases and allowsincreased current flow.

The control unit detects the increased cur-rent flow and illuminates the red FIRE PULLor ENG FIRE PULL light in the T-handlewhen current flow equates to 890°F at the hotsection sensor, or 410°F at the engine acces-sory and/or firewall sensors. DC essential bus elec tr ical power for the system is suppl ied through the L and R FIRE DET ci rcui t breakers on the p ilot and copi lo t CB panels.

8

F I R E

P R O T E

ENG FIRE

PULLCONTROL

UNIT

FIRE PULL OR ENG FIRE PULLLIGHTS

The red FIRE PULL or ENG FIRE PULL warn




The red FIRE PULL or ENG FIRE PULL warn-ing lights are part of the T-handles; one is ateach of the glareshield annunciator panel(Figure 8-2). In the event of an engine f ire, thewarning light in the T-handle will flash untilthe fire or overheat condition ceases to exist.Operation of the T-handles is explained under Engine Fire Extinguishing.

A thermal relief valve on each container is plumbed to a common discharge port (red disc)on the outside of the fuselage below the left en-gine pylon. The thermal relief valves will re-

valve) circuit breakers on the pilot and copilotCB panels, respectively.

There are two ARMED lights above each T-han-




g py

lease bottle pressure at approximately 220°F.

FIRE OR ENG FIRE T-HANDLES AND ARMED LIGHTS

When a FIRE PULL or ENG FIRE PULL light begins to flash, it indicates a fire or overheatcondition in the respective engine cowling.

Following AFM procedures, the pilot should first place the affected engine thrust lever toCUT-OFF and then pull the cor responding T-handle. Pulling out on the T-handle closes themain fuel, hydraulic, and bleed air shutoff valves for that engine. DC essential bus electrical power to close these valves is provided through the L and R FW SOV (f irewall shutoff

g

dle. Pulling either T-handle arms the fire-ex-tinguishing system, which is indicated by illumination of the two ARMED lights abovethe handle pulled. Depressing an illuminated ARMED light momentarily supplies DC power to the explosive cartridge, which discharges thecontents of one f ire- extinguisher bottle and al-lows it to flow into the affected engine nacelle.When the ARMED light is depressed, a holdingrelay is also engaged that extinguishes theARMED light to indicate the associated bottle discharged. Either ARMED light may be depressed to extinguish the fire. Should onecontainer control the fire, the other container isstill available to either engine (Figure 8-4).

8

F I R E

P R O T E A RMED ARME D ARME D ARM ED

NOTE

If the red warning light goes out, thecontinuity of the detection circuit

PORTABLE FIREEXTINGUISHERS




should be tested using the rotary system test switch.

EXTERIOR EXTINGUISHERDISCHARGE INDICATORS

Two colored disc indicators are flush-mounted in the side of the fuselage below the left engine

pylon (Figure 8-5). The red disc covers thethermal discharge port. It will be ruptured if one or both thermal relief valves have released bottle pressure. The ye llow disc will be ruptured if either bottle is discharged by depressing an illuminated ARMED light. Theintegrity of the two discs is checked during theexternal preflight inspection.

One (standard) or two (optional) hand-held fire extinguishers (Figure 8-6) provide for in-terior fire protection. Location of the extin-guisher(s) varies with aircraft conf iguration.




8

F

I R E

P R O T E


l. Engine fire ext inguisher bott les are 4. When an engine fire occurs, the control


QUESTIONS



g g

located in:

A. The nacelles

B. The engine pylons

C. The tai l cone

D. The baggage compartment

2. The power-off preflight check of the enginefire extinguishers includes:

A. Checking the condition of one yellowand one red blowout disc

B. Checking the condition of two yellowand two red blowout discs

C. Checking blowout discs and extin-guisher charge gages, all on the leftside of the fuselage

D. Activating the system TEST switch

to FIRE DET

3. When the left FIRE PULL or ENG FIRE

g

unit:

A. Arms the fire-extinguishing system

B. Il luminates the MSTRWARN lightand sounds the warning horn

C. A u t om a t ic a l ly d i s c h a rg e s t h e respective fire-extinguishing system

D. Causes the respective FIRE PULL or ENG FIRE PULL light in the T-han-

dle and both MSTR WARN lights toflash

5. The fire-extinguishing agent is dis-charged by:

A. A temperature switch

B. A mechanically fired pin at the baseof the supply cylinder

C. The FIRE PULL or ENG FIRE PULLT-handle electr ical circuits

D. Pushing an illuminated ARMED light

CHAPTER 9

PNEUMATICS




CONTENTS

Page

INTRODUCTION................................................................................................................... 9-1

GENERAL .............................................................................................................................. 9-1

DESCRIPTION AND OPERATION ...................................................................................... 9-3

Bleed Air Shutoff and Regulator Valves.......................................................................... 9-3

BLEED AIR Switches ..................................................................................................... 9-4

Bleed Air Check Valves................................................................................................... 9-5

Bleed Air Manifold.......................................................................................................... 9-5

Bleed Air Warning Lights................................................................................................ 9-5

ILLUSTRATIONS

Figure Title Page




9-1 Pneumatic System—SNs 35-002 through 35-112 and 36-002 through 36-031 ...... 9-2

9-2 Pneumatic System—SNs 35-113 and Subsequent and 36-032 and Subsequent ..... 9-3

9-3 BLEED AIR Switches ............................................................................................. 9-4

9-4 Engine Bleed Air Conditioning System (SNs 35-082, 35-087 through 35-112,36-023 through 36-031, and Earlier Aircraft Incorporating AMK 76-7) ................ 9-6

CHAPTER 9PNEUMATICS




stabilizer anti-icing, and hydraulic reservoir pressu riza tion. HP air is used for fan sp inner anti-icing and Aeronca thrust reversers (if installed).

Control of pneumatic bleed air is accom- pl ished with the L and R BLEED AIR switches on the copilot lower right switch panel and by the engine FIRE PULL T-han-

dl i i i d f d i f




On current aircraft, HP air is used for the alcoholanti-icing system, operation of the emergency pressurization valves, as servo pressure for thecabin pressurization and temperature controlsystems, and for control of modulating valveson aircraft with AAK 85-6.

dles. Provision is made for detection of over-heat conditions within the engine pylon, the py lon bl eed-air duct itself, and, on someaircraft, manifold overpressure. Visual indica-tion is given by illumination of warning lightson the glareshield annunciator panel.

BLEED-AIR SHUTOFF

AND REGULATOR VALVE

PYLON TEMP

SENSOR

HP BLEED AIR

LP

BLEED

AIR

FAN SPINNER ANTI-ICE*

T/R AIR MOTOR**

NACELLE ANTI-ICE

DESCRIPTION AND OPERATION

are electrically controlled by the BLEED AIR switches and may also be closed by pullingthe respective engine FIRE PULL T-handle.When open, the valves operate pneumatically

i i d i h




BLEED AIR SHUTOFF AND REGULATOR VALVES

The bleed air shutoff and regulator valves, oneon each engine, are often called “mod valves” because they modulate air pressure. The valves

to maintain downstream pressure in the man-ifold of 27–35 psi.

Both HP and LP bleed air are available to thevalves. As long as enough LP air is available tomeet system demands, the valves will use only LP

FAN SPINNER ANTI-ICE *

T/R AIR MOTOR**

HP BLEED AIR

HP

SOLENOID

VALVE

LP

BLEED

AIR

BLEED-AIR SHUTOFF

AND REGULATOR VALVE

NACELLE ANTI-ICE

BLEED

AIR R

PYLON TEMP

SENSOR

air. If there is not enough LP air available to meetsystem demands, the valves automatically useHP air to maintain the required pressure.

Th h t ff f ti f h h t ff d

On SNs 35-113 and subsequent and 36-032 and subsequent, the L and R BLEED AIR switchesare located on the copilot lower right switch panel). The three-posi tion, OFF–ON –EMER

it h t l th i ti bl d i




The shutoff function of each shutoff and regulator valve is provided by a solenoid- operated shutoff valve that is spring-loaded open; DC power is required to c lose it. Withloss of electri cal power, the shutoff and reg-ulator valves fa il open. However, on SNs 35-1 1 3 a n d s u b s e q u e n t a n d 3 6 -0 3 2 a n d subsequent, an HP solenoid valve, which isspringloaded closed, is installed. On these

aircraft, if electrical power is lost, the shut-off and regulator valve fails open, but the HPsolenoid valve fails closed so that only LP air will be available.

BLEED AIR SWITCHES

On SNs 35-002 through 112 and 36-002through 031, the L and R BLEED AIR switchesare located on the copilot lower right switch pa ne l (Figu re 9-3) . They are two-po si tion ,ON-OFF, switches, powered by the AIR BL

switches control their respective bleed air shutoff and regulator valves and their respec-tive emergency pressurization valves:

• In OFF, the bleed air shutoff and reg-ulator valve is closed, and the emer-gency pressurization valve is in itsnormal position

• In ON, the bleed air shutoff and regula-

tor valve is open, and the emergency pressuriza tion valve remains in i ts nor-mal position (see Figure 9-2)

• In EMER, the bleed air shutoff and reg-ulator valve is open, and the emergency pressuriz ati on va lve is moved to theemergency position. At the same time,the HP solenoid valve is closed, which

restricts the shutoff and regulator valve output to LP air.

On SNs 35-113 through 658 and 36-032th h 063 t difi d b AMK 90 3 th

On SNs 35-659 and subsequent, 36-064 and subsequent, and earlier aircraft modified byAMK 90-3, the L and R BLEED AIR switchesuse DC electrical power from the L and R

BLEED AIR d L d R EMER PRESS i

A temperature sensor in each engine pylonoperates when pylon structure temperatureexceeds 250°F and illuminates the respectivered L or R BLEED AIR light on the glareshield

( A i t P l ti )




BLEED AIR and L and R EMER PRESS cir-cuit breakers on the left and right main DC buses.The BLEED AIR circui t breakers pro-vide power for cont rol of the bleed air shutoff and regulator valves. The EMER PRESS cir-cuit breakers provide power for control of theemergency pressurization valves.

See Chapter 12, Pressurization, for additional

information on the emergency pressurizationvalves.

BLEED AIR CHECK VALVES

A check valve is installed in the bleed air ducting from each engine. Each check valveallows airflow in one direction and blocks airflow applied in the opposite direction. The

check valves prevent loss of bleed air duringsingle-engine operation.

(see Annunciator Panel section).

A temperature sensor installed in each engine pylon bleed air duct causes the respect ive red L or R BLEED AIR light to illuminate if ducttemperature is excessive. SNs 35-002 through35-064 and 36-002 through 36-017 use 590°Fsensors. Later production aircraft use 645°Fsensors.

On SNs 35-082, 087 through 112, and 36-023through 031, a pressure sensor in the regu-lated bleed air manifold causes both BLEEDAIR warning lights to illuminate if pressure inthe manifold exceeds 47 psi (Figure 9-4). Thisalso applies to earlier aircraft incorporatingAMK 76-7 (relocation of cabin air distributionflow control valve).

HP SERVO AIR


CREW OUTLETS

INTERNAL DEFOGOUTLETS

BLEED AIR

LEGEND



LESSBUS

CABHT

AUTO

CABIN AIR DIFFUSERS(TYPICAL)

TO SENSORBLOWER MOTOR

BLEED AIR(LEFT ENGINE)

AIR DISTRIBUTIONTO LOWER CABIN DOOR

CREW OUTLETS

FOOTWARMER OUTLET

CABIN TEMPSENSOR

BAGGAGE COMPARTMENTAIR DIFFUSER(35A AIRPLANE ONLY)

AIR DISTRIBUTIONCHECK VALVES

BLEED AIR(RIGHT ENGINE)

PRESSURE SWITCH

AIRBLEED

BLEED AIR

RAM AIR

CONDITIONED BLEED AIR

LMAIN

BUS

l. Pneumatic air is extracted from:

A The LP compressor

4. The temperature of the bleed air in the

duct between the engine and the bleed


QUESTIONS



A. The LP compressor

B. The HP compressor

C. Ram air

D. Both A and B

2. With loss of DC electrical power, theshutoff and regulator valves:

A. Fail closed

B. Fail open

C. Remain in their last position

D. Can be closed only by pulling a FIREPULL/ENG FIRE PULL T-handle

3. The L and R BLEED AIR ON–OFFswitches are located:

A. On the copilot lower right switch panel

B. On the left side panel

C. On the pilot lower left switch panel

D O th h d l

duct between the engine and the bleed air manifold is monitored by the:

A. Pylon overheat thermostat

B. Aft fuselage equipment section ther-mostat

C. Duct temperature sensor

D. Duct overheat thermostat

5. The BLEED AIR L annunciator illumi-nates:

A. When the temperature in the left pylonor the left bleed air duct is too high

B. When the pressure in the left pylon is below the system’s operat ional l imit

C. When the left half of the bleed-air system is operating

D. When the left half of the bleed-air system has failed

CHAPTER 10

ICE AND RAIN PROTECTION




CONTENTS

Page

INTRODUCTION................................................................................................................. 10-1

GENERAL ............................................................................................................................ 10-1

ICE DETECTION................................................................................................................. 10-2

Windshield Ice Detection .............................................................................................. 10-2

Wing Ice Detection ........................................................................................................ 10-2

ANTI-ICE SYSTEMS .......................................................................................................... 10-5

Engine Anti-ice System (Nacelle Heat) ........................................................................ 10-5Exterior Windshield Defog, Anti-ice, and Rain Removal System ................................ 10-8

ILLUSTRATIONS

Figure Title Page

10-1 Anti ice Control Panel 10-3




10-1 Anti-ice Control Panel ........................................................................................... 10-3

10-2 Wing Ice Inspection Light Control ........................................................................ 10-4

10-3 Wing Ice Inspection Light...................................................................................... 10-4

10-4 Nacelle and Fan Spinner Anti-ice Flow ................................................................. 10-5

10-5 Windshield Anti-ice System (SNs 35-002 to 086 [except 082]

and 36-002 to 022, without AAK 76-7A or AMK 91-2)....................................... 10-9

10-6 Defog Control Knob .............................................................................................. 10-8

10-7 Windshield Anti-ice System (SNs 35-082, 087 to 112, 36-023 to 031,and Earlier Aircraft with AAK 76-7A )............................................................... 10-11

10-8 Windshield Anti-ice System (SNs 35-113 to 662 and 36-032 to 063

without AMK 91-2) .............................................................................................10-13

10-9 Windshield Anti-ice System (SNs 35-663 and Subs.; 36-064 and Subs ; SNs 35-113 to 662 and 36-032 to 063 with AMK 91-2) 10-15

CHAPTER 10ICE AND RAIN PROTECTION




Engine bleed air is used to heat the wind-shields, wing and horizontal stabilizer lead-ing edges, nacelle inlets, and the engine fanspinners on earlier aircraft with elliptical—

dome shaped—spinners

ICE DETECTION

During daylight operation, ice accumulationcan be visually detected on the windshield,wing leading edges and tip tanks




dome shaped spinners.

An alcohol system is installed for radome anti-icing and as a backup to the pilot windshield bleed a ir anti-icing.

On SNs 35-643 and subsequent and 36-058 and subsequent, an auxiliary windshield defogheat system is installed.

All anti-icing equipment must be turned on before icing condit ions are encountered. Todelay until ice buildup is visually detected onaircraft surfaces constitutes an unacceptablehazard to safety of flight.

If anti-ice systems are required during take-off, they should be turned on prior to setting

takeoff power. Appropr iate takeoff power and performance charts must be used.

i di i i h h i i ibl

wing leading edges, and tip tanks.

WINDSHIELD ICE DETECTION

During night operations, the windshield ice de-t e c t i o n l i g h t s i n d i c a t e i c e o r m o i s t u re formation on the windshield. Two probes, oneon the pilot side of the glareshie ld and one on

the copilot side, contain red lights that con-tinuously shine on the inside of the windshield surface. The ice detection lights normallyshine though unseen; however, they will reflectred spots approximately 1.5 inches in diame-ter if ice or moisture forms on the windshield.

The ice detection light on the pilot side is inside the anti-ice airstream; the light on the

copilot side is located outside the anti-iceairstream. For this reason, the copilot lightshould be monitored when flying in icingconditions (anti icing equipment on) The ice


WSHLD HEAT

ON AUTO

WSHLD &

RADOME

STAB

WING

HEAT

OHO

R

WSHLD WSHLD/

HT ON RADOME

STAB

WING

HEAT



35-002 THROUGH 35-112

36-002 THROUGH 36-030

PITOT HEAT

L R

OFF OFF

NAC HEAT

L R

OFF OFF

OFF MAN RADOME OFF

O

F

F

OLD

A

D

35-113 THROUGH 35-642

36-031 THROUGH 36-063

PITOT HEAT

L R

OFF OFF

NAC HEAT

L R

OFF OFF

OFF OFF OFF

On some aircraft, an optional wing iceinspection light is installed on the forward rightside of the fuselage and is focused on a three-inch black dot on the wing leading edge next to

the tip tank. The light is operated by a switchl t d th il t id ll l (Fi




e p a . e g s ope a ed by a sw clocated on the copilot sidewall panel (Figures10-2 and 10-3).

Figure 10-2. Wing Ice Inspection LightControl

ANTI-ICE SYSTEMS

ENGINE ANTI-ICE SYSTEM

(NACELLE HEAT)

Nacelle Heat Switches

Each engine anti-ice system is independentlycontrolled by the L and R NAC HEAT switches

located on the anti-ice control panel (see Fig-ure 10-1).




(NACELLE HEAT)The engine anti-ice system provides anti-icingfor the engine nacelle inlet lips, the ellipticalfan spinners, and the PT2TT2 probes. The nacelle lips are heated with regulated bleed air.The PT2TT2 probe is heated electrically.

On SNs 35-002 through 244 and 36-002

through 044, not incorporating AAK 79-4,the elliptical spinner is anti-iced by high pres-sure bleed air. On aircraft SNs 35-245 and subsequent, 36-045 and subsequent, and ear-lier aircraft incorporating AAK 79-4, a coni-cal spinner replaces the elliptical spinner; noanti-icing is required.

p ( gure 10 1).

When the NAC HEAT switch is turned on (Lor R position), electrical power is supplied toheat the P T2TT2 p robe . The swi tch a lso energizes the fan spinner shutoff valve open(if applicable) and deenergizes the nacelle lipshutoff valve open. Selecting the OFF position

deenergizes the fan spinner shutoff valveclosed and energizes the nacelle shutoff valveclosed. Figure 10-4 is a schematic portrayalof the engine anti-ice systems.

DC electrical power to operate the systems is provided through the L and R NAC HT circuit breakers on the lef t and right main buses.

PT2TT2

Bleed air for nacelle lip anti-icing is takenf rom the regu la t ed b l eed -a i r l i ne ju s t downstream from the bleed-air shutoff and regulator valve (Figure 10-4). It is ducted

through the nacelle heat shutoff valve to a dif-fuser tube that distributes it around the inner

Engine Ice Lights

The amber L and R ENG ICE lights on theglareshield annunciator panel (see Annunci-

ator Panel section) provide a visual indicationof fan spinner or nacelle lip anti-ice system




gfuser tube that distributes it around the inner surface of the nacelle lip and then exhausts itoverboard through a hole at the bottom of thenacelle lip.

The source of fan-spinner heat is high-pressure(HP) bleed air.

) pof fan spinner or nacelle lip anti ice system malfunction. The lights are operated by pres-sure switches in the associated fan spinner and nacelle lip bleed air plumbing. Illumina-tion of an ENG ICE light with the associated NAC HEAT switch on indicates that bleed air pressu re to ei ther the fan sp inner or to thenacelle lip is not sufficient to provide satis-

factory anti-ice protection.

When a NAC HEAT switch is turned on or off,the respective ENG ICE light illuminates mo-mentarily until bleed air pressure at the pres-sure switch agrees with the switch command.

Under some conditions, bleed air pressure maynot be suff icient at idle rpm to keep the pres-

GREEN NAC HT ON LightSNs 35-634 and Subsequent, andSNs 36-058 and Subsequent

A single green NAC HT ON annunciator lighti i t ll d th l hi ld i t




p ynot be suff icient at idle rpm to keep the pres-sure switches from illuminating the ENG ICElight; in this event, advance the thrust leversto check proper nacelle heating operation.

Illumination of either ENG ICE light NACHEAT switches in the OFF position indicatesthe presence of bleed-air pressure in the nacelle

lip or fan spinner plumbing due to a malfunc-tion of the nacelle lip or fan spinner anti-iceshutoff valve. Cycling the NAC HEAT switchon and back to OFF may close the open valve.

g g gis installed on the glareshield annunciator panel. The light illuminates when either NACHEAT switch is on as a reminder that the nacelle heat system is operating.

EXTERIOR WINDSHIELDDEFOG, ANTI-ICE, ANDRAIN REMOVAL SYSTEM

There are five different systems used in thej / id i i d hi ld




Learjet 35/36 to provide exterior windshield anti-icing, defogging, and rain removal. Theywill be covered individually. All systems op-erate on DC power from the WSHLD HT cir-cuit breaker on the left main bus.

SNs 35-002 to 086, except 082,

and 36-002 to 022, without AAK 76-7A or AMK 91-2

The exterior windshield heat/defog system can be controlled either automatical ly or manually(Figure 10-5). It is also used to supplementcockpit heating through the pilot footwarmers,and to provide an alternate bleed air source for emergency pressurization.

An IN–NORMAL/OUT–DEFOG knob, lo-cated below the instrument panel to the left of Figure 10-6. Defog Control Knob




CONTROL

UNIT

DEFOG PRESSURE

FOOTWARMERS

OVERBOARD

DRAIN

WINDSHIELDIN

NORMAL

OUTDEFOG

WSHLD

OV HT

The pressure-regulator valve is solenoid- operated and is deenergized closed. Its functionis to regulate the engine bleed air from the man-ifold to 16 psi. It is energized open when DC

electrical power is applied to the aircraft and will be deenergized and closed to shut off wind-

Manual Operation

Selecting MAN enables the spring-loaded ON–OFF switch to control the shutoff valveand, therefore, the amount of bleed air supplied

to the windshields.




gshield anti-ice in case of windshield overheat.

Automatic Operation

The flow of bleed air to the windshields is con-trolled in the automatic mode by the high(2 5 0 ° F ) a n d l o w (2 1 5 ° F ) t e m p e ra t u re

thermoswitches installed in each windshield outlet nozzle.

NOTE

AAK 77-6 provides for changing thehigh and low limit thermoswitches to290°F and 250°F, respectively.

For ground operation, when the low limitthermoswitch senses 215°F, it will close theshutoff valve; this extinguishes the green

On the ground in manual mode, a low limit ther-moswitch will illuminate the red WSHLD OVHT light, but will not close either the regulator valve or the shutoff valve. However, the highlimit thermoswitch does close the pressure-regulator valve. Therefore, an overheat condi-tion is indicated by illumination of both the

green and red lights, regardless of which limitis exceeded. In flight, the low limit ther-moswitch is disabled.

SNs 35-082, 087 to 112, 36-023to 031, and Earlier Aircraft with

AAK 76-7A

The exterior windshield heat/defog systemcan be controlled either automatically or man-ually (Figure 10-7). It is also used to supple-




WSHLDOV HT

WSHLD

SERVOPRESSURELINE

CHECKVALVE

FOOTWARMERS

TO WING/STABHEAT

TOCABIN

OVERBOARDDRAIN

INNORMAL

OUTDEFOG

WINDSHELD

CONTROLUNIT

ice control panel. It takes four to f ive secondsto cycle fully. Selecting AUTO will open thecontrol valve and illuminate the green WSHLDHT light. If MAN is selected, the control valve

may be opened or closed with the ON–OFFswitch. Since this switch is spring-loaded to

If either outlet nozzle temperature reaches the250°F limit (ground) or 290°F limit (airborne),the thermoswitch will i l luminate the red WSHLD OV HT light on the glareshield an-

nunciator panel and cause the solenoid shut-off valve to close. The anti-ice control valve




neutral, it must be held in the ON positionwhile the valve drives toward the fully open posi tion. The switch may be re leased beforethe valve reaches full open. The control valvewill then stop and remain in an intermediate posi tion. The control valve can be closed only by holding the ON–OFF switch to OFF (with

MAN selected) for at least four seconds.

Operation

With windshield anti-ice on, bleed air flowsthrough the open shutoff valve and anti-icecontrol valve, and through a heat exchanger from which it is ducted to the outlet nozzlesat the base of each windshield.

The anti-ice heat exchanger cools the bleed air with ram air. A ram air modulating valve op-erates to maintain a 300°F duct temperature

will remain in the position it was in, but thegreen WSHLD HT light will be extinguished while the solenoid shutoff valve is closed. TheWSHLD OV HT light will extinguish and theshutoff valve will open again when the temperature at th e thermoswitch coo ls. If thewindshield heat has not been turned off, air-

flow will resume to the windshield, the greenWSHLD HT light will illuminate, and the red WSHLD OV HT light will extinguish.

Through the squat switch relay box, the lowlimit thermoswitches are disabled for 10 sec-onds after touchdown. This prevents auto-matic shutoff of bleed air at the moment of touchdown, which could restrict the pilot’s

vision due to loss of rain-removal capability.

SNs 35-113 to 662 and 36-032 to




TO WING/STAB

HEAT

TO

CABIN

CONTROLUNIT

WINDSHIELD

SERVO

PRESSURE

LINE

For reduced airflow to the windshield, thecontrol valve may be stopped at any interme-diate position by positioning the WSHLD HTswitch to HOLD.

With both valves open, bleed air flows throughh h f hi h i i d d h

Through the squat switch relay box, the lowlimit thermoswitches are disabled for 10 seconds after touchdown. This prevents automatic shutoff of bleed air at the moment

of touchdown, which could restrict the pilot’svision due to loss of rain-removal capability.




a heat exchanger from which it is ducted to theoutlets at the base of each windshield. Theanti-ice heat exchanger cools the bleed air with ram air. A ram air modulating valve op-erates to maintain a 300°F duct temperaturedownstream of the heat exchanger by using aduct temperature sensor and a regulated bleed

air servo line. The subsequent heat loss oc-curring in the duct as the bleed air reachesthe outlet nozzles keeps the outlet airflowtemperature within the limits of windshield heat operation. During ground operation, ramair is not available to cool the bleed air.

Under normal conditions, the windshield heat bleed ai r temperature is au tomatica lly con-

trolled. However, an overheat warning sys-tem alerts the pilot and automatically shuts off windshield heat in the event of an overheat con-di i A l li i ( i l 250°F) d

Bleed air is not available for windshield anti-icing with both the left and right emergency pressurization valves in the emergency position.

SNs 35-663 and Subs.; 36-063

and Subs.; SNs 35-113 to 662and 36-032 to 062 with AMK 91-2

The exterior windshield defog, anti-ice, and rain removal system is shown in Figure 10-9.

With the engines running and the BLEED AIR switches ON, engine bleed air from the regu-lated bleed air manifold is available to twowindshield anti-ice system valves: the anti-iceshutoff valve and the anti-ice control valve.The shutoff valve is solenoid-operated and isnormally energized open whenever electrical


HIGHLIMITS

LOW

LIMITS



LIMITS

TO WING/STAB

HEATSERVO

PRESSURE

LINE

TO

CABIN

CONTROL

UNIT

WINDSHIELD

and positions itself automatically to maintainan air temperature of approximately 300°F.From the heat exchanger, the temperature con-trolled bleed air is directed forward and dis-

pensed over the outs ide of both the pi lot and copilot windshields through outlets at the basef h i d hi ld

which could restrict the pilot’s visibility dueto loss of rain-removal, if the outlet temper-ature is between the inflight and ground limitsat the moment of touchdown.

With loss of electrical power, the windshield anti icing system will be inoperative since the




of each windshield.

Normally, the windsh ie ld anti -ice bleed-ai r temperature is maintained at a safe level by theram air modulating valve. However, an automaticshutdown and warning system has been provided to prevent windshield damage from an overheat

condition. The system uses signals from four thermoswitches, two under the windshield heatair outlets at the base of each windshield.

One thermoswitch on each side operates onlyon the ground while the other operates on theground and in the a i r . H igh l imi t t he r -moswitches are located on the left side and lowlimit thermoswitches are on the right.

If the bleed-air temperature at the windshield reaches a low limit (250°F in flight or 215°F

h d) h i i h ff l i

anti-icing system will be inoperative since theanti-ice shutoff valve will be deenergized and will close. The control valve will remain in itslast position.

Bleed air is not available for windshield anti-icing with both the emergency pressurization

valves in the emergency position.

SNs 35-002 to 112 and 36-002 to031 with AAK 76-7A and AMK 91-2

The exterior windshield heat/defog systemcan be controlled either automatically or manually (Figure 10-10). It is also used to supplement cockpit heating through the pilotfootwarmers and to provide an alternate bleed air source for emergency pressurization

loaded to the center (neutral) position. Theother switch has two positions: AUTO and MAN.

Bleed air from the regulated bleed air mani-fold is routed through two valves: the anti-iceshutoff valve and the anti ice control valve

The con t ro l va lve i s moto r -d r iven and controlled by either of the two switches onthe anti-ice control panel. It takes four to fiveseconds to cycle fully. Selecting AUTO will

open the control valve and illuminate the greenWSHLD HT light. If MAN is selected, thecontrol valve may be opened or closed with the




shutoff valve and the anti-ice control valve.

The shutoff valve is solenoid-operated and isdeenergized closed. Its function is to regulatethe engine bleed air from the manifold to 16 psi. It is energized open when DC electr ical power is applied to the aircraft and will be

deenergized and closed to shut off windshield anti-ice in case of windshie ld overheat.

control valve may be opened or closed with theON–OFF switch. Since this switch is spring-loaded to neutral, it must be held in the ON position while the valve drives toward the fullyopen position. The switch may be released before the valve reaches full open. The controlvalve will then stop and remain in an interme-

diate position. The control valve can be closed only by holding the ON–OFF switch to OFF(with MAN selected) for at least four seconds.

HIGH

LIMITSLOW

LIMITS

WINDSHIELD

INNORMAL

Operation

With both valves open, regulated engine bleed air flows through a heat exchanger in whichit is cooled by ram air. The ram air flow is

cont ro l led by a pneumat ica l ly ac tuated modulating valve. The modulating valve senses

If the bleed air temperature at the windshield reaches a low limit (250°F in flight or 215°Fon the ground), the anti-ice shutoff valve isdeenergized closed and the green WSHLD HT

light is extinguished. When the overheat cools,the thermoswitches will reset and the anti-iceshutoff valve will reopen If the anti ice control




g g bleed ai r temperature downstream of the heatexchanger through a temperature sensor and posi tions itse lf automatica lly to mainta in anair temperature of approximately 300°F. Fromthe heat exchanger, the temperature controlled bleed air is directed forward and dispensed over the outside of both the pilot and copilot wind-

shields through outlets at the base of eachwindshield.

Normally, the windshie ld anti -ice bleed ai r temperature is maintained at a safe level by theram ai r modula t ing valve . However , an automatic shutdown and warning system has been provided to prevent windshield damagefrom an overheat condition. The system uses

signals from four thermoswitches, two under the windshield heat air outlets at the base of each windshield

shutoff valve will reopen. If the anti-ice controlvalve is still open, the green WSHLD HT lightwill illuminate and windshield anti-ice airflowwill be restored.

If the bleed air temperature at the windshield reaches a high limit (270°F in flight or 250°F

on the ground; 215°F on the ground on aircraftwith electrically heated windshields), the anti-ice shutoff valve is deenergized closed, thegreen WSHLD HT light is extinguished, and thered WSHLD OV HT light illuminates. When theoverheat cools, the thermoswitches will reset,the red WSHLD OV HT light extinguishes, and the anti-ice shut-off valve will reopen. If theanti-ice control valve is still open, the green

WSHLD HT light will illuminate and wind-shield anti-ice airflow will be restored.

Th d li i h i h di bl d

With loss of electrical power, the windshield anti-icing system will be inoperative since theanti-ice shutoff valve will be deenergized and will close. The control valve will remain in its

last position.

INTERNAL WINDSHIELD

Positioning the switch to CKPT applies DC power to the coil , heat ing al l the air cominginto the cockpit.

Positioning the switch to W/S AUX DEFOGHEAT again applies DC power to the coil,heating all the air coming into the cockpit It




INTERNAL WINDSHIELDDEFOG

All aircraft use conditioned engine bleed air for internal windshield defog (see Chapter 11, Air Conditioning, for additional infor-mation). On late model aircraft, auxiliary

interna l windshield defog systems have been provided.

SNs 35-643 to 670

The internal windshield defog system on theseaircraft uses an electrically heated coil in the bleed air duct leading into the cockpit, and theFreon air conditioning system. It is controlled by a three-posi tion (OFF–CKPT–W/S AUXDEFOG HEAT) switch on the anti-ice controlpanel

heating all the air coming into the cockpit. Italso arms the Freon air conditioning systemso it will turn on automatically as the airplane descends through 18,000 ft. When theFreon air conditioning system turns on, elec-trically actuated diverter doors on the cabin blower assembly open and direc t the cold ai r

into the space between the cabin headliner and the fuselage skin. This dehumidifies thecabin air without lowering the cabin temper-ature excessively (see Chapter 11 for addi-t i o n a l i n fo rm a t i o n o n t h e F re o n a i r conditioning system).

DC electrical power to heat the auxiliary windshield defog coil is provided by the

battery charging bus through two, 20 A currentlimiters. DC control power for the auxiliarywindshield defog system is provided by theAUX DEFOG i i b k h l f

SNs 35-671 and Subsequent;36-064 and Subsequent

The internal windshield defog system on these

aircraft is shown in Figure 10-11. It uses 163VAC power from the auxiliary and secondaryinverters and is controlled by a two position

annunciator light. If an overheat condition exists, the relay box will also remove electrical power from the he ating el ement in the af-fected windshield.

The d i f fe rence be tween an overhea t o r underheat temperature condit ion may be




inverters and is controlled by a two-position(OFF–WSHLD DEFOG) switch located onthe anti-ice control panel (see Figure 10-1).

When the switch is positioned to WSHLDDEFOG, DC control power is applied to awindshield defog relay box. The relay box

receives 163 VAC power from the auxiliary and secondary inverters—through 5 A current lim-iters—and directs it to the heating elements inthe windshield. Each heating element is a thin,gold film laminated in the windshield. Theauxiliary inverter powers the element on theleft side, and secondary inverter powers the element on the right side.

Both heating elements are turned on and off together; once operating, however, the two elements are controlled separately by the

underheat temperature condit ion may be determined by touching the windshield. If anoverheat temperature condition is suspected,and the windshield does not cool off, the relay box has not removed elec tr ical power fromthe heating element; the system should beturned off.

A windshield temperature of 90°F or below iscommon when the defog system is f irst turned on, and the annunciator light will illuminate.However, the light should soon extinguish asthe windshield warms up.

The WSHLD DEFOG annunciator l ight , located to the left of the left ENG FIRE PULL

T-handle, consists of three separate lights and is controlled by the windshield defog relay box. The upper WSHLD DEFOG light wil l il-l i h i h f h l li h


WSHLDDEFOG

L R



HEATING ELEMENT(GOLD FILM) HEATING ELEMENT(GOLD FILM)

WINDSHIELD DEFOG RELAY BOX

BELOW 90/ ABOVE 150° F

BELOW 90/ ABOVE 150° F

WINDSHIELD/RADOME ALCOHOL ANTI-ICE SYSTEM

Methyl alcohol from a reservoir located in the

left side of the nose compartment is provided to prevent ice formation on the radome and,if necessary the pi lot windshield as a backup

In this case, a fully serviced reservoir should dispense alcohol for approximately 1.5 hours.

When the switch is positioned to WSHLD &RADOME, the pump is energized and thesolenoid valve in the windshield supply line isenergized open so that alcohol is delivered to




if necessary, the pi lot windshield as a backupfor the windshield anti-ice—defog—system.The systems are operated by DC power fromthe right main bus.

There are two different systems in use.

SNs 35-002 to 112 and36-002 to 031

A DC motor-driven pump supplies filtered alcohol from a 2.25 gallon reservoir to theradome only, or to the radome and pilot wind-shield, depending on the position selected onthe WSHLD/RADOME switch on the pilotanti-icing control panel.

When the switch is positioned to RADOME,the pump is energized and alcohol is delivered

g p both surfaces. Flow to the windshield is dispensed through an orifice assembly integrated with the pilot defog outlet. In this case, dura-tion is reduced to approximately 43 minutes.

A pressure switch installed in the radome supply line actuates the amber ALC AI lightwhen the reservoir is empty or if the pumpfails. The light will extinguish when the controlswitch is turned off (Figure 10-12).

The reservoir is vented through an open venttube located in the same area as the pitot-staticdrains on the left side of the nose compartment.A pressure relief valve operates to relieve ex-

cessive supply line pressure by returning it tothe reservoir. Some aircraft are equipped witha siphon-break valve to prevent the siphoning

f fl id f h k f h h b


LEGEND



PILOT'S EXTERNALDEFOG OUTLET

ORIFICE ASSEMBLY

ANTI-ICEVALVE (NC)

WSHLD &RADOME

O

RADOME

LOW-PRESSURESWITCH

ALC AI

ALCPMP

SIPHON-BREAK

LEGEND

SUPPLY

PRESSURE

RETURN

AMBIENT

ELECTRICAL CIRCUIT

EFFECTIVE WITH

35-076, 36-021

*


SNs 35-113 and Subs.;36-032 and Subs.

