cessna citation ii tm [pwd] dec 1999

8/20/2019 Cessna Citation II TM [PWD] Dec 1999

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Cessna Citation II Technical Manua

Chapter 1Aircraft General

Table of Contents

Overview ...............................................1-1

Publications ..........................................1-1

Airplane Equipment ...............................1-1

Airframe Structure .................................1-2

Fuselage...........................................1-3

Wings...............................................1-4

Empennage.......................................1-6

Nose Section ........................................1-7

Pressurized Center Section .................... 1-9

Flight Compartment .........................1-12

Passenger Cabin .............................1-16

Cabin Door and Stair Assembly ....... 1-20

Cabin/Cargo Door ........................... 1-26

Emergency Exit Door ........................... 1-29

Hand-Held Fire Extinguishers ............1-30

Aft Fuselage Section ........................... 1-31

Limitations..........................................1-32

Emergency Procedures.........................1-32



Cessna Citation II Technical Manual

AIRCRAFT GENERAL 12/99 FOR TRAINING PURPOSES ONLY 1-1

Overview

The Cessna Citation II is a high performance, twin-turbofan,pressurized, eight to twelve place airplane certificated in accor-

dance with FAR Part 25 airworthiness standards. The standardairplane is approved for operation in day, night, VFR, IFR, andknown icing conditions as defined by the FAA.

Publications

Cessna Aircraft Company publishes documentation providingdetailed airplane systems information and operating proce-dures. This Technical Manual is not intended to supersede the

Operating Manual, FAA approved Airplane Flight Manual(AFM), the Pilot’s Check List, and/or related publications spe-cific to your airplane.

Airplane Equipment

Airplane systems and equipment provided by the manufactureras standard from the factory, as well as manufacturer installed

optional systems or equipment will be covered in this publica-tion. Vendor supplied Supplemental Type Certificated (STC)accessories or equipment will not be covered. This chapterprovides a general description of the airplane structure, acces-sories, and equipment.



FOR TRAINING PURPOSES ONLY CITATION II CHAPTER 1 12/991-2

Airframe Structure

The Cessna Citation II is a low-wing airplane of primarilyall-metal construction with retractable landing gear andaft-mounted engines. The airframe is a conventionalsemi-monocoque design of aluminum alloy with

composite structures used in specific areas. Flushriveting, fairings, and aerodynamic joint sealants areused where appropriate to minimize drag.

Protection against electromagnetic interference (EMI) and radiofrequency interference (RFI) is accomplished primarily by theincorporation of bonding jumpers throughout the airframe. Aconductive finish applied to the inner surface of wing skin

panels provides additional protection against EMI and RFI.Protection against lightning strikes and accumulation of staticelectricity is accomplished by lightning strips installed on thenose cap (radome), and by static wicks installed on the trailingedges of the wings, flight control surfaces, and tailcone stinger.

FRONT

SPAR

STRINGER

RIB

FORWARD

SPAR

AFT

PRESSURE

BULKHEAD

WINGCARRY-THRU

SPARS

DORSAL

FIN

FORWARD

PRESSURE

BULKHEAD

FRAME

FRAME

STRINGER

RADOME

FLOOR PANEL/

RAIL ASSEMBLIES

NOSE

WHEEL WELL WINDOW FRAME

REAR

SPAR

ENGINE

PYLON

LIGHTNING

STRIPS

DOOR FRAME

STRINGER

ENGINE

CARRY-THRU

BEAMS

STATIC

WICKS

AILERON

SPEED BRAKE

FLAP

AFT

SPAR

MAIN

WHEEL WELL

ELEVATOR

RUDDER

AILERON

TRIM

RUDDER

TRIM

ELEVATOR

TRIM





Fuselage

The fuselage is an aluminum alloy, semi-monocoque structureconsisting of transverse frames and bulkheads, longitudinal

stringers, and external skin panels. The nose wheel well struc-ture, integral to the fuselage, provides attachment points for thenose gear assembly and related components. Frames andfittings are also provided for the attachment of doors and win-dows. Composite nose and tailcone fairings provide aerody-

namic smoothness and access to avionics components. Carry-thru spars pass laterally through the lower fuselage for attach-ment of the wings. Carry-thru beams pass laterally through theaft fuselage for attachment of the engines.

Chordwise ribs, spanwise stringers, and external skin panelsare fastened to the outboard front (main) and rear wing carry-thru spars to form the stub wings. The interior of each stub wingis sealed for fuel storage forward of the rear spar, between theinboard and outboard ribs. The fuel storage area is chemicallytreated and finished with an epoxy primer for corrosion resis-

tance. Aluminum alloy fairings provide aerodynamic smooth-ness between the fuselage and stub wing, as well as access towire bundles and various air and fluid lines.




Spanwise bulkheads, angles, and stringers; chordwise ribs,

and external skin panels are fastened to the outboard fore andaft engine carry-thru beams to form the engine pylons. Theoutboard ribs are constructed of stainless steel and sealed toform a firewall and vapor barrier through which control cables,wire bundles, and various air and fluid lines are routed to and

from the engines.

Access panels on the lower surface of the fuselage, stub

wings, and engine pylons facilitate inspection and mainte-nance. Drain holes are provided on the lower surface of thefuselage, stub wings, and pylons where fluids and/or moisturecollect. Drainage from unpressurized areas is continuous.Drainage from pressurized areas is regulated by check valveseals which are open only when the airplane is unpressurized.

Note: Moisture drain holes must be clear and free of obstruc-tions for proper operation.

The nose section, pressurized center section, and aft section ofthe fuselage are further described individually in this chapter.

Wings

The wings are aluminum alloy, semi-monocoque structuresconsisting of front (main) and rear spars, spanwise stringers,chordwise ribs, and external skin panels. The wings and stub

wings are mated and secured by threaded fasteners at attach-

ment points on the upper and lower front and rear spars. Mainwheel wells, integral to each wing structure, provide attachmentpoints for the main gear assemblies and related components.The skin panel directly above each main wheel well consists ofa honeycomb core material bonded between aluminum skinpanels. The interior of each wing, excluding the main wheelwell, is sealed for fuel storage forward of the rear spar, betweenthe inboard and outboard ribs. The fuel storage area is chemi-

cally treated and finished with an epoxy primer for corrosionresistance.





The wing structures also provide attachment points for the

ailerons, flaps, speed brakes, and their associated actuators.The ailerons and flaps are of aluminum alloy, semi-monocoqueconstruction incorporating spanwise spars, chordwise ribs, andexternal skin panels. The speed brakes are of aluminum-reinforced, magnesium alloy construction.

The outboard end of each wing is enclosed by a wing tip ofaluminum alloy, semi-monocoque construction incorporatingchordwise ribs, spanwise stringers, and external skin panels.Wing tip fairings may be of aluminum alloy or composite con-struction.

The inboard leading edge of each wing is formed by an electri-cally-heated anti-ice panel. A Kevlar insulation shield provides

a thermal barrier between the heated leading edge panel andthe wing structure. Pneumatic deice boots are installed on theoutboard leading edge of each wing. Access panels on thelower surface of the wings facilitate inspection and mainte-

nance of control surface actuators and fuel systemcomponents.




Empennage

The empennage is an aluminum alloy, semi-monocoque struc-ture consisting of the vertical stabilizer, horizontal stabilizer,and associated control surfaces. The vertical stabilizer is con-structed primarily of fore and aft spars, chordwise ribs, andexternal skin panels. The fore and aft spars are secured by

threaded fasteners to the fuselage structure. The horizontalstabilizer is constructed primarily of fore and aft spars,spanwise auxiliary spars, chordwise ribs, and external skin

panels. The fore and aft spars are secured by threaded fasten-ers to the vertical stabilizer. Attachment points are provided forthe rudder, elevators, and their associated actuators. Therudder and elevators are also constructed of spars, ribs, andexternal skin panels.

A composite dorsal fin and saddle fairing provide aerodynamicsmoothness between the upper fuselage and the vertical stabi-lizer. The emergency locator transmitter is housed within thesefairings. Vertical and horizontal stabilizer tip fairings (caps) may

be of aluminum alloy or composite construction.

Pneumatic deice boots are installed on the leading edge of thevertical and horizontal stabilizers. Access panels facilitate

inspection and maintenance of control surface actuators andsome navigational components. The access panels located onthe vertical stabilizer directly below the horizontal stabilizerconsist of a honeycomb core material bonded between alumi-num skin panels.





Nose Section

The unpressurized nose section contains the avionics bay andthe nose baggage compartment.

The avionics bay is locatedjust aft of the fiberglass nosecap (radome) which housesthe weather radar antenna. An

avionics access panel, con-structed of honeycomb corematerial bonded betweenaluminum skin panels, en-closes the width of the uppersurface of the nose sectionbetween the nose cap and thenose baggage compartment.The access panel and nose

cap form a single unit attached by quick-disconnect “Tridair”

fasteners and secured by two key locks, one installed on eachside of the panel.