Methyl alcohol is stored in a 1.75 gallon reservoir. When the cockpit control switch is po sit ioned to WSHLD/R AD OM E or RAD,circuits are completed to position a three-way

The pressure relief valve, which is set at 2.6 psi,relieves any overpressure in the reservoir should the pressure regulator fail; it also bleeds off residual pressure when the control switch isturned off.

The float switch in the reservoir illuminates



valve in the fluid supply line (Figure 10-13)and to open the shutoff valve and pressureregulator in the servo bleed air supply line.

Servo bleed air tapped from the high pressure bleed air manifold passes through the shutoff valve and pressure regulator where i t is regulated to 2.3 psi and sent to pressurize thealcohol reservoir.

The alcohol is forced through a filter to thethree-way valve that is positioned accordingto the selected switch position.

the ALC AI annunciator when the tank isempty. The light stays on even if the switch isoff as a reminder to service the reservoir.

If the RAD position is selected, a fully serviced reservoir supplies only the radome with approximately 2 hours and 9 minutes of alco-h o l . W h e n s e l e c t e d t o t h e W S H LD /RADOME position, alcohol is also dispensed to the pilot defog outlet via the three-wayvalve;, duration of the supply is reduced toapproximately 45 minutes. This system is stilloperational if both emergency pressurization

WSHLD/ RADOME

R

valves are in emergency (provided DC power is available).

WING AND HORIZONTAL

STABILIZER ANTI-ICE SYSTEMBleed air is used to prevent ice formation on the

Controls and Indications

The STAB WING HEAT switch located on the pilot anti-icing control panel controls the valve.When the switch is moved up (on), the valve is

energized open. With the switch off, or if DC power fails, the valve deenergizes closed.




Bleed air is used to prevent ice formation on thewing and horizontal stabilizer leading edges.The bleed air is directed from the regulated en-gine bleed air manifold through a solenoid-op-erated pressure regulator valve (Figure 10-14)to the respective leading edge surfaces.

With the valve open, manifold bleed air isrouted through the wing-stabilizer pressureregulator valve—where it is regulated to 16 psi—to piccolo tubes in the leading edges of the wing and the horizontal stabilizer. After

W

I

N

G

T

E

M

P

35°

215°

WING

OV HT215°

the bleed air has heated its respective leadingedge, it continues outboard where it ventsoverboard; each wing has a scupper vent on theunderside of the leading edge, while the hor-izontal stabilizer has holes at each tip.

On the glareshield annunciator panel, red WING OV HT d STAB OV HT li h




WING OV HT and STAB OV HT lights are illuminated should their respective sensors(see Figure 10-14) detect 215°F.

Separate WING TEMP and STAB TEMP indicators on the center instrument panel(Figure 10-15) indicate leading-edge skin tem- pera ture and are color-coded as fol lows:

• Red—Temperature below 35°F (danger of icing in visible moisture)

• Green—Temperature between 35–215°F(normal operation)

• Yellow—Temperature above 215°F (pos-sible overheat)

When either overheat light comes on and thesystem is turned off, the light will remain on

Figure 10-15. WING TEMP and STAB

TEMP Indicators

PITOT, STATIC, AND ANGLE-OF-ATTACK VANE ANTI-ICE SYSTEM

Pitot and Angle-of-Attack Vane Anti-icing

Static Port Heating(FC 200 Only)

There are f ive static ports: two on the left sidefu s e l a g e a n d t h r e e o n t h e r i g h t . P i l o t instruments are supplied static pressure bythe interconnected left front and right center

t hi h h t d Th i t t d




The left and right pitot tubes and angle-of- attack (AOA) vanes contain electrical heatingelements. The L and R PITOT HEAT switcheslocated on the pilot anti-icing control panel (seeFigure 10-1) each supply essential bus power to both respective heating elements. Even

though each se t o f hea t ing e l emen t s i s controlled by the same switch, separate circuit protection for the AOA vane heater is provided;the L and R PITOT HT circuit breakers (for pitot heaters) and S WRN HT circuit breakers(for the AOA vane heaters) are all located onthe left and right essential buses, respectively.

On FC 530 aircraft, one heating element in

each pitot-static probe heats all of the pitot and static ports.

ports , which are heated. The interconnected left rear and right front static ports supplycopilot static pressure and are also heated.

The right rear port—interconnected with an al-ternate port inside the nose compartment—isused by the altitude controller and does not require heat.

Two additional shoulder-static ports forward of the windshield are also heated. These por tsare used by the air data sensor.

All static por t heating elements are connected directly to their respective L or R PITOT HTcircuit breakers. Consequently, they are heated whenever aircraft DC power is available, provided the CBs are closed (i .e ., in).





1. Bleed air is used for anti-icing on:

A. Pitot tubes and static ports

B. PT2/TT2 sensors

5. Anti-icing equipment must be turned on:

A. When in icing conditionsB. Before entering icing conditions

C Before takeoff


QUESTIONS



C. W in g a n d h o r i z on t a l s t a bi l i ze rleading edges

D. Conical fan spinners

2. The L or R PITOT HEAT switches alsosupply heating element power for:

A. The angle-of-attack vanes

B. The shoulder static ports

C. The instrument static ports

D. PT2/TT2 probe heater

3. The crew action required when the red WING OV HT light illuminates is :

A. No action is required; the system isautomatic

B. Pos i t ion the STAB WING HEAT

C. Before takeoff

D. During climbout

6. With the loss of aircraft electrical power,anti-icing will be lost on:

A. All systems

B. Pitot, static, and PT2/TT2 probes onlyC. All systems except the nacelle inlet

lips

D. All systems except the windshield and radome alcohol system

7. The L NAC HEAT switch in the up (on) posi tion provides ant i- icing to a ll of thefollowing except the:

A. Nacelle l ip

B Dome spinner (early models)

CHAPTER 11

AIR CONDITIONING

CONTENTS

P




Page

INTRODUCTION................................................................................................................. 11-1

GENERAL ............................................................................................................................ 11-1

ENGINE BLEED AIR CONDITIONING AND DISTRIBUTION ..................................... 11-2

General .......................................................................................................................... 11-2

Flow Control Valve........................................................................................................ 11-6

Hot Air Bypass Valve (H-Valve) ................................................................................... 11-6

Ram Air Heat Exchanger............................................................................................... 11-7

Ram Air Ventilation....................................................................................................... 11-7

ILLUSTRATIONS

Figure Title Page

11-1 Engine Bleed Air Conditioning System (SNs 35-002 to 35-086[except 35-082] and 36-002 to 36-022) ................................................................. 11-3

11 2 E i Bl d Ai C diti i S t (SN 35 082 35 087 t 35 112




11-2 Engine Bleed Air Conditioning System (SNs 35-082, 35-087 to 35-112;36-023 to 36-031; and Earlier Aircraft Incorporating AMK 76-7) ....................... 11-4

11-3 Engine Bleed Air Conditioning System (SNs 35-113 and Subsequent and 36-032 and Subsequent)......................................................................................... 11-5

11-4 CABIN AIR Switch ............................................................................................... 11-6

11-5 Temperature Control Indicator............................................................................... 11-7

11-6 Conditioned Bleed Air Distribution....................................................................... 11-8

11-7 CABIN CLIMATE CONTROL Panel ................................................................... 11-9

11-8 CABIN TEMP Indicator...................................................................................... 11-11

11-9 Evaporator and Blower Assembly ....................................................................... 11-13

CHAPTER 11 AIR CONDITIONING




ENGINE BLEED AIRCONDITIONING

AND DISTRIBUTION

GENERAL

Regulated engine bleed air, supplied to a manifold located in the tail cone section, isducted to the flow control valve. From the flowcontrol valve, there are three slightly differentcabin and cockpit distribution configurations;

each performs the same basic functions, but differs in component arrangement. Figures 11-1, 11-2, and 11-3 depict the three basic con-


1 1

A I R C O N D I T I O N I N G



This sec t ion addresses the condi t ioning process that the engine bleed air is subjected to before it enters the cabin area, beginning atthe flow control valve. Chapter 9, Pneumatics,describes the bleed air supply system. Chapter 12, Pressurization, descr ibes how conditioned

bleed a ir is used for cabin pressuriza tion.

1, 11 2, and 11 3 depict the three basic conf igurations by aircraft serial number.

All three configurations use the flow controlvalve to control the flow of bleed air througha hot air bypass valve and an air-to-air heat exchanger before it enters the cabin and cock-

pi t d istr ibut ion ducting.





LESSBUS

CABHT

AUTO



BLEED AIR( G )


CREW OUTLETS

FOOTWARMER OUTLET

CABINTEMP

SENSOR



AIRBLEED

BLEED AIR

LEGEND

RAM AIR



1 1


CREW OUTLETS


FOOTWARMEROUTLET

BLEED AIR

LEGEND

RAM AIR



LESSBUS

CABHT

AUTO





CABIN TEMPSENSOR

BAGGAGE COMPARTMENTAIR DIFFUSER(35A AIRCRAFT ONLY)



AIRBLEED

CONDITIONEDBLEED AIR

L

MAINBUS


BLEED AIR

LEGEND

RAM AIR


REGULATED SERVO AIR


CREW OUTLETS

FOOTWARMER OUTLET



LESSBUS

AIRBLEED

REGULATED SERVO AIR


EMERGENCYPRESSURIZATIONVALVE

ON

AUXDEFOG/CREWHEATER *


CABINTEMP

SENSOR

CABIN AIR DIFFUSERS

(TYPICAL)

BAGGAGE COMPARTMENTAIR DIFFUSER

(35A AIRCRAFT ONLY)

CHECK VALVES

FLOW CONTROL VALVE

The flow control valve is a solenoid-operated valve that controls and regulates the flow of bleed air into the cabin. The posi tion of the

valve is determined by the CABIN AIR switch(Figure 11-4). The most current aircraft (seeFigure 11-3) use a two-position OFF–ON

i h li i f ( i d

Airflow through the venturi is measured by pneumatic sensing lines connected to a mod-ulating mechanism in the flow control valvewhich ensures that airflow remains constantwhen engine power changes occur.

HOT AIR BYPASS VALVE(H VALVE)


1 1




switch. Earlier aircraft (see Figures 11-1 and 11-2) use a three-position OFF–NORM–MAXswitch. When the CABIN AIR switch is inOFF, the valve is energized and closes. Whenthe switch is in ON or NORM, the valve isdeenergized and opens. In MAX, the valve

opens fully to provide an increase in airflowto the cabin. DC power for valve operation i s provided through the AIR BLEED circuit breaker on the left essential bus .

A venturi, located downstream of the flowcontrol valve, adjusts the valve to smooth outthe flow of bleed air as it enters the cabin.

(H-VALVE)

A butterfly bypass valve, more commonly referred to as the “H-valve,” is located in the bleed-air duct upstream of the heat exchanger.Its function is to split the flow of bleed air,

directing some to the heat exchanger for cool-ing and some to bypass the heat exchanger. Theresult is a mixture of the two airflows, therebyconditioning the bleed air before it enters thecabin area. The position of the H-valve is in-dicated on the TEMP CONT indicator located in the lower center instrument panel (Figure11-5).

air then exhausts overboard through a port inthe lower left s ide of the fuselage.

The cooled bleed air flowing out of the heatexchanger core is ducted back to the bypass

side of the H-valve where it mixes with hot bypassed bleed air. The result ing condit ioned air is then directed into the cabin and cockpitdi ib i




On SNs 35-002 through 35-112 and 36-002

through 36-031, the H-valve butterfly is positioned by a DC electric motor operated byh li l A i l 25

distribution system.

When the aircraft is on the ground, do not per-form extended engine operation above idlew i t h t h e C A B IN A IR a n d B LEED A IR switches positioned to ON. Since there is no

ram air for cooling of the bleed air, possibledamage to the air conditioning componentscould result. Damage might also occur to in-terior cabin furnishings, as well as overheatingthe tail cone area.

On SNs 35-082, 35-087 through 35-112, 36-023 through 36-031, and earlier aircraft in-corporating AMK 76-7, the flow control valve

is located downstream of the hea t exchanger.Engine bleed air is available to the heat ex-h h i i i d

Figure 11-5. Temperature ControlIndicator

CABIN AND COCKPIT AIR DISTRIBUTION

Conditioned airflow distribution to the cabinand cockpit areas is essentially the same for all

aircraft (Figure 11-6). The conditioned air isrouted from the tail cone into the cabin areathrough two ducts, one on each side of the

On SNs 35-113 and subsequent and 36-032 and subsequent (see Figure 11-3), distribution of air changes when either (or both) emergency pressurization valves are posi tioned to emer-gency.

If only one emergency valve is positioned toemergency, all bleed air from that engine is

t d di tl i t l th bi di t ib ti


1 1




cabin. The left duct ends at the entry door, and the right duct continues forward to the cockpit.

Cabin Air Distribution

Cabin air distribution is furnished by diffusersinstalled at intervals along the two ducts; thediffusers direct airflow toward the floor.

A one-way distribution check valve is located at the aft end of each cabin duct. These valvesare functionally related to the pressurizations y s t e m , a s d e s c r i b e d i n C h a p t e r 1 2 ,Pressurization.

routed directly into only the cabin distributionduct on that side; temperature control of thatair is los t. However, bleed air from the oppo-site engine is still subject to the normal con-ditioning process. One-way check valves in thenormal distribution ducting prevent the emer-

gency airflow from being lost through the nor-mal distribution system.

If both emergency valves are positioned toemergency, all bleed air from both engines isrouted directly into the respective left and right distribution ducts. Temperature controlis then sacrificed for pressurization.

Cockpit Air Distribution

Cockpit air distribution is provided by theducting connected to the forward end of theright hand cabin duct. Four WEMAC outlets

—two on each s ide of the cockpit and located on the sidewall panels adjacent to the outboard rudder pedals—enable the pilots to con-trol and direct the airflow as desired A

position results in improved interior defoggingfor the sides of the windshield.

Interior windshield defogging can be maximized by closing the four WEMAC outlets to divert

the maximum amount of conditioned air tothe windshield piccolo tubes.

On SNs 35 002 to 35 112 and 36 002 to 36




t rol and direct the airflow as desired. Afootwarmer diffuser, which is below the in-strument panel just forward of the center pedestal , directs continuous condit ioned ai r along the center floor. Two piccolo tubes in-stalled vertically on each side of the wind-

shie ld center support s t ructure d i rec t acontinuous flow of conditioned air across theforward section of each pilot windshield for interior windshield defogging.

On SNs 35-328 and subsequent and 36-050 and subsequent, increased continuous interior wind-shield defogging capability has been provided.Two additional piccolo tubes are installed, one

for each windshield. They are positioned hori-zontally along the lower edge and extend for-ward from the aft corner of the windshield This

On SNs 35-002 to 35-112 and 36-002 to 36-031, additional heat is available to the cockpit via separate footwarmers that operate fromthe windshield heat/defog system discussed inChapter 10, Ice and Rain Protection.

TEMPERATURE CONTROL

Temperature control of the engine bleed-air entering the cabin area is accomplished byvarying the position of the H-valve butterfly.As the valve opens, less bleed air is directed to the heat exchanger for cooling, while more bleed air is bypassed and mixed with the cooled air. Manual and automatic operation of the H-valve is achieved by controls on the CABINCLIMATE switch panel, located on the copilot

On SNs 35-002 to 35-112 and 36-002 to 36-031, the climate control system is operatedelectrically. System control is accomplishedwith a rheostat and a HOT–COLD toggleswitch that is spring-loaded to cente r. Other

system components include:

• A temperature sensor located behind thecopilot seat

On SNs 35-107, 35-113 and subsequent, and 36-032 and subsequent, the H-valve is posi-t ioned pneumatical ly by servo bleed air (Chapter 9, Pneumatics), and no electrical cir-cuits are involved.

The CLIMATE CONTROL panel (see Figure11-7) incorporates two control knobs. TheAUTO MAN knob is actually a servo bleed air


1 1




p

• A duct temperature sensor and duct temperature limiter; both are in the air ductdownstream of the H-valve (see Figure11-1 or 11-2, as applicable)

• A cont ro l uni tIf the rheostat is turned fully counterclock-wise to the MAN detent, the cabin temperature sensor and duct temperature sensor are both off.The H-valve is then controlled manually by ac-tuating the spring-loaded HOT-COLD switch.The TEMP CONT indicator (see Figure 11-5)displays the position of H-valve. DC power for manual operation is provided by the CABIN

HT MAN circuit breaker on the right main bus.The TEMP CONT indicator is powered from

AUTO–MAN knob is actually a servo bleed air selector valve. The COLD–HOT knob is aneedle valve that controls the servo air pres-sure applied to the H-valve butterfly (spring-loaded to the full cold position). Other systemcomponents include a temperature sensor

located in the upper forward cabin, a duct tem- pera ture sensor , and a duct temperature lim-iter located in the air duc t downstream of theH-valve (see Figure 11-3). The control systemconsists of an interconnected servo bleed air network.

With the AUTO–MAN knob in MAN, the se-lector valve isolates the control system from the

influences of the cabin temperature sensor and the duct temperature sensor. Servo air pres-

i d di l h h h dl

For manual or automatic operation in case of a duct overheat, the duct temperature limiter causes the control system to shut off servo air pressure to the H-valve butterf ly. This a llowsit to spring to the full cold position and direct

all bleed air through the heat exchanger.

A CABIN TEMP indicator may be installed onthe center pedestal or instrument panel to

AUXILIARY AIRCONDITIONINGSYSTEMS

GENERAL

Additional air circulation is provided by a




the center pedestal or instrument panel toindicate the temperature in the cabin from aremote sensor (Figure 11-8).

Additional air circulation is provided by acabin blower and a cockpit fan, ducted thoughdistribution networks also used by the Freonrefrigeration system (auxiliary cooler) and the optional electrical heating system (auxil-iary heater).

The cabin blower and cockpit fan may be used simply to recirculate air within the cabin and cockpit areas, or the auxiliary cooler or aux-iliary heater can be used to cool or heat the recirculated air.

For operational requirements on the ground (subject to certain limitations), it is possible to

precool or preheat the cabin prior to engine star t.Figure 11-8. CABIN TEMP Indicator

DISTRIBUTION SYSTEM

The heart of the distribution system is theevaporator and blower assembly, which is inthe cabin ceiling above the baggage com-

partment (Figure 11-9). The assembly houses:• Ducting

• Cabin blower assembly

CLOSE, the doors are flush with the bottom of the ducting; the airflow from the cabin blower is directed into the cabin.

On aircraft with the electric diverter doors,

the doors are controlled by a two-position,ON–OFF, rocker switch below the cabin blower ai r outlet . When the switch is posi -tioned to OFF the diverter doors are lowered


1 1




• Cabin blower assembly

• Cockpit fan assembly

• Freon system evaporator

• Optional electrical heating elements

(when installed)

Cabin Blower Distribution

The cabin blower assembly consists of twosquirrel-cage blowers driven by a single DCmotor. The blowers draw air from the bag-gage compartment area though the evaporator and force it through separate ducts to a lou-

vered grille at the front the ducts. The air isdiffused as it blows out directly into the cabin.When installed the optional heating elements

tioned to OFF, the diverter doors are lowered into the airf low from the cabin blower, whichdirects the air into the space between the cabinheadliner and the fuselage skin. When theswitch is positioned to ON, the diverter doorsare flush with the top of the ducting; the air-

flow from the cabin blower is directed into thecabin. On SNs 35-643 to 35-670, the doors mayalso be controlled by the auxiliary internalwindshield defog system (see Chapter 10, Iceand Rain Protection).

When used simply for additional air circula-tion, the cabin blower is turned on by selectingFAN on the three -position FAN–OFF–COOL

switch on the CABIN CLIMATE CONTROL panel ( see Figure 11-7). DC electr ical power i id d b h CAB BLO i i b k




CABIN BLOWER rheostat on the copilot side-wall panel (Figure 11-11). Earlier aircraft donot have this feature unless AMK 77-4 is incorporated.

Cockpit Fan Distribution

Between the two ducts fed by the cabin blowersis another duct that encloses the axial cockpit

The cockpit fan is controlled by the COCK-PIT AIR rheostat on the copilot sidewall panel(Figure 11-11) using DC power from the CABBLO circuit breaker on the left main bus.

• On SNs 35-002 to 35-112, and 36-002to 36-031, the OFF detent is at the fullc l o c k w i s e p o s i t i o n ; f a n s p e e d i s increased by rotating the rheostat in a

t l k i di ti


1 1




is another duct that encloses the axial cockpitfan. This fan draws air from the baggage compartment area through the evaporator, but i tsoutput is furnished directly to four smaller ducts concealed in the cabin overhead panel-ing. Two of these ducts run direc tly to the two

louvered overhead outlets in the cockpit(Figure 11-12). On SNs 35-092 and 36-025 and subsequent, two additional ducts—one on eachside—are connected to the individual overhead WEMAC outlets above each of the passenger s e a t s (F i g u re 1 1 -1 3 ) . A i r v o l u m e a n d directional control is provided at each outlet.The fan motor is cooled by the air it movesthrough the ducting.

counterclockwise direction

• On SNs 35-113 and subsequent and 36-032 and subsequent, the OFF detent isat the full counterclockwise position;speed is increased by rotating the rheo-

stat in the clockwise direction

If all the cockpit and overhead outlets areclosed, the cockpit fan must not be operated because no cooling airf low for the motor isavailable; the motor will overheat.




Figure 11-11. COCKPIT AIR and CABIN BLOWER Rheostats

AUXILIARY COOLING SYSTEM

A Freon refrigeration system—an auxiliarycooler—is installed to provide supplementalcooling for ground and inflight operations; it

can also be used for dehumidification.

System components, identified schematically(Figure 11-14), are conventional. The com-

Cool air is drawn through the evaporator and circulated as already described in CabinBlower Distribution, except that the blower motor runs continuously at its maximum speed;the CABIN BLOWER rheostat, if installed, is

inoperative. The compressor motor is pow-ered from the battery charging bus through a150 A current limiter and a control relay pow-ered from the FREON CONT circuit breaker

LEARJET 30 SERIES PILOT TRAINING MANUAL1 1




( g ) pressor (belt-driven by a 3.75 hor sepowe r motor), the condenser, and the dehydrator arelocated inside the tail cone. The compressor motor is cooled by air from the tail cone ventilation airscoop on the left side of the

fuselage. The evaporator and expansion valveare located inside the evaporator and blower assembly above the baggage compar tment.

Operation

Electrical power for system operation must be suppli ed by ei ther a GPU or an engine-driven generator. The system is turned on by

selecting the COOL position on the FAN– OFF–COOL switch. DC power is applied

on the left main bus.

The diverter doors may be positioned as desired to control airflow into the cabin through the louvered grille above the divan seat. If desired,

the cockpit fan may also be used to providewider circulation of the cooled air to the cockpitand passenger WEMAC outlets.

When the cooling system is being powered bya GPU, it is possible in some conditions for the aircraft batteries to be depleted if GPUfailure occurs.

The compressor motor i s au tomat ica l ly deenergized when the START–GEN switch is

l d START H l i




AUXILIARY HEAT SYSTEMS(OPTIONAL)

Two optional electric auxilia ry heat systemsare available: one for the cabin and one for the

cockpit. Both systems may be used to provideadditional heating on the ground or in flight.

A ili C bi H t S t m

on the Freon air conditioning system turnsoff the auxiliary cabin heat system.

Operation

On SNs 35-002 to 35-670 and 36-002 to 36-063, the auxiliary cabin heat system is con-trolled by a three-position (LO–OFF–HI)AUX HT switch on the copilot lower rights itch panel Selecting LO po ers the cabin





Auxiliary Cabin Heat System

General

The auxiliary cabin heat system uses the cabin b lower to c i r cu la t e h ea te d a i r. I t a l s o

incorporates two, dual coil heating elements,one located in each of the cabin blower ducts(Figure 11-15). Each heating element containsa thermoswitch set for high and low limits (150°F and 125°F) and a thermal fuse for over-heat protection.

On SNs 35-002 to 35-646 and 36-002 to 36-063, if the manual diverter doors are open

(i.e., air being diverted into the baggage compar tment), the cabin heat system is inoperative.

switch panel. Selecting LO powers the cabin blower and one heating coi l on each element;the HI position powers the cabin blower and all four coils.

On SNs 35-671 and subsequent and 36-064and subsequent, the cabin auxiliary heat sys-t e m i s c o n t r o l l e d b y a t h r e e - p o s i t i o n(OFF–CREW–CAB & CREW) AUX HTswitch on the copilot lower right switch panel. The CREW position energizes thecrew auxiliary heater (explained later in this section). Selecting CAB & CREW energizesthe cabin blower and all four auxiliary cabin

heating coils.

Initially the cabin blower runs at one tenth its




Auxiliary Cockpit Heat System(SNs 35-643 and Subsequentand 36-064 and Subsequent)

General

The auxiliary cockpit heat system providesadditional heat for crew comfort and interior windshield defogging. It includes an electrich t i th f d d f th i ht bi

On SNs 35-643 to 35-670, the auxiliary cockpit heating system is controlled by a three- po-sition (OFF–CKPT–W/S AUX DEFOG HEAT)swi tch on the ANTI–ICE cont ro l panel .Selecting either CKPT or W/S AUX DEFOG

HEAT will power the heater element (seeChapter 10, Ice and Rain Protection, for addit ional information on the W/S AUXDEFOG HEAT function).





heater in the forward end of the right cabin bleed air duct , where it connects to the cockpi t air distribution duct ing; it uses condit ion bleed airflow to circulate heated air (see Figure11-15).

Operation

The heating element for the auxiliary cockpitheat system requires bleed air flow for cool-ing. Because of this, on SNs 35-671 and sub-sequent and 36-064 and subsequent , theCABIN AIR switch must be ON, at least oneengine must be running and its bleed air shut

off and regulator valve must be open beforeelectrical power can be applied to the heatingelement If only the left engine is running the

On SNs 35-671 and subsequent and 36-064 and subsequent, the auxiliary cockpit heating sys-t e m i s c o n t ro l l e d b y a t h r e e -p o s i t i o n(OFF–CREW–CAB & CREW) AUX HT

switch, located on the copilot lower, rightswitch panel. Selecting either CREW or CAB& CREW powers the heater element as longas the CABIN AIR switch is ON and the other conditions described above are met.

With the heater e lement powered, all the air coming through the bleed air outlets in thecockpit are heated. A thermoswitch between

the windshield defog diffusers and the center footwarmer monitors the temperature of thei fl Th h i h l l i l

1. The manual diverter doors must be fullyclosed:

A. To operate the cockpit fan

B. To operate the Freon system

C. To opera te the aux il i a ry heat ing

6. The Freon system should not be usedabove:

A. 5,000 f tB. 8,000 ft

C. 18,000 f t

D 35 000 ft


QUESTIONS



p y gsystem

D. The aircraft does not have manual di-verter doors

2.Equipment that can be operated with air-craft battery power only is:

A. The auxiliary defog system

B. The Freon air conditioning system

C. The cabin blower and cockpit fan

D. The auxiliary heating system

3. When the aircraft is unpressurized on the

ground, air circulation is provided by:A. Ram a ir

D. 35,000 f t

7. The F reon sys t em au tomat i ca l lydisengages:

A. During engine start

B. Upon touchdown

C. When unpressurized

D. If the main door is opened

8. When the Freon system is operating, itcools:

A. Ram air

B. Cabin a ir

C. Outside air

D Bl d i

11. If DC power fails, the flow control valve:

A. Closes completely

B. Modulates from open to closed

C. Remains open

D. Is bypassed

12. The temperature control indicator shows:

A. Cabin air temperature

B. Cockpit air temperature

C. The temperature of the bleed air inthe plenum chamber

D. The position of the H-valve





CHAPTER 12

PRESSURIZATION

CONTENTS

Page

2




INTRODUCTION................................................................................................................. 12-1

GENERAL ............................................................................................................................ 12-1

MAJOR COMPONENTS ..................................................................................................... 12-2

Cabin Outflow Valve ..................................................................................................... 12-2

Vacuum Jet Pump and Regulator Assembly .................................................................. 12-2

Pressurization Control Module...................................................................................... 12-2

Cabin Safety Valve ........................................................................................................ 12-6

CABIN AIR Switch....................................................................................................... 12-6

ILLUSTRATIONS

Figure Title Page

12-1 Pressurization System Control............................................................................... 12-3

12-2 Pressurization Control Module .............................................................................. 12-2

12-3 HORN SILENCE and Test Control ....................................................................... 12-5




12-4 CABIN ALT and DIFF PRESS Indicator.............................................................. 12-7

12-5 Engine Bleed Air Conditioning System(SNs 35-002 to 35-086 [except 35-082] and 36-002 to 36-022) ......................... 12-10

12-6 Engine Bleed Air Conditioning System(SNs 35-082, 35-087 to 35-112, and 36-002 to 36-022) ..................................... 12-11

12-7 Engine Bleed Air Conditioning System(SNs 35-113 and Subs. and 36-032 and Subs.) ................................................... 12-12

12-8 Emergency Pressurization Override Switches..................................................... 12-15

TABLES

CHAPTER 12PRESSURIZATION




During climbs and descents the controller regulates the outflow discharge rate. This ratecontrol is necessary to maintain a cabin changerate that is comfortable regardless of aircraftrate of climb or descent.

Chapters 9 and 11—Pneumatics and Air Conditioning, respectively—describe how thecabin and cockpit are pressurized, heated, and cooled. This chapter deals primarily with how

PRESSURIZATIONCONTROL MODULE

General

The pressurization control module is located onthe copilot lower instrument panel. The controlson the front of the module are located on whatis referred to as the pressurization control panel.Figure 12-2 illustrates a typical aircraft pres-


1 2



cooled. This chapter deals primarily with howthe pressure is regulated.

MAJOR COMPONENTSThe pressurization control system (Figure 12-1)incorporates the following major components:

• Cabin outflow valve

• Vacuum jet pump and regulator assembly

• Pressurization control module

• Cabin safe ty valve

Figure 12 2 illustrates a typical aircraft pressurization control module configuration.

P R E S S U R I Z A T

I O N

CURRENT

SOL VALVE (NC)

ENERGIZED OPEN

ON GND

SOL VALVE (NO)

ENERGIZED CLOSED

IN MANUAL, ON GROUND

OR ABOVE 8 750 ±250 FT

SOL VALVE (NO)

ENERGIZED CLOSED

ON GND

CONTROL PRESS

(VACUUM) SOURC

CAB

LIMI11,5

±1, 5

CABIN

PRESS

STATIC PORT

UP

DN

SOL VALVE (NC)

ENERGIZED OPEN

ON GND WITH

CAB AIR OFF7

CAB ALT LIM5

11,500 FT ±1,500 FT6



F OR T RAI NI N G P URP O SE S

ONL Y

OR ABOVE 8,750 ±250 FT

CABIN ALT LIGHT4

AUTO

MAN

AUTO

FILTER

NCR

UP

DN

CABIN

AIR

ON

OFF

UP

RATE

CURRENT AIRCRAFTSOLENOID VALVES

IN FLIGHT AUTO PO

NOTE:

STATICFIL

PRESS DIFF

RELIEF 9.7 PSID3

PRESS DIFF

RELIEF 9.4 PSI2

OUTFLOW

VALVE

9.2 PSID1

STATIC

PRESS

DN

ALTERNATE

STATIC

PORT

NOSE CABIN

CAB

PRE

C O N T RO L L E R C AB I N

C A B

I N

C A B

I N

3

2

1 0

3 0

2 5 X 1

0 0 0

Cabin Controller

In AUTO mode, the cabin controller regulatescabin pressure in relation to the altitude thatis set on the altitude selector knob. Rotating theknob on the face of the cabin controller either turns a dial or aligns a window to indicate twoscales with a f ixed index between them. Theouter scale represents cabin altitude, and theinner scale represents aircraft altitude.

UP–DN lever is released to neutral, the cabincontroller will return the cabin pressure to theoriginal value.

Differential PressureRelief Valve (Primary)

The primary differential pressure relief valvefunctions in association with the CABINCONTROLLER Its purpose is to relieve


1 2

P



For the current ECS, the cabin controller is capable of maintaining the cabin pressure at sealevel with aircraft altitudes up to approximately24,000 ft. If the aircraf t is flown to an altitudehigher than 24,000 ft, the cabin altitude mustincrease in order to maintain the same pressure differential. At an altitude of 45,000 ft, thecabin altitude will normally be approximately6,700 ft.

Rate Control

A RATE knob is installed to the lower left of the CABIN CONTROLLER to control the ratehi h h bi li b d d d Th

CONTROLLER. Its purpose is to relieve excessive control pressure to the outflowvalve in the event that cabin pressure should exceed the normal limit when operating inAUTO mode.

NOTE

The primary differential pressure relief valve does not function inMAN control.

On SNs 35-002 to 35-112 and 36-002 through

36-031, the relief valve is set for 8.9 psid.

PR E S S U R I Z A T I O N

sure into the control line, thereby restrictingcabin altitude to the listed level.

Controller Solenoid Valves

Three solenoid-operated valves installed inthe controller are used to control the routingof pneumatic control pressure to the outflowvalve. All three valves are energized on theground by the squat switch re lay box, which

The outflow valve—now isolated—holds itslast atta ined posit ion. When cabin altitude de-creases to 8,000 ft or below, the aneroid rese tsand deenergizes the solenoid valve open, pro-vided the AUTO MAN switch is in AUTO.

On SNs 35-113 and subsequent and 36-032 and subsequent, the description of operation isthe same as early SNs, except that the aneroid switch actuates at 8,750 ±250 ft, resets at




ground by the squat switch re lay box, whichcauses the outflow valve to open, thereby de- pressurizing the cabin.

One of the valves is used in flight to effect

manual control of the outflow valve, and is referred to as the “manual-mode solenoid valve”(see Flight Operation-Manual).

For normal automatic inflight operation, allthree solenoid valves are deenergized.

On the early SNs, valve actuation requires DC power from the AIR BLEED circuit breaker on

the left essential bus. Later SNs require DC power from the CAB PRESS circui t breaker

7,200 ft, and—when actuated—causes theamber CAB ALT annunciator to illuminate (seeAnnunciator Panel section). When the aneroid resets, the annunciator extinguishes.

Should the above cabin altitudes be reached or exceeded, the cherry picker is the only wayto control the outf low valve.

Cabin Altitude Warning Horn Aneroid Switch

Early SNs use the manual pressurization

aneroid just described, while later SNs use aseparate 10,100-foot cabin aneroid to sound bi ltit d i h A i l d d

approximately 30 seconds after being silenced with the HORN SILENCE switch.

The horn will continue to reactivate after eachuse of the HORN SILENCE switch until theaneroid resets at a cabin altitude of 8,000 ft(early SNs) or 8,590 ft (later SNs).

The rotary system TEST switch on the center switch panel (Figure 12-3) is used to test thecabin altitude warning horn Rotating the

10 seconds after the CABIN AIR switch isturned to ON to close the safety valve. Thesolenoid is deenergized in flight regardlessof CABIN AIR switch position.

On earlier SNs, the safety valve does not openon the ground.

Differential PressureR li f V l (S d )


1 2

P



cabin altitude warning horn. Rotating theswitch to CAB ALT and depressing the TEST button provides a ground, which simulates al-titude warning horn aneroid switch actuation.This test does not illuminate the CAB ALT

light (if installed). During the test sequence,HORN SILENCE switch operation should also be checked.

CABIN SAFETY VALVE

General

A pneumatically operated cabin safety valveis installed in the aft pressure bulkhead. Itspurpose is to re lieve a cabin overpressure or

Relief Valve (Secondary)

The secondary pressure relief valve functionsin association with the safety valve. Should the prim ar y pressur e rel ie f va lve no t func ti on

properly, the secondary pressure re lief valveforces the safety valve open to limit cabin pressure. The safety valve will relieve pressureat 9.2 psid on SNs 35-002 to 35-112 and 36-002 to 36-031. On SNs 35-113 and subsequentand 36-032 and subsequent, the pressure isrelieved at 9.7 psid.

Cabin Altitude Limiter(Secondary)


is on the ground. Early aircraft have a MAX posi tion that opens the valve to fu ll flow for smoke and fume elimination. Current aircrafthave no position equivalent to MAX; increased airflow is achieved by positioning the BLEED

AIR switches to EMER. The CABIN AIR switch uses DC power from the AIR BLEEDcircuit breaker on the left essential bus.

INDICATORS

The cabin differential pressure is indicated by a circular scale on the inner por tion of theindicator and a single pointer. The scale represents differential pressure from 0–10 psiand is divided into three bands:

• On early aircraft

° A green band from 0–8.9 psi

° A yellow band from 8.9–9.2 psi




INDICATORS

CABIN ALT and DIFFPRESS Indication

Cabin altitude and differential pressure are indicated on a single indicator incorporatingtwo scales and two pointers (Figure 12-4).

Cabin altitude is indicated by a single pointer and a circular scale on the outer edge withCABIN ALT markings from 0 to 50,000 ft.

° A red band above 9.2 psi

• On current aircraft

° A green band from 0–9.4 psi

° A yellow band from 9.4–9.7 psi

° A red band above 9.7 psi

Cabin altitude should always be equal to or lessthan the aircraft altitude; therefore, cabin pres-sure should always be equal to or greater thanatmospheric pressure at the aircraft altitude.The indicator should normally read approxi-

mately 0.2 psi below the yellow arc.

Cabin Vertical Speed Indicator

The cabin vertical speed indicator (see Figure

to maintain a comfortable cabin altitude climbrate of approximately 600 fpm. As the air-craft climbs to cruise altitude, the cabin con-troller automatica lly adjusts the outflow valveto give the desired cabin climb rate until the

cabin altitude reaches the altitude set on thecabin controller dial. As the aircraft contin-ues its climb, the differential pressure increaseswhile the cabin altitude remains constant untilthe aircraft ar rives at the selected ACFT alti-t d If it i b d th t th DIFF PRESS i


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P



12-4) is to the right of the cabin altimeter. It provides an indication of cabin climb or de-scent rates of between 0 and 6,000 fpm.

NORMAL SYSTEMOPERATION

BEFORE TAKEOFF

During ground operation, the CABIN AIR

switch is normally not turned on until just prior to take off unle ss engine bl eed air isd i d f bi h ti

tude. If it is observed that the DIFF PRESS in-dicator is riding on the yellow/red line, aslightly higher cabin altitude should be se-lected. Adjust the cabin controller as neces-

sary when changing cruise altitude.Moni tor cabin pressure and d i fferent ia l pressure throughout the f light.

FLIGHT OPERATION—MANUAL

If the cabin controller is not functioning properly, fol low the Manual Mode Operation

procedures in Sect ion 2 of the approved AFM

.

M l d ti i t bli h d h th


the CABIN CONTROLLER dial. The aircraftrate of descent should be controlled so that thedescent rate is comfortable (approximately600 fpm).

LANDINGAs the aircraft descends and reaches the preselec ted cabin al ti tude, the outf low valvemodulates toward the open position. The cabin

On SNs 35A-082, 35A-087 to 35A-112; 36A-023 to 36A-031; and earlier aircraft incorpo-rating AMK 76-7, the flow control valve islocated downstream of the heat exchanger.Engine bleed air is available to the heat ex-changer whenever an engine is operating and the BLEED AIR switches are on. Because of this, a pressure switch is installed in the tailcone ducting prior to the heat exchanger.Should this pressure switch actuate (whichocc rs at appro imatel 47 psi) both red




should be unpressurized at landing.

At touchdown, the squat switch relay box actuates the three pneumatic solenoid valves in

the controller; this causes the outflow valve toopen completely to ensure cabin depressur-ization. In addition, when the CABIN AIR switch is placed to OFF, the flow control valvecloses, and—on SNs 35-099 and subsequent and 36-029 and subsequent—an additional solenoid valve is energized open, which causes the safetyvalve to open.

EMERGENCY

occurs at approximately 47 psi), both red BLEED AIR L and R annunciator lights illu-minate to indicate the overpressure condition.

To deactivate emergency pressurization, selectMAN and toggle the spring-loaded WSHLDHT switch to OFF until the valve is closed.

SNS 35-113 AND SUBSEQUENT AND 36-032 AND SUBSEQUENT

Emergency pressuriza tion is accomplished byrouting bleed air directly into the cabin from

either (or both) engine(s) through the emer-gency pressurization valves This air com-


1 2

P R

CREW OUTLETS


FOOTWARMER OUTLET

LEGEND



RE S S U R I Z A T I

O N

LESSBUS

CABHT

AUTO





CABINTEMP

SENSOR



AIRBLEED

BLEED AIR

RAM AIR


L

MAXNORM

OFF



CREW OUTLETS


FOOTWARMEROUTLET

CABIN TEMP

BLEED AIR

LEGEND

RAM AIR

CONDITIONEDBLEED AIR



LESSBUS

CABHT

AUTO




CABIN TEMPSENSOR

BAGGAGE COMPARTMENTAIR DIFFUSER(35A AIRCRAFT ONLY)



PRESSURE SWITCH

AIRBLEED

LMAINBUS


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P R

BLEED AIR

LEGEND

RAM AIR


REGULATED SERVO AIR

TO SENSOR


CREW OUTLETS

FOOTWARMER OUTLET



RE S S U R I Z A T I

O N

LESSBUS

AIRBLEED


EMERGENCYPRESSURIZATIONVALVE

AUXDEFOG/CREWHEATER

*

BLOWER MOTOR

CABINTEMP

SENSOR


BAGGAGE COMPARTMENTAIR DIFFUSER

(35A AIRCRAFT ONLY)

CHECK VALVES

The emergency pressurization valves are alsocontrolled by two cabin aneroid switches (onefor each valve). The aneroids are set to operateat 9,500 ±250 ft cabin altitude. Should thecabin altitude reach 9,500 ±250 ft, the aneroid switches deenergize the solenoids on the emer-gency pressurization valves, and the valvesmove to the emergency position. The aneroidsreset when the cabin altitude decreases to approximately 8,300 ft ; however, the approved AFM requires that the cabin altitude be at or

On SNs 35-659 and subsequent; 36-064 and subsequent; and earlier aircraft modified byAMK 90-3, the emergency pressurizationvalves are powered by the L and R EMER PRESS circuit breakers on the left and rightmain DC buses. On these aircraf t, the bleed air shutoff and regulator valves are powered byseparate circuit breakers labeled L and R BLEED AIR, also located on the left and rightmain DC buses. With an EMER PRESS cir-cuit breaker open, the emergency pressuriza-




AFM requires that the cabin altitude be at, or below 7,200 f t before a ttempt ing to reset theemergency pressurization valves.

To reset the emergency pressurization valvesafter they have been positioned to emergency,the BLEED AIR switches—one at a time— must be positioned to OFF momentarily, then back to ON.

On SNs 35-113 to 35-658, and 36-032 to 36-063 not incorporating AMK 90-3, the emer-gency pressurization valves are powered by theL and R MOD VAL circuit breakers on theleft and right main DC buses. These circuitb k l id l t i l t th

cuit breaker open, the emergency pressurization valve positions to emergency and the bleed air shutoff and regulator valve remainsopen. In this case, positioning the BLEEDAIR switch to OFF will stop air flow into thecabin since DC electrical power, from theBLEED AIR circuit breaker, will be availableto close the bleed air shutoff and regulator valve.

See Chapter 9, Pneumatics, for additional information on the bleed air shutoff and regulator valves.

During the first engine start, the valves willi ll hif i i f


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P R

Table 12-1. AUTOMATIC PROTECTION AND WARNING FEATURES—SNs 35-002

TO 35-112 AND 36-002 TO 36-031

PROTECTION AND WARNINGCABIN ALTITUDE

10,000 ±250 ft

11,000 ±1,000 ft

14,000 ±750 ft

• Pressurization aneroid automatically switches the system to manual control.

• Cabin altitude warning horn sounds—initiate emergency descent.

• Cabin altitude limiters actuate.

• Passenger oxygen masks are deployed and cabin overhead lights are illuminated.



E S S U R I Z A T I O N

* The differential pressure relief for the outflow valve is 8.9 psid, and the differential pressure relief for the safety valve is 9.2 psid.

Table 12-2. AUTOMATIC PROTECTION AND WARNING FEATURES—SNs 35-113 AND SUBS. AND 36-032 AND SUBS.

PROTECTION AND WARNINGCABIN ALTITUDE

8,750 ±250 ft

9,500 ±250 ft

10,100 ±250 ft

• Pressurization aneroid automatically switches the system to manual control. • CABIN ALT caution light illuminates.

• Emergency pressurization valves are activated by aneroid switches, directingengine bleed air directly into the cabin.

• Cabin altitude warning horn sounds—initiate emergency descent.

EMERGENCY PRESSURIZATIONOVERRIDE SWITCHES

On SNs 35-605 and subsequent; 36-056 and

subsequent; and earlier SNs incorporatingAAK 84-4, two emergency pressurization over-ride switches (Figure 12-8) allow the crew to override the 9,500-foot cabin aneroids to facilitate landing at high elevation airports.

swi t ches to the OVERRIDE pos i t ion disconnects the 9,500-foot aneroids from thesystem. The switches can also be used:

• To reset an emergency valve that has inadvertently positioned to emergency

due to a malfunctioning aneroid • To reset the emergency valves in order

to restore windshield and stab/wing anti-icing (at any altitude)

I ith l ti OVERRIDE t b




The guarded switches are labeled L and R EMER PRESS and have positions labeled OVERRIDE and NORMAL. With the guards

down, the switches are in the NORMAL po-sition. Lifting the guards and moving the

In either case, selecting OVERRIDE must befo l lowed by cyc l ing the BLEED AIR switch(es) to OFF and then to ON, provided DC power is available and the MOD VAL (or

EM ER P R ES S , a s a p p l i c a b l e ) c i r c u i t breaker(s) are in.


1 2

P R E



E S S U R I Z A T I O N


1. To regulate cabin pressure, the cabin con-troller modulates the:

A. Cabin safety valveB. Flow control valve

C. Outflow valve

D. Primary differential pressure relief valve

5. To dump residual cabin pressure on touch-down:

A. The outflow valve opens automati-cally

B. The cabin safety valve opens auto-matically

C. The flow control valve closes auto-ti ll


QUESTIONS



2. Illumination of the amber CABIN ALTlight (if installed) indicates:

A. Cabin altitude is at or above 8,750±250 ft, and the pressurization controlsystem is in manual mode

B. Cabin altitude is at or above 8,750±250 ft, and the pressurization controlsystem may be in either AUTO or MAN mode

C. Cabin altitude is at or above 9,500±250 ft, and the emergency pressur-

ization mode has activated D Th CABIN AIR i t h i i th

matically

D. The bleed air shutoff and regulator valves close automatically

6. On aircraft without the emergency pres-surization valves, if DC power fails:

A. The windshield anti-ice/defog systemcan be used in the event of a pressur-ization failure

B. The cabin will remain pressurized, but emergency pressuri zation capa- bi li ty will be lost

C. The flow control valve fails closed D Th bl d i h t ff d l t

CHAPTER 13

HYDRAULIC POWER SYSTEMS

CONTENTSPage

INTRODUCTION................................................................................................................. 13-1




GENERAL ............................................................................................................................ 13-1

HYDRAULIC SYSTEM OPERATION ............................................................................... 13-3

HYDRAULIC SUBSYSTEMS............................................................................................. 13-4

QUESTIONS......................................................................................................................... 13-5

ILLUSTRATIONS

Figure Title Page

13-1 Controls and Indicators.......................................................................................... 13-2

13-2 Hydraulic System Schematic................................................................................. 13-3




CHAPTER 13HYDRAULIC POWER SYSTEMS




An accumulator precharged with dry air or nitrogen dampens pressure surges and helpsmaintain system pressure. A direct-readingindicator on the center instrument panel displays system pressure. An amber annun-ciator light warns of low pressure.

There are three filters in the system: one ineach pressure line and one in the return line.

A system relief valve set to open at 1,700 psipre ents s stem damage b porting e cessi e

reversers (if installed). A check valve preventsauxiliary pump actuation of the spoilers/spoilerons.

The reservoir and the accumulator are located in the tail cone. Reser voir fluid level should

be jus t above the s ight glass with zero system pressure . Fluid is low if the level can be seenin the glass or if fluid is not visible.

Accumulator precharge, indicated by the gageon the accumulator should be 750 psi when




prevents system damage by porting excessive pressure in to the re turn l ine.

Pressure from the engine-driven pumps is

available to actuate the spoilers/spoilerons,flaps, landing gear, brakes, and TR 4000 thrust

on the accumulator, should be 750 psi whenhydraulic pressure is zero.

Controls and indicators for the system are

shown-in Figure 13-1.

1 3

H Y D R A

U L I C

S Y

S T E M

LOWHYD

(STANDARD)

HYDRAULIC SYSTEMOPERATION

Unless there is residual hydraulic system pressure , the auxiliary hydraulic pump must

be opera ted to provide pressure for setting the parking brakes pr ior to engine star t. Placingthe HYD PUMP switch in the on (HYD PUMP) posi tion star ts the auxil iary pump, assuming both engines a re shut dow n and systempressure is below 1 125 psi (Figure 13 2)

As pressure increases, a pressure switch actuates at 1,250 psi to extinguish the amber LOW HYD light on the annunciator panel.(see Annunciator Panel section. ) At approxi-mately 1,250 psi, the pressure switch stopsthe auxiliary pump. The HYD PUMP switch

should then be positioned to OFF, where itnormally remains unless flap operation is re-quired prior to engine start. The LOW HYDlight will illuminate if pressure drops below1,125 psi.




pressure is be low 1,125 psi (Figure 13-2).If the HYD PUMP switch is left on, the pressure switch will cycle the pump between1,125 psi and 1,250 psi.

FILTER

ENGINE

BLEED

AIR

REGULATOR

PRESSURE

RELIEF VALVE

(20 PSI)

OVERBOARD

VACUUM

RELIEF

In the event of engine fire or when mainte-nance is to be performed, the DC motor-drivenshutoff valves can be closed by pulling theappropriate FIRE handle on the glareshield.Pull ing ei ther handle also arms the f i re-extinguisher system; therefore, if valve clos-

ing is to facilitate maintenance, the applicableFIRE EXT circuit breaker(s) should be pulled to prevent accidental discharge of the bottles;the shutoff valves are opened by pushing in theappropriate handle(s). The shutoff valves op-erate on DC power supplied through the L and

Loss of fluid due to a system leak is the most probable cause of complete loss of hydraulic pressure . If al l hydraulic system pressure islost, the LOW HYD light will illuminate as pressure decreases below 1,125 psi. Do not operate the auxiliary pump until alternate

landing gear extension procedures have beenaccomplished, as directed by the approved AFM . Otherwise, the auxiliary pump may dis-charge the 0.4 gallon of reserve fluid throughthe same leak.




erate on DC power supplied through the L and R FW SOV circuit breakers on the left and right essential buses, respectively.

A f t e r s t a r t i n g t h e f i r s t e n g i n e , t h e HYDRAULIC PRESSURE indicator should bechecked to verify engine-driven pump opera-tion. Pressure should stabilize at 1,550 ±25 psi,indicating that the engine-driven pump is operating properly.

When the second engine is started, there is nochange in pressure indication, but capacity is

doubled. There is no positive indication that thed i i l h f

There is no CB protection in the cockpit for the auxiliary pump; it is powered directly fromthe bat tery charging bus through a 50 A

current limiter.

HYDRAULICSUBSYSTEMS

Operation of hydraulic subsystems is pre-sented in Chapter 14, Landing Gear and

Brakes; Chapter 15, Flight Controls (flaps and il / il ) d Ch t 7 P l t

1 3

H Y D R A

U L I C

S Y

S T E M

1. Normal hydraulic system pressure withthe engine-driven pumps operating is:

A. 1,400 ±50 psiB. 1,550 ±25 psi

C. 1 ,650 psi

D. 1 ,700 psi

2 The hydraulic shutoff valves are closed:

6. The approved fluid for the hydraulic system is:

A. MIL-H-5606B. MIL-O-M-332

C. Skydrol

D. MIL-H-2380

7 The operational time limit of the auxiliary


QUESTIONS



2. The hydraulic shutoff valves are closed:

A. By pulling the engine FIRE handles

B. Automatically when the FIRE warn-

ing light comes onC. By the GEN switch in the OFF posi-

tion

D. By the BLEED AIR switches

3. In the event of hydraulic system pressurefailure in flight:

A. Immediately turn the HYD PUMP

switch onB Turn the HYD PUMP switch on when

7. The operational time limit of the auxiliary pump is :

A. 5 minutes on, 15 minutes off

B. 5 minutes on, 25 minutes off C. 3 minutes on, 20 minutes off

D. 2 minutes on, 30 minutes off

8. The auxiliary hydraulic pump will provideapproximately:

A. 1 ,200 psi

B. 1 ,550 psi

C. 1 ,700 psiD 1 250 i

CHAPTER 14

LANDING GEAR AND BRAKES

CONTENTSPage

INTRODUCTION................................................................................................................. 14-1

GENERAL 14 1




GENERAL ............................................................................................................................ 14-1

LANDING GEAR................................................................................................................. 14-2

Indicating System .......................................................................................................... 14-2

Main Gear Components................................................................................................. 14-4

Nose Gear Components................................................................................................. 14-4

Operation ....................................................................................................................... 14-6

BRAKES............................................................................................................................. 14-12

ILLUSTRATIONS

Figure Title Page

14-1 Gear Position Indicator Lights ............................................................................... 14-2

14-2 Gear Position Indications....................................................................................... 14-3

14-3 Main Gear .............................................................................................................. 14-4

14-4 Nose Gear .............................................................................................................. 14-5




14-5 Nose Gear Centering Cams.................................................................................... 14-5

14-6 Landing Gear Retracted ......................................................................................... 14-7

14-7 Landing Gear Extended ......................................................................................... 14-9

14-8 Emergency Air Pressure Indicator....................................................................... 14-10

14-9 Alternate Extension Controls .............................................................................. 14-10

14-10 Alternate Landing Gear Extension ...................................................................... 14-11

14-11 Brake System Schematic ..................................................................................... 14-13

CHAPTER 14LANDING GEAR

AND BRAKES




The antiskid system provides maximum deceleration without skidding the tires. Whent h e s y s t e m i s o p e r a t i n g , w h e e l s p e e d t ransducers—generators—furnish wheelspeed information to a control box that sig-nals the antiskid servo valves to modulate

braking pressure . The parking brake is se t by pull ing a handle on the th ro tt le quadrant and depressing the brake pedals; this traps hy-draulic pressure in the brake assemblies.

The variable authority, electric nosewheel steer -ing system operates only on the ground When

The nose gear red UNSAFE light is illuminated when the nose gear is in transit (i.e., neither down-and-locked nor up-and-locked). Whenthe nose gear is locked in either the up or thedown position, the light extinguishes.

The two main gea r red UNSAFE l igh t s illuminate whenever the respective main gear door is unlocked. As each inboard door latchesu p d u r i n g e x t e n s i o n o r r e t r a c t i o n , t h e corresponding red light extinguishes.




ing system operates only on the ground. Whenthe system is engaged, a computer determinesthe amount of nosewheel deflection allow-

able—based on rudder pedal movement and taxi speed—and uses a DC electric motor todeflect the nosewheel accordingly. Maximumauthority is 45° either side of center at slowspeeds, which decreases as speed increases.

LANDING GEAR

INDICATING SYSTEM

Indications for gear down-and-locked, up-and-locked, and in-transit conditions areshown in Figure 14-2.

If the gear is extended with the alternate— the pneumatic—system, all three green lightsand the two main gear red lights illuminate; bo th main gear doors remain ful ly extended.

The position lights are tested by holding theTEST/MUTE switch on the LANDING GEAR panel in the TEST position. All six lights will il-luminate and the warning horn will sound. Thelights can be dimmed with the dimming rheostat(Figure 14-1) if the navigation lights are on; oth-

erwise, they will be at maximum intensity.

Circuitry related to the left and right maingear green position lights is common with thelanding/taxi light for that side. Conf irmationof main gear downlocking—after bulb test-

Landing Gear Warning System

The aural warning horn will sound and threered UNSAFE lights will illuminate when allof the following conditions are present:

• Landing gear is not down-and-locked

• Altitude is less than 14,500 ±500 ft

• Either thrust lever is retarded below approximately 55–60% N1.

• Airspeed is below 170 KIAS (FC 530 air-ft l )




ing—can be made by switching on the re-spective LDG LTS switch.

Nose gear green light circuitry is common withthe engine synchronizing system (if installed).Confirmation of nose gear downlocking (after bulb tes ting) is made by posit ioning the ENGSYNC switch on the pedestal to ENG SYNC(on) and observing that the amber ENG SYNClight on the annunciator panel illuminates.

craft only)

At altitudes above 14,500 ±500 ft, the horn willnot sound when the thrust levers are retarded,

and the UNSAFE lights may illuminate. Thehorn also sounds when the flaps are extended beyond 25° if the landing gear is not down-and-locked, regardless of thrust lever position or altitude.

Holding the TEST/MUTE switch in TEST illuminates all six position indicator lightsand sounds the horn. Momentarily positioningthe switch to MUTE silences the horn when

MAIN GEAR COMPONENTS

Each main gear consists of the following:

• Conventional air-hydraulic shock strut

• Dual wheels

• Scissors

• Squat switch

• Main gear actuator

• Inboard and outboard doors

The main gear is hydraulically held in the retracted position and enclosed by an outboard door and an inboard door. The outboard door is mechanically linked to, and travelswith, the gear. The inboard door is hydrauli-cally actuated, electrically sequenced by mi-

croswitches, and held retracted by hydraulic pressure and a spring-loaded, overcenter up-latch that is released by a hydraulic actuator.

Proper shock strut inflation is an importantconsideration. When the aircraft weight is on




• Inboard door actuator (Figure 14-3)

The main gear hydraulic actuator also serves asa side brace when the gear is extended. It fea-tures an integral downlock mechanism that can be unlocked only by hydraulic pressure on theretract side; therefore, no downlock pins are provided. Each main gear scissors link actuatesa squat switch.

the gear, the amount of strut extension will varywith the aircraft load. With a full fuel load and no passengers or baggage aboard, 3 to 3.5

inches of bright surface should be visible onthe lower portion of the main gear strut.

Main Gear Wheel and Tires

Each main gear wheel incorporates a fusible plug that prevents ti re blowout due to exces-sive heat resulting from hard braking. Tiresmust be changed when the tread has worn tothe base of any groove at any location or if the

Because the cams cannot center the wheel if it is swiveled 180° from the normal position,the nose gear should be checked on the exterior inspection to ascertain that the gear uplatchroller is facing forward.




The nose gear actuator incorporates an integraldownlock mechanism to maintain a positivedownlocked condition; therefore, a downlock pin is no t requi re d. As wi th the main gear

actuator, the locking mechanism can bereleased only by hydraulic pressure on the

Figure 14-4. Nose Gear

OPERATION

The landing gear system incorporates twosolenoid operated hydraulic control valves:one for operation of the main gear inboard doors and one for gear operation. Both in-

board doors must be fully open before the gear can be extended or retracted.

The door control valve is energized to the door-open position when the landing gear selector switch is placed in either the UP or DN posi-ti It i i d t th d l iti

Normal Retraction

Positioning the landing gear selector switch toUP energizes the door control valve to the open posi tion; this directs pressure to release themain gear inboard door uplatches and to open

the doors. The two red main gear UNSAFElights illuminate simultaneously with uplatchrelease.

When both inboard doors are fully open, thedoor-open switches are actuated. When bothdoor-open switches are actuated and both squat




tion. It is energized to the door-close position by main gear operated switches when both gear are fully retracted or down-and-locked.

The gear control valve is energized to the extend or retract position by switches sensing the fullopen position of both main gear inboard doors.During retraction, the circuit is routed throughboth squat switches to ensure that the aircraftis off the ground before the valve can be ener-gized to the retract position.

Normal landing gear operat ion requires DCl i d h h h GEAR i i

door-open switches are actuated and both squatswitches are in the airborne position, the gear control valve energizes to the retract position;

hydraulic pressure is directed to retract thelanding gear (Figure 14-6). The three greenLOCKED DN lights extinguish, and the red nose gear UNSAFE light illuminates.

When the nose gear fully retracts, the red nosegear UNSAFE light extinguishes. When bothmain gear fully retract, two gear-up trunnionswitches actuate to energize the door control

valve to the closed position. Pressure closes thegear inboard doors which lock in position by

F OR T

SO

RETRACTEXTEND

SOL

TO

EMERGENCY

BRAKES

OVERBOARD

GEAR ALTERNATE EXTENSION CONTROL VALVE



TRAI NI N G P URP O SE S

ONL Y

SO L SO

EMERAIR

BOTTLE

DOOR

ACTUATOR

TO

BRAKE

SYSTEM

PRIORITY

VALVE

UPLATCH

UPLATCH

ACTUATOR

NOSE

GEARACTUATOR

Normal Extension

Positioning the landing gear selector switch toDN energizes the door control valve to the open posi tion; this directs pressure to release themain gear inboard door uplatches and to open

the doors. The two red main gear UNSAFElights illuminate simultaneously with uplatchrelease.

When both inboard doors are fully open, thedoor-open switches actuate to energize thegear control valve to the down position; this

When the gear is fully down-and-locked, thethree green LOCKED DN lights illuminate and the red nose gear UNSAFE light extinguishes.Circuitry is completed by both main gear downlock switches to energize the door con-trol valve to the closed position. Pressure

closes the gear inboard doors (Figure 14-7),which lock in position by spring tension on thedoor uplatches; the two red main gear UN-SAFE lights extinguish.




gear control valve to the down position; this directs pressure to release the nose gear up-latch and extend the nose and main gear (Figure

14-7). The red nose gear UNSAFE light illu-minates.

F OR T

OVERBOARD


S

RETRACTEXTEND

SOL

TO

EMERBRAKES

TO

BRAKE

SYSTEM



TRAI NI N G P URP O SE S

ONL Y

PRIORITY

UPLATCH

VALVE

SO L S

EMER

AIRBOTTLE

UPLATCH

ACTUATOR

NOSE

GEAR

ACTUATOR

MAIN

GEAR

ACTUATOR

UPLATCH

DOOR

ACTUATOR

UPLATCH

Alternate Extension

The a l t e rna te gea r ex tens ion sys t em i s pneumatically operated by a bottle charged to1,800–3,000 psi with dry air or nitrogen. Bottle pressure i s show n on the di rec t- reading

EMERGENCY AIR indicator on the center in-strument panel (Figure 14-8). The bottle also provides pressure for emergency braking.

Operation

Pushing the emergency gear lever down opens avalve to release air bottle pressure to position thegear control and door control valves to the ex-tend position (Figure 14-10). This provides areturn flow path for fluid in the retract side of the gear and door actuators. The air pressurealso repositions the shuttle valves to accomplish the following:

• Release the nose gear uplatch and themain gear door uplatches




• Open the main gear inboard doors

• Extend all three gear

Since no provision is made to close the maininboard doors, the two main gear red UNSAFElights remain illuminated. The three greenLOCKED DN lights illuminate.

In a hydraulic failure situation, after the gear is down-and-locked, air pressure must be bled from the gear system by lifting the release tab

(see Figure 14-9) and raising the emergencygear lever to the normal position This closes0

180 300

500

F OR T

OVERBOARD


SO L

RETRACTEXTEND

SOL

TO

EMERGENCY

BRAKE VALVE

TO



RAI NI N G P URP O SE S

ONL Y

UPLATCH

SO L SO L

TO

BRAKE

SYSTEM

EMER

AIR

BOTTLE

PRIORITY

VALVE

UPLATCH

ACTUATOR

NOSE

GEAR

ACTUATOR

UPLATCH

ACTUATOR

UPLATCH

UPLATCH


If alternate extension is required due to anelectrical fault, the emergency gear lever mustremain in the down position to prevent subsequent inadvertent retraction of the gear.

BRAKES

The brake system (Figure 14-11) is powered by hy draulic pressure from the nose gear down—extend—line. The brakes can be ap-plied by either pilot The system has four mul-

assemblies). The f irst set of shuttle valves de-termines whether the pilot or copilot has con-trol of the brakes (i .e . , highest pressure predominating).

P is tons in each brake assembly move a

pres su re plat e, which forces the st at ionaryand rotating discs together against a backing plate to produce braking action . Depressingo n e p e d a l a p p l i e s b o t h b ra k e s o n t h e corresponding main gear; therefore, differ-ential braking is available, if required.



plied by either pilot. The system has four multidisc, self-adjusting brake assemblies—onefor each main gear wheel—operated by power

brake valves linked to the top sect ion of therudder pedals. The left pedals control both brake assemblies on the left gear ; the right peda ls cont ro l the brake as sembli es on theright gear. Braking force is in direct propor-tion to pedal application unless modulated bythe antiskid system.

The antiskid system, monitored by the red an-

tiskid gen warning lights permits stopping

Releasing pedal pressure repositions the brakevalve; springs in the brake assembly force

fluid back through the brake valves to thereservoir, thereby releasing the brakes.

During gear retraction, a restrictor in the nosegear return line creates back pressure on the brakes that is suff icient to stop the wheels fromrotating prior to their entry into the wheel wells.

A priority valve, also in the nose gear downline,

ensures proper gear sequencing during retractionb i i h d li li d h


COPILOT

BRAKE VALVE

PILOT

BRAKE VALVE

TO

RESERVOIR

FROM NOSE

GEAR DOWN

LINE

PILOT

BRAKE VALVECOPILOT

BRAKE VALVE



OVERBOARD

PARKING

VALVES

GEAR

ALTERNATE

EXTENSION

CONTROL VALVE

BRAKE

AIR BOTTLE

EMERG

ANTISKID

DISCONNECT

SWITCH

WARN

LIGHT

CB

ON

ANTI

SKID

PARK

BRAKE


ANTISKID

One of two antiskid systems may be installed.The early system was standard on SNs 35-002to 35-066 and 36-002 to 36-017. The later sys-tem is standard on SNs 35-067 and subsequent

and 36-018 and subsequent; it may also beretrofitted to early aircraft by AAK 76-4. Thetwo systems are similar and are discussed to-gether with the differences being noted.

The antiskid system limits braking on eachmain gear wheel independently to allow max-

Operation

The fol lowing condit ions must exist for operation of the antiskid system:

• The ANTISKID switch must be on

• Both squat swi tches must be in theground mode (left for outboard, rightfor inboard)

• The parking brake must be released

• Taxi speed must be above 8 to 10 kt(wheel speed, 150 rpm)



g p yimum braking under all runway conditionswithout tire skidding.

The system consists of the following:

• Four wheel speed transducers (one oneach main wheel)

• Two antiskid control valves

• Control box

• Monitor l ights

• Lever locking ANTISKID switch on the

At high speed, with the ANTISKID switch onand brakes applied, the control box receives and analyzes wheel speed inputs from the transducer on each main wheel (see Figure 14-11). If anywheel deceleration rate reaches a predeter-mined limit, the applicable servo valve willmodulate braking force on the corresponding brake by diverting pressure into the return line.

With the gear extended in flight, the brakingsystem is disabled. When the main gear squatswitches go airborne all braking pressure is

Four red ANTI-SKID GEN lights monitor cir-cuitry from each wheel speed transducer and individually illuminate if a fault is detected.Cycling the ANTISKID switch to OFF then back to ON may clear the faul t. All four lightsilluminate if power to the control box is lost

or if the ANTISKID switch is off.

EMERGENCY BRAKES

Pneumatic emergency brakes are provided for use in the event of normal brake system fail-ure. Antiskid protection, differential braking,

to the stop. If the PARKING BRAKE handleis not pushed in to the stop, the parking brakesmay be released, but the antiskid disconnectswitch may not actuate to enable the antiskid system. The ANTI-SKID GEN lights will notilluminate, and subsequent heavy braking will

result in wheel skids.

On SNs 35-626, 35-627, 35-630 and subsequent, 36-056, and 36-059 and subsequent,an additional PARK BRAKE light is just abovethe ANTI-SKID GEN lights. The PARK BRAKE light illuminates if the parking brake




ure. Antiskid protection, differential braking,and parking brakes are not available whileusing the emergency brakes.

To apply brakes with the emergency system,the EMER BRAKE handle must be pulled outof its recess (see Figure 14-11) and pressed downward. This meters pressure from theemergency air bottle through four shuttlevalves to the brake assemblies in proportionto handle movement. Releasing the handlestops flow from the bottle and allows applied

air pressure to vent overboard which releases

handle is not in the completely forward–re-leased—position.

NOSEWHEEL STEERING

The electrically actuated nosewheel steeringsystem has variable authority, as determined bysignals from the left inboard and both rightwheel speed transducers. System components

also include a rudder pedal fol low-up a


28 VDC STEER LOCKBUTTON

CONTROL WHEEL

MASTER SWITCHES

STEER ON

NOSE GEAR

UPLOCK SWITCH

(RELEASED)



C

115 VAC

RUDDER PEDAL

FOLLOW-UP

LEFT INBOARD

Steering authority varies from a maximum of 45° either side of center at speeds below 10 ktand decreases as groundspeed increases. Atthe maximum steering speed of 45 kt, author-ity falls to approximately 8°.

OPERATIONWith the squat switches in ground mode,nosewhee l s t ee r ing can be engaged by momentarily depressing the STEER LOCK switch or by depressing and holding the con-trol wheel master switch (MSW) on either con-

When steering engages, the green STEER ONannunciator illuminates. A rudder pedal fol-lowup p rov ides the d i sp lacemen t and direc tional signals modified by the computer- a m p l i f i e r i n p u t f ro m t h e w h e e l s p e e d transducers. The computer-amplifier drives

the s teer ing ac tuator in the appropria te direction until it is stopped by a signal froma follow-up located in the drive gearbox.

If the nosewheel steering system is inopera-tive, differential power and braking can beused to taxi the aircraft.




( )trol wheel (Figure 14-13). STEER LOCK isdisengaged by momentarily depressing ei-

ther control wheel master switch.

Since variable authority steering is depen-

dent upon wheel speed transducer signals,steering should not be used above 10 kt if anytwo of the following three ANTI-SKID GENlights are illuminated: two inboard and rightoutboard.





l Emergency air pressure can be used for:

A. Gear extension and parking brake

B. Gear, flaps, spoilers, and brakes

C. Gear extension and brakes

D. Gear extension, flaps, and brakes

2. Prior to takeoff, the EMERGENCY AIR pressure indicators should indicate:

A. 1,800 to 3,000 psi

6. Three gear UNSAFE lights will be on and the gear warning horn sounds when the:

A. G e a r i s re t r a ct e d a nd n o g r e en

LOCKED DN lights are onB. Gear is down, thrust levers are above

approximately 70% N1, and altitudeis below 14,500 ±500 ft

C. Gear is up, thrust levers are below approximately 55–60% N1, altitude isbelow 14 500 ±500 ft and on FC 530


QUESTIONS



, , p

B. Minimum 1,700 psi

C. 3,000 to 3,350 psi

D. Maximum 1,750 psi

3. During normal gear operation, main gear inboard doors and the main gear are se-quenced by:

A. Microswitches

B. Emergency air pressure

C Mechanical linkageD B th A d B

below 14,500 ±500 ft and, on FC 530aircraft, airspeed is below 170 KIAS

D. Flaps are extended below 25°, re-gardless of altitude

7. With the flaps extended beyond 25° and the gear not down-and-locked, the warn-ing horn:

A. Will sound, but can be muted

B. Will not sound

C Will sound but cannot be mutedD N f th b


10. Parking brakes can be set with the:

A. Pilot brake pedals only

B. Copilot brake pedals only when theANTISKID switch is on

C. Pilot or copilot brake pedals

D. Pilot or copilot brake pedals only withthe ANTISKID switch off

11. If the fir st three ANTI-SKID GEN lightsare illuminated:

A. Takeoff weight is limited to 17,000 lb

B. Nosewheel steering should not be en-

14. If the green main gear LOCKED DN lightis burned out, positive down-and-locked condition can be confirmed by:

A. GND IDLE light illuminated

B. ENG SYNC light illuminated

C. Illumination of the corresponding land-ing light when the switch is turned on

D. Red UNSAFE lights illuminate

15. The electrical requirements for nosewheelsteering are:

A. 24 VAC and 28 VDC



ggaged above 10 kts

C. Takeoff (VR

) will be affected

D. Both A and B are correct

12. Normal brake pressure is provided by:

A. Main hydraulic system pressure fromthe nose gear down line

B. Brake accumulator

C. Emergency air bottle through the an-tiskid control valves

D E i b ttl

B. Only 28 VDC

C. Only 115 VAC

D. 28 VDC and 115 VAC

16. When STEER LOCK is engaged:

A. Nosewheel steering is engaged and full steering is available up to 45 kt

B. The nosewheel is locked in whatever posi tion it is in at the t ime

C. Up to 45° left or right steering is avail-able with decreasing authority at

CHAPTER 15

FLIGHT CONTROLS

CONTENTS

Page

INTRODUCTION................................................................................................................. 15-1

GENERAL ............................................................................................................................ 15-1

15 3




PRIMARY FLIGHT CONTROLS........................................................................................ 15-3

Elevators ........................................................................................................................ 15-3

Ailerons ......................................................................................................................... 15-3

Rudder ........................................................................................................................... 15-4

TRIM SYSTEMS.................................................................................................................. 15-4

General .......................................................................................................................... 15-4

General ........................................................................................................................ 15-22

Operation..................................................................................................................... 15-24

MACH OVERSPEED WARNING/STICK PULLER ........................................................ 15-25

General ........................................................................................................................ 15-25Operation..................................................................................................................... 15-25

QUESTIONS ...................................................................................................................... 15-27




ILLUSTRATIONS

Figure Title Page

15-1 Flight Control Surfaces.......................................................................................... 15-2

15-2 Flight Controls Gust Lock ..................................................................................... 15-2

15-3 Aileron Tabs........................................................................................................... 15-4

15-4 Trim Systems Controls and Indicators................................................................... 15-5

15-5 Pitch Trim System Schematic (FC 200 AFCS) ..................................................... 15-8




15-6 Pitch Trim System Schematic (FC 530 AFCS) ..................................................... 15-9

15-7 Mach Trim System Schematic............................................................................. 15-11

15-8 Flap System ......................................................................................................... 15-13

15-9 Spoiler System..................................................................................................... 15-16

15-10 Spoiler Operation................................................................................................. 15-17

15 11 Spoileron Operation (Left Aileron Up) 15 18

CHAPTER 15FLIGHT CONTROLS




A dual stall warning system provides an in-dication of impending stall by vibrating thecontrol column and, if no corrective action istaken, induces a forward control column move-ment to reduce the aircraft angle of attack.

A Mach overspeed warning system warns of overspeed and induces an aft control columnmovement to raise the nose of the aircraft.

A Mach trim system provides automatic pitchtrim to compensate for Mach tuck.

All flight control surfaces appear in Figure

control surfaces. When installed, the lock holds full left rudder, full left aileron, and fulldown elevator displacement (Figure 15-2).




g pp g15-1.