Note: The avionics access panel and nose cap are removedand installed as a single unit. The double row of fastenersshould not be disturbed unless separation of the nose cap fromthe access panel is required for maintenance.

AVIONICS BAY




The nose baggage compart-

ment is located between theavionics bay and forwardpressure bulkhead. Access isthrough left and right sidedoors attached to the fuselage

structure by two hinges each,and secured by two latchesand one key lock each. Each

door is opened by disengag-ing the key lock and latchesand lifting the door to the openposition. When fully open, aspring-loaded stop assembly holds the door in position. Clos-ing the doors is accomplished by releasing the stop, loweringthe door, and reengaging the latches and key lock.

Fore and aft dividers separate the nose baggage compartment

from the avionics bay and forward pressure bulkhead respec-tively. A hinged access panel on the aft divider, accessiblethrough the right baggage door, facilitates fluid and pneumaticservicing. Five inspection windows on the access panel arepositioned to permit viewing the sight gages on the brake fluidand windshield anti-ice fluid reservoirs, as well as pneumaticpressure gages for the emergency braking and gear extensionstorage cylinder and anti-skid accumulator(s). On airplanes

550-0254 and earlier (not incorporating SB550-35-2) the oxy-

gen storage cylinder is also serviced and accessed throughthe right baggage door. In some installations, a hinged accesspanel may be located below this door to facilitate oxygenservicing.

Though limits vary with equipment installation, the maximumvolume and load capacity of the nose baggage compartmentare 17 cubic feet and 350 pounds respectively.

Indication of nose baggage door security is provided by theamber [DOOR NOT LOCKED] annunciator. The annunciator iscontrolled by a microswitch integral to each forward latchassembly. With electrical power applied to the airplane andeither door unsecured, the door warning circuit is complete andthe annunciator illuminates. When both doors are properlysecured, the door warning circuit is interrupted and the annun-

ciator is extinguished.

NOSE BAGGAGE COMPARTMENT





Pressurized Center Section

The airplane center section is reinforced and sealed for pres-surization to the skin between the forward and aft pressurebulkheads. Included in the center section are the flight com-partment, passenger cabin, standard cabin door, optionalcabin/cargo door (if installed), and the emergency exit.

A two-piece windshield, two

side windows, and a foulweather window provide flightcompartment visibility. Thewindshield is a Plexiglaslaminate of stretched acrylicouter and inner layers with avinyl core.

PILOT'S WINDSHIELD




On airplanes 550-0681 and earlier, both side windows consist

of prestressed Plexiglas outer and inner panes separated bythe window frame and spacers to form a thermal air barrier. Onairplanes 550-0682 and after, both side windows consist ofprestressed Plexiglas outer and middle panes with an acrylicinner “frost” pane. A thermal air barrier is formed between the

outer pane and middle pane, separated by the window frameand spacers; and between the middle pane and frost pane,separated by spacers and seals.

Six windows are located on each side of the passenger cabinincluding those located in the cabin door and the emergencyexit. Each includes a Plexiglas laminate outer pane of stretchedacrylic outer and inner layers with a vinyl core, and an acrylicinner “frost” pane. The outer pane and frost pane are sepa-rated by the window retainer and a spacer to form a thermal air

barrier. All passenger cabin windows, excluding those locatedin the passenger door and the emergency exit, incorporate

integral sliding shades. All windows, forming part of the pres-sure vessel, are fixed except for the foul weather window.

The foul weather window,located forward of the pilot’sside window, is also aPlexiglas laminate of stretchedacrylic outer and inner layers

with a vinyl core. The window

is hinged at the bottom,latched at the top, and incor-porates a peripheral sealwhich makes the windowairtight when properly se-cured. The window is openedby disengaging the latch andpulling the window inward. Closing and securing the window is

accomplished by repositioning the window in the frame andreengaging the latch.

During ground operation, the foul weather window may beopened to supplement flight compartment ventilation. Duringcold weather operation, the flight compartment may be warmedbefore flight by routing a preheater hose through the foulweather window. For pressurized operation, the foul weather

window must be properly secured.

FOUL WEATHER WINDOW




Flight Compartment

The flight compartment provides for a crew of two, with fullinstrumentation for the pilot and copilot. Layout is conventionalin that all controls, switches, and instruments are accessible tothe pilot for single pilot operation.

The pilot’s and copilot’s instrument panels contain primary andnavigational flight instruments and controls. Flood and panellight controls, light switches, and the standby gyro switch/

light(s) are located on the pilot’s lower instrument panel. Con-trols for the parking brake, control surface lock, emergencybraking, and auxiliary/emergency gear extension are locatedbelow the pilot’s instrument panel. Windshield bleed air con-trols, fan switches, and gyro switches are located on thecopilot’s lower instrument panel. The windshield rain removalaugmenter control is located below the copilot’s instrumentpanel.