A flight controls gust lock is provided to pre-vent wind gust damage to the primary flight

Figure 15-2. Flight Controls Gust Lock

PRIMARY FLIGHTCONTROLS

ELEVATORS

The elevators are hinged to the aft edge of thehorizontal stabilizer and are positioned byfore-and-aft movement of the control column.Three scuppers are located near the aft edgeof each elevator for moisture drainage; threestatic dischargers are attached to the trailingedge of each elevator.

On FC 200 AFCS aircraft, the electric clutchmust be engaged to couple the servo to the el-evator linkage. The clutch engages when anyone of the following switches is in ON:

• L STALL WARNING

• R STALL WARNING• AUTOPILOT master

With all three of the above switches in OFF,the electric clutch is disengaged, which dis-connects the servo from the elevators. This en-ables the pilot to gain manual control of theelevator by eliminating the servo in the event




The elevators can also be positioned by an

electrically actuated pitch servo.

A bob weight attached to the control columnand a downspring assembly in the elevator control linkage are incorporated to enhance pi tch s tabi li ty.

Pitch Servo

The pitch servo (torquer) is DC operated It is

elevator by eliminating the servo in the eventof a malfunction.

By exerting sufficient force on the controlcolumn to slip the mechanical clutch, the pilotcan also override any undesirable servo inputs to the e levators , i f necessary.

On FC 530 AFCS aircraft, the electric clutchremains deenergized until the servo is signalled by either the autopilot, L or R stall warning

system or overspeed puller system On these

Roll Servo(Autopilot Function Only)

The ailerons can also be positioned by the au-topilot roll servo. The roll servo is similar tothe pitch servo, but it does not incorporate anelectric clutch. A mechanical slip clutch allowsthe pilot to override undesired roll servo inputs;the servo can also be disconnected by disen-gaging the autopilot.

Balance Tab

The balance tab on each aileron (Figure 15-3)

Aileron Follow-ups

Aileron follow-up mechanisms, which aredriven by the aileron control linkage, provideai leron displacement information to thespo i l e ron compute r , yaw damper , and autopilot.

RUDDER

The rudder can be manually positioned with either set of rudder pedals, or by either of twoyaw damper servos: primary or secondary.The crew can manually override the yaw




( g ) provides aerodynamic assistance in moving

the aileron, thus reducing control wheel forces.

Trim Tab

The electrically operated aileron trim tab at-tached to the inboard trailing edge of the leftaileron (Figure 15-3) is positioned by either the pilot or copilot control wheel trim switch.Aileron trim tab position is indicated on the

cockpit center pedestal

damper through a mechanical slip clutch in the

event of a malfunction. The yaw damper can be disengaged by depressing either wheel mas-ter switch or the corresponding yaw damper OFF button.

Rudder Trim Tab

A trim tab mounted on the bottom trailingedge of the rudder is controlled by a trim

s itch on the center pedestal Trim position


ARMING

BUTTON

CONTROL WHEEL

TRIM SWITCH



PILOT CONTROL WHEEL

(COPILOT SIMILAR)

PITCH TRIM

WHEEL

MASTER

SWITCH

SECONDARY PITCH

TRIM SWITCH

RUDDER (YAW) TRIM

Control

Rudder (yaw) trim is controlled by the rudder trim switch on the center pedestal (see Figure15-4), which is spring-loaded to OFF.

The switch knob is split into an upper and alower hal f . Both halves must be ro ta ted simultaneously to initiate rudder trim tab motion. This is a safety feature to reduce the possibility of inadver tent trim actuation. Therudder trim system is DC powered through theYAW circuit breaker on the left essentia l bus

Aileron Trim Indicator

Aileron trim tab position indication is provided by the AIL TRIM indicator (see Figure 15-4).

PITCH TRIM

General

Pitch trim is accomplished by repositioning thehorizontal stabilizer to the desired trim settingwith a dual-motor—primary and secondary— actuator that operates in four modes:

1 Primary pitch




YAW circuit breaker on the left essentia l bus.

Rudder Trim Indicator

Rudder trim tab position indication is pro-vided by the RUDDER TRIM indicator (seeFigure 15-4).

AILERON TRIM

C t l

1. Primary pitchtrim mode

}Primary trim motor

2. Mach trim mode

3. Secondary pitchtrim mode } Secondary trim

4. Autopilot pitchmotor

trim mode

The pilot-operated primary pitch trim and sec-

d it h t i t l t i ll

NOTE

The PITCH TRIM selector switchmust be in the PRI position to enablethe Mach trim system. It may be ineither the PRI or SEC position duringautopilot operation.

Horizontal stabilizer position is displayed onthe PITCH TRIM indicator (see Figure 15-4).

Pitch Trim Actuator

The pitch trim actuator is operated by either

pr imary pi tch tr im is available from both of the control wheel tr im switches and from theMach trim system. In OFF, both tr im motorsand control circuits are deenergized. In theSEC (aft) position, secondary pitch trim isavailable from the secondar y trim switch(Figure 15-4); this renders the pilot’s primarytrim and Mach trim inoperative. The secondary pitch trim switch is spring-loaded to the OFF posi tion.

The autopilot always uses the secondary trimmotor whether the PITCH TRIM selector switch is in PRI or SEC; however, if either

t l h l t i it h i t t d ith




The pitch trim actuator is operated by either of two DC powered motors, either of which can

move the horizontal stabilizer.

• On FC-200 AFCS aircraft, the primarytrim motor and control circui ts are pow-ered through the PITCH circuit breaker on the left essential bus

• On FC-530 AFCS aircraft, the motor is powered by the ba tt er y charging bus,and the PITCH circuit breaker on thel f i l b l l i h

control wheel trim switch is actuated withthe arming button depressed (Figure 15-5 or 15-6) or if the secondary trim switch is ac-tuated, the autopilot disengages.

In the event of inadvertent primary pitch trimoperation (runaway trim), depressing and hold-ing the wheel master switch will:

• Stop only the primary pitch trim motor (aircraft with FC 200 AFCS)


CONTROL LTRIM SWITCH

WHEELMASTERSWITCH

(MSW)



PITCHTRIMPRI

NOSEDN

NOSE

UPSEC

PRI

SEC

OFF


CONTROL WHEELTRIM SWITCH

FAST

SLOW

WHEELMASTERSWITCH(MSW)

3 FLAP SWITCHo



ANNUNCIATORPITCHTRIM

PRIMARY TRIM

PITCHTRIMPRI

NOSEDN

NOSEUP

SEC

TRIMMONITOR

SECONDARY TRIM

PRI

SEC

O

F

F

Pitch Trim Monitor System(FC 530 AFCS)

General

A monitor system incorporated in these aircraft provides a visual indication of certa in faul ts

in the primary trim system.

Though not physically a part of the monitor system, a clicker provides audible evidence of trim in motion—primary or secondary trimsystem—when the flaps are up.

Operation

The monitor system and trim-in-motion clicker are tested in accordance with procedures out-lined in Section 2 of the approved AFM . Either a t h r e e -p o s i t i o n s w i t c h d e c a l e d TR IMOVSP–OFF–TRIM MON and spring-loaded toOFF or the TRIM OVSP and TRIM MON po-sitions of the rotary systems test switch areused to perform the test.

MACH TRIM

General

The Mach trim system is an automatic pitchtrim system that uses the primary trim motor




Operation

The monitor system monitors the primary trimsystem, 3° flap switch, and horizontal stabi-lizer actuator mechanism. Faults are indicated by il lumination of the amber PITCH TRIMlight.

With flaps up (slow trim required), the mon-itor system illuminates the PITCH TRIM lightif it senses that primary trim is running at the

f i d

trim system that uses the primary trim motor to enhance longitudinal stability during ac-ce l e ra t ions /dece le ra t ions a t h igh Mach numbers to compensate for Mach tuck. Thereis no switch to engage the system; it auto-matically becomes active at approximately0.69 MI if the autopilot is not engaged.

Since the Mach trim system requires the useof the primary pitch trim motor, the PITCH

TRIM l i h b i PRI f

Operation

During flight, the air data sensor receivesstatic pressure inputs from the left and rightshoulder static pressure ports (FC 200 AFCS)and a pitot pressure input from the right pitottube (Figure 15-7). On FC 530 AFCS aircraft,

static pressure is provided by the right static1 and left static 2 lines. This will be shown inChapter 16, Avionics.

The air data sensor electrically transmits thisinformation to the Mach trim computer. Withthe autopilot disengaged, the Mach trim sys-tem becomes active at approximately 0.69 MI.

Mach trim is interrupted whenever the aircraftis manually t rimmed. The system resyn-chronizes to function about the new hori-zontal stabilizer position when manual trimis released. In flight, synchronization mayalso be accomplished by selecting the MACHTRIM position on the SYS TEST switch and depressing the TEST button (applies to SNs35-247 and subsequent, 36-045 and subse-quent, and earlier aircraft incorporating SB35/36 22-4).




pp y IThe Mach trim computer commands the ap-

propriate pitch tr im changes—noseup tr im for increasing Mach and nosedown for decreas-ing Mach—through the primary motor of the pi tch trim actuator. The follow-up on the hor-izontal stabilizer provides the nulling signalto the computer.

OVERSPEED WARNING

HORN

MACH

MACH TRIM

FOLLOW UP

Mach Trim Monitor

The Mach trim monitor circuit continuouslymonitors input signals and power to the Machtrim computer, and compares signal inputsfrom the air data sensor (Mach) and Mach trimfollow-up on the horizontal stabilizer. A mal-

function exists if the Mach trim monitor doesnot receive a corresponding signal change fromthe Mach trim follow-up when the air data sen-sor signals change (Mach change).

A malfunction is also indicated in the event of power loss to the Mach trim computer, loss of input signals, or a Mach number/horizontal

• On SNs 35-067 and subsequent, SNs36-018 and subsequent, and earlier air-craft incorporating AAK 76-4, the flap posi tion switches actuate at 3°, 13°, and 25° of flap extension

• On earlier aircraft, the switches actuate

only at 13° and 25°• On aircraft with the preselect flap sys-

tem, flap limit switches automaticallymaintain flap position at the selected setting

If hydraulic system pressure is lost, the flapswill probably remain in their last position




p g stabilizer trim position error. In either case, the

Mach trim monitor disengages Mach trim and illuminates the MACH TRIM light. If speed isabove 0.74 MI, the Mach overspeed warninghorn also sounds. If the fault clears or power isrestored, the system can be resynchronized byselecting the MACH TRIM position on the SYSTEST switch and depressing the TEST button(applies to SNs 35-247 and subsequent, 36-045and subsequent, and earlier aircraft in-

i SB 35/36 22 4) If h i

will probably remain in their last position.However , i f the f laps are extended and hydraulic pressure is lost due to a leak in theflap downline, airloads on the flaps may causesome flap retraction.

The flaps can also be operated from EMER BAT 1 (ON position) in the event of electri-cal failure; however, the flap indicator is not powered by the emergency battery.


SWITCHES

FLAP

LIMIT19 /217 /9

FLAP POSITION

SWITCHESFLAP POSITION

SWITCH

INTERCONNECT

CABLE

FLAP

ACTUATOR

**



FLAP CONTROL

VALVE

POSITION

TRANSMITTER

SWITCHES

(PRESELECT)

Flap Position Indicator

A vertical-scale FLAP position indicator ismounted on the center switch panel (see Fig-ure 15-8).

Left flap position is electrically transmitted to

the indicator. The indicator is DC powered by the TAB FLAP POSN circuit breaker on theright essential bus. The indicator indicatesDN with loss of electrical power, regardlessof actual flap position.

Operation (Preselect Flaps)

switches energize the down solenoid to returnthe flaps to the selected position when air-speed is reduced appropriately.

When the selector switch is moved from DNtoward UP, an intermediate stop is encountered at the 20˚ position to facilitate retraction in a go-

around situation. Further movement of the se-lector switch toward UP or 8˚ requires that theswitch lever be pulled out to clear the stop.

When the flap selector switch is placed in UP,the up solenoid positions the flap control valveto direct pressure to the retract side of both flapactuators. In the fully retracted position, the




When the flap selector switch is placed in

DN, the down solenoid positions the flap con-trol valve to direct pressure to the extend sideof both flap actuators. The down solenoid re-mains energized, and the control valve main-tains down pressure on the flap actuators tohold the flaps full down (40°). A check valveat the control valve inlet prevents flap re-traction in the event of upstream hydraulicsystem failure.

y p ,up solenoid remains energized and the controlvalve maintains positive pressure on the retractside of both flap actuators.

Operation (Nonpreselect Flaps)

When the flap selector switch is placed in DN,the down solenoid positions the flap controlvalve to direct pressure to the extend side of

b h fl h fl b d

If the flap selector switch is left in DN, thedown solenoid remains energized, and the con-trol valve maintains extend pressure on the flapactuators. A check valve at the control valve inlet prevents flap retraction in the event of an up-stream hydraulic system failure.

Placing the selector switch in UP energizes theup solenoid; the control valve repositions todirect pressure to the retract side of both ac-tuators. In the fully retracted position, the upsolenoid remains energized, and the controlvalve maintains retract pressure on the flap ac-tuators. Returning the selector switch to theneutral position deenergizes the up solenoid

buses. If either CB is pulled or either power source is lost in flight, the spoilers will slamdown (if extended) and will be inoperative in both modes. Spoiler mode operation does notrequire 115 VAC on the ground.

A spoiler annunciator light illuminates during

normal spoiler deployment or when an uncom-manded unlocked condition exists on either spoiler. On FC 200 AFCS models, the light isred; on FC 530 AFCS models, the light is amber.

In the event of main system hydraulic failure,the spoilers, if extended, blow down and areinoperative. Spoilers cannot be operated with




p g pand the control valve repositions to neutral.

SPOILERS

The spoilers, which are on the upper surface of the wings forward of the flaps, may be extended symmetrically for use as spoilers (i.e., spoiler mode) or asymmetrically for aileron augmen-tation when the flaps are extended beyond 25°

p p phydraulic pressure from the auxiliary hydraulic pump.

The spoiler mode, when selected, overrides thespoileron mode (if operating).

While airborne, flaps and spoilers should not be exte nded simul tane ously. To do so maycause damage to the flaps and create excessivedrag and loss of lift; this results in inc reased

Operation (Spoiler Mode)

The spoilers can be symmetrically extended or retracted with the SPOILER switch (Fig-ure 15-9).

When the SPOILER switch is in EXT, the

spoiler selector valve is energized, the servovalves meter pressure to the extend side of the spoiler actuators, and the SPOILER lightilluminates steady. Full extension is limited toapproximately 40°. Returning the switch toRET deenergizes the spoiler selector valve; thisvalve repositions to route pressure to the re-tract side of the actuators, and the servo valves

t li Th SPOILER li ht ti i h

Spoiler extension and retraction times vary depending on whether the aircraft is airborneor on the ground, and which AFCS is installed (FC 200 or FC 530). Ground deploy and re-tract times (all aircraft) is 1–2 seconds and 3–4seconds, respectively. Inflight deploymenttimes are 3–4 seconds (FC 200) and 5–7 sec-

onds (FC 530). Retract times are 3–4 secondsfor all aircraft.

Spoiler deployment and retraction causes sig-nificant nosedown and noseup pitching, re-spectively. This should be anticipated and offset by application of elevator control pres-sure and pitch trim, as necessary.




neutralize. The SPOILER light extinguishes

when both spoilers are locked down by lockswithin the actuators (Figure 15-10).


SPOILER SPOILERON

(FC 530)

R AC BUS

SQUAT SWITCHRELAY BOX

(FC 200)

R ESS BUS

SPOILER

SPOILER



SPOILER SWITCH

13 FLAP SWITCHSPOILERONCOMPUTERAMPLIFIER

AUG

AIL

Operation (Spoileron Mode)

During the spoileron—aileron augmentation— mode of operation, the spoilers are indepen-dently extended and retracted in a one-to-oneratio with the upgoing aileron to increase lat-eral control in the landing configuration.

Aileron augmentation—spoilerons—increasesroll control authority up to 50%.

The spoileron mode is automatically engaged when the flaps are lowered beyond 25° and theSPOILER switch is RET. The spoileron computer continuously monitors aileron position.When the ailerons are displaced from neutral,the computer signals the servo valve to ex

SPOILER light will not illuminate duringspoileron operation.

Spoileron operation is shown in Figure 15-11.

Spoileron Monitor System

The computer monitors spoiler and spoileronmodes of operation by a followup in eachspoiler and each aileron. In flight, if a splitof more than 6° occurs between the two spoil-ers (spoiler mode) or between the aileron and spoiler (spoileron mode), the amber AUG AILlight illuminates and the spoi lers slam down.Both modes remain inoperative in flight aslong as the AUG AIL light is illuminated;




the computer signals the servo valve to ex-

tend the spoiler on the wing with the raised aileron. The spoiler on the opposite wing isheld retracted by its servo valve. Spoiler ex-tension is limited to approximately 15° dur-ing spoileron operation (full up aileron). The

long as the AUG AIL light is illuminated;

however, the spoiler mode may be opera tiveon the ground.

SPOILER SPOILERON

R ESS BUS R AC BUS

Spoileron Reset Switch

The SPOILERON RESET switch (see Figure15-9) is spr ing-loaded to OFF. If a malfunc-tion occurs in either mode (AUG AIL light on),moving the SPOILERON RESET switch mo-men ta r i ly to RESET may res to re

spoiler/spoileron operation, provided the mal-function has cleared. If the AUG AIL lightdoes not extinguish, both modes are inopera-tive in flight.

The SPOILERON RESET switch is also used during the spoileron/spoiler preflight check of monitor circuit operation. On the ground withflaps down, holding the switch in RESET

Each system consists of:

• Yaw rate gyro

• Lateral accelerometer

• Computer-amplifier,

• Aileron follow-up

• DC rudder servo-actuator

Additionally, FC 530 AFCS models use a yawdamper force sensor, a calibration assembly,and a three-axis disconnect box.

The rudder servo actuator incorporates a cap-stan mechanism (slip clutch) that allows the




flaps down, holding the switch in RESET

induces a fault that inhibits spoileron movement.Therefore, if the control wheel is turned whileholding the switch in RESET, the AUG AIL lightshould illuminate after the aileron deflectsapproximately 6°. The system can be reset byreleasing the SPOILERON RESET switch toOFF and then momentarily moving it back toRESET. Refer to the approved AFM for thecomplete spoileron/spoiler check.

pi lot to over ride the yaw damper at any time,if required, by applying sufficient rudder pedalforce.

When the stall warning indicators are in theshaker range, yaw damper effectiveness is re-duced. The reduction signal for the primaryyaw damper comes from the left stall warningsystem; for the secondary yaw damper, itcomes from the right stall warning system.

YAW DAMPERCONTROL PANEL

The yaw damper control panel on the center pedestal (Fi gure 15-12) provides the yawdamper selection, test, and indicating func-tions. The dual systems are independent, butshare a common control panel.

On FC 200 AFCS models, two PWR/TEST buttons—one for each yaw damper—are used to apply power to the respective controller-am- pl if ier, and for system test ing. The two greenPWR/TEST lights illuminate to indicate thatthe associated system is powered. The twoENG b tt id th f

A single servo force indicator provides indi-cation of the amount of rudder force beingapplied by whichever yaw damper happens to be engaged, with clockwise deflection indi-cating a right rudder force.

On FC 530 AFCS models, a single TST button

provides si mult aneous te sti ng of bo th yawdamper systems. Two PWR buttons—one for each yaw damper—are used to apply and re-move power to their respective controller-am- plif iers. Two ENG buttons, one for each yawdamper, are used to engage and disengage theselected yaw damper. The two green ON an-nunciators illuminate to indicate that theassociated system is powered The two green




ENG buttons provide the means of engage-

ment. The two green ENG lights illuminate toindicate an engaged yaw damper. Yaw damper disengagement may be accomplished by depressing the associated inboard OFF button,while power may be removed from the systems by depressing the associated outboard OFF button.

associated system is powered. The two green

ENG annunciators illuminate to indicate anengaged yaw damper. A servo force indicator is provided for each yaw damper, providing in-dication of rudder force being applied by its respect ive yaw damper, wi th c lockwise deflection indicating right rudder force.

OPERATION (FC 200 AFCS)

When the AUTOPILOT master switch is on,electrical power is applied to both yaw damper amplifiers, which causes both green PWR/TESTlights to illuminate. However, if the AUTO-PILOT master switch is off, the PWR/TEST

buttons, when indiv idually depressed, apply power to their respective systems, which causesthe associated PWR/TEST light to illuminate.

With power on (PWR/TEST lights illuminated),depressing either ENG button engages the cor-responding yaw damper and illuminates the as-sociated green ENG light. If one yaw damper is engaged depressing the opposite ENG button

OPERATION (FC 530 AFCS)

On these aircraft, the PWR buttons must be depressed in order to apply power to the indi-vidual amplifiers. Depressing a PWR buttona second time removes power from the am- pl if iers .

With power on (PWR annunciators illuminated),depressing either ENG button the first time en-gages the corresponding yaw damper and illu-minates the associated ENG annunciator.Depressing the ENG button a second time dis-engages the yaw damper. If one yaw damper isengaged, depressing the opposite ENG buttonautomatically disengages the first yaw damper




is engaged, depressing the opposite ENG button

automatically disengage the first yaw damper and engages the second.

Disengagement of either yaw damper may beaccomplished by depressing the correspond-ing OFF button or by momentarily depressingeither control wheel’s master switch (MSW).On these aircraft, there is no audible annun-ciation of disengagement.

automatically disengages the first yaw damper

and engages the second.

Disengagement of either yaw damper mayalso be accomplished by momentarily depressing either contro l wheel’s master switch(MSW). On these aircraft, the audible au-topilot disconnect tone always sounds to sig-nal yaw damper disengagement.

STALL WARNINGSYSTEMS

GENERAL

One of two stall warning systems may be in-stalled on the airplane. SNs 35-067 and sub-sequent, 36-018 and subsequent, and earlier aircraft incorporating AAK 76-4, have theAlpha Dot system. Earlier unmodif ied aircrafthave the non-Alpha Dot system.

Both are dual systems that provide visual and tactile warning of an impending stall and are

therefore, each system can be powered evenwhen the battery switches are off. The L and R STALL warning lights are the only compo-nents that do not take power directly from the battery buses.

ANGLE OF ATTACK IndicatorsThe computers translate signals from the stallvane transducers into visual indications of stallmargin on the ANGLE OF ATTACK indicators.The face of the indicators is divided into threecolor segments: g reen, yellow, and red:

• Green—Represents the normal operat-ing range




g p g

equipped with the following dual (left and right) components:

• Stall vane/transducer assemblies

• Computer-amplifiers

• Red STALL warning lights

• Stick shaker motors

• ANGLE OF ATTACK indicators

ing range

• Yellow—Warns of an approaching stallcondition; tactile warning occurs in thisarea, alerting the pilot to take positiveaction

• Red segment—Signifies that aerody-namic stall is imminent or has occurred;the stick pusher is engaged in this area,thereby forcing a reduction in angle of


L

STALL

R

STALL

L STALL

WARNING

STALL WARNING

VANES

STALL WARNING

ANGLE OF ATTACK ANGLE OF ATTACK

R STALL

WARNING



STALL STALLWARNING

COMP/AMP

OFF

L R

BIAS INPUTS:

FLAP POSITION

ALTITUDE

RATE SENSOR

WARNING

COMP/AMP

**

Pusher

The stick pusher function utilizes the elevator pi tch servo to reduce angle of a t tack by decreasing pitch attitude. Pusher activation provides elevator down motion, causing a sud-den abrupt forward movement of the control col-

umn. The mechanical slip clutch on the pitchservo allows the pilot to override an inadver-tent pusher actuation due to malfunction.Additionally, on aircraft with the FC 530 AFCS,depressing and holding the wheel master switch cancels an inadvertent pusher. See the approved AFM for appropriate corrective action.

OPERATION

During flight, the stall warning vanes alignwith the local airstream. Vane-operated trans-ducers produce a voltage proportional to air-craft angle of attack. These signals, biased byinformation from the flap position switches,

altitude switches, and rate sensors (as appli-cable) are sent to the respective computer.

As angle of attack increases, the indicator pointer moves to the right. As it crosses thegreen/yellow line, activation of the flashingSTALL lights, stick shaker, and stick nudger (if installed) begins. If angle of attack is al-lowed to increase further, the pusher is activated




Nudger (FC 530 AFCS)On these aircraft, a nudger is incorporated into the stall warning system. As angle of attack increases slightly beyond the point of shaker motor operation (but prior to pusher operation), a gentle pulsating forward pushcommand is applied to the pitch servo (thesame servo that operates the pushers).

lowed to increase fur ther, the pusher is activated

as the pointer crosses the yellow/red line.

Assuming an unaccelerated entry to a stallcondition at altitudes below 22,500 ft, thegreen/yellow line approximates 7 kt or 7%above pusher speed, whichever is higher.The yellow/red line approximates 5% abovestall speed (non-Alpha Dot); 1 kt above stallspeed (Alpha Dot, except FC 530 AFCS air-


MACH OVERSPEEDWARNING/STICK PULLER

GENERAL

The Mach overspeed warning system providesaudible overspeed warning in the event aircraftspeed reaches VMO or MMO . The stick puller function signals the pitch servo to torque theelevator nose up if MMO is exceeded. On FC530 AFCS models, the puller also operates if high-altitude VMO is exceeded.

OPERATION

The overspeed warning horn is functional whenever the aircraft electrical system is powered and either WARN LTS circuit breaker is engaged (essential buses). The stick puller system becomes functional when the L STALL

WARN switch i s positioned to the on (STALLWARN) position.

The STALL WARN switches should remain onat all times in flight except as directed by theapproved AFM Emergency Procedures and Abnormal Procedures sections.

With the stick puller inoperative, speed is



The stick puller utilizes the autopilot pitchaxis circuitry to control the elevator servoforce applied. The resultant noseup force onthe control column during puller actuation isapproximately 18 lb.

If the autopilot is engaged, puller actuationcancels any selected flight director pitch modesand inhibits autopilot use of the pitch servo

With the stick puller inoperative, speed is

limited to 0.74 MI. The mechanical slip clutchon the pitch servo allows the pilot to overridea n i n a d v e r t e n t p u l l e r a c t u a t i o n d u e t o malfunction. Additionally, on FC 530 AFCSaircraft, depressing and holding the wheelmaster switch cancels an inadvertent puller.See the approved AFM for appropriate cor-rective action.




1. The aircraft systems that use the pitchservo to position the elevator are:

A. Autopilot, Mach trim, stick puller

B. Autopilot, stick pusher, stick puller C. Pusher, stick puller, Mach trim

D. Yaw damper, stick pusher, stick puller

2. The aircraft is trimmed in the pitch axis by:

A. The elevator trim tab

B. Canards

6. In the event of runaway trim, both trimmotors can be disabled by:

A. Depressing and holding either controlwheel master switch

B. Moving the PITCH TRIM selector switch to OFF

C. Moving the PITCH TRIM selector switch to EMER

D. A or B

7. The MACH position on the rotary systemtest switch is used to test:


QUESTIONS



C. The movable horizontal stabilizer D. The elevator downspring

3. To enable pitch trim through the controlwheel trim switches, the PITCH TRIMselector switch must be in:

A. PRI or SEC

B. PRI, OFF, or SEC

A. Mach trim and Mach trim monitor

B. Mach overspeed warning horn and stick puller

C. Mach monitor

D. The HORN SILENCE switch

8. In the event of aircraft electrical failure,the flap position indicator will:


10. The SPOILERON RESET swi t ch i sused to:

A. Retract the spoilers in the event of amalfunction

B. Extend the spoilers in the event of amalfunction

C. Reset the spoiler/spoileron systemwhen the AUG AIL light illuminates

D. Test the monitor system in flight

11. If one yaw damper is found inoperative pr ior to takeoff:

A. The aircraft may be flown, but altitudeis restricted to 20,000 ft

B The aircraft may be flown but altitude

13. The electrical power source for the stallwarning system is provided by:

A. Battery buses

B. Battery-charging bus

C. Main DC buses

D. Emergency battery

14. If either L or R stall warning system isfound to be inoperative before takeoff:

A. The aircraft can be flown provided the STALL WARN circuit breaker is pulled for the inoperat ive system

B. The aircraft can be flown provided the pilot has an ATP rating



B. The aircraft may be flown, but altitudeis restricted to 41,000 ft

C. The aircraft may be flown, but theYAW DAMP circuit breaker for the in-operative system must be pulled

D. The aircraft must not be dispatched

12. When the ANGLE OF ATTACK indica-tor pointers are in the yellow segment:

C. The aircraft may be flown provided theautopilot and yaw damper systems areoperating

D. The aircraft must not be flown

15. The switch used to turn the stick puller system on and off is the:

A. STICK PULLER switch

CHAPTER 16

AVIONICS

CONTENTS

Page

INTRODUCTION................................................................................................................. 16-1

GENERAL ............................................................................................................................ 16-1

NAVIGATION SYSTEM...................................................................................................... 16-2

Pitot-Static System (FC 200 AFCS) 16-2




Pitot Static System (FC 200 AFCS).............................................................................. 16 2

Pitot-Static System (FC 530 AFCS).............................................................................. 16-4

Air Data ......................................................................................................................... 16-6

Ram Air Temp Indicator................................................................................................ 16-7

AUTOFLIGHT SYSTEM ..................................................................................................... 16-7

ILLUSTRATIONS

Figure Title Page

16-1 Pitot-Static System (FC 200 AFCS) ...................................................................... 16-3

16-2 Pitot Head (Typical) ............................................................................................... 16-2

16-3 Static Ports (Typical).............................................................................................. 16-2

16-4 ALTERNATE STATIC SOURCE Valve................................................................ 16-4

16-5 Pitot-Static Head (Typical)..................................................................................... 16-4

16-6 Pitot-Static System (FC 530 AFCS) ...................................................................... 16-5




16-7 STATIC PORT Switch ........................................................................................... 16-6

16-8 RAM AIR TEMP Indicator ................................................................................... 16-7

16-9 ADI and HSI (Typical)........................................................................................... 16-8

16-10 Remote Heading and Course Selector (Typical).................................................... 16-8

16-11 Autopilot and Flight Director Control Panels........................................................ 16-9

16-24 Rosemount Pitot and Static Probe ....................................................................... 16-30

16-25 Shoulder Static Port ............................................................................................. 16-30

16-26 Pitot-Static System Schematic for AFCS FC 200 Aircraft.................................. 16-31

16-27 Pitot-Static System Schematic for AFCS FC 530 Aircraft.................................. 16-32

16-28 Static Port/Source Switch .................................................................................... 16-33

16-29 West Star Air Data Computer.............................................................................. 16-33

16-30 West Star Learjet 35/36 RVSM Avionics Block Diagram ................................... 16-36

16-31 West Star Pilot Altimeter..................................................................................... 16-36

LEARJET 30 SERIES PILOT TRAINING MANUAL1

6 A V I O N I C S



16-32 West Star Copilot Altimeter................................................................................. 16-37

16-33 Altitude Alerter.................................................................................................... 16-37

16-34 Standby Altimeter................................................................................................ 16-38

16-35 Right Airspeed Static Valve................................................................................. 16-38

CHAPTER 16 AVIONICS




NAVIGATION SYSTEM

PITOT-STATIC SYSTEM(FC 200 AFCS)

The pitot-static system supplies pitot and static

air pressure for operation of the airspeed and Mach indicators, the high- and low-altitudeoverspeed switches, the air data sensor, and thestatic defect correction module. Static pressureis also supplied to the copilot vertical veloc-ity indicator, both altimeters, the pressuriza-tion control module, and the aft differential pressure relief valve (Figure 16-1).

The normal static system provides independentsources of static pressure to the pilot and copi-lot instruments. Each static source (pilot or copilot) has one static port on each side of the aircraft nose (Figure 16-3). The dual static ports are provided for redundancy and to re-duce sideslip effects on the instruments that

use static air.


6 A V I O N I C S



A heated pitot head is located on each side of the fuselage just forward of the cockpit (Figure16-2). Pitot heat switches are located on the pi lot ant i- icing control panel. They also supply heat to both sta ll warning vanes. Refer toChapter 10, Ice and Rain Protection, for additional information.


L SHOULDER STATIC PORT

DRAIN VALVE

L PITOT HEAD ALTITUDE

PRESSURE

SWITCH*

FLAP BLOWUP AIRSPEED SWITCH **

R SHOULDER STATIC PORT

R PITOT HEAD

DRAIN VALVE

R FWD

STATIC PORT

DRAIN VALVEDRAIN VALVE

DRAIN VALVE

L FWD STATIC PORT

AIR DATA

SENSOR



R CENTER

STATIC PORT

R AFT STATIC

PORT

ALTERNATE STATIC PORT

(IN NOSE COMPARTMENT)

ALTITUDE

PRESSURE

SWITCH*

L AFT

STATIC PORT

STATIC DEFECT

CORRECTION

An ALTERNATE STATIC SOURCE valve islocated below the pilot instrument panel(Figure 16-4). For normal operation, the lever remains down (CLOSED); for alternate air, thelever is moved up (OPEN).

Four drain valves located near the aft end of the nose gear doors—two on each side—are in-stalled at the system’s low points to drainmoisture from the system.


6 A V I O N I C S



The pitot systems (Figure 16-6) are indepen-dent. The left probe provides pitot pressure for the pilot Mach/airspeed indicator; the right probe he ad provide s pitot pr essure for the

Figure 16-4. ALTERNATE STATICSOURCE Valve

Figure 16-5. Pitot-Static Head (Typical)


GEAR WARNING

AIRSPEED SWITCH

GEAR WARNINGALTITUDE SWITCH

OPTIONALEQUIPMENT

AIR DATA UNIT

MACHSWITCH

LEGEND

PILOT PITOT

COPILOT PITOT

PILOT STATIC

COPILOT STATIC

OTHER



RATE-OF-CLIMBINDICATOR (COPILOT)

PRESSURIZATIONMODULE

RATE-OF-CLIMBINDICATOR (PILOT)

The source of static pressure is controlled with the static port switch located on the pilotswitch panel. The static port toggle switch hasthree positions: L (left), BOTH, and R (right).This switch is normally set to both except inthe event one of the pitot-static heads be -comes inoperable or unreliable (Figure 16-7).

In BOTH, the pilot instruments receive static pressure from the forward port on the left head and the aft port on the right head. The copilotinstruments, the Mach switch, the gear warn-ing altitude switch (14,500 ft), the gear warn-ing airspeed switch, the air data unit, and other optional equipment receive static pressurefrom the front port on the r ight head and the

f h l f h d Thi i

is provided to all user systems only from thetwo static ports on the right pitot-static head.

The shutoff valves operate on DC power suppl ied through the STATIC SOURCE circui t breaker on the left main bus. In the event of electrical failure, all shutoff valves will be

open regardless of the STATIC PORT switch posi tion.

A separa te unheated s ta t ic port i s f lushmounted on the right side of the nose sectionto provide static pressure to the pressurizationcontrol module. Refer to Chapter 12, Pres-surization, for additional information.

AIR DATA


A V I O N I C S



aft port on the left head. This cross connectioneliminates yaw error.

When the STATIC PORT switch is placed inL or R, solenoid-operated shutoff valves areenergized to shut off the static source from theopposite side static ports (see Figure 16-6).

When the STATIC PORT switch i s in L, stat ic

AIR DATA The air data sensor provides air data to the autopilot computer and to the Mach trim computer.On aircraft equipped with the FC 200 automaticflight control system, static input to the air datasensor is from the shoulder static air ports. TheFC 530 equipped aircraft use the copilot static air system for air data unit input. On all aircraft, the

RAM AIR TEMP INDICATOR

Ram-air temperature is displayed on the RAMAIR TEMP indicator located on the center in-strument panel (Figure 16-8). The indicator iscalibrated in degrees Celsius and requires DC power from the ram air temp circui t breaker

on the left essential bus. For conversion tooutside air temperature (OAT), refer to theRam Air To Outside Air Temperature Con-version (RAT to OAT) f igure in Section V of the approved AFM .

installed on SNs 35-408, 35-447, 35-468, 35-506 and subsequent, and 36-054 and subse-quent, and earlier SNs incorporating AAK 83-2.

NOTE

The yaw axis is controlled by the dualyaw damper system, which operatesindependently of the autopilot and flight director.

Both systems incorporate a dual-channel AFCScomputer that integrates the autopilot pitchand roll axes with the customer-specified flightdirector system. The AFCS control panel,which is located in the center of the glareshield,

id il t t th t il t d t th




provides pilot access to the autopilot and to theAFCS compute r fo r t he f l i gh t d i rec to r programming (i.e. , mode selection).

The AFCS computer processes informationreceived from the primary vertical and direc-tional gyros, horizontal situation indicator (HSI), the NAV 1 receiver, and the air data sen-sor. The resulting computed roll and/or pitch

FLIGHT DIRECTOR SYSTEMS

General

Several different flight directors are availablefor installation on the Learjet 35/36. The mostcommon installations are the Collins FD 108,

FD 109, FDS 84, and FDS 85. Either systemincludes an ADI and an HSI that provide con-ventional raw-data attitude and heading ref-erence and glide slope and course deviationdisplays. The basic aircraft attitude and head-ing references are energized whenever DCand AC power is applied to the aircraf t.

The flight director system is connected to the

AFCS h th AUTO PILOT t it hi d


A V I O N I C S



AFCS when the AUTO PILOT master switchis turned on.

When the autopilot master switch is posi-tioned to auto pilot (on), the PWR annuncia-tor illuminates on the AFCS control panel,which indicates that power is available to theautopilot and flight director. The AFCS con-trol panel provides for flight director mode se-

perform the ro ll and pi tch maneuvers neces-sary to align the aircraft symbol with the com-mand bars. Figure 16-12 illustrates the visualindications provided by the ADI and HSI. TheADI also provides for indication of localizer and glide-slope deviation and turn and slip.

Horizontal Situation Indicator(HSI)

The HSI provides a pictorial presentation of aircraft position relative to VOR radials and localizer and glide-slope beams. Heading ref-erence with respect to magnetic north is pro-vided by a remote directional gyro that isslaved to a remote fluxgate compass. The

SLAVE FREE switch on the lower instrument

AUTOPILOT/FLIGHT DIRECTOR

General

The autopilot will automatically fly the aircraftto, and hold, desired heading, attitudes, and al-titudes. The autopilot system can also cap-ture and track VOR/LOC/ILS radio beams.The system provides modes for speed controland vertical rate control as well.

On Learjet 35/36 aircraft with the standard avionics installation, the flight director is in-tegrated with the autopilot by a computer through the AFCS cont ro l panel on theglareshield. Autopilot and flight director

modes are engaged by depressing the applibl d l t b tt th t l




SLAVE-FREE switch on the lower instrument panel allows unslaved opera tion by selectingFREE, in which case the magnetic reference— the flux-gate compass—is removed.

The HSI provides the AFCS computer infor-mation regarding existing heading, headingmarker reference, selected course, and coursedeviation The heading marker—the bug—is

modes are engaged by depressing the appli-cable mode selector buttons on the control panel. Fl ight director only mode se lect ion isaccomplished by depressing the desired modeselectors on the control panel (Figure 16-11), but with the autopilot d isengaged.

When the autopilot is not engaged, the ADIcommand bars indicate the deviation from the

Description

Aircraft SNs 35-462, 35-447, 35-506 and sub-sequent and 36-054 and subsequent areequipped with the FC 530. Earlier aircraft havethe FC 200. Both are manufactured by J.E.T.AAK 83-2 is available to retrofit earlier aircraft

with the FC 530. Both systems include: an au-topilot/flight director computer, an electric box, and interface—all under the pilot seat; theAFCS control panel in the center glareshield;the roll and pitch servoactuators and follow-ups; the customer specified flight director sys-tem; a roll-rate gyro; the NAV 1 receiver; the primary (pilot) ver tical gyro, directional gyroand HSI; and the air data sensor. The FC 530

also uses the altitude alerter and pilot altime-ter for its altitude preselect feature.

AFCS Control Panel

The control panel (see Figure 16-11) in the cen-ter of the glareshield is accessible to both pilots.It provides the switches required for autopilotengagement and flight director mode selection.Annunciator lights—green, amber, blue, or white—appear above the mode select switches.The legend (white lettering) on the panel is backlit. On FC 200 models, annunciator inten-sity and legend lighting is controlled by thePEDESTAL lights rheostat on the copilot side-wall. On FC 530 models, annunciator intensity


A V I O N I C S



20 20

1010 FAST

DECISION HEIGHTANNUNCIATOR

SPEED DEVIATIONDISPLAYGLIDE SLOPE

HORIZON

ATTITUDE TAPE DH

is f ixed so that they are legible in daylight, whilethe NAV LTS switch must be turned on for fixed illumination of the legend lighting.

The autopilot engage (ENG) pushbutton isused only to engage the autopilot; all other pushbutton switches operate wi th al te rnate

action. The fir st depression engages a mode;a second depression cancels it. Automatic can-cellations also occur. Annunciation of themode selected appears above the pushbutton.Any operating mode not compatible with anewly selected mode is automatically can-celed in favor of the latest selection. This al-lows the pilot to advance along the flightsequence without the inconvenience of having

to deselect modes manually

ture and a vertical accelerometer that moni-tors G forces.

When a pitch mode is selected on the AFCS con-trol panel, the computer positions the flight di-rector V-bars accordingly. If the autopilot isengaged, a signal is also applied to the elevator pitch servo, which adjusts elevator position.

Feedback of elevator movement is provided bythe servo follow-up. When the new pitch atti-tude is established, the computer zeroes theservo effort by applying horizontal stabilizer trim via the secondary pitch trim motor, thereby preventing any aircraft pitching motion whendisengaging the autopilot. Pitch changes canalso be induced by either pilot wheel trim switchwithout depressing the center button.

The computer uses the servo follow up to con




to deselect modes manually.

Computer

The two-channel—roll and pitch—computer continuously monitors input signals from allAFCS component sensors. The computer is programmed by depress ing the desi red modeselector button(s) on the AFCS control panel.The computer computes the roll and pitch at

The computer uses the servo follow-up to con-trol pitch changes to a rate of 1° per second,and limits pitch attitudes to ±25° (FC 200) or +20° and –10° (FC 530).

Roll Axis Control

The computer roll channel processes infor-mation from the primary (pilot) vertical gyro,

The computer uses the roll rate gyro to controlroll rates to 6° per second (FC 200), and 4–5º per second (FC 530). Bank angles are limited to a maximum of 30°.

The FC 200 uses a 13° flap position switch toincrease autopilot roll authority when the air-craft is configured for approach. This pro-

vides more lateral author ity at slower speedsand is annunciated by the green APPR light onthe AFCS control panel. The FC 530 uses a 3°flap position switch to desensitize VOR and LOC signals, which enhances close-in stabil-ity during approaches. It does not affect autopilot rol l authori ty, nor is it annunciated.

Electrical Requirements

The autopilot requires DC and AC electrical power

manual autopilot controller when moved in any of the four directions without depressing the trimarming button (Figure 16-13). When an attitudechange is made this way, the appropriate servochanges the attitude of the aircraft and disengagesany modes previously selected in the affected axisexcept NAV ARM, G/S ARM, and ALT SEL

ARM. The autopilot reverts to basic attitude hold in the affected axis when the switch is released.


A V I O N I C S



The autopilot requires DC and AC electrical power.DC power is via the AFCS, AFCS PITCH, and AFCS ROLL circuit breakers on the left essential bus; 115 VAC is via the AFCS PITCH and AFCSROLL circuit breakers on the left AC bus. Allautopilot circuit breakers are on the pilot CB panel;however, on FC 200 AFCS aircraft, there are threecircuit breakers on the front side of the autopilotelectric box under the pilot seat for autopilot and Figure 16-13. Control Wheel Switches

On FC 200 aircraft, depressing and holding either the pilot or copilot MANEUVER switch (Figure16-13) temporarily releases autopilot access tothe pitch and roll servos, biases the command bars out of view, and cancels the ROLL and PITCH modes if engaged previously. Thisenables either pilot to change the aircraft attitude

in both pitch and roll axes manually. When theswitch is released, the autopilot assumes basicattitude hold functions.

During flight director only operation, themaneuver switch simply cancels all selected flight director modes and biases the command bars out of view.

On FC 530 aircraft, depressing and holding either

the pilot or copilot MANUV/RP switchil l il h i h

• Synchronizes the command bars to theexisting pitch attitude

In the case of a dual flight director installation,the copilot pitch SYNC switch synchronizes onlythe copilot command bars to the existing attitudeand cancels the copilot G/A mode, if selected. Itdoes not affect the autopilot in any way (as the

maneuver switch does).

Autopilot Engagement

The AUTO PILOT master switch must be placed on to accomplish system ground checks prior toflight and normally remains on throughout theflight. When the PWR annunciator is illuminated,the autopilot can then be engaged at any time(except during takeoff and landing) by depressing

the ENG button Illumination of the PITCH and




the pilot or copilot MANUV/RP switchtemporarily releases autopilot access to the pitchand roll servos and extinguishes the green ROLLand PITCH annunciators, but does not cancel any previously selected flight director roll or pitchmodes. This enables either pilot to change theaircraft attitude in both pitch and roll axesmanually. When the switch is released, theautopilot resynchronizes to and holds the original

the ENG button. Illumination of the PITCH and ROLL annunciators indicate engagement of therespective axes.

On FC 200 aircraft, initial autopilot engagementcancels all previously selected flight director modes (if bank angle happens to be more than5°), the command bars disappear, and theautopilot holds the existing roll and pitchattitudes (if within normal limits) If bank angle

wings level at the existing pitch attitude. Theautopilot will not engage at bank angles in excessof 38 ±2° regardless of pitch attitude; however, if bank angle happens to be between 30 and 38 ±2°and/or pitch angle is greater than –10° or +20°,the autopilot—at normal rates—rolls and/or pitches the aircraft to the normal limit(s).

If the pitch trim selector switch is in off, theautopilot will not engage.

Attitude Hold Mode

The autopilot is in pitch attitude hold when thePITCH annunciator is illuminated and all other pitch axis annunciators are extinguished (exceptG/S ARM and, for FC 530, ALT SEL ARM).

The autopilot is in roll attitude hold when theROLL i t i ill i t d d ll th

Autopilot/Flight Director ModeSelection

Autopilot and flight director modes are engaged by depressing the appl icable mode se lector but ton on the aut op ilo t con trol pan el . Theengaged modes may be disengaged by depressingthe selector button (except for the SPD mode on

the FC 530 AFCS) a second time or by selectinganother pitch mode.

Flight director only mode selection is made bydepressing the applicable mode selector with theautopilot disengaged.

The roll axis modes are LVL (level), HDG(heading), NAV (navigation), VOR or LOC (used

in conjunction with the NAV mode) BC (backFC 530) REV (b k FC 200)


A V I O N I C S



The autopilot is in roll attitude hold when theROLL annunciator is illuminated and all other roll axis annunciators are extinguished (except NAV ARM). When the autopilot is in both pitchand roll attitude hold, the flight director command bars will be out of view. Autopilot roll(bank) limit is a nominal 30°, while pitch limitsare ±25° (FC 200) or +20° and –10° (FC 530).

in conjunction with the NAV mode), BC (back course, FC 530), REV (back course, FC 200),and 1 ⁄ 2 BNK (half bank, FC 530).

The pitch modes are SPD (speed), V/S (verticalspeed), G/S (glide slope), ALT SEL (altitudeselect, FC 530), ALT HOLD (altitude hold), and SFT (soft). The SPD submodes of IAS and MACH, and the V/S, G/S CAPT, ALT SEL


NOITCNUFROTAICNUNNAEDOM

SOFT SOFT When depressed, the autopilot provides softer response in the pitchand roll axes for flying through turbulence. No function during flightdirector only operation.

NOTE

SOFT mode is locked out when an ILS frequency istuned on NAV 1.

HDG ON When selected, flight director commands are generated tomaneuver the aircraft to fly a heading selected with the pilot HSIheading bug using up to 25° of bank.

NOTEThe turn will be commanded in the shortest direction.It is recommended that the heading bug initially beset to not more than 135° in the direction of the

desired turn when the turn is more than 135°.

Table 16-1. FC 200 AUTOPILOT SYSTEM MODES AND ANNUNCIATORS (Cont)



des ed tu e t e tu s o e t a 35

serutpactahtnoitcnufrotceridthgilfehtsetavitcati,detcelesnehWVANand tracks VOR and LOC. Functional only when the NAV 1 receiveris tuned to the appropriate frequency, NAV flag is out of view, anddesired course is set on the pilot HSI. The HDG mode may beused to intercept the course provided the intercept angle is lessthan 90°.

ARM Illuminates when NAV mode is selected. Goes out when the CAPTlight illuminates The ARM light will flash if NAV CAPT disengages


A V I O N I C S


ylnorotceridthgilfehtybdednammocsilevelsgniw,degagnenehWLVL.degagnesitolipotuaehtfi)LEVEL(

ON Indicates the level mode is engaged. It is also a function of G/A

mode, but has no other flight director only functions. tahtedutittahctipadnammoclliwrotceridthgilfeht,detcelesnehWDPS

.noitcelesedomfoemitehttagnitsixedeepsriaehtniatniamlliw)DEEPS(Power must be set by the pilot.

IAS Illuminates at altitudes up to approximately 29,000 ft.

MACH Illuminates at altitudes above approximately 29,000 ft.

tahtedutittahctipasdnammocrotceridthgilfeht,detcelesnehW-REV(S / V ehtybtesebtsumrewoP.deepslacitrevgnitsixeehtniatniamlliwLACIT

.tolip)DEEPSON Ill i t h V/A d i l t d




p)ON Illuminates when V/A mode is selected.

NOTEBefore engaging this mode, maintain the desiredrate long enough (approximately 15 seconds) forvertical speed indicator lag to diminish.

serutpactahtnoitcnufrotceridthgilfehtsetavitca,detcelesnehWS / G.epolsedilgehtEDILG(

SLOPE)



PWR Indicates electrical power is available for autopilot/flight directoroperation (circuit breakers are in and AUTO PILOT master switch isin ON position).

TST (TEST) When depressed, all autopilot controller annunciators illuminate

(light test only). When depressed simultaneously with ENG button,a system self-test is performed.

MON (MONITOR) Illuminates during self-test. Flashes if fault is detected.

ENG ROLL When depressed, the autopilot engages and the ROLL and PITCHPITCH annunciators illuminate.

SFT SOFT When depressed, the autopilot provides softer response in the pitchand roll axes for flying through turbulence. No function during flightdirector only operation.

NOTE

SFT mode is locked out when in NAV localizerCAPT NAV VOR APPR d ALT SEL CAPT

Table 16-2. FC 530 AUTOPILOT SYSTEM MODES AND ANNUNCIATORS



CAPT, NAV VOR APPR, and ALT SEL CAPT.

HDG ON When selected, flight director commands are generated tomaneuver the aircraft to fly a heading selected with the pilot HSIheading bug using up to 25° of bank.

NOTEThe turn will be commanded in the shortest direction.It is recommended that the heading bug initially beset to not more than 135° in the direction of thedesired turn when the turn is more than 135°


A V I O N I C S


esruockcabrezilacolrofdetcelesedomVANhtiwylnolanoitcnuF-KCAB(CBsirotceridthgilfehtotnoitamrofniesruoc,detcelesnehW.hcaorppa)ESRUOC

reversed and the glide-slope signal is locked out. The publishedinbound (front) course must be set in the pilot HSI course window.

ON Indicates that the backcourse mode is selected. Is also a function ofG/A mode.

NOTE

BC may also be used to fly outbound on an ILS frontcourse.

eht,)tonrodegagnetolipotua(desserpedsinottubLVLehtnehWLVLdetcelesylsuoiverpynadna,levelsgniwdnammoclliwrotceridthgilf)LEVEL(

roll mode will be canceled. If a pitch mode happens to be engaged,

pitch commands for that mode will not be affected; otherwise, thed b ill th i ti it h ttit d




command bars will assume the existing pitch attitude.

ON Indicates the level mode is engaged.

NOTE

During flight director only operation, selecting SPD,V/S, or ALT HLD without a prior roll mode selectionwill automatically engage the LVL mode.

dihidlliidhilfhdlhWDPS



serutpactahtnoitcnufrotceridthgilfehtsetavitca,detcelesnehWS / G.epolsedilgskcartdnaEDILG(

SLOPE)Functional only when the NAV 1 receiver is tuned to an ILS fre-

quency, an active glide-slope signal is present, the G/S flag is out ofview, and the BC mode is not selected.

ARM Illuminates when the G/S mode is selected and the aircraft is noton the glide-slope beam. Goes out when the aircraft captures thebeam.

CAPT Illuminates when the aircraft captures the glide-slope beam.

FNL (FINAL) Illuminates during an ILS or a localizer approach when the LOC andG/S beam signals are being desensitized for close-in stability.

NOTE




If the radio altimeter signal is valid, the FNL light willilluminate at approximately 1,200 ft AGL. If theradio altimeter is not valid, the FNL mode will beactivated when passing over the outer marker. If theradio altimeter and outer marker are not valid,depressing the NAV 1 TEST button will activate theFNL mode. This should be accomplished at the finalapproach fix. The flaps must be down 3° or more to

Autopilot Disengagement

Whenever the autopilot and/or rol l axes disengage, the applicable PITCH and/or ROLLannunciators will extinguish and the autopilotdisengage tone will sound, as def ined below:

• Either control wheel trim switch, with

arming button depressed and moved inany of the four directions (NOSE UP, NOSE DN, LWD, or RWD), will disen-gage both autopilot axes

• Ei ther cont ro l wheel master switch(MSW), when depressed, will disengage both autopilot axes and the yaw damper

• The AUTO PILOT master switch, whenset to OFF, will disengage both autopilotaxes

at a wings level 9° noseup pitch position.

NOTE

On SNs 35-002 to 35-009 and 36-002to 36-006, the G/A mode is coupled tothe autopilot if engaged when power isadvanced to approximately 80%

N1.

Servo Force Meters

Two servo force meters are located in the cen-ter of the control panel. The indicators providean indication of what autopilot servo forcesare present when the autopilot is engaged. Theleft one indicates roll force and the right, pitchforce. If the force meter(s) are deflected, the ap-

propriate axes should be trimmed to center themeter(s) prior to engaging the autopilot If the


A V I O N I C S



• The PITCH TRIM selector switch, whenmoved to the OFF position, will disen-gage both autopilot axes, but only whenit attempts to trim the horizontal stabi-lizer and cannot (FC 200); on FC 530 air-craf t , au topi lo t d isengagement i simmediate

meter(s) prior to engaging the autopilot. If theautopilot is engaged and the meter(s) indicatea steady deflection, the autopilot should be dis-engaged and the appropriate axis retrimmed.Small deflections before and after engagementare normal.

Roll Monitors

Out-of-trim Monitors (FC 530)

With the autopilot engaged, the out-of-trim monitors cause the applicable PITCH or ROLLannunciator to flash if an out-of-trim conditionexists to a degree that servo force is continuouslyapplied for more than approximately 20 seconds.The light continues to flash until either the trim is

restored or the axis is disengaged.

G-force Monitor (FC 530)

G forces are sensed by the vertical accelerometer with the autopilot engaged. The G-force monitor causes the elevator to streamline whenever the Glevel reaches 1.6 G or 0.6 G. The pitch axis remainsengaged, but keeps the elevator streamlined.Previously engaged pitch modes also remain on.When the aircraft is within the G limits, the pitch


Figure 16 14 Altitude Display



axis resumes normal elevator inputs.

Autopilot/Stick Nudger/Pusher/ Stick Puller Interface

If the autopilot is engaged and the stick nudger (FC 530), pusher, or puller actuates, any selected pi tch mode di sengages The au topi lo t then

The altitude alerter located in the center in-strument panel functions in conjunction withthe pilot altimeter. An OFF flag adjacent to thealtitude display will be in view whenever power is not available to the aler ter. Duringfli ht th i ft ithi i

Figure 16-14. Altitude Display

COMMUNICATIONSYSTEM

STATIC DISCHARGE WICKS

A static electrical charge (commonly referred to

as P static or precipitation static) builds up onthe surface of an aircraft while in flight and causes interference in radio and avionicsequipment operation. The charge may bedangerous to persons disembarking after landingas well as to persons performing maintenanceon the aircraft. The static wicks are installed onall trailing edges (Figure 16-15) to dissi patestatic electricity.

RVSM SYSTEM

GENERAL

In the late 1950s, vertical separation for air-craft in upper airspace was 1,000 ft. However,in the early 1960s, as more and more aircraftwere entering the airspace above 29,000 ft, adetermination was made to increase the ver-tical separation above 29,000 ft to 2,000 ft.Starting in the late 1970s, a series of studieswas conducted to determine the feasibility of reducing the current 2,000-foot vertical sep-aration between FL290 and FL410 to 1,000 ft.These studies continued through the late 1980s.The studies concluded that the reduction to a1,000 foot separation was feasible, providingh i f i d i h l i


A V I O N I C S



the aircraft were equipped with an altimetrysystem with increased accuracy, which would also produce increased accuracy in the altitudereporting system.

The f irst implementation of Reduced VerticalSeparation Minimum (RVSM) began in theN h A l i R i i M h 1997 Si

After January 20, 2005, aircraft not equipped with special RVSM equipment must be granted special permission to transition through block altitudes FL290 to FL410, or maintain an al-titude of FL290 or lower.

All Learjet models 35-35A/36-36A are eligi- ble for RVSM modif ication. However, in somecases specific aircraft modif ications must have been al ready successfully completed and doc-umented in the aircraft log book, or complied with concurrent with the RVSM modif ication.A maintenance log check must be accomplished to ensure a ll necessary modif icationshave been completed or scheduled.

There are currently two Supplemental TypeCertificate (STC) holders that can accomplishh i f difi i f RVSM

With the implementation of D-RVSM, the fol -lowing are areas of significant importanceand checks should be closely monitored:

1. Altimeter Checks—Prior to takeoff for flights planned into RVSM airspace, pri-mary altimeters must be within 75 ft of a known elevation. Whi le within RVSMairspace, primary altimeters must bewithin 200 ft of each other.

2. Altitude Awareness—To preclude er-rors in hearing clearances and/or incor-rectly setting the altitude pre-select, thefollowing technique/SOP is suggested:

a. Pilot flying is manually flying the air-craft, and pilot monitoring sets altitude pre-selector; both pilots point to the al-

i d i h l i d l d




the necessary aircraft modifications for RVSMfor the Learjet 35-35A/36-36A group. One isAero Mech, Inc. (AMI) under their STC Number s ST 00952SE, ST 00952SE-D, ST01199NY and ST 01199NY-D. To simplify future discussion, this will be referred to as theLearjet RVSM Installation. The other is WestS /H ll d h i STC N b ST

titude set in the altitude pre-selector, and both verbally state that altitude.

b. Pilot flying is flying the aircraft on au-topilot, and pilot flying sets the alti-tude pre-selector; both pilots point to thealtitude set in the altitude pre-selector,and both verbally state that altitude

LEARJET RVSM INSTALLATION

Rosemount Pitot-Static Probes

For FC 200 equipped aircraft, the traditional pitot tubes and static ports are removed and replaced by Rosemount pi tot s tatic probe (Fig-ure 16-16). Earlier FC 200 aircraft that have

already been modified with the Rosemount pi tot stat ic probe system and FC 530 autopi-lot are described under the FC 530 modif ica-tion (see AFMS W1266 ). The alternate staticsource valve at the bottom of the left side of the instrument panel is removed. The pres-surization static port installation has notchanged.

Static pressure is sensed by two sources on each probe; static 1 (S1) and static 2 (S2). Static 1on the left probe is cross-connected to static 2on the right probe, and static 1 on the right probe is cross connected to static 2 on the lef t probe.

Four solenoid-operated isolation shutoff valvesenable the pilot to select the source of static pressure . The source of stat ic pressure is con-trolled by a static source/static port switch(Figure 16-17) located on the top of the throt-tle quadrant or on the anti-ice control panel.


A V I O N I C S



deformation has occurred in that area. Also,check the pitot static probe heads for any de-formation or obstruction around the inlet or static ports.

IS&S Altimeter/ADC System

General

On the FC 200 and FC 530 autopilot aircraft, the pilot (servo pneumatic or pneumatic) and copi-lot (pneumatic) altimeters are replaced with theIS&S combination self sensing altimeter

(ADDU–Air Data Display Unit/ADC–Air DataComputer) (Figure 16-19).




Static Source/Static Port On FC 200 aircraft an analog interface unit

Figure 16-18. Right SidePitot-Static Probe

Figure 16-19. Pilot and Copilot Altimeters

On both FC 200 and FC 530 aircraft, the alti-tude alerter panel is removed and replaced with an air data switch panel (Figure 16-20).This panel consists of green ADC1 and ADC2 pushbutton switchlights and a red AIU FAILannunciator light.

NOTE

The autopilot must be disengaged when swi tching from one ADCsource to another.

To toggle between IN.HG or hPa, press the

BARO select knob located to the lower right onthe altimeter. If the BARO knob is held depressed for longer than four seconds, unit se-lection mode is entered and each additional press of the knob for four seconds will togglethe altimeter display between IN.HG and hPa.

If the BARO select knob is depressed and held for eight seconds or longer, the altitude unit

display will toggle between feet and meters.

B i i b i h


A V I O N I C S

Figure 16-20. Air Data Switch Panel



The altimeter/ADC combines the function of the basic altimeter with those of the traditional al-titude alerter and is also a self-sensing unit with pitot and static connections. A new standby al-timeter is installed and plumbed to the copilot

Barometric pressure is set by rotating theBARO select knob. Momentarily depressing theBARO knob for less than two seconds will set29.92 IN.HG or 1013 hPa. Note that the mas-ter A and the slave baro set knobs are totallyindependent and different units (IN.HG or hPa)and different baro settings are possible.

Figure 16 20. Air Data Switch Panel

Power Source/Failure

Elect r ica l power for the p i lo t a l t imeter (Figure 16-21) is su pplied by the ALTM or PRI ALTM circuit breaker located on the lef tessential bus (L ESS BUS). It may also be powered by the emergency battery through theEMER ALTM circuit breaker located on the

left circuit breaker panel.

If normal elec trical power is lost to the pilotaltimeter (ADDU) and it is being powered bythe emergency battery, the pilot ADDU willfunction using the emergency battery power, but the PWR and COM indication will il lu-minate on the pilot ADDU (altimeter) display(Figure 16-22).

On FC 200 aircraft, a pilot altimeter (ADDU)emergency lighting (PLT ALTM EMER LTG)switch may be installed on the pilot side panel.If the switch is installed and normal electri-cal power is lost to the pilot altimeter, theADDU back lighting will remain ON and the pi lot may select desired intensity of the digi-tal display by using this switch.

If this switch is not installed, the ADDU back lighting will remain on if the pilot INSTR PNL dimmer knob (pilot side panel) is turned ON (out of detent) and the altitude displaywill be dimmed.

For daylight conditions, the INSTR PNL dim-mer knob should be turned OFF (in the OFFdetent position), which will cause the back lighting to be off and the altitude display to be




lighting to be off and the altitude display to be bright .

RESSBBUS

GEAR

For FC 530 aircraft, this switch is not in-stalled; the ADDU back lighting will remainON if the pilot INSTR PNL (pilot side panel)dimmer knob is turned on (out of detent),and the altitude display will be dimmed.

For daylight conditions, the INSTR PNL dim-mer knob should be turned off (in the OFF de-

tent position), which will cause the back lighting to be off and the altitude display to be bright. If normal electrical power is lost to thecopilot altimeter, the copilot ADDU will beinoperative.

Altitude Alerter Operation

Select the desired alerter altitude by rotatingthe ALT SEL knob on the face of the altime-ter (ADDU) (see Figure 16-19). Clockwiserotation causes the selected altitude to in-crease and counter-clockwise to decrease.Knob sensitivity is 100 ft per detent (30 me-

ters in metric mode). As long as the same units(feet or meters) are selected, rotating the ALTSEL knob on the master ADDU (A illumi-nated) changes the selected altitude on boththe master and the slave ADDU. If differentunits are selected, the display on the slave unit bl anks and it s ALT SEL knob is disabled.Momentarily depressing the ALT SEL knobextinguishes the altitude alarms until the ap-

propriate approach condi tions are met again.


A V I O N I C S



PRI ALTML ESS BUS A

1

ANALOGINTERFACE UNIT

(AIU)

EMER ALTML ESS BUS

1

PILOT ALTIMETER

Altitude Reporting

Altitude reporting data may be supplied fromeither air data display unit (ADDU). SelectingADC-1 on the air data switch panel (seeFigure 16-20) provides altitude informationfrom the pilot ADDU for either transponder.Selecting ADC-2 on the switch panel provides

altitude information from the copilot ADDUfor either transponder.

The TFR 1-2 switch—if installed—is located on the transponder control panel. SelectingTFR-1 transmits altitude information fromthe LEFT transponder supplied by the selected ADDU. Selecting TFR-2 transmits altitudeinformation from the RIGHT transponder sup-

pl ied by the se lected ADDU.

Standby Altimeter

The standby altimeter—a pure static altime-ter—is plumbed to the copilot static system(Figure 16-23). Electrical power for the al-timeter lighting and vibrator is supplied fromthe aircraft emergency bat tery when theswitch is placed in ON. The standby altime-

ter is not powered when the EMER BATswitch is in STBY. There is an OFF flag onthe left lower corner to indicate that the vibrator is no t operat ing.




System Checks/Tests

System Operational Check

An operational check of the altimeter/ADCsystem is outlined in the appropriate AirplaneFlight Manual Supplement Refer to your sup

For FC530 aircraft, the altitude position cor-rection charts supplied with the Airplane

Flight Manual must be applied to the standbyaltimeter. The AFM corrections must also beapplied to the indicated altitude when theSTBY indicator light is illuminated on theIS&S ADDU (primary altimeters). The cor-rection charts supplied with AFMS W1484 are

used for altitude correction on the primaryaltimeter (ADDU) when the STBY indicator light is not illuminated.

WEST STAR RVSMINSTALLATION

Rosemount Pitot-Static Probes

For FC 200 autopilot equipped aircraft, the tra-ditional pitot tubes and static ports are re

The Rosemount pitot static probes on each sideof the nose section provide both pitot and static pressures to designated systems. The probes alsocontain heating elements for anti-icing and arecontrolled by the L and R PITOT HEAT switches

(see Chapter 10, Ice and Rain Protection). Drainvalves are located near the end of the nose gear


A V I O N I C S

Figure 16-25. Shoulder Static Port



ditional pitot tubes and static ports are re-moved and replaced with Rosemount pitot and static probes (Figure 16-24).

doors and are installed at the system low pointsto drain moisture during preflight.

The left pitot pressure is plumbed to the pilotairspeed indicator and the Honeywell AZ-252air data computer (ADC). The right pitot pres-sure is plumbed to all other systems that use

F OR T RAI NI N G

P URP

EXISTING NOSE

LOWER SKIN

L ROSEMONT

P/S PROBE

EXISTING COPILOT

STALL WARNING

22,500 FT SWITCH

F.S.

SHOULDERSTATIC PORTS

NEW AZ-252 AIR DATA COMPUTER

EXISTING DRAINS5 PLACES

E

W



P O SE S

ONL Y

PILOT AIRSPEED

STATIC VALVE

F.S.160.77FR 5

CO

STA

W

1 6

- 3 2

F OR T RAI NI N

G P U

NOSE

LOWER SKIN

L ROSEMONT

P/S PROBE

ISOLATIONVALVES (REF)

NEW AZ-252AIR DATA COMPUTER

DRAINS4 PLACES

MACH SWITCH



RP O SE S

ONL Y

F.S.160.77FR 5

REFERENCE DESTINATIONS

Static Port/Source Switch

On FC 530 aircraft, including earlier FC 200aircraft that have been modified with theRosemount pitot static probe system, a static port /source switch is installed (Figure 16-28).This switch is installed either during produc-tion or is installed by STCs: ST 00321WI or ST

00321WI-D-FC 200 autopilot retrofit withRosemount pitot-static probes. The functionof this switch does not change with the West Star installation.

Air Data Computer

The West Star RVSM installation chose to useHoneywell equipment. The Honeywell AZ-252 advanced air data computer (ADC) systemconsists of a RVSM capable advanced digitalair data computer with analog outputs for boththe FC 200 and FC 530 (Figure 16-29).




Figure 16-29. West Star Air DataComputer

Refer to Tables 16-3 and 16-4 for a partial listof cockpit indications should the ADC fail.The AZ-252 air data computer requires 115VAC electrical power and it is supplied from theleft AC bus.

See Figure 16-30 for the West Star AvionicsBlock Diagram.


A V I O N I C S

ADC FAILURE

EQUIPMENT INDICATION REMARKS

PILOT MACH/AIRSPEED – – M/ASI IS OPERATIVE, AURAL OVER-SPEED WARNING INOPERATIVE

Table 16-3. WEST STAR ADC FAILURE INDICATIONS CHART FOR FC 200 AIRCRAFT



PILOT ALTIMETER

PILOT VERTICAL SPEED

OR

**PILOT IDC VSI

DASHES IN ALL LCD DISPLAY

FIELDS, POINTER PARKS AT 8

– –

OR

VSI POINTER PARKED AT 0

ALTIMETER IS INOPERATIVE, USE

STANDBY ALTIMETER OR CROSS

SIDE ALTIMETER

PNEUMATIC VSI IS OPERATIVE

OR

IDC VSI IS INOPERATIVE


ADC FAILURE

EQUIPMENT INDICATION REMARKS

PILOT MACH/AIRSPEED

PILOT ALTIMETER

PILOT VERTICAL SPEED

OR

**PILOT IDC VSI

ALTITUDE ALERTER

**SAT/TAS/TAT

– –

DASHES IN ALL LCD DISPLAY

FIELDS, POINTER PARKS AT 8

– –

OR

VSI POINTER PARKED AT 0

0 DISPLAYED

DASHES DISPLAYED

M/ASI IS OPERATIVE, THE 300 KIAS

AURAL OVERSPEED WARNING

INOPERATIVE

ALTIMETER IS INOPERATIVE, USE

STANDBY ALTIMETER OR CROSS

SIDE ALTIMETER

PNEUMATIC VSI IS OPERATIVE

OR

IDC VSI IS INOPERATIVE

INOPERATIVE

INOPERATIVE

Table 16-4. WEST STAR ADC FAILURE INDICATIONS CHART FOR FC 530 AIRCRAFT



AUTOPILOT

MACH TRIM

SAT/TAS/TAT DASHES DISPLAYED

MACH TRIM ILLUMINATED

INOPERATIVE

VERTICAL MODES ARE INOPERATIVEVERTICAL MODES ARE CANCELED

LIMIT MACH NO. TO ≤MO.74


A V I O N I C S

LEAR 35/36 FC 200/FC 530 AFCS

AIR DATA SIGNALS AIR DATA LOGIC

LEARJET 35/36

AZ-252 AIR DATA COMPUTER

ALTITUDE ALERT LIGHT

PS

PTATC #2

IDC VSI*

*OPTIONAL

DISPLAY AND

SELECT DATABAROSET A429

PS

PT

BA-250

BAROMETRIC

INDICATOR

AL-800

ALTITUDE

ALERTER

AM-250

BAROMETRIC

INDICATOR

ALTITUDE

ALERT HORN

ATC #1



LEAR 35/36 FC 200/FC 530 AFCS

MACH TRIM

LANDING GEAR WARNING

OTHER EQUIPMENT

Figure 16-30. West Star Learjet 35/36 RVSM Avionics Block Diagram

Copilot Altimeter—AM-250Barometric Altimeter

The AM-250 barometric altimeter, installed inthe copilot posit ion, is a fully RVSM capablealtimeter with an integrated air data computer (Figure 16-32). It is a self-contained unit and is not connected to the AZ-252 air data com-

puter. It incorporates an analog/LCD display of baro-cor rected pressure altitude, baro-corrected displays, and an amber altitude alert light.

AL-800 Altitude Alerter

The AL-800 altitude alerter system provides both visual and aural signals for altitude aware-ness (Figure 16-33). The desired altitude isselected by slewing the displayed altitude tothe des i red va lue . Dur ing f l i gh t , when approaching the preselected altitude, at 1,000

ft prior to reaching that altitude, the amber al-titude alert light in each altimeter is illuminated and an aural alert is sounded. The altitude alertlight remains illuminated until the aircraft iswithin 200 ft of the selected altitude where itextinguishes.





A V I O N I C S

Standby Altimeter

The standby altimeter is a pure static altimeter (Figure 16-34). On FC 200 aircraft, i t is plumbed to the shoulder static ports that areheated anytime there is elect rical power on theaircraft.

Airspeed Static Valves

There are two airspeed static valves installed below the instrument panel on each side (Figure16-35). These manual valves are provided tosupply an alternate static source to the re-spective airspeed indicator. The valves havetwo positions—NORMAL and ALTERNATE.



On FC 530 aircraft, the standby altimeter is plumbed to the copilot Rosemount probe staticsystem; a vibrator installed in the standby al

When NORMAL is selected, the respectiveairspeed indicator receives static pressure

Figure 16-34. Standby AltimeterFigure 16-35. Right Airspeed Static Valve

Altitude Position and AirspeedCorrection Charts

The new Rosemount pitot static probe instal-lation changes the static source position error for the basic aircraft. New charts are included the the AFM Supplement, Document Number 30A04002, and have been developed from flight

test calibrations. The chart numbering systemin the supplement matches the basic aircraft

AFM to the maximum extent possible.

The new charts include aircraft weights up to19,600 lb to accommodate the increased grossweights that may be applicable to some Learjet35/35A and 36/36A aircraft altered by AvconDivision gross weight increase modificat ions.

The pilot and copilot altimeters are electrical,with the pilot BA-250 altitude display being

source correction curves incorporated into thedisplay, so the pilot and copilot altimeters havenegligible errors in cruise flight.

The standby altimeter is connected to the shoul-der ports and has a static source error. Whenusing the standby altimeter, the static source cor-rection factor must be applied to obtain the

prop er indicat ion. When an ai rspeed st ati csource valve, which is located under the in-strument panel, is selected to ALTERNATE, itapplies shoulder port static pressure to the applicable airspeed indicator. Airspeed indicator and Mach position correction chart values must be applied. These correction charts are located in the AFM Supplement . A cross-reference between Figures in the AFM Supplement and the

Learjet AFMs (AFM-019—Model 35/36) and AFM-102—Model 35A/36A wi th FC200Autopilot) is presented in the AFM Supplement




with the pilot BA-250 altitude display beingdriven by the AZ-252 air data computer and thecopilot having an AM-250 barometric altime-ter. The AZ-252 air data computer and the copilot AM-250 barometric altimeter have static

Autopilot) is presented in the AFM Supplement .In some cases, charts in the West Star Sup- plement are new and did not exist in the Learjet AFM .


A V I O N I C S




NAVIGATION SYSTEM

FC 200 Autopilot Aircraft

1a. The static ports for flight instrument op-eration are located:

A. In the unpressurized nose section

B. On the top and bottom of the pitot-static heads

C. Flush mounted on the left and rightsides of the fuselage nose section

D. On both sides of the aft fuselage

2a. The pilot controls the static pressuresource for the pilot flight instrument

FC 530 Autopilot Aircraft

1b. The static ports for flight instrument operation are located:

A. In the unpressurized nose sectionB. In the pitot-static heads

C. Flush mounted on the left and rightsides of the nose section

D. On both sides of the aft fuselage

2b. The pilot controls the static pressuresource for the pilot flight instrument op-

eration:A. Electrically with the STATIC PORT

switch


QUESTIONS



p g operation:

A. Electrically with the STATIC PORTswitch

B. Mechanically with the STATIC PORTswitch

switch

B. Mechanically with the STATIC PORTswitch

C. Electrically with the ALTERNATESTATIC SOURCE switch

D Mechanically with the ALTERNATE

CHAPTER 17

MISCELLANEOUS SYSTEMS

CONTENTS

Page

INTRODUCTION................................................................................................................. 17-1

GENERAL ............................................................................................................................ 17-1

OXYGEN SYSTEM ............................................................................................................. 17-2

Oxygen Cylinder ........................................................................................................... 17-3

Overboard Discharge Indicator...................................................................................... 17-3

Crew Distribution System 17 4




Crew Distribution System ............................................................................................. 17-4

Passenger Distribution System ...................................................................................... 17-6

DRAG CHUTE ..................................................................................................................... 17-8

ILLUSTRATIONS

Figure Title Page

17-1 Oxygen System...................................................................................................... 17-2

17-2 Oxygen Cylinder and Overboard Discharge Indicator .......................................... 17-3

17-3 OXYGEN PRESSURE Gage ................................................................................ 17-4

17-4 Crew Oxygen Mask ............................................................................................... 17-4

17-5 OXY-MIC Panel (Typical)...................................................................................... 17-5

17-6 Passenger Distribution System .............................................................................. 17-6

17-7 Passenger Mask...................................................................................................... 17-7

17-8 Drag Chute Components Location ........................................................................ 17-8




CHAPTER 17MISCELLANEOUS SYSTEMS




OXYGEN SYSTEM

The oxygen system components include ano x y g e n s t o r a g e c y l i n d e r a n d a s h u t o f f

valve-regula tor assembly, a n overboarddischarge indicator, an oxygen pressure gage,and distribution systems for the crew and pass engers . Figure 17 -1 depict s the oxygensystem.


1 7

MI S C E L L A N E O U S

S Y S T E M S

DISCHARGEINDICATOR

FILLERVALVE

OXYGENCYLINDER

030

PSI X 10

155195

200



PILOTMASK

PRESSURE REGULATORAND SHUTOFF VALVE

OXYGENPRESSURE

OXYGEN CYLINDER

The system is supplied with oxygen from astorage cylinder located in the right nose section on SNs 35-002 to 35-491 and 36-002to 36-050 (Figure 17-2). On SNs 35-492 and subsequent and 36-051 and subsequent, thecylinder is located in the dorsal fin. An optional

long-range installation incorporating two cylin-ders is available; location of the cylindersvaries.

Each oxygen cylinder has a storage capacityof 38 cu ft at 1,800 psi. The shutoff valve and pressure regulator assembly is at tached to thestorage cylinder and provides for pressureregulation, pressure indication, and servic-

ing. Oxygen pressure for the passenger and crew distribution system is regulated at 60–80 psi. The cyl inder, along with its shutoff valve

d l bl b h d h h

sition; this is a specified item on the exterior preflight inspection. The pilot should be awarethat if the oxygen cylinder shutoff valve isclosed, oxygen pressure will still be read onthe OXY PRESS gage in the cockpit. Duringthe interior preflight inspection, ensure that theshutoff valve is open by checking for oxygenflow through both crew oxygen masks, using

the 100% (EMER) position.

OVERBOARD DISCHARGEINDICATOR

The overboard discharge indicator (green blowoutdisc) (Figure 17-2) provides the pilot with a visual indication that there has not been an overpressure condition in the oxygen storage cylin-der. The disc blows out if the cylinder pressurereaches 2,700–3,000 psi, releasing all oxygenpressure System pressure should normally be be-




and regulator assembly, can be reached throughan access door. Under normal conditions, thisvalve should always be left in the on (open) po-

pressure. System pressure should normally be between 1,550 and 1,850 psi. The green blowoutdisc is located on the right side of the dorsal f inor the lower right side of the nose section.

OXYGEN PRESSURE Gage

The OXYGEN PRESSURE gage (Figure 17-3) provides a direct reading of oxygen cylin-der pressure, which is necessary to ensure thatan adequate supply of oxygen is aboard. Thegage is marked as follows:

• Yellow arc ................................0–300 psi

• Green arc ......................1,550–1,850 psi

• Red line ....................................2,000 psi

The gage is located on the pilot side panel onlate model aircraft; on early models, it ismounted on the instrument panel.

The crew masks (Figure 17-4) are stowed onthe pilot and copilot sidewalls. The mask oxy-gen lines are connected to quick-disconnect re-ceptacles located on the cockpit sidewalls.Optional oxygen-flow detectors may be in-stalled in the mask oxygen lines.

NOTE

Headsets, eyeglasses, or hats worn bycrewmembers may interfere with thequick-donning capabilities of theoxygen mask.


1 7


S Y S T E M S

* LATE MODELS

**EARLY MODELS



***

• T h e R o b e r ts h aw di l u te r - de m a nd mask/r egulator has two controls: the NORMAL–EMERGENCY selector and the 100% lever. With NORMAL selected,the regulator delivers diluted oxygen ondemand, up to 30,000 ft cabin altitude.Above 30,000 ft, the regulator delivers100% oxygen under a slight positive pres-

sure. Depressing the 100% lever will deliver 100% oxygen at any time. WithEMERGENCY selected (at any altitude)and the 100% lever depressed, the regu-lator delivers 100% oxygen and maintainsa slight positive pressure for respiratory protection from smoke and fumes.

• The Puritan-Bennett pressure demand mask/regulator incorporates a three-po-

sition selector knob labeled NORM, 100%,and EMER. With NORM selected, the reg-ulator delivers diluted oxygen on demand,

t 33 000 ft bi ltit d Ab

• The Scott ATO MC 10-15-02 mask, inthe normal pressure regulator positionwith the 100% lever extended, will de-liver diluted oxygen up to 30,000 ft cabinaltitude, 100% oxygen above 30,000 ftcabin altitude, and automatic pressure breathing above approximately 37,000ft cabin altitude. To obtain 100% oxy-

gen at any time, depress the 100% lever on the mask pressure regulator. WithEMERGENCY selected, the mask willdeliver 100% oxygen and maintain a positive pressure in the mask cup at alltimes for respiratory protection fromsmoke and fumes.

Each mask assembly includes a microphoneand has an electrical cord that is plugged into

the OXY-MIC jack on the respective OXY-MIC panel (Figure 17-5) on each side panel. To op-erate the mask microphone, the OXY-MIC




up to 33,000 ft cabin altitude. Above33,000 ft, the regulator automatically de-livers 100% oxygen. At 39,000 ft, it provides positive-pressure breathing. To obtain 100% oxygen at any time, 100%

b l d h l

switch must be in ON and the microphonekeyed, using the microphone switch on theoutboard horn of the control wheel. Commun-ication between crewmembers can be ac-complished by using the INPH function of the

PASSENGER DISTRIBUTIONSYSTEM

The passenger distribution system (Figure17-6) is used to provide oxygen to the pas-sengers in case of a pressurization system

failure or any other t ime that oxygen is required. Oxygen is available in the crew oxy-gen distribution lines whenever the oxygencylinder shuto ff valve is open; however, oxy-gen is not available to the passenger distribution syst em unti l required .


1 7


S Y S T E M S

FROMCREW

OXYGENSYSTEM

PASS OXY VALVENORM—OFF

NORMALLY OPEN (NORM)



PASS MASK VALVEMAN—AUTONORMALLY CLOSED (AUTO)

SOLENOID VALVEO C OS

Oxygen supply to the passenger system is con-trolled with three valves. Two valves are man-ually operated with control knobs on the pilotsidewall, and the third is solenoid-operated byan aneroid switch. The manually controlled PASS OXY valve is normally in the NORM(open) position, which allows oxygen up and to the manually controlled PASS MASK valve

and to the aneroid-controlled solenoid valve.Oxygen can be admitted to the passenger dis-tribution system through either of these pas-senger mask valves, both of which are normallyclosed.

With the PASS OXY valve in the OFF (closed) posi tion, oxygen will not be available to the passenger distribution system in any event .

This position may be used only when no pas-sengers are being carried.

With the PASS OXY valve in the NORM (open)

In the event of aircraft electrical failure, au-tomatic deployment of the passenger masks isnot possible. The oxygen solenoid valve re-quires DC power through the OXY VAL cir-cuit breaker on the left essential bus for automatic mask deployment.

With the PASS OXY valve in the NORM (open)

position, rotating the PASS MASK valve fromAUTO to MAN admits oxygen into the pas-senger distribution system and causes the pas-senger oxygen masks to drop. This positioncan be used to deploy the passenger masks atany altitude, but will not cause the cabin over-head lights to illuminate.

The passenger oxygen masks (Figure 17-7)

are stowed in compartments in the convenience panels above the passenger seats. The compar tments may contain as many as three masks,depending on the aircraft seating configuration




With the PASS OXY valve in the NORM (open) posi ti on , oxygen wi ll be au tomati ca lly ad -mitted to the passenger distribution systemthrough the aneroid-control led solenoid valveif the cabin reaches 14,000 ±750 ft. The

id i h h l id l d

depending on the aircraft seating configuration.There will be at least one spare mask.

The passenger mask storage compartmentdoors are held closed by latches. When oxy-

i d i d i h di ib i

masks fall free and are available for passen-ger use. As the passenger pulls down on hismask to don it, an attached lanyard withdrawsa pin from the supply valve that releases oxy-gen into the mask breather bag at a restricted,constant flow rate. The rebreather bag mayseem to inflate slowly, but this is normal.When inhaling, 100% oxygen is delivered to

the mask cup. The breath is then exhausted intothe rebreather bag.

Should the doors be inadvertently opened fromthe cockpit, oxygen pressure must be bled from the passenger distribution system beforethe masks can be restowed. This is accomplished by pulling one of the passenger mask lanyards after ensuring that the PASS MASK

valve is closed (AUTO). If the doors open dueto malfunction of the solenoid-operated valve,the PASS OXY valve must be turned off topermit stowage of the passenger masks

DRAG CHUTE

GENERAL

The optional drag chute may be used to shortenstopping distances. The greatest decelerationrate is produced at the highest speed; however,

the chute is stil l effective at speeds below 60kt. The chute is stored in a removable canis-ter that is mounted inside the tail cone accessdoor. The canister lid is released from the can-ister when the drag chute handle is pulled, al-lowing the pilot chute to deploy. The pilotchute then pulls the main chute canopy out of the canister.

The main chute riser attaches to the aircraft atthe chute control mechanism just forward of the canister (Figure 17-8). The loop at the end of the main riser slips over a recessed metal


1 7


S Y S T E M S



permit stowage of the passenger masks.

The compartment doors can be opened manually for mask cleaning and servicing.

of the main riser slips over a recessed metal pin that is held in posit ion by spring pressurewhen the drag chute handle is stowed. There-fore, if the chute should inadvertently deploy(handle in stowed position), the main chutei ill li f f h i d f

When the drag chute handle is pulled, the pinis mechanically locked in position to retain thechute riser while the mechanical canister con-trol mechanism operates to release the canis-ter lid, thereby deploying the chute.

The drag chute can be used:

• When landing on a wet or icy runaway

• During any landing emergency involvingno-flap hydraulic or brake failure, or lossof directional control

• During takeoff if the decision is made toabort

Do not deploy the drag chute under the fol-lowing conditions:

• In flight

• If the nose gear is not on the ground

Wh h i di d i d i b

SQUAT SWITCHSYSTEM

GENERAL

Some aircraft systems operate only on theground while others operate only in the air. The

squat switch system is designed to providethe necessary ground or airborne signals tothese systems. The squat switch system con-sists of two squat switches—one on each mainlanding gear strut scissors—and a relay box lo-cated under the cabin floor. When the aircraftis on the ground, and the main landing gear struts are compressed, the squat switches closeto provide a ground mode signal. When the air-craft lifts off the ground and the main landinggear struts extend, the squat switches open,which interrupts the ground mode signals,




• When the indicated airspeed is above150 kt

• With thrust reversers deployed

thereby shifting to air mode.

SQUAT SWITCHES

Each squat switch provides ground or air sig

• Gear control valve

° The switches disable the gear-upsolenoid on the ground to prevent inadvertent landing gear retraction.

° Either squat switch in ground modewill disable the gear-up solenoid. Bothsquat switches must be in the air mode

to allow landing gear retraction.• Squat switch relay box

° Either squat switch in the ground mode puts the relay box in ground mode.

° Both squat switches must go to air mode to put the relay box in air mode.

The position of the SQUAT SW circuit breaker

has no effect on landing gear, antiskid, or stallwarning system operation. These systems receive signals directly from the squat switchesas explained previously

• Cabin pressurization

• Safety valve vacuum solenoid closes inthe air (SNs 35-099 and subsequent and 36-029 and subsequent only)

• Amber CAB ALT light (if installed) isdisabled on the ground

• Control module solenoids shift from

ground to air mode

• Amber TO TRIM light—Disabled in theai r

• Windshield heat system—Shifts fromground to air mode (see Chapter 10, Iceand Rain Protection, for additional in-formation)

• Hourmeter and Davtron clock flight timefunction (if installed)—Disabled on theground

h i O l h


1 7


S Y S T E M S



as explained previously.

SQUAT SWITCH RELAY BOX

The squat switch relay box is necessary because

• Mach trim test—Operates only on theground

• Thrust reversers—Operate only on theground

1. During preflight, the pilot can determineif the oxygen bottle is turned on by:

A. Reading the pressure indicated on theoxygen pressure gage in the cockpit

B. Selecting 100% on the mask regula-

tor and taking several deep breathsthrough the mask

C. Placing the OXY-MIC switch to theOXY position

D. Visually checking for the green flowindicator on the mask supply hose

2. With the PASS OXY valve in the NORM

posit io n, se le ct in g MA N on th e PASSMASK valve:

A. Causes passenger masks to drop and turns on the cabin overhead lights

4. The OXY PRESS gage reads:

A. Direct pressure of the cylinder

B. Electrically derived system high pres-sure

C. Direct pressure of the pilot supplyline

D. Electrically derived system low pres-sure

5. The maximum demonstrated crosswind component for drag chute deployment is:

A. 10 kt

B. 15 kt

C. 20 kt

D. 25 kt


QUESTIONS



turns on the cabin overhead lights

B. Prevents oxygen from entering the passenger oxygen d istr ibution l ines

C. Disarms the 14,000 ft cabin aneroid

D Ad i t t th

6. The drag chute is deployed by:

A. Squeezing the control handle

B. Rotat ing the contro l handle fu llyclockwise and pulling it up to its full

CHAPTER 18

MANEUVERS AND PROCEDURES

CONTENTS

Page

INTRODUCTION................................................................................................................. 18-1

GENERAL ............................................................................................................................ 18-1

ABBREVIATIONS ............................................................................................................... 18-1

STANDARD OPERATING PROCEDURES........................................................................ 18-2

General .......................................................................................................................... 18-2

Responsibilities.............................................................................................................. 18-2




Checklist Procedures ..................................................................................................... 18-2

Briefing Guides ............................................................................................................. 18-2

Engine Failure Above V1 Speed.................................................................................. 18-11

Steep Turns.................................................................................................................. 18-12

Unusual Attitude Recovery—Nose High, Low Speed................................................ 18-14

Unusual Attitude Recovery—Nose Low, High Speed ................................................ 18-15

Slow Flight .................................................................................................................. 18-16Approach to Stall......................................................................................................... 18-18

Emergency Descent..................................................................................................... 18-20

Visual Traffic Pattern—Two Engines ......................................................................... 18-21

Visual Traffic Pattern—Single Engine ....................................................................... 18-21

Flaps Up Landing ........................................................................................................ 18-22

Precision Instrument Approach................................................................................... 18-23




pp

Nonprecision Instrument Approach ............................................................................ 18-24

Circling Instrument Approach..................................................................................... 18-26

1 8

MA N E U V

A N D P R O C E D

ILLUSTRATIONS

Figure Title Page

18-1 Normal Takeoff...................................................................................................... 18-9

18-2 Rejected Takeoff.................................................................................................. 18-10

18-3 Engine Failure at or above V1 Speed................................................................... 18-11

18-4 Steep Turns .......................................................................................................... 18-13

18-5 Unusual Attitude Recovery—Nose High, Low Speed ........................................ 18-14

18-6 Unusual Attitude Recovery—Nose Low, High Speed......................................... 18-15

18-7 Slow Flight........................................................................................................... 18-16

18-8 Slow Flight—Takeoff Configuration................................................................... 18-17

18-9 Slow Flight—Landing Configuration.................................................................. 18-17




g g g

18-10 Approach to Stall—Clean Configuration ............................................................ 18-18

18-11 Approach to Stall—Takeoff Configuration ......................................................... 18-19


INTRODUCTION

The general pilot information in this chapter is intended to supplement and expand upon

information in other sources. It is not intended to supersede any off icial publication. If there is any conflict between the information in this chapter and that in any offic ial pub-lication, the information in the official publication takes precedence.

GENERALGeneral pilot information includes Standard Operating Procedures and the maneuvers normallyencountered during Learjet training and operations. The following abbreviations are used in this

chapter.

ABBREVIATIONS

CHAPTER 18MANEUVERS AND PROCEDURES



ABBREVIATIONS

AFM Airplane Fl ight Manual

AGL Above Ground Level

MMO Mach, Maximum Operational

MSL Mean Seal Level

STANDARD OPERATINGPROCEDURES

GENERAL

Standard Operating Procedures (SOPs) areused to supplement the information in the

AFM and Federal Air Regulations. Adherenceto SOPs enhances individual and crew situa-tional awareness and performance. SOPs mayinclude assignment of responsibilities, brief-ing guides, and procedures to be followed dur-ing specif ic segments of flight. The SOPs inthis section are not intended to be mandatoryor to supersede any individual company SOPs.They are simply provided as examples of good

operating practices.

RESPONSIBILITIES

CHECKLIST PROCEDURES

Normally, the PF initiates all checklists.However, if the PM thinks a checklist should be accomplished, and the PF has not ca lled for it, the PM should prompt the PF. For example,“Ready for the Approach checklist, Captain?”

FlightSafety International recommends theuse of the checklist challenge and responseconcept. Using Normal Procedures checklists,the PM challenges the PF and the PF responds.Using Abnormal or Emergency Procedureschecklists, the PM challenges the PF and, asa memory aid, also gives the checklist item re-sponse. The PF then responds.

The PF may elect to have the PM accomplishsome Abnormal or Emergency Procedurechecklists on the PF’s command. In this case,the PM gives the checklist item and response.Th PF li ith th d th PM


A



PIC —The Pi lot in Command is designated by the company for fli gh ts re qu ir ing morethan one pilot. Responsible for conduct and safety of the flight Designates pilot flying

The PF replies with the response, and the PMaccomplishes the action.

When a checklist has been completed, the PMh h kli i l d h h / h

1 8

MA N E U V

AN D P R O C E D

be low should be used when flying wi th unfa-miliar crewmembers or any other time thePIC believes they are necessary.

It should be noted that many of these items can,and should, be briefed well before engine star t.Many of them can be discussed before arriv-ing at the aircraft.

Pretakeoff Briefing

The pretakeoff briefing should address thefollowing items:

• Type of takeoff; rolling or standing, flapsetting, etc.

• Review takeoff data to include power

setting and speeds

• Procedures to be used in the event of anemergency before or after V1 speed including emergency return procedures

• Special procedures to be used duringthe approach (i.e., circling approach procedures, interception of a radial froman arc, VDP)

• Altitudes of IAF, FAF, stepdowns, sec-tor and obstacles

• Minimums (DH, MDA), HAT, HAA,

radio altimeter setting• Missed approach point and procedures,

timing to MAP/VDP

• Radio (COM/NAV) setup desired

• Anti-icing requirements

• Specific PM duties and callouts (seeApproach Procedures, this chapter)

• The procedure for transitioning to visualflight

• A request for “Any questions?” directed




including emergency return procedures

• Headings and altitudes to be flown dur-ing the departure including restrictions,if any

q y qto all cockpit crewmembers

At the completion of the Approach briefing,th PF “A h b i fi

With positive rate of climb, the PF calls“Positive rate, gear up, yaw damper on.” ThePM positions the gear handle to up and calls“Gear selected up, yaw damper engaged.” ThePM monitors the gear while it is retractingand reports “Gear up,” when retraction is complete.

Before VFE (V2 + 30 kt, minimum), the PFcalls, “Flaps up, After Takeoff checklist.” ThePM positions the flap handle to up and calls“Flaps selected up.” The PM monitors theflaps while they are retracting and reports“Flaps up,” when retraction is complete. PMaccomplishes the After Takeoff checklist.

CLIMB AND CRUISE

PROCEDURESThe PM announces all assigned altitudes and sets them in the altitude alerter. The PM also

by the pilot assuming control. Specif ic targetvalues are provided to the pilot assuming con-trol. For example, the PF announces, “Take theaircraft for a minute. We’re climbing at 250knots to 7,000 on a vector to the 045 radial.”PM acknowledges, “I’ve got the aircraft,climbing at 250 to 7,000 on this heading untilintercepting the 045 radial.”

APPROACH PLANNING

Approach planning and briefing should be ac-complished during cruise. Review hazardousterrain, MEAs, and minimum sector altitudes.Complete and review performance data to in-clude VREF speed, landing distance, approachclimb speed, and power setting.

The PF directs the PM to obtain destinationweather or obtains it himself. If the PM ob-tains the weather, the PF normally assumesA C i i hil h i b


A



calls out 1,000 ft above—or below—all as-signed altitudes and altitude restrictions. Thesecalls normally are made by stating the exist-ing altitude and the assigned altitude or re-

ATC communications while the PM is ob-taining weather. In either case, after check-ing weather, the pilot who did so briefs theother pilot on the des tination weather, the ex-

1 8

MA N E U V

AN D P R O C E D

The Descent checklist should be started be-fore, or early in, the descent to permit proper windshield heat and pressurization systemoperation.

Descent below FL 180 will not be started beforeobtaining a local area altimeter setting.

DESCENT PROCEDURESThe same procedures used during climb and cruise are used during descent. The PM ac-complishes the Descent checklist, as directed by the PF, and makes al ti tude ca llouts to in-clude the transition level and 10,000 ft.

APPROACH PROCEDURES

The PF initiates the Approach checklist whendescending out of 18,000 ft or when within 50miles of the destination airport. The checklist

if an altitude change is required. For example,“Time’s up, right turn now to 225° and cleared down to 3,000.”

Approaching the final approach course, the PMmonitors the CDI or bearing pointer and repor ts “CDI alive,” or “Within 5° of the inbound course.”

Established on final approach, the PF callsfor flaps 20°, slows the aircraft to VREF + 20kt (minimum), and begins a descent, if nec-essary. Prior to the FAF, the PF calls “Gear down, Before Landing checkl ist.” The PM ex-tends the landing gear, completes the BeforeLanding checklist up to flaps down and reports,“Before Landing checklist complete to full

flaps.”

Over the FAF, on a two-engine, straight-inapproach, the PF calls for flaps 40°, slows theaircraft to V (minimum) and begins a de




pis accomplished so as to not interfere with thevisual lookout for other traffic.

Configuration changes during the approach

aircraft to VREF (minimum), and begins a de-scent. (For a single-engine, or circling approach, the flaps remain at 20°.) The PM begins timing, if necessary, extends the flaps

d l h f di h kli

“Minimum descent altitude” or “Decisionheight.” The PM also reports visual contactwith the ground such as, “Visual contact, norunway yet,” “Approach ligh ts in sight a t 11o’clock,” or “Runway in sight straight ahead.”If the PM does not call, “Runway in sight,”at the MAP or DH, or repor ts, “No contact,”the PF will initiate a go around.

Approaching minimums, or the missed approach point, the PF begins cross-checkingoutside the aircraft for visual references. Whensatisfied that visual references are adequatefor landing, the PF announces, “I’m going visual,” or “Going outside.” At this poin t, thePM directs his attention primarily inside theaircraft, while cross-checking outside, and calls airspeed, descent rate, and altitude. The purpose is to provide the PF, verbally, the sameinformation he/she would have if still on in-struments.

MANEUVERS

GENERAL

This section contains a description of most of the maneuvers that are likely to be encountered during Learjet training and operational flying.While there is always more than one way to fly

an aircraft, these procedures have been de-veloped over many years of Learjet opera-tions. They have proven to be safe, efficient,and readily manageable. These procedures areconsistent with the AFM . However, if a con-flict should develop between these proceduresand those in the AFM , the AFM proceduresshould be used.

PERFORMANCE STANDARDS

The performance standards in Table 18-1should be maintained during all Learjet flight


1 A



Airspeed should be called as plus or minusVREF, descent rate as up or down and altitudeabove the ground. For example, “Plus 5, down

00 100 f ” i di h i d i

should be maintained during all Learjet flightoperations.

MINIMUM MANEUVERING

1 8

MA N E U V

AN D P R O C E D


Steep Turns

Bank angle: 45°, ±5°

Altitude: ±100 ftAirspeed: ±10 KIAS

Heading: ±10°

Approach to Stall

Initiate recovery at stick shaker onset.

Recover with minimum altitude loss.

Holding

Altitude: ±100 ft

Airspeed: ±10 kt

Instrument Approaches

Initial: Altitude: ±100 ft

Table 18-1. PERFORMANCE STANDARDS



Airspeed: ±10 kt

Final: Airpseed: -0, +5 kt

Localizer: ±1 dot

Spoilers deployed ....................VREF + 40 KT

Flaps up......................................VREF + 30 KT

Flaps 8° ......................................VREF + 20 KT

Flaps 20° ....................................VREF + 10 KT

Flaps 40° ..................................................VREF

POWER SETTINGS

Actual power settings vary depending upon thetemperature, pressure altitude, and aircraftgross weight. The following target settingsare approximate, but may be used to providea starting point to determine the actual power setting:

• Below 10,000 MSL—60% N1 to main-tain 200 KIAS, 70 to 75% N1 to main-tain 250 KIAS

B 10 000 MSL d FL 250 75

specified in the AFM . If the runway availableis at least 10% longer than the planned take-off distance, a rolling takeoff may be used. The procedures are the same except for a standingtakeoff, power is set before brake release. For a rolling takeoff, the brakes are released be-fore the power is set. During a rolling takeoff,takeoff power must be set be fore the runwayremaining equals the takeoff distance.

Normally, before VFE (V2 + 30 kt minimum),the flaps are retracted and the After Takeoff checklist is accomplished. However, if trafficconditions warrant, the Afte r Takeoff check-list may be delayed until the aircraft is clear of local traffic.

Approaching 200 kt, the PF should adjust power and pitch attitude if necessary, to main-tain 200 kt or less within the ATA (Class DAirspace). For passenger comfort and ease of aircraft control, it is recommended that the


1 A N



• Between 10,000 MSL and FL 250—75to 80% N1 to maintain 250 KIAS

TAKEOFF

aircraft control, it is recommended that the pi tch a tt itude not exceed 20° noseup.

The maximum continuous climb power settingi i bl d di d

1 8

MA N E U V

ND P R O C E D


BEFORE VFE

BEFORE TAKEOFF

1. PF HOLDS BRAKES AND ADVANCES POWER2. PM SETS TAKEOFF POWER

INITIAL AIRSPEED INDICATION

1. PM CALLS "AIRSPEED"2. PF DISENGAGES NOSEWHEEL STEERING **

POSITIVE RATE OF CLIMB

1. PF CALLS "GEAR UP, YAW DAMPER ON"2. PM RETRACTS THE LANDING GEAR

AND ENGAGES THE YAW DAMPER

APPROACHING 200 KIAS

1. PF ADJUSTS PITCH AND POWER

TO REMAIN BELOW 200 KIAS IN

AIRPORT TRAFFIC AREA (ATA)

CLASS D AIRSPACE

CLEAR OF ATA/CLASS D AIRSPACE

1. PF SETS MAXIMUM CONTINUOUS CLIMB POWER AND ACCELERATES AIRPLANE TO 250 KIAS



BEFORE VFE(V2 + 30 KT MIN)80 KIAS

1. PM MONITORS AND ADJUSTS TAKEOFF POWER

1. PF CALLS "FLAPS UP, AFTER TAKEOFF CHECKLIST"2 PM ACCOMPLISHES AFTER

ENGINE FAILURE BELOW V 1SPEED

If an engine fails below V1 speed (Figure18-2), the takeoff must be aborted. The PF simultaneously reduces power to idle, appliesmaximum braking and deploys the spoilers.

The drag chute or thrust reversers (if installed)are deployed if necessary.

Takeoffs may be aborted for malfunctionsother than engine failure; however, the same procedures should normally be used.


1 8

A N

TO 80 KIAS

INITIAL TAKEOFF ROLL

1. STANDING OR ROLLING

TAKEOFF PROCEDURES

INITIAL AIRSPEED INDICATION

1. PM CALLS "AIRSPEED"2. PF DISENGAGES NOSEWHEEL STEERING

1. PM MONITORS AND ADJUSTS

ABORT TAKEOFF

1. POWER—IDLE2. WHEEL BRAKES—APPLIED3. SPOILERS—DEPLOYED4. DRAG CHUTE/THRUST REVERSERS (IF INSTALLED)—DEPLOY IF NECESSARY



8

MA N E U V

ND P R O C E D

TAKEOFF POWER

ENGINE FAILURE ABOVE V 1SPEED

If an engine fails above V1 speed (Figure18-3 ), the takeoff is normally continued. ThePF maintains directional control with aileronsand rudder and keeps the nosewheel on the run-way until reaching rotate speed. After lif toff,

the initial climb is made at V2 speed with take-off flaps until the aircraft is clear of obstaclesor, if there are no obstacles, to 1,500 f t AGL.The PF then accelerates the aircraft to V2 + 30kt (minimum) and directs the PM to retract theflaps. The PF then accelerates the aircraft to

single-engine climb speed (normally 200 kt)and climbs to the assigned altitude.

At a safe altitude above the g round (normally,no lower than 400 ft), the memory items for the Engine Failure/Fire Shutdown in Flightchecklists are completed. The rest of theEngine Failure During Takeoff checklist alongwith the Engine Failure/Fire Shutdown inFlight checklists (as appropriate), and theAfter Takeoff checklist are normally completed at , or above, 1,500 ft AGL. The crewthen elects to obtain clearance to return to thedeparture airport for landing or proceeds to analternate airport.


AT SAFE ALTITUDE

1. ENGINE FAILURE DURING TAKEOFF CHECKLIST



INITIAL TAKEOFF ROLL

1. STANDING OR ROLLINGTAKEOFF PROCEDURES

INITIAL CLIMB

1. V2 SPEED2. TAKEOFF FLAPS

STEEP TURNS

Steep turns (Figure 18-4) are used to build confidence in the aircraft and improve instru-ment cross-check. They may be accomplished at any altitude above 5,000 ft AGL. The higher the altitude, the more difficult the maneuver is to perform correctly. Steep turns are ac-complished without flight director steeringcommands since the flight director does notcommand 45° of bank.

Power must be increased approximately 2% N1to maintain airspeed during steep turns. Theaircraft should be kept in trim and the bank angle should be held constant. If altitude cor-rections are necessary, they should be made in pi tch only. It is not necessary to shal low the bank to cl imb during a steep turn in a Learjet .

Steep turns of at least 180°, preferable 360°,should be practiced in each direction.


1 8

A N



8

MA N E U V

DP R O C E D


ENTRY

1. ROLL INTO 45˚ OF BANK

2. INCREASE POWER TO MAINTAIN AIRSPEED

3 TRIM AS REQUIRED



3. TRIM — AS REQUIRED

UNUSUAL ATTITUDERECOVERY—NOSE HIGH, LOWSPEED

Recovery from a nose-high, low-speed unusualattitude (Figure 18-5) should be made whilemaintaining positive G forces and withoutstalling the aircraft. It is accomplished by

increasing power while simultaneously in-creasing the angle of bank, not to exceed 90°,to allow the nose of the aircraft to descend tothe horizon without negative G forces. The at-titude indicator should be used during the re-covery and the angle-of-at tack indicator cross-checked to maintain the pointer in thegreen band.


1 8

A N D

NOSE THROUGH THE HORIZON

1. AIRSPEED > 180 KIAS2. ROLL WINGS LEVEL

3. REDUCE POWER — AS REQUIRED



8

MA N E U V

DP R O C E D PITCH AND ROLL ATTITUDE DETERMINED

UNUSUAL ATTITUDERECOVERY—NOSE LOW, HIGHSPEED

Recovery from a nose-low, high-speed un-usual attitude (Figure 18-6) should be madewith minimum loss of altitude while keepingthe airspeed below VMO or MMO . It is ac-

complished by simultaneously reducing power to idle and rolling the wings level. When the bank is less than 90°, elevator and pi tch t rim

(if required) are used to raise the nose to thehorizon. Spoilers should not be used during re-covery from a nose low unusual attitude.

During training, nose-low, high-speed unusualattitudes are always presented so the aircraftcan be recovered without exceeding any lim-itations. However, during recovery from anactual, inadvertent, nose-low, high-speed un-

usual attitude, an overspeed condition maydevelop. In this case, the overspeed recovery procedures in the AFM should be used.


PITCH AND ROLL ATTITUDE DETERMINED



1. SIMULTANEOUSLY ROLL WINGS LEVEL AND REDUCE

POWER TO IDLE

2. WHEN BANK ANGLE IS LESS THAN 90˚:

ELEVATOR AND PITCH TRIM NOSE-UP AS REQUIRED

SLOW FLIGHT

Slow flight is used to develop the pilot’s senseof feel for the aircraft’s low-speed handlingcharacterist ics and to improve the pilot’s coor-dination and instrument cross-check. Slow flightis accomplished in the clean, takeoff, and land-ing configurations (Figures 18-7, 18-8 and18-9), and is normally accomplished between

12,000 and 15,000 ft MSL. Slow flight should not be accomplished below 5,000 ft AGL.

Slow flight may be practiced while maintain-ing a constant altitude and heading or whilemaintaining a constant altitude and makingturns to preselected headings. Slow flight mayalso be practiced while making constant rateclimbs and descents to preselected altitude.Slow flight practice may be terminated by a re-

covery to normal cruise or an approach to stall.


1 8

A N D

ENTRY

1. GEAR—UP

2. FLAPS—UP3. AIRSPEED—VREF + 20 KTDURING SLOW FLIGHT

1. MAINTAIN ALTITUDE AND HEADING

ALTITUDE: 12,000' - 15,000' (AIRPLANE) AS DESIRED (SIMULATOR)



MA N E U V

DP R O C E D


OPTIONAL

1. 15° BANK TURNS TO PRESELECTED HEADINGS2. CONSTANT RATE CLIMBS

AND DESCENTS

ENTRY

1. GEAR—UP OR DOWN2. FLAPS—8° OR 20°3. AIRSPEED—VREF + 10 KT (FLAPS 8) —VREF (FLAPS 20)

DURING SLOW FLIGHT

1. MAINTAIN ALTITUDE AND HEADING




AND DESCENTS

Figure 18-8. Slow Flight—Takeoff Configuration

APPROACH TO STALL

Approaches to stall are accomplished in theclean, takeoff, and landing configurations(Figures 18-10, 18-11, and 18-12), and arenormally accomplished between 12,000 and 15,000 ft MSL. Approaches to stalls should not b e a c co mpl i s h ed be low 5 , 000 f t AGL.Approaches to stalls may be made from level

or turning flight with 15 to 30° of bank.Approaches to stalls may also be combined with slow flight practice. All recoveries aremade with power and minimum loss of altitude.

Approach to stall recovery is initiated at thefirst indication of an impending stall. This indication is provided by the stick shaker and stall warning annunciator ligh ts that activate

as the angle-of-attack indicator needle movesinto the yellow band.

Power shou ld be advance d to takeoff power.However, the angle-of-attack indicator should be moni to red and the p it ch at ti tude re duced,if necessary, to keep the needle at the line be -tween the green and yellow bands.

To set takeoff power in minimum time, the PFshould move the thrust levers smoothly forward to the stop. The PM should monitor and adjustthe power setting if necessary. Approaches tostall from the landing conf iguration are nor-mally terminated by a simulated missed approach (Figure 18-12).


1 8

A N D

FIRST INDICATION OF STALL

SIMULTANEOUSLY:




MA N E U V

P R O C E D

BEFORE ENTRY

1. DETERMINE V2 FOR AIRPLANE WEIGHT

SIMULTANEOUSLY: 1. ROLL WINGS LEVEL 2. LOWER PITCH ATTITUDE TO REDUCE ANGLE OF ATTACK


FULL RECOVERY

1. ADJUST PITCH ATTITUDE TO MAINTAIN ALTITUDE2. ADJUST AIRPSEED TO 180 KIAS,

FIRST INDICATION OF STALL

SIMULTANEOUSLY: 1. ROLL WINGS LEVEL 2. LOWER PITCH ATTITUDE TO REDUCE ANGLE OF ATTACK 3. THRUST LEVERS TO TAKEOFF POWER 4. ACCELERATE 5. MINIMIZE LOSS OF ALTITUDE

START RECOVERY

1. AIRSPEED INCREASES2. ABOVE V2 —FLAPS 203. POSITIVE RATE—GEAR UP4. V2 + 30 KT—FLAPS UP

ENTRY

1. REDUCE POWER TO 65% N12. SIMULTANEOUSLY PITCH UP' TO 20° AND ROLL INTO 20° BANK

BEFORE ENTRY

1. DETERMINE V2 FOR AIRPLANE WEIGHT WITH FLAPS AT TAKEOFF SETTING2. GEAR—UP OR DOWN3. FLAPS—TAKEOFF SETTING




JUS S O 80 S, OR AS INSTRUCTED

Figure 18-11. Approach to Stall—Takeoff Configuration

EMERGENCY DESCENTEmergency descents are accomplished in accordance with AFM procedures as shownin Figure 18-13. The PF should accomplish thechecklist memory items and allow the aircraftto pitch down to a 10 to 15° nosedown pitchattitude. This pitch attitude is maintained until

the aircraft accelerates to MMO /VLE . Thenthe pitch attitude is adjusted to maintainMMO/VLE.

After the emergency descent has been estab-lished, the crew should determine the desired level-off altitude.


1 8

M

A N D

ENTRY

1. CREW OXYGEN MASKS — ON

2. POWER — IDLE



MA N E U V

PR O C E D

3. AUTOPILOT — DISENGAGED

4. SPOILERS — EXTENDED5. LANDING GEAR — DOWN

(BELOW M /V )

DESCENT

1 MAINTAIN PITCH ATTITUDE

VISUAL TRAFFIC PATTERN—TWO ENGINES

A two-engine visual traffic pattern is shownin Figure 18-14. The airspeeds indicated on thediagram are minimums. Traffic pattern altitudefor jet aircraft is normally 1,500 ft AGL.During gusty wind conditions, 1/2 the gustvelocity should be added to V

REF

on final approach. If a c rosswind exis ts, f inal approachshould be flown with a drift correction angle(crab) to maintain alignment with the runwaycenterline. Approaching touchdown, rudder should be applied to align the aircraft withthe runway centerline and the upwind winglowered with aileron to prevent drift.

VISUAL TRAFFIC PATTERN—SINGLE ENGINE

A single-engine visual traffic pattern is flownexactly the same as a two-engine pattern ex-cept for the flap setting on f inal approach. For single-engine approach, maintain flaps 20°and VREF + 20 kt (minimum) when maneu-vering. When established on final approach, setflaps 20° and VREF + 10 kt (minimum).


ENTRY LEG

1. GEAR AND FLAPS—UP2. AIRSPEED—VREF + 40 KT (MIN)3. APPROACH CHECKLIST—COMPLETE



FINAL APPROACH **

1 FLAPS 40°

FLAPS UP LANDINGThe corrected landing distance for a flaps uplanding (Figure 18-15) is determined by mul-tiplying the normal landing distance by 1.35.Considerations should be given to reducing theaircraft weight, if poss ible, to lower the land-ing speed and reduce landing distance, if theavailable runway length is marginal.

To avoid excessive floating during the landingflare, the PF should establish the landing atti-tude as power is reduced to idle, maintain theat t i tude, and al low the aircraft to touchdown. The use of the drag chute, or thrust re-versers, (if installed) is recommended duringa flaps up landing.


1 8

M

A N D P

ENTRY LEG

1. GEAR AND FLAPS — UP

2. AIRSPEED — VREF + 40 KT (MIN)

3. APPROACH CHECKLIST — COMPLETE

FINAL APPROACH

1. AIRSPEED — VREF + 30 KT (MIN)

2. YAW DAMPER — DISENGAGED

BEFORE TOUCHDOWN



MA N E U V

PR O C E D

PRECISION INSTRUMENT APPROACH

A typical, precision instrument approach isshown in Figure 18-16. All accepted instrumentflying procedures and techniques should beused while making instrument approaches inthe Learjet.

Two-engine, precision approaches should beflown with a stabilized airpseed and config-uration from the f inal approach f ix (FAF) inbound. Single-engine, precision approachesshould be flown with flaps 20° at VREF + 20kt (minimum) if maneuvering is required.When established on final approach, set flaps20° and VREF + 10 kt (minimum).


APPROACHING INITIAL APPROACH FIX (IAF)

1. GEAR AND FLAPS—UP2. AIRSPEED—VREF + 40 KT (MIN)3. APPROACH CHECKLIST—COMPLETE

IAF OUTBOUND *

1. FLAPS—8°2. AIRSPEED—VREF + 30 KT (MIN)3. DESCEND, IF REQUIRED

ON COURSE INBOUND



1. FLAPS—20°2. GEAR—DOWN3. AIRSPEED—VREF + 20 KT (MIN)

NONPRECISION INSTRUMENT APPROACH

A typical, nonprecision instrument approachis shown in Figure 18-17. All accepted in-strument flying procedures and techniquesshould be used while making instrument approaches in the Learjet.

Two-engine, nonprecision approaches should be flown with a stabil ized airspeed and con-f iguration from the final approach f ix (FAF)i n b o u n d . S i n g l e - e n g i n e , n o n p re c i s i o n approaches should be flown with flaps 20° atVREF + 20 kt (minimum) if maneuvering is re-quired. When established on final approach,set flaps 20° and VREF + 10 kt (minimum).


1 8

M

A N D P



MA N E U V

PR O C E D



1. GEAR AND FLAPS—UP2. AIRSPEED—VREF + 40 KT3. APPROACH CHECKLIST—COMPLETE

IAF OUTBOUND *

1. FLAPS—8°

2. AIRSPEED—VREF + 30 KT3. DESCEND, IF REQUIRED

ON COURSE INBOUND

1. FLAPS—20°2. GEAR—DOWN3. AIRSPEED—VREF + 20 KT4. BEFORE LANDING CHECKLIST— COMPLETE TO FLAPS 40°



FINAL APPROACH FIX **

CIRCLING INSTRUMENT APPROACH

Any instrument approach that requires a head-ing change of 30° or more to line up with thelanding runway is a circling approach. Anidentifiable part of the airport must be dis-tinctly visible to the pilot during the circlingapproach, unless the inability to see an iden-

tif iable part of the airport results only from anormal bank of the aircraft. The circling MDAand weather minima to be used are those for the runway to which the approach is flown.

The Learjet is an approach category C air-craft. However, category D minimums should be used if the ai rcraft will be maneuvered atspeeds over 141 kt (the minimum for category

D aircraft) during the circling approach.

There are two types of circling approaches.The firs t type of circling approach positionsthe a i rcraf t wi th in 90°—or less—of the

runway heading on a base leg for landing.With two engines, this type of approach isnormally flown with the gear d own and 40°of flaps at VREF + 10 kt (minimum) fromthe FAF inbound. When landing is as sured,airspeed may be reduced to VREF minimum.

The second type of circling approach (Figure18-18) requires a heading change of more than

90° to line up with the landing runway. Withtwo engines, this type of approach is normal lyflown with the gear down and 20° of f laps atVREF + 20 kt (minimum) from the FAF inbound. On final approach, flaps should be ex-tended to 40° and airspeed reduced to VREFminimum.

All single-engine circling approaches should

be f lown wi th f laps 20° at VREF + 20 kt (min-imum) if maneuvering is required. When es-tablished on final approach, set flaps 20° and VREF + 10 kt (minimum).


1 8

M

A N D P R



A N E U V

R O C E D



1. GEAR AND FLAPS—UP2. AIRSPEED—VREF + 40 KT3. APPROACH CHECKLIST—COMPLETE

IAF OUTBOUND *

1. FLAPS—8°2. AIRSPEED—VREF + 30 KT (MIN)3. DESCEND, IF REQUIRED

1. FLAPS—20°2. AIRSPEED—VREF + 20 KT

ON COURSE INBOUND

1. FLAPS—20°2. GEAR—DOWN3. AIRSPEED—VREF + 20 KT4. BEFORE LANDING CHECKLIST— COMPLETE TO FLAPS 40°



GO-AROUND/BALKEDLANDING

The Learjet go-around/balked landing pro-cedure, shown in Figure 18-19, should beused for all missed approaches. Generally, if a missed approach is started at—or above— MDA or DH, it is considered a go-around. If a missed approach is started below MDA or

DH, it is considered a rejected, or balked,landing. During training, rejected, or balked landings will normally be initiated over therunway threshold at an altitude of approxi-mately 50 ft.

In either case, use of the flight director go-around mode is recommended to provide a tar-

get 9° nose-high pitch attitude. After the aircraftis clear of obstacles and the flaps have been re-tracted, the pitch attitude and power may be adjusted to maintain the desired airspeed.

If the go-around/balked landing is made froman instrument approach, the published missed approach procedure should be accomplished un less o the rwise in s t ruc ted . I f t he go -

around/balked landing is made during a circling approach, the initial turn to the missed approach heading must be made toward thelanding runway. The turn may then be con-tinued until the aircraft is established on themissed approach heading.


1 8

MA

A N D P R

GO AROUND

SIMULTANEOUSLY:DISENGAGE AUTOPILOT *

CLEAR OF OBSTACLES

1. ACCELERATE TO VREF + 30 KT (MIN)2 FLAPS UP



AN E U V

R O C E D

DISENGAGE AUTOPILOT * ESTABLISH 9˚ NOSE-UP PITCH ATTITUDE SET TAKEOFF POWER, OR AS REQUIRED

2. FLAPS — UP

SINGLE-ENGINE DRIFT DOWNThe single-engine drift down procedure shownin Figure 18-20 is used to cover the greatest possible dis tance while descending to single-engine cruise altitude after an engine failureat high altitude.

As the note on the chart explains, the speed

schedule depicted also approximates the best

single-engine, rate-of-climb speed below thesingle-engine service ceiling. This speed sched-ule may then also be used to climb to single-engine cruise altitude after an engine failureat low altitude.


ENGINE FAILURE

1. SET MAXIMUM CONTINUOUS THRUST

2. MAINTAIN ALTITUDE UNTIL AIRSPEED

REACHES 170 KIAS

1. DESCEND AT 170 KNOTS UNTIL

AIRSPEED REACHES .50 MACH45,000 FEET



CHAPTER 19WEIGHT AND BALANCE

CONTENTS

Page

INTRODUCTION................................................................................................................. 19-1

GENERAL ............................................................................................................................ 19-2

PLANNING DATA ............................................................................................................... 19-4

Example Conditions ...................................................................................................... 19-4

WEIGHT AND BALANCE COMPUTATION .................................................................... 19-9




ILLUSTRATIONS

Figure Title Page

19-1 Basic Empty Weight Moment Sources (Weight and Balance Data) ...................... 19-3

19-2 Sample Weight and Balance Worksheet—Model 35 ............................................. 19-5

19-3 Weight and Balance Worksheet—Model 35 .......................................................... 19-6

19-4 Weight and Balance Worksheet—Model 36 .......................................................... 19-7

19-5 Configuration Diagram and Provisions Loading Tables........................................ 19-8




CHAPTER 19WEIGHT AND BALANCE




GENERALBefore any weight and balance computationscan be made, the basic empty weight and resulting moment must be ascertained. This information is available on the first and last pages of the Weight and Balance section of the AFM (Figure 19-1). Any changes to the aircraftthat affect weight and balance must be entered in the aircraft records, and a new AircraftWeighing Record must be prepared. It is advisable to check both pages to make certainthat the weights and moments agree; the CG

percentage is not required to compute the CGfor a given loading situation. The Weight and Balance Data section also contains all chartsand tables necessary for CG computations.

It should be noted that there are eight basicinterior configurations for the 35 model and sixfor the 36 model. Diagrams of each configu-ration are provided in the Weight and Balance

Data section, which should help in selecting thecorrect loading tables for provisions, baggage, passengers, and fuel and in verifying stationlocations of the various seating arrangementsand storage compartments.





ONL Y



Y

PLANNING DATA The following information is included to pro-vide an example of the process used to computeweight and balance.

EXAMPLE CONDITIONS

Aircraft (Model 35):• Basic empty weight—9,858 lb

• CG—28.94%

• Moment—3,806,239

• Aircraft configuration:

° Executive door

° Standard seating (swivel seats)

° Right-hand recognition light only

• Fuel, kerosene

° Fuselage tank—1,206 lb

° Wing tanks—2,508 lb

° Tip tanks—2,390 lb

° Planned fuel reserve at destination—

1,500 lb

A typical weight and balance computation isdescribed in this chapter with the exampledata entered on Figure 19-2.

Two airplane loading forms for weight and balance computations are provided in theWeight and Balance section of the AFM , one

each for the 35 model and the 36 model.Sample worksheets (Figures 19-3 and 19-4) areadaptations for training purposes.

The interior configuration diagram (Figure19-5) for the example aircraft used in the sam-




Load:

) p ple problem that follows has been included for ill i










1 9 - 8

F OR T RAI NI N G P

URP O SE S

ONL

1 9 W E I G H T A N D B A L A N C E



LY

WEIGHT AND BALANCECOMPUTATION

1. The first step in computing weight and balance is to determine the basic emptyweight and moment from the AirplaneWeighing Record in the AFM . However,if the aircraft has been altered, determine

the basic empty weight and momentfrom the aircraft records.

The moment may be listed as a seven-digit figure, as shown in Figure WB-1.In this case, the decimal point must bemoved three digits to the left when en-tering the moment on the worksheet.This is because all weight and balance

charts and tables are based on moment per 1,000. This reduces the f igures in thenumerical data to a more manageablesize.

Example

ExampleUsing the information given for the exampleconditions, enter the weights for each item inthe appropriate block on the worksheet. Selectthe correct loading tables from the Weight and Balance section, beginning with provisions. Notice that there are several provisions tables provided from which selection of the proper data must be made. Since the example aircrafthas a standard inter ior, all data shown for mid-cabinet and club configurations can be elimi-nated, leaving only the tables pertaining to theexample aircraft (Figure 19-6).

The correct water load (for the wash basin) isselected by comparing the listed station loca-tions with the configuration diagram. Notethat the wash basin is directly opposite the aft provisions cabinet (S ta . 253) ; therefore the correct water-loading entries for weight and moment/1,000 are 15 and 3.80, respectively.

A similar process is used to select the appro-priate crew passenger baggage and fuel




Enter the basic empty weight (9 858) and mo- priate crew, passenger, baggage, and fuelentries Enter each moment/1 000 table with the

ExampleCompare subtotal results with those given onthe sample worksheet. The wing bendingweight is below the maximum limit.

4. Compute the takeoff gross weight bysubtracting fuel burned during start,taxi, and takeoff. Determine the equiv-alent moment of the fuel burned by re-

ferring to the Fuel Used Vs MomentLoss table in the approved AFM . As anaverage, 3.5 ppm per engine may beused for simplification, which would be fu rni shed by the tip tanks (unlessthey are empty). The takeoff grossweight must not exceed the certified maximum takeoff weight.

ExampleAssuming a 15 minute burn on two engines (7 ppm), the fuel burn is es timated at 105 lb.Referr ing to the Fuel Used Vs Moment Losstable, enter the chart with 112 lb (for simpli-f ication) and read 42.84 moment loss from the

The center of gravity, expressed as a per-centage of MAC, can be read at the bottom or top of the chart, whichever is closer to the point at which the weight and moment linesintersect.

Since the gross weight and moment lines crossat a rather shallow angle, a small error in plot-ting the intersect point can result in a signif-

icant error in computing CG. The point on theenvelope charts can be more accurately plot-ted by mathematically computing CG (%MAC) and then finding the point on the en-velope chart where gross weight and CG (%MAC) intersect. This is more accurate sincethe weight and CG (% MAC) lines cross morenearly at right angles.

The formula to calculate the CG in % MACis:


CG (% MAC) =

× 100(Fuselage Station CG) – LEMAC)

MAC



)tip tank column Note that the entries on the Where:

The Center-of-Gravity table in the Weight and Balance section of the AFM may be used asan alternate means of determining whether the aircraft load is within the weight and CGlimits.

Enter the table with aircraft gross weight (100- pound increments). A forward limit momentand an aft limit moment are listed. If the com-

puted moment fa ll s between those li sted atthe forward and aft limits, the aircraft is withinlimits for flight.

The tables may be used to identify the CGlimits more accurately than the CG charts.However, the charts provide a more graphic depict ion of the aircraft weight and moment inrelation to the limits.

Example

Compute CG in percent MAC for takeoff weight and moment using the formula givenabove.

6. Landing weight and moment may be cal-culated by subtracting the weight and re-sulting moment loss of fuel burned enroute out of each tank, the resultingsubtotals being the planned landingweight and moment. The CG (% MAC)can then be determined using the same process described for the takeoff con-ditions. Check to ensure that the certi -

f ied landing weight is not exceeded.

Example

Given an estimated 1,500 lb of fuel remainingat destination, for operation, the fuel must belocated in the wing tanks because all of the fuelloaded in the fuselage and tip tanks was burned.The 112 lb burned out of the tip tanks prior to

takeoff and the associated 42.84 moment lossalready accounted for leaves 2,278 lb of fueland a moment loss of 885.91. Since every-thing loaded in the fuselage tank was burned,the appropriate form entries are 1,206 lb and 530.22 moment loss.




If 1 500 lb f f l i i th i t k

Use the Weight-Moment-CG Envelope chartto ensure that the landing weight (12,946) and the CG (% MAC) intersect within the flight en-velope. In this example, the lines intersectwithin the envelope and the aircraft is withinlimits for landing.

If the aircraft is not within CG limits, the load must be adjusted before takeoff or the fuelload adjusted in flight to remain within the en-velope. The Center-of-Gravity table can also be used to determine whether the ai rcraft iswithin CG limits for landing.




CHAPTER 20PERFORMANCE

CONTENTS

Page

INTRODUCTION................................................................................................................. 20-1

GENERAL ............................................................................................................................ 20-1

PERFORMANCE ................................................................................................................. 20-1

General .......................................................................................................................... 20-2

Definitions..................................................................................................................... 20-2

Flight Planning Data...................................................................................................... 20-6

Takeoff Performance...................................................................................................... 20-7

Thrust .......................................................................................................................... 20-18




ILLUSTRATIONSFigure Title Page

20-1 Sample Takeoff Worksheet ................................................................................... 20-9

20-2 Takeoff Profile Example...................................................................................... 20-14

20-3 Sample Operational Planning Form..................................................................... 20-21

20-4 Sample Landing Worksheet................................................................................. 20-26

TABLE

Table Title Page

20-1 Configurations ....................................................................................................... 20-6





CHAPTER 20PERFORMANCE



PERFORMANCEGENERAL

Assumed Conditions

The performance data presented for each phaseof operation is based on certain assumed con-ditions. For example, the takeoff distance chartassumes that takeoff power is set before brakerelease. Assumed conditions, along with thedescription of the corresponding charts, aregiven in this chapte r.

Standard Conditions

Standard conditions that apply to all perfor-

mance calculations are:• Cabin a ir—on

• Factors for 50% headwind componentsand 150% tailwind components have been applied to takeoff and landing dataas prescribed in pertinent regulations

• Runway gradient

• Anti-ice—on or off

• Antiskid—on or off

• Flaps 8° or 20° for takeoff and 40° for landing

DEFINITIONS

The following definitions apply to terms used throughout this manual.

Airspeeds

CAS— Calibrated Airspeed The airspeed indicator reading cor-rected for instrument and position

error. KCAS is calibrated airspeed expressed in knots.

IAS —Indicated Airspeed The airspeed indicator reading as in-stalled in the aircraft. KIAS is indi-cated airspeed expressed in knots.




p p g p ph i f i i hi l i

VSO

— Stall Speed, Landing VSO is the stal l ing speed in the landing configuration.

VS1 —Stall Speed, Gear/FlapsVS1 i s the s ta l l ing speed in the appropriate gear/flap configuration.

VMCA — Minimum Control Speed, Air The minimum flight speed at which

the aircraft is control lable with 5° of bank when one engine suddenly becomes inoperative and the remain-ing engine is operating at takeoff thrust.

VMCG — Min imum Co ntro l Sp eed ,Ground

The minimum speed on the ground

at which control can be maintained using aerodynamic controls alone,when one engine suddenly becomesinoperative and the remaining engineis operating at takeoff thrust.

V1—Critical Engine Failure Speed

(VMCA

), or less than the rotationspeed (VR ) plus an increment inspeed attained prior to reaching a 35ft height above the runway surface.

VAPP — Approach Cl imb Speed The airspeed equal to 1.3 VS1 (air-craft in approach configuration).

VREF — Landing Approach Speed

The airspeed equal to 1.3 VS0 (air-craft in landing configuration).

Weights

Maximum Allowable Takeoff Weight

The maximum allowable takeoff weight at thestart of takeoff roll is limited by the most

restrictive of the following requirements:

• Maximum certified takeoff weight.

• Maximum takeoff weight (climb or brakeenergy limited) for altitude and tem- pera ture as determined from the appli-cable figure entitled Takeoff Weight




V1 Critical Engine Failure Speed cable figure entitled Takeoff Weight

• Maximum landing weight (approachclimb or brake energy limited) for alti-tude and temperature as determined fromthe applicable figure entitled LandingWeight Limits

Distances

Accelerate-Stop Distance

The accelerate-stop distance is the horizontaldistance traversed from brake release to the point at which the aircraft comes to a completestop on a takeoff during which one enginefails at V1 and the pilot elects to stop.

Engine-Out Accelerate-Go Distance

The engine-out accelerate-go distance is thehorizontal distance traversed from brake re-lease to the point at which the aircraft atta insa height of 35 ft above the runway surface, ona takeoff during which one engine fails at V1and the pilot elects to continue.

T k ff Fi ld L h

Actual Landing DistanceThe actual landing distances presented in thissection are based on a smooth, dry, paved run-way. The landing field length is equal to thehorizontal distance from a point 50 ft above therunway surface to the point at which the air-craft would come to a full stop on the runway.

Factored Landing Distance

The factored landing distances presented inthis section are equal to the actual landingdistance divided by 0.60 (multiplied by 1.67).

Meteorological

ISA— International Standard Atmosphere

OAT— Outside Air TemperatureThe free air static temperature ob-tained from either ground meteo-rological sources or from inflighttemperature indications adjusted for instrument error and compress-ibility effects




Demonstrated CrosswindThe demonstrated crosswind veloc-ity of 24.7 kt is the velocity of the reported tower winds (measured at a20-foot height) for which adequatecontrol of the aircraft during takeoff and landing was actually demon-strated during certification tests.The value shown is not considered to

be limiting.

Miscellaneous

Position Correction— Static PositionCorrection

A correction applied to indicated airspeed or altitude to eliminate theeffect of the location of the static pressure source on the inst rument

reading. Since all airspeeds and al-titudes in this section are presented as indicated values, no position cor-rections need be made when readingfrom the charts. Any change in theairspeed-altitude system external tothe aircraft or locating any external

1.1% enroute. This conservatism isrequired by FAR 25 for terrain clear-ance determination to account for variables encountered in service.

First Segment ClimbClimb from the point at which theaircraft becomes airborne to the poin t at which the landing gear isfully retracted. The gross climb gra-d ient must be posi t ive , wi thoutground effect. This requirement issatisfied by observing the Takeoff Weight Limits chart. Velocity in-crease is from VLOF to V2 with gra-dient calculated at liftoff velocity(VLOF).

Second Segment Climb

Climb extending from the end of thefirst segment to a height of at least400 ft. The gross climb gradient maynot be less than 2.4%. This require-ment is satisfied by observing theTakeoff Weigh t L imi t s cha r t .Velocity for this segment is V




Approach ClimbClimb from a missed or aborted approach wi th approach (20°) fl aps,landing gear retracted, and takeoff thrust on one engine. The gross climbgradient may not be less than 2.1%.This requirement is satisf ied by ob-serving the Landing Weight Limitschart. Velocity for this segment is

1.3 VS1.

Landing ClimbClimb from an aborted landing withlanding flaps DN (40°), landing gear extended, and takeoff thrust on bothengines. The gross climb gradientmay not be less than 3.2%. This re-quirement is satisfied by observingthe Landing Weight Limits chart.Velocity for this segment is 1.3 VSO.

ConfigurationsThe configurations referred to byname in the charts correspond to thesettings in Table 20-1.


Table 20-1. CONFIGURATIONS

No. of Engines Flap

Configuration Operating Thrust Setting Gear

1st Segment 1 Takeoff 8° or 20° DN

Takeoff Climb



FLIGHT PLANNING DATA The following conditions are provided to compute takeoff, climb, cruise, descent, and land-ing data for a flight from airport A to airpor tB. This should not be construed as a completeexample of flight planning procedures. The purpose is to introduce charts and tables mostcommonly used in flight planning. Sampleworksheets for takeoff and landing data and asample operational planning form for the flight planning problems and solutions are included in this chapter.

Example Conditions

Departure Airport A

Aircraft• Basic Operating Weight (Lb) 9,800

• Payload +200

• Zero-Fuel Weight 10,000

• Fuel Weight +5 200

Airport A Runway

• Runway 01R, Length—7,300 ft

• Dry

• Gradient—0%

• Obstacle—20,100 ft from departure of RWY 01R, 1,500 ft above runway

Climb Conditions• Climb Schedule—250 KIAS/0.70 MI

• Maximum Continuous Thrust (N1)

• Climb on course

• Climb unrestricted to FL 430 (Long-range Cruise)

• Standard Atmospheric Conditions (ISA)

• Average headwind components—20 kt

Cruise Conditions

• Distance—1,000 miles from Airport Ato Airport B




TAKEOFF PERFORMANCEWind Components

Headwind, tailwind, and crosswind compo-nents can be calculated by using the Wind Component chart found in the General sectionof the AFM Performance Data chapter.

Problem

Using runway 01R at airport A, with a reported surface wind of 330/20, the wind di-rection is 40° from the runway heading. Usethe Wind Components chart to determine theheadwind and crosswind components.

Solution

Find the point on the chart at which the 20-knot

arc crosses the 40° line. From this point pro-ceed horizonta lly to the left margin to read theheadwind component (15 kt) or proceed straightdown to the bottom margin to read the cross-wind component (13 kt). Enter headwind component on Takeoff Worksheet (Figure 20-1).

• Maximum takeoff weight to meet min-imum single-engine climb gradient re-quirements and not exceed brake energylimits (climb or brake energy limited)

• Maximum takeoff weight for runwaylength available

• Maximum takeoff weight for obstacleclearance

Maximum Certificated TakeoffWeight

The maximum certif icated takeoff weight for most Learjet 30 series aircraft is 18,300 lb.However, some earlier production aircraft may be limited to 18,000 or 17,000 lb takeoff weight.

• If the aircraft records indicate that theaircraft does not incorporate ECR 1495,ECR 2234, AAK 77-8, or AAK 80-2.the certif icated maximum takeoff weightis 17,000 lb.

• If the aircraft records indicate that the










Maximum Takeoff Weight (Climb orBrake Energy Limited)

The takeoff weight limit charts found in theTakeoff sec t ion o f t he AFM chap te r ,Performance Data, provide the maximum take-off weight for a given temperature and pres-sure altitude (PA) which will allow: (1) theaircraft to meet minimum climb gradients if an engine fails at or after V1 speed and take-

off is continued (left side of charts) or (2) braking to a full stop without exceeding brakeenergy limits if takeoff is rejected at or belowV1 speed (right side of charts).

If the temperature and pressure altitude linesintersect to the left of the Engine Temp Limitline, takeoff should not be at tempted at anygross weight.

NOTE

There are separate charts for takeoff with flaps at 8 or 20° and anti-ice off or on.

at which the aircraft can meet the minimumclimb gradients established by FAR 25 should an engine fail at V1.

Using the example of 60°F and 1,300 ft pres-sure altitude, the 60°F line and the 1,300 al-titude lines do not intersect on the takeoff climb portion of the chart. This indicates thatthe takeoff weight is not limited due to take-off climb. Enter maximum certificated take-off weight (18,300 lb) under CLIMB WT onthe Takeoff Worksheet (Figure 20-1).

Now, determine if the takeoff weight is lim-ited due to brake energy. Enter the Takeoff Weight Limit chart at the left margin, again at60°F, and proceed right until intersecting the1,300-foot altitude line on the brake energyside of the chart. From this point, proceed down to the zero-wind reference line and thendiagonally parallel to the guidelines to a pointopposite 15 kt wind velocity. Directly belowthis point, read the brake energy weight.

In this example, the gross weight is found tob b 18 300 lb hi h i h d i k




Maximum Takeoff Weight forRunway Available

If the computed takeoff field length deter-mined from the AFM Takeoff Distance char tis less than the runway length available, take-off weight is not limited due to runway length.However, if the computed takeoff distance ex-ceeds the runway length available, the aircraftgross weight must be reduced or takeoff de-

layed until atmospheric conditions change(e.g., cooler temperature, increased wind ve-locity, or wind shift to a longer runway).

The maximum takeoff weight limited by avail-able runway can be determined by enteringthe Takeoff Distance chart on the right sidewith the runway length available and working backward to the Gross Weight sect ion. Then

enter the chart at the left with the temperatureand pressure altitude and proceed to the GrossWeight section. Read the gross weight directly below the point at which these two entr ies in-tersect in the Gross Weight section. This is thegross weight that will permit takeoff within therunway length available

the runway length available. Takeoff distance

is discussed in greater detail later in this chap-ter under Takeoff Field Length.

Problem

Determine the maximum takeoff g ross weightfor the runway available.

Solution

To determine the maximum takeoff grossweight for the runway available, work back-ward through the Takeoff Distance char t. Enter the chart on the right margin at the actual f ield length (7,300 ft) and proceed horizontally tothe Wind section. (The Antiskid, Anti-ice,Runway Gradient, and Altitude sections donot apply in this example.) Intersect the 15-kt headwind velocity line and follow the wind lines to the zero-wind reference line. Fromthis point, draw a light pencil line horizontallyacross the Gross Weight section.

Now, enter the char t at the bottom, left mar-gin with the temperature (60°F) and proceed

i ll h l i d (1 300 f )




Maximum Takeoff Weight forObstacle Clearance

Commercial operators (FAR 121 and 135) of U.S.-registered aircraf t are required to deter-mine the maximum takeoff weight that will en-able the aircraft to clear obstacles in the takeoff flight path in the event an engine fails at or after V1 speed. Although not specifically listed asa requirement for other operators, it would be

prudent for al l operators to make these computations to ensure safe opera tion.

Takeoff flight path cha rts are provided in theTakeoff sec t ion o f t he AFM chap te r ,Performance Data, to enable the operator to de-termine the net climb gradient required toclear an obstacle in the takeoff flight path.Additionally, climb gradient charts are pro-

vided in the same section that enable the op-erator to determine the net climb gradient possible (one engine inoperat ive) for aircraftgross weight and existing atmospheric con-ditions.

In the event that the computed climb gradient

Reducing the gross weight increases climb

gradient possible. At the same time, climbgradient required also decreases because thetakeoff distance is reduced, providing more dis-tance from the obstacle. Therefore, an inter- polative process i s required to f ind the exactminimum gradient and maximum weight for obstacle clearance. This process will be de-scribed further in the example.

Takeoff Flight Path

Takeoff flight path char ts are provided for 8°and 20° flap settings and also for close-in and distant obstacles.

The close-in charts are used to determine re-quired climb gradients for obstacle clearancewithin 10,000 ft of Reference Zero, and the dis-

tant charts are used to determine climb gradientrequirements for obstacles up to 40,000 ftfrom Reference Zero.

The origin for each climb gradient line isReference Zero. This point is a point 35 ftabove the runway at the computed takeoff dis-




Horizontal distance from Reference Zero is

calculated by adding the runway remaining be-yond Reference Zero to the distance betweenthe end of the runway and the obstacle (Figure20-2). The appropriate Takeoff Flight Pathchart (Close In—Flaps 8°, Distant—Flaps 8°,Close In—Flaps 20°, or Distant—Flaps 20°) isentered at the bottom margin with the calculated horizontal distance from Reference Zero and atthe left margin with obstacle height above the

runway. Commercial opera tors must enter thechart at the right margin with obstacle heightabove Reference Zero.

Problem

Determine the climb gradient required to clear the obstacle using the previously listed ex-ample conditions.

Solution

First, the horizontal distance from ReferenceZero must be determined. In order to calculatethis, determine the takeoff distance, the lengthof the runway, and the distance of the obsta-l f h d f h h k ff di

to be 3,400 ft. Computation of takeoff distance

is described under Takeoff Field Length in thischapter.

Calculate horizontal distance from ReferenceZero by first subtracting the takeoff distancefrom the runway length of f ind the runway re-maining beyond the takeoff point (7,300–3,400= 3,900 ft). Then add the runway remaining be-yond takeoff point to the distance the obs ta-

cle is from the end of the runway (3,900 +20,100 = 24,000 ft) (Figure 20-2). Enter 24,000ft under DIST FROM REF ZERO on theTakeoff Worksheet.

Now, use the Distant Takeoff Flight Path (Flaps8°) chart in the AFM to determine the climbgradient required. Enter the chart at the bot-tom margin with horizontal distance fromReference Zero (24,000 ft) and proceed ver-tically. (Do not apply winds on this chart.) Now enter the char t at the left margin (non-commercial operator) with obstacle heightabove the runway (1,500 ft) and proceed to theright. The two lines intersect between the 6.0%

d 7 0% di li I l h h




lines. In this example, the climb gradient re-

quired is 6.5% (Figure 20-2).

In addition to finding the climb gradient re-quired, note whether the obstacle falls withinthe first or second segment. If the intersect point is to the left of the Gear Down—gear upline, the obstacle is in the firs t segment. If theintersect point is to the right of the line, theobstacle is in the second segment. It is im-

por tant to note this in order to select the proper climb gradient chart (fir st segment or second segment) to find the climb gradient possiblefor this example. Note also that the climb gra-dient lines on the chart have a different valuein the first and second segments.

Climb Gradients

First , Second, and Final Segment ClimbGradient charts are provided to determine theclimb gradient possible for aircraft grossweight and atmospheric conditions. First and Second Climb Gradient charts (Flaps 8 or 20°) are used in conjunction with the Takeoff Flight Path charts which show required net

read second segment net climb gradient (pos-

sible).

If anti-ice systems are to be turned on for take-off, anti-ice system guidelines must be fol-lowed to the right margin. In this example,anti-ice systems are not necessary for takeoff,so the climb gradient possible is found to beapproximately 9.8%.

It was previously determined that only a 6.5%gradient was required to clear the obstacle.Therefore, the planned takeoff weight of 15,000 lb is acceptable for obstacle clearance.

If the climb gradient poss ible was found to beless than the climb gradient required to clear the obstacle, takeoff should not be attempted under the existing conditions.

As previously mentioned, reducing takeoff gross weight reduces climb gradient required and increases climb gradient possible. As a re-sult, f inding the maximum takeoff gross weightthat allows obstacle clearance becomes an in-

l i A d h d f




With this trial takeoff weight, a new takeoff

distance is computed, a new distance fromReference Zero is calculated, and a new climbgradient required is determined. If the newclimb gradient required is less than the climbgradient possible for the trial gross weight(3.5%), the obstacle can be cleared at the trialgross weight.

In most cases, this process will provide a new

gross weight that will provide obstacle clear-ance. If, however, the new required gradientis still greater than the possible gradient, thetwo gradients (possible and required) can beaveraged again and the entire process repeated.

Takeoff Speeds (V 1,V R, and V 2 )

These speeds are found in the Critical EngineFailure Speed (V1), Rotation Speed (VR ), and Takeoff Safety Speed (V2) charts in the AFM .Separate charts are provided for 8 and 20°flap settings. For a review of these abbrevia-tions (V1, VR , and V2), see Def initions in thischapter.

Problem

Using the Critical Engine Failure Speed (V1),(Flaps 8°) chart in the AFM , determine V1 for the example conditions.

Solution

Enter the chart at the bottom left margin withthe temperature (60°F) and proceed verticallyuntil intersecting the field pressure altitude

(1,300 ft). Then proceed horizontally to theright to the gross weight reference line.

Parallel the guidelines until intersecting thetakeoff gross weight (15,000 lb). From this point, proceed horizontally to the right to thezero-wind reference line. Follow the wind guidelines to the right until intersecting thewind velocity (15 kt headwind).

From this point, proceed horizontally to theright margin and read V1 (118 KIAS). If theanti-ice systems are on, the antiskid system isoff, or if there is a gradient, follow the guide-lines in those two sections of the chart.




Takeoff Safety Speed (V 2 )

Takeoff safety speed (V2), like rotation speed,is affected only by aircraft gross weight and flap setting.

Problem

Determine V2 from the Takeoff Safety Speed (V2), (Flaps 8°) chart in the AFM for the ex-ample takeoff gross weight.

Solution

Enter the chart at the left margin with the take-off gross weight (15,000 lb). Proceed hori-zontally right to the reference line and thenstraight down the margin and read V2 (133KIAS). Enter the V2 value on the Takeoff Worksheet.

Takeoff Field Length

Takeoff fi eld length data assumes a smooth,dry, hard-surface runway.

The takeoff distances computed from the take

The takeoff field length data presented in the

AFM is governed by the accelerate-stop or theengine-out accelerate-go distance, whichever is greater. Generally, unless V1 is limited byVR or VMCG , the takeoff field lengths are bal-anced, and the accelerate-s top distance equalsthe accelerate-go distance.

The Takeoff Distance charts in the AFM are presented for 8 or 20° flaps sett ings. These

charts may be used to determine either of thefollowing:

1. Runway length required for a given air-craft weight.

2. Maximum aircraft takeoff weight cor-responding to a specific runway length.The process for finding the maximum

aircraf t weight for a given runway lengthwas previously described in this sectionunder Maximum Takeoff Weight for Runway Available .

Problem




If takeoff with a gradient is planned with the

antiskid system off or the anti-ice systemson, follow the guidelines through the corr e-sponding section of the chart while proceed-ing to the right margin.

Pressure altitude is compensated for on theright side of the chart. If takeoff is planned ata pressure altitude above 11,000 ft, an addi-tional factor must be applied in the alti tude sec-

tion on the far right side of the chart. For normaltakeoffs below a pressure altitude of 11,000 ft,the altitude section can be disregarded.

NOTE

Certif ication for U.S.-registered air-craft limits takeoffs and landings to10,000 ft pressure altitude.

THRUST

Takeoff Thrust

Takeoff performance is based on the assump-

If N1 is below that specified in the takeoff

power sett ing charts for the existing temper-ature and pressure altitude, aircraft takeoff performance will not meet the takeoff per-formance specified in the performance charts.If N1 is above computed takeoff power, air-frame or engine limits may be exceeded. Thus,it is necessary to compute takeoff power and adjust the power levers as necessary to se t N1equal to chart value. In addition, operation at

a specific N1 should always be within ITTlimits.

NOTE

During takeoff, N1 may decreaseslightly from the initial static read-ing. Therefore, N1 should be contin-uously monitored and adjusted until

reaching 80 KIAS.

Separate takeoff power setting charts are pro-vided for aircraft equipped with standard noz-zles and those equipped with thrust reverser nozzles Takeoff power setting charts for stan




Maximum Climb Thrust

The climb performance data in the Pi lo t’s Manual is predicated on adjusting thrust (N1)a f t e r t akeof f t o the va lue found in theMaximum Continuous Thrust (N1) tables in the

AFM . As with takeoff thrust, continuous thrustdata is presented for standard nozzles (in thePerformance section) and thrust reverser noz-zles (in the thrust reverser supplement). In

addition, maximum continuous thrust data is presented for single-engine opera tion.

The maximum continuous thrust (N1) settingmay be determined before takeoff using esti-mated temperature and altitude at start of climb. Since the Maximum Continuous Thrust(N1) table is based on ram-air temperature indegrees Celsius, the reported or estimated

OAT must be converted to RAT before enter-ing the chart.

It is more practical to set power at 795°C ITTafter takeoff at the beginning of the climb.Later when crew workload permits, compute

i i h d N

Many operators prefer to simply set the engines

to 795°C on the ITT gages and adjus t power levers as necessary during the climb to main-tain 795°C (recommended continuous ITT).This eliminates the need to compute a climb N1 setting and also possibly extends enginecomponent life due to operating at lower en-gine temperatures. This power managementtechnique does not guarantee the climb per-formance presented in the Pilot’s Manual .

Problem

Assume the pilot elects to set 795°C on the ITTgages at the beginning of climb and computemaximum continuous thrust (N1) passingthrough 15,000 ft. The RAT indicator readingat 15,000 ft is –7°C.

SolutionEnter the Maximum Continuous Thrust (N1)(All Engines, Standard Nozzle) chart in the

AFM and determine power setting (N1) at15,000 ft and –7°C.

h bl k hi h d l i d




CLIMB, CRUISE, ANDDESCENT PLANNING

An Operational Planning Form is provided inthe Flight Planning Data section of the Pilot’s

Manual . See Figure 20-3 for a sample form.

Problem

Determine fuel at star t of climb.

Solution

Enter example conditions on the OperationalPlanning Form. Zero-fuel weight (10,000 lb) plus fuel (5,200 lb) equals ramp we ight(15 ,200 lb) . Subt rac t 200 lb of fuel for warmup and takeoff to find takeoff/start climbweight (15,000 lb). Subtract 200 lb from fuelload (5,200 lb) to find fuel at a start climb(5,000 lb).

Climb Performance

A set of climb performance tables is pro-vided in the Pilot’s Manual to determine time

Problem

Using the example conditions, determine time,distance, and fuel used in the climb from 1,300ft to planned long-range cruise altitude. For this example, planned cruise altitude is de-termined to be FL 430. The method of deter-mining this altitude will be described under Long-range Cruise in this chapter.

To determine the required time, distance, and fuel, refer to the Climb Performance (Two-engine, 15,000-pound) chart in the Pi lo t’s

Manual . The example gross weight at startclimb is 15,000 lb. For intermediate grossweights, two charts are required for interpo-lation. To simplify this example, however,only the 15,000-pound table is used.

SolutionUsing the ISA column on the 15,000-pound table, interpolate data listed for 1,000 and 3,000 ft to find time, distance, and fuel for 1,300 ft (start climb altitude). The result should be approximately 0.3 minutes, 1.2 NM, and










Thus, the climb performance data with wind

applied is found to be 17.2 minutes, 101 NM,and 512 lb. Enter this data on the OperationalPlanning Form.

Cruise Performance

Cruise performance tables are provided in the Pilot’s Manual for normal cruise, high-speed cruise, and long-range cruise.

Normal Cruise

Normal cruise tables provide fuel flow and true airspeed for constant 0.77 MI cruise.Engine power is adjusted to maintain the con-stant Mach as weight decreases. Enter the appropriate table for the average aircraft grossweight for each cruise segment.

High-speed Cruise

High-speed cruise tables provide fuel flow,indicated Mach or airspeed, and true airspeed for a MMO/VMO or VMAX cruise. Power for maximum speed cruise is set for the l imiting

portion of the fl ight shou ld be divided in to

segments, with an appropriately higher cruiseal t i tude planned as the gross weight de-creases. As a rough guide in planning for changes in cruise altitude, increase cruisealtitude 1,000 ft for each 1,000 lb decreas ein gross weight (i.e., fuel used).

The specific range chart assumes zero wind.If winds are significant, it may be advanta-

geous to select a different altitude to avoid headwinds or take advantage of tailwinds.

Once the initial cruise altitude has been de-termined, refer to the appropriate long-rangecruise chart to determine the indicated Machor airspeed, true airspeed, and fuel flow for theinitial cruise segment. Each chart providesthe above data for a different aircraft gross

weight. The gross weight is specified in thetop, left corner of each chart and represents theaverage gross weight for a cruise segment in500-pound increments.

Problem




After establishing the planned cruise altitude,

the climb data can be extracted from the climb performance charts as previously described inthis chapter under Climb Performance.

Using the Long-Range Cruise (Two-Engine,Weight—13,500-Pound) char t, extract cruisedata. In the ISA column, opposite 43,000 ft,f ind: Mach Ind.—0.736; KTAS—415; and fuel pph—905.

Enter this data on the planning form. The ex-ample conditions specified an average 20-knot headwind component; therefore, 20 ktshould be subtracted f rom the TAS to deter-mine ground speed (415 – 20 = 395 GS).

Now, f ind the cruise distance by subtract ingclimb and descent distances from total trip

length. Descent distance is 82 NM, and computation of this distance is described in thischapter under Descent (1,000 – [101 + 82] =817 NM). Enter 817 distance on the example planning fo rm. Now use a navigation computer to f ind cruise time using 395 kt GS for 817 NM Th i f 2 04 (2 h d 4 i

Solution

Enter the chart at the bottom margin with planned cruise Mach (0 .70) and proceed ver-tically to the altitude line (41,000). Estimate41 between the 40 and 45 lines. From this point, proceed horizontal ly to the right unti lintercepting the aircraft weight line on theright side of the chart (17,000), halfway be-tween the 16,000 and 18,000-foot line. Directly

below this point on the bank angle scale, read the maximum bank angle to avoid buffet in alevel turn (approximately 48˚ of bank).

Now, return to the gross weight line (17,000)above this point and fo llow it diagonally downto the left to the vertical 1.0 G reference line.Draw a line horizontally left of this point untilit intersects the curved 41,000 altitude line (be-

tween 40 and 45 lines). On the Mach scale di-rectly below the first point, read the Machnumber at which low-speed airframe buffetmay be encountered (0.575 MI).

Low Speed Buffet Boundary at 1.5G (FC 200)




Note

1. 5 G = 48 1 ⁄ 2˚ bank in level flight.

Descent Performance

Two Descent Performance Schedules are pro-vided in the Pilot’s Manual to provide time,distance (no wind), and fuel used for descentto sea level: one for minimum fuel descent

and one for normal descent. The tables as-sume an average descent weight of 12,000 lb.Subtraction of performance values for two al-titudes results in time, distance, and fuel re-quired for descent between the two altitudes.The descent speed schedules presented at the bo ttom of the table sho uld be followe d toachieve the desired results. The power settingfor descent is IDLE thrust.

Problem

U s i n g t h e M i n i m u m F u e l D e s c e n tPerformance Schedule, extract descent datafor descent from FL 430 to 1,300 ft (destina-tion elevation). The descent might be planned

20 kt in the descent; therefore, 20 kt should

be subtrac ted from the average no wind speed (TAS) to f ind an average ground speed (GS)in the descent. Use a navigation computer todetermine no wind speed (TAS):

KTAS = 87 KTAS = 37860 13.8

378 KTAS – 20 kt headwind = 358 kt GS

358 kt GS for 13.8 minutes = 82 NM

Thus, the descent data with wind applied isfound to be 13.8 minutes, 82 NM, and 162 lb.Enter this data on the planning form and sub-tract fuel used in descent from the start descentgross weight and fuel remaining at start de-scent. The end descent data is 12,466 lb gross

weight and 2,466 lb of fuel remaining.

Fuel Reserve

FAR Part 91 requires a fuel reserve (IFR con-ditions) of 45 minutes at destination or at thealternate airport if an alternate is required




APPROACH AND LANDING

PERFORMANCE

Approach and performance data are provided in chart form in the AFM performance stat ionand in tabular form in the Learjet Pi lot’s

Manual and checklist (Figure 20-4).

Maximum Allowable Landing

WeightThe maximum allowable landing weight islimited by the most restrictive of the fol-lowing requirements:

• Maximum certificated landing weight(14,300 or 15,300 lb)

• Maximum landing weight (approach

climb or brake energy limited)

• Maximum landing weight for the runwaylength available




Maximum Landing Weight

(Approach Climb or BrakeEnergy Limited)

The Landing Weight Limit charts (ApproachClimb and Brake Energy Limited) provide amaximum approach/landing weight that al-lows (1) the aircraft to meet a minimum climbgradient (single engine) in the event of missed approach or (2) braking to a full stop without

exceeding brake energy limits.

If landing distance for existing gross weightis computed to be greater than the runwayavailable, the gross weight must be reduced be-fore using that runway. Landing weight for runway length available may be determined byworking through the Landing Distance chart backward . Use the same procedure as previ-

ously described for f inding maximum takeoff gross weight for a given runway length.

Problem

Use the Landing Weight Limit (Anti-ice-Off)chart in the AFM to determine the maximumlanding weight for the example conditions.

Solution

Enter the chart on the left margin with temper-ature (60°F) and proceed horizontally right untilintersecting the pressure altitude (1,300 ft) line.

If the temperature line intersects the altitude line,read the approach climb weight limit directly below the point at which they intersect.

In this example, the 60˚F temperature linedoes not intersect the 1,300 ft altitude line.This indicates that a safe missed approachcould be made on one engine at any grossweight up to 18,300 lb at this temperature and

pressure al titude. Enter 18,300 lb under APPR CLIMB WT on the Landing Worksheet.




Using the same chart, determine if landing

weight is limited by brake energy. Enter theleft side of the chart again with temperature(60˚F) and proceed horizontally until inter-cepting the altitude line (1,300 ft). Again thelines do not intersect, indicating that the air-craft can be stopped at gross weights up tomaximum certificated g ross weight withoutexceeding brake energy limits. Enter 18,300on the Landing Worksheet under BRAKE EN-

ERGY WEIGHT.

If the temperature and pressure altitude lineshad intercepted, the wind and runway gradi-ents are accounted for at the lower section of the chart.

Maximum Landing Weight forRunway Available

This computation is made using the samemethod as that used to f ind maximum takeoff weight for runway available. Enter the LandingDistance chart on the right and work back-ward in the chart t o the Gross Weight section

d d li ht li th h th G W i ht

Operations) chart can be achieved when the

following procedures are used:

1. Approach through the 50-foot point over the end of the runway at VREF with flapsand gear down, using a 2 1⁄2 –3˚ glideslope.

2. After passing through the 50-foot point, progress ively reduce thrust unt il thrust

levers are at IDLE prior to touchdown.

3. After touchdown, extend spoilers im-mediately.

4. Apply wheel brakes as soon as practi-cal and continue maximum braking ac-tion until the airplane stops.

5. After landing, move the control col-umn full aft and maintain that positionuntil the aircraft stops.

NOTE

Pulling the control column aft will




When the runway is other than dry, the fol-

lowing factors should also be applied to theLanding Distance chart.

• Wet—Apply a 1.4 factor to the computed landing distance

• Wet (in the process of freezing)—Applya factor of at least 1.7 to the computed landing distance

Those operators governed by FAR Part 121 or 135 first determine landing distance from theLanding Distance (FAR Part 91 Operations)chart and then apply the appropriate abnormallanding factor if required. Next, enter theLanding Distance (FAR Part 121 and 135Operations) chart to compute landing field length for scheduled and alternate stops.

When the runway is wet, commercial opera-tors must apply a 1.15 factor to the landingfield length.

NOTE

Next , follow the guidel ines diagonally up and

to the right until intersecting the weight(12,466 lb) from the Operational PlanningForm (see Figure 20-3). Move horizontally tothe right to the zero-wind reference line.

Follow the wind guidelines until intersecting15 kt headwind velocity. From this point pro-ceed horizontally through the runway gradi-ent section (zero gradient), antiskid section

(antiskid on), and through the altitude sec-tion (below 11,000 ft) to the right marginand read landing distance (2,550 ft). Enter 2,550 ft on the Landing Worksheet under LANDING D ISTANCE.

To determine maximum landing gross weightfor the runway available, enter the LandingDistance chart on the r ight with runway length

(13,300 ft) and work backward through thechart to the Gross Weight section.

In this example, the runway available exceedsthe chart values for all conditions, indicatingthat there is no limitation in landing gross

i h f il bl E 18 300 lb




Landing Approach Speed

(V REF )V R EF i s d e t e rm i n e d f ro m t h e La n d i n gApproach Speed (VREF) chart in the AFM .Since VREF is determined strictly by aircraftgross weight, VREF speeds listed in tabular form in the Pilot’s Manual and checklist may be used with equal accuracy.

ProblemUse the Landing Approach Speed (VREF) chartin the AFM to determine VREF for the planned landing weight in the example (12,466 lb).

Solution

Enter the chart at the left margin with grossweight (12,466 lb) and proceed horizontally

until intersecting the reference line, thenstraight down to the bottom margin of thechart and read VREF (117 KIAS). Enter thisvalue under VREF on the Landing Worksheet.

Approach and Landing Speeds

ular data in the Pilot’s Manual and checklist

i s a s accura t e a s the cha r t i n the AFM .Approach climb and landing climb speeds are provided on the same char t in the AFM .

Problem

Use the Approach and Landing Climb Speedschart in the AFM to determine these speeds for the example conditions.

Solution

Enter the chart on the left margin with thegross weight (12,466 lb) and proceed hori-zontally to the first reference line. Then movestraight down to the bottom margin of thechart to read landing climb speed (117 KIAS).It should be noted that landing climb speed isthe same value as landing approach speed

(VREF). Therefore, if VREF is known, it is notnecessary to compute landing climb speed.

Using the same chart and the approach climbspeed reference line, f ind approach climb speed (123 KIAS). It should be noted that approachli b d b 6 k h




CHAPTER 21

CREW RESOURCE MANAGEMENT

CONTENTS

Page

CREW CONCEPT BRIEFING GUIDE ............................................................................... 21-3

Introduction ................................................................................................................... 21-3

Common Terms ............................................................................................................. 21-3

Pretakeoff Briefing (IFR/VFR) ..................................................................................... 21-4

Crew Coordination During the Approach Sequence ..................................................... 21-4

ALTITUDE CALLOUTS...................................................................................................... 21-5

Enroute .......................................................................................................................... 21-5

Approach—Precision .................................................................................................... 21-5




ILLUSTRATIONS

Figure Title Page

21-1 Situational Awareness in the Cockpit .................................................................... 21-1

21-2 Command and Leadership ..................................................................................... 21-1

21-3 Communication Process ........................................................................................ 21-2

21-4 Decision Making Process ...................................................................................... 21-2

21-5 Error Management................................................................................................. 21-3




CHAPTER 21CREW RESOURCE MANAGEMENT


CLUES TO IDENTIFYING:• Loss of Situational Awareness

• Links In the Error Chain

O P E R A T I O N A L 1. FAILURE TO MEET TARGETS

2. UNDOCUMENTED PROCEDURE

3. DEPARTURE FROM SOP

4. VIOLATING MINIMUMS OR LIMITATIONS

5. FAILURE TO MONITOR

PILOTMONITORING

(PM)

Events thatmay happen

Events thathave

happened

Events thatare

happeningnow

PILOTFLYING

(PF)

COLLECTIVES/A

SA

SITUATIONAL AWARENESS IN THE COCKPIT




2 1

C R E W

R E S O U R C E MA N A G E ME N T

— THINK—

— REMEMBER —

Questions enhance communication flowDon’t give in to the temptation to ask questions when Assertion is required

Use of Inquiry or Advocacy should raise a “red flag”.

• Solicit and give feedback

• Listen carefully

• Focus on behavior, not people

• Support Conclusions with Facts

• State Position, Suggest Solutions

• Clear, Concise Questions

ADVOCACY:• Increase collective S/A

INQUIRY:• Increase individual S/A

ASSERTION:• Reach a conclusionNEED SEND RECEIVE

OPERATIONALGOAL

FEEDBACK

• Maintain focus on the goal

• Verify operational outcome is achieved

• Be aware of barriers to communication

COMMUNICATION PROCESS

Figure 21-3. Communication Process



CREW CONCEPTBRIEFING GUIDE

INTRODUCTION

To a large extent the success of any aircrew de- pends on how effect ively crewmembers coor-dinate their actions using standardized and

COMMON TERMS

PIC Pilot in Command

Designated by the company for flightsr e q u i r i n g m o re t h a n o n e p i l o t .Responsible for conduct and safety of the flight. Designates pilot flying and pi lot not flying duties .


• IDENTIFY AREAS OF

VULNERABILITY

• USE SOPs, CHECKLISTS AND

EFFECTIVE MONITORING TO

ESTABLISH LAYERS OF

DEFENSE

MITIGATE

DETECT & TRAP

ANTICIPATE & AVOID

ERROR

PREVENTION

ERROR

CONTAINMENT

ERROR MANAGEMENT

Figure 21-5. Error Management



PRETAKEOFF BRIEFING

(IFR/VFR)

NOTE

The following briefing is to be completed during item 1 of the Pretakeoff checklist. The pilot flying will ac-complish the briefing.

1. Review the ATC clearance and de- par ture procedure (route and alti tude,type of takeoff, significant terrainfeatures, etc.).

2. Review those items that are not stan-dard procedure to include deferred or MEL items (if applicable).

3. Review required callouts, unless stan-dard calls have been agreed upon, inwhich case a request for “Standard Callouts” may be used.

4. Review the procedures to be used incase of an emergency on departure.

PF Requests the pilot monitoring to ob-

tain destination weather. (Transfer of communication duties to the pilot fly-ing may facilitate this task.)

PM Advises the pilot of current dest ina-tion weather, approach in use, and special information pertinent to thedestination.

PF R eq ue st s t he pi lo t m on it or in g t o

perform the approach se tup.

PM Accomplishes the approach setup and advises of frequency tuned, identified and course set.

PF Transfers control of the aircraft to the pilot monitoring, advising, “You havecontrol, heading , altitude ” and special instructions.(Communications duties should betransferred back to the pilot moni-toring at this point.)

PM Responds, “I have control , heading , altitude .”


2 1

C R E W




ALTITUDE CALLOUTS

ENROUTE

1,000 ft prior to level off

PM PF

State altitude leaving and assigned “CHECKED”

level off altitude

“200 above/below” “LEVELING”

APPROACH—PRECISION

PM PFAt 1,000 ft above minimums

“1,000 feet above” “DH _________”

At 500 ft above minimums

“500 feet above minimums” “NO FLAGS”




APPROACH—NONPRECISION

PM PFAt 1,000 ft above MDA

“1,000 feet above” “MDA _________”

At 500 f t above MDA

“500 feet above.” “NO FLAGS”

At 100 f t above MDA

“100 feet above.”

At minimum descent altitude (MDA)

“MDA” “MAINTAINING MDA”

At or prior to the missed approach point (MAP)

“A h li h ( l k i i )” “CO G”


2 1

C R E W




SIGNIFICANT DEVIATION CALLOUTS

PM PFIAS ±10 KIAS

“VREF ± __________” “CORRECTING TO _________”

Heading ±10° enroute, 5° on approach

“Heading __________ degrees lef t/r ight” “CORRECTING TO _________”

Altitude ±100 ft enroute, +50/-0 ft on f inal approach

“Altitude __________ high/low” “CORRECTING TO _________”

CDI left or right one dot

“Left/right of course__________ dot” “CORRECTING”





WA-1FOR TRAINING PURPOSES ONLY

WALKAROUND

The following section is a pictorial walkaround.It shows each item called out in the exterior

po wer -o ff pr ef li gh t in sp ec ti on . Th e fo ld -o ut pa ges , WA-2 an d WA-1 5, sho uld be unf ol ded bef ore sta rti ng t o re ad.

The general location photographs do not specifyevery checklist item. However, each item is por-trayed on the large-scale photog raphs that follow.





FC 200 AND FC 530

8. NOSEWHEEL AND TIRE—CONDITION AND NOSE GEAR

UPLOCK FORWARD12. RIGHT STALL WARNING VANE—FREEDOM OF

MOVEMENT, LEAVE IN DOWN POSITION





W

A L K A R O U N D 15. RIGHT SHOULDER STATIC PORT—CLEAR OF

OBSTRUCTIONS (FC0-200)

16. COPILOT WINDSHIELD DEFOG OUTLET—CLEAR OF

OBSTRUCTIONS

20. ROTATING BEACON LIGHT AND LENS (ON VERTICAL

FIN)—CONDITION



23. RIGHT MAIN GEAR AND WHEEL WELL—

HYDRAULIC/FUEL LEAKAGE AND CONDITION

27. RIGHT WING ACCESS PANELS (UNDERSIDE OF

WING)—CHECK FOR FUEL LEAKAGE





W

A L K A R O U N D 30. RIGHT WING HEAT SCUPPER (UNDERSIDE OF WING

FORWARD)—CLEAR OF OBSTRUCTIONS

33. RIGHT TIP TANK SUMP DRAIN VALVE—DRAIN



40. RIGHT SPOILER AND FLAP—CONDITION37. RIGHT TIP TANK FUEL JETTISON TUBE—CLEAR OF

OBSTRUCTIONS





W

A L K A R O U N D 43. RIGHT ENGINE THRUST REVERSER—CONDITION AND

STOWED (AERONCA)

45. RIGHT ENGINE FUEL BYPASS VALVE INDICATOR—

CHECK, NOT EXTENDED



48. TAIL CONE ACCESS DOOR—OPEN 49B. TAIL CONE INTERIOR—CHECK FOR FLUID LEAKS,

SECURITY, AND CONDITION OF INSTALLED

EQUIPMENT





WA L K A R O U N D 51. TAIL CONE ACCESS DOOR—CLOSED AND SECURE 55. RIGHT FUEL COMPUTER DRAIN VALVE—DRAIN (DRAIN

VALVES ARE RECESSED ON AIRCRAFT EQUIPPED WITH

DRAG CHUTE.)



62. LEFT FUEL COMPUTER DRAIN VALVE—DRAIN (DRAIN

VALVES ARE RECESSED ON AIRCRAFT EQUIPPED WITH

DRAG CHUTE.)

65A. LEFT ENGINE THRUST REVERSER—CONDITION AND

STOWED (DEE HOWARD)





WA L K A R O U N D 67. LEFT ENGINE FUEL BYPASS VALVE INDICATOR—

CHECK, NOT EXTENDED

70. LEFT AILERON—CHECK FREE MOTION, BALANCE,

AND TRIM LINKAGE, AND BRUSH SEAL CONDITION




73. LEFT TIP TANK FIN AND STATIC DISCHARGE

WICKS (2)—CONDITION

77. LEFT TIP TANK RECOGNITION LIGHT AND LENS

(IF INSTALLED)—CONDITION




WA L K A R O U N D 80. VORTEX GENERATORS OR BOUNDARY LAYER

ENERGIZERS—CONDITION

83. STALL STRIP (IF INSTALLED) AND WING LEADING

EDGE—CONDITION

LEARJET 30 SERIES PILOT TRAINING MANUAL LEARJET 30 SERIES PILOT TRAINING MANUAL



WFOR TRAINING PURPOSES ONLYFOR TRAINING PURPOSES ONLY

86. LEFT MAIN GEAR LANDING LIGHT—CONDITION

87. LEFT MAIN GEAR WHEELS, BRAKES, AND TIRES—

CONDITION

88. LEFT ENGINE INLET AND FAN—CLEAR OF

OBSTRUCTIONS AND CONDITION57

58

59

60

56 48

49

50

51

55 47 43

45

42

46

39

38

37 36

4144

5452

53

40

75 73 80

68 67 66

74 71 72 70 69 64 63 65 62

61




WA-16 FOR TRAINING PURPOSES ONLY

NOTE:

THE NUMBERS ON THIS

DIAGRAM CORRESPOND TO

THE PREFLIGHT POSITIONS

DEPICTED IN THE AIRPLANE

FLIGHT MANUAL.



APPENDIX

CONTENTS

Page

CONVERSIONS............................................................................................................... APP-1

ANSWERS TO QUESTIONS........................................................................................... APP-6




TABLES

Table Title Page

APP-1 Conversion Factors ............................................................................................. APP-1

APP-2 Fahrenheit and Celsius Temperature Conversion ............................................... APP-2

APP-3 Inches to Millimeters.......................................................................................... APP-3

APP-4 Weight (Mass): Ounces or Pounds to Kilograms ............................................... APP-4 APP-5 Weight (Mass): Thousand Pounds to Kilograms................................................ APP-5





MULTIPLY

centimeterskilograms

kilometers

kilometers

litersliters

meters

meters

millibars

feetgallons

inches

in. Hg (32°F)

nautical milesnautical miles

poundsquarts (liquid)

statute miles

BY

0.39372.2046

0.621

0.539

0.2641.05

39.37

3.281

0.02953

0.30483.7853

2.54

33.8639

1.1511.852

0.45360.946

1 609

TO OBTAIN

inchespounds

statute miles

nautical miles

gallonsquarts (liquid)

inches

feet

in. Hg (32°F)

metersliters

centimeters

millibars

statute mileskilometers

kilogramsliters

kilometers

Table APP-1. CONVERSION FACTORS

A P P E N D I X



A P P -2


ONL Y




INCHES 0.0000 0.0001 0.0002 0.0003 0.0004 0.0005 0.0006 0.0007 0.0008 0.0009

MILLIMETERS

0.000 0.0025 0.0050 0.0076 0.0101 0.0127 0.0152 0.0177 0.0203 0.0228

0.001 0.0254 0.0279 0.0304 0.0330 0.0355 0.0381 0.0406 0.0431 0.0457 0.0482

0.002 0.0508 0.0533 0.0558 0.0584 0.0609 0.0635 0.0660 0.0685 0.0711 0.0736

0.003 0.0762 0.0787 0.0812 0.0838 0.0863 0.0889 0.0914 0.0939 0.0965 0.0990

0.004 0.1016 0.1041 0.1066 0.1092 0.1117 0.1143 0.1168 0.1193 0.1219 0.1244

0.005 0.1270 0.1295 0.1320 0.1346 0.1371 0.1397 0.1422 0.1447 0.1473 0.1498

0.006 0.1524 0.1549 0.1574 0.1600 0.1625 0.1651 0.1676 0.1701 0.1727 0.1752

0.007 0.1778 0.1803 0.1828 0.1854 0.1879 0.1905 0.1930 0.1955 0.1981 0.2006

0.008 0.2032 0.2057 0.2082 0.2108 0.2133 0.2159 0.2184 0.2209 0.2235 0.2260

0.009 0.2286 0.2311 0.2336 0.2362 0.2387 0.2413 0.2438 0.2463 0.2489 0.2514

INCHES 0.000 0.001 0.002 0.003 0.004 0.005 0.006 0.007 0.008 0.009

MILLIMETERS

0.00 0.025 0.050 0.076 0.101 0.127 0.152 0.177 0.203 0.228

0.01 0.254 0.279 0.304 0.330 0.355 0.381 0.406 0.431 0.457 0.482

0.02 0.508 0.533 0.558 0.584 0.609 0.635 0.660 0.685 0.711 0.736

0.03 0.762 0.787 0.812 0.838 0.863 0.889 0.914 0.939 0.965 0.990

0.04 1.016 1.041 1.066 1.092 1.117 1.143 1.168 1.193 1.219 1.244

0.05 1.270 1.295 1.320 1.346 1.371 1.397 1.422 1.447 1.473 1.498

0.06 1.524 1.549 1.574 1.600 1.625 1.651 1.676 1.701 1.727 1.752

0.07 1.778 1.803 1.828 1.854 1.879 1.905 1.930 1.955 1.981 2.006

Table APP-3. INCHES TO MILLIMETERS




A P P E N D

0 1 2 3 4 5 6 7 8 9

kg kg kg kg kg kg kg kg kg kg

oz

0 – 0.028 0.057 0.085 0.113 0.142 0.170 0.198 0.227 0.255

10 0.283 0.312 0.340 0.369 0.397 0.425 0.454 0.482 0.510 0.539

lb

0 – 0.45 0.91 1.36 1.81 2.27 2.72 3.18 3.63 4.08

10 4.5 5.0 5.4 5.9 6.4 6.8 7.3 7.7 8.2 8.6

20 9.1 9.5 10.0 10.4 10.9 11.3 11.8 12.2 12.7 13.2

30 13.6 14.1 14.5 15.0 15.4 15.9 16.3 16.8 17.2 17.7

40 18.1 18.6 19.1 19.5 20.0 20.4 20.9 21.3 21.8 22.2

50 22.7 23.1 23.6 24.0 24.5 24.9 25.4 25.9 26.3 26.8

60 27.2 27.7 28.1 28.6 29.0 29.5 29.9 30.4 30.8 31.3

70 31.8 32.2 32.7 33.1 33.6 34.0 34.5 34.9 35.4 35.880 36.3 36.7 37.2 37.6 38.1 38.6 39.0 39.5 39.9 40.4

90 40.8 41.3 41.7 42.2 42.6 43.1 43.5 44.0 44.5 44.9

100 45 46 46 47 47 48 48 49 49 49

0 10 20 30 40 50 60 70 80 90

(1 oz = 0.028 349 52 kg) (1 lb = 0.453 592 4 kg)

Table APP-4. WEIGHT (MASS): OUNCES OR POUNDS TO KILOGRAMS





lb 0 100 200 300 400 500 600 700 800 900

(000)* kg kg kg kg kg kg kg kg kg kg

1 454 499 544 590 635 680 726 771 816 862

2 907 953 998 1043 1089 1134 1179 1225 1270 1315

3 1361 1406 1451 1497 1542 1588 1633 1678 1724 17694 1814 1860 1905 1950 1996 2041 2087 2132 2177 2223

5 2268 2313 2359 2404 2449 2495 2540 2585 2631 26766 2722 2767 2812 2858 2903 2948 2994 3039 3084 3130

7 3175 3221 3266 3311 3357 3402 3447 3493 3538 35838 3629 3674 3719 3765 3810 3856 3901 3946 3992 4037

9 4082 4128 4173 4218 4264 4309 4354 4400 4445 449110 4536 4581 4627 4672 4717 4763 4803 4853 4899 4944

11 4990 5035 5080 5126 5171 5216 5262 5307 5352 5398

12 5443 5488 5534 5579 5625 5670 5715 5761 5806 5851

13 5897 5942 5987 6033 6078 6123 6169 6214 6260 630514 6350 6396 6441 6486 6532 6577 6622 6668 6713 675915 6804 6849 6895 6940 6985 7031 7076 7121 7167 7212

16 7257 7303 7348 7394 7439 7484 7530 7575 7620 7666

17 7711 7756 7802 7847 7893 7938 7983 8029 8074 811918 8165 8210 8255 8301 8346 8391 8437 8482 8528 8573

19 8618 8664 8 09 8 4 8800 884 8890 8936 8981 9026

(1 lb = 0.453 592 4 kg)

Table APP-5. WEIGHT (MASS): THOUSAND POUNDS TO KILOGRAMS





A P P E N D

CHAPTER 2

1. D2. B3. D4. B5. D6. A or B

7. D8. B9. A10.C11.C12.A13.D14.C

CHAPTER 3

1. C2. C3. C

CHAPTER 5

1. B2. D3. D4. C5. D6. D

7. A8. D9. A10.C11.B12.A13.D14.A

CHAPTER 7

1. A2. C3. C

CHAPTER 9

1. D2. B3. A4. C5. A

CHAPTER 101. C2. A3. C4. C5. B6. C7. D

8. D

CHAPTER 11

1. C2 C

CHAPTER 13

1. B2. A3. D4. A5. A6. A

7. C8. D9. D

CHAPTER 14

1. C2. A3. A

4. D5. B6. C7. C8. C

ANSWERS TO QUESTIONS




CHAPTER 15

1. B2. C3. C4. A5. D6. B or D7. B8. C9. D10.C11.D12.C13.A14.D15.C

CHAPTER 16

1a.C1b.B2a.D2b.A3a B






ANNUNCIATORS

The Annuncia tor Sect ion presents a colorrepresentation of all the annunciator lights inthe aircraft.

Please unfold pages ANN 3 or ANN 5 and leave

LEARJET 30 SERIES PILOT TRAINING MANUAL LEARJET 30 SERIES PILOT TRAINING MANUAL

CUR

LIM



AFOR TRAINING PURPOSES ONLYFOR TRAINING PURPOSES ONLY

CLOSECROSSFLOW

PITCH

TRIM

LIM

L CUR

LIMITER

R CUR

LIMITER

LOW

HYD

FUEL

XFLO

R LO

OIL

L LO

OIL

FMIZ

MMOM

MSTR

WARN

EMER

PWR 1

EMER

PWR 2

FIRE PULL FIRE PULL

ARMED ARMED ARMED ARMED

LH ENG

CHIP

RH ENG

CHIP

L PITOT

HEATR PITOT

HEAT

RIGHTTHRUST REVERSER

ARM ARM

OFF

TEST TEST

DEPLOY DEPLOYARM ARM

UNLOCK DEPLOYBLEED

VALVEUNLOCK DEPLOY

BLEED

VALVE

TEST THRUST NORM E ME R STO W R EV ERS ER EMER S

AERO

HDG REV GA FNL

GS

CAPT

GS

ARM

NAV

CAPT

NAV

ARM

FM/Z

OM MM

MSTR

WARN

ANTI-SKID

GEN

L R

TRK

APPR

HDG NAV REV LVL TEST ENG SOFT SPD V/S G/S ALT

AMR CAPT. FNL

G/A

PITCH IAS MACHPW R RO LLARM CAPT

BRT UPTEST

MUTE DN

LOCKED DNUNSAFE

EMPTYXFER

OFF

FILL

FULL

OPEN

CLOSE

FUS VALVE

FUS

TANK

A IR I GN L O N

OFF OFF OFF

BAT 2

START 1 START R

GEN

OFF

OFFOFF

OFF

GEN

RBUS

A UX I NV ER TE R I NV ER TE R

L BUS PRI SEC AIR IGN R

RGENRESET

LGENRESET

OFF

OFF

LEFT

DEE HOWARD TR 4000

NOTE: FOR FC-530 AUTOPILOT/FLIGHT DIRECTOR PANEL SEE CHAPTER 16

OR

OR

OR

AMK 80-17

FUEL

QUANTITY

LBSI x 1000

L ON R

L ON R

L ON R

JET PUMP

OPEN

FUEL TSN

EMPTY

XFER

OFF

FILL

FULL

F

U

S

T

A

N

K

0 0 0 03

2

1

0

45

6

7

8

L WING1254

RWING1254

L TIP1215

RTIP1175

TOTAL6238LBS

FUS1340

OR

DH

PRI

INV

LOW

FUEL

SEC

INV

L FUEL

PRESS

AUX

INV

R FUEL

PRESS

LO OIL

PRESS

SPOILER

STAB

OV HT

DOOR

WSHLD

OV HT

AUG

AIL

STEER

ON

PITOT

HT

BLEED

AIR L

FUEL

FILTER

BLEED

AIR R

L ENG

ICE

L

GEN

R ENG

ICE

R

GEN

L FUEL

CMPTR

CAB

ALT

R FUEL

CMPTR

WING

OV HT

L

STALL

WSHLD

HT

R

STALL

ALC

AI

L VG

MON

BAT

140

R VG

MON

BAT

160

MACH

TRIM

ENG

SYNC

DH

TO

TRIM

BAT 1

Figure ANN-1. Annunciators—FC200 Aircraft Only



Documents

Pilot Training Manual (FS)