The upper left instrument panel contains the pilot’s audio con-trol panel, outside air temperature (OAT) indicator, clock, andoptional angle-of-attack (AOA) indicator. The lower left instru-

ment panel contains electrical system switches and indicators,fuel system switches and controls, engine switches, ice protec-tion switches, exterior light switches, and the test selector

switch. The right instrument panel contains the copilot’s audiocontrol panel, battery temperature indicator (if installed), gyropressure gage (550-0626 and earlier), oxygen pressure gage,and flight hours meter.

The center instrument panel contains engine instrumentation,

the annunciator panel, avionics control panels, weather radar,supplemental navigation equipment, and the landing gearcontrols and position indicators. Circuit breaker panels arelocated on the left and right sidewalls. Crew oxygen outlets and

audio jacks are located on the left and right side consoles.Oxygen system controls are located on the left side console.Controls and indicators for the engine fire protection systemand thrust reversers are located on a panel directly below the

glareshield.

The center pedestal contains the engine control levers as wellas controls for the flaps, speed brakes, manual trim, autopilot,and navigation equipment. The environmental panel containscabin pressurization switches, controls, and indicators. Refer tocorresponding systems chapters for specific detail.




1. Height Adjustment Handle

2. Fore and Aft Position Handle

3. Seat Back Tilt Handle

The pilot’s and copilot’s seats are installed on independent railassemblies attached to longitudinal floor beams integral to thefuselage structure. Fore and aft position, height, and tilt angleare manually adjustable. Each seat moves fore and aft alongthe rails on roller and guide assemblies attached to its frame.Stop pins engage the rails to secure the seat in position. Fore

and aft adjustment is accomplished by lifting the handle lo-cated below the forward center of the seat to disengage thestop pins from the rails, and sliding the seat to the desiredposition. Height adjustment is accomplished by lifting thehandle located below the inboard forward corner of the seat

and weighting or unweighting the seat to the desired position.A shock cord (bungee) and pulley arrangement provides ap-proximately 100 pounds of lift assistance when the seat is

unweighted. Tilt angle adjustment is provided by a pneumaticactuator attached to each seat frame, and a handle located onthe aft inboard corner of the seat. Pushing down on the forwardend of the handle releases air pressure within the actuatorallowing the seat to be tilted to the desired position. When thehandle is released, air pressure trapped within the actuatorholds the seat in the selected position.

Armrests are installed on each inboard seat back. When not in

use, the armrests are stowed in an upright position behind andflush with the seat backs. For use, each armrest is pulled in-board from its stowed position and lowered by pushing downon its forward end. On airplanes 550-0222 and after, the low-ered position of the armrest may be selected by means of anadjustable stop. This adjustment, however, cannot be madeduring flight.

1

2 3





Each crew seat is equipped

with a restraint system incor-porating an inertia reel typeshoulder harness. Airplanes550-0431 and after areequipped with a five-point

restraint system consisting oftwo lap belts, a dual-strapshoulder harness, forward

restraint strap, and paddedrotary buckle. Airplanes 550-0356 through 550-0430 areequipped with a four-pointrestraint system consisting of two laps belts, a dual-strap shoul-der harness, and padded rotary buckle. In four-point and five-point installations, the inertia reel is attached to the seat frame

and all restraints engage the rotary buckle. The quick-releasedesign of the rotary buckle permits simultaneous disengage-

ment of all restraints except the outboard lap belt, to which thebuckle is attached, by rotating the release mechanism counter-clockwise. Airplanes 550-0355 and earlier are equipped with athree-point restraint system consisting of two laps belts, asingle-strap shoulder harness, and conventional buckle. In thisinstallation, the inertia reel is attached to the overhead airframestructure aft and outboard of each crew seat. The shoulderharness engages a link on the outboard lap belt which en-

gages the buckle on the inboard lap belt.

Other flight compartmentequipment and furnishingsinclude a navigation chartcase located behind thecopilot’s seat, a relief tubestorage case located behindthe pilot’s seat, sun visors,

overhead directional air vents,and overhead flight compart-ment lighting. The relief tubeincorporates an electrically-heated drain/vent which issupplied with 28 VDC rightmain bus power through the 7.5-amp TOILET circuit breaker(not accessible from the flight compartment). On airplanes 550-

0627 and after, optional tinted sun visors positionable along amonorail track may be installed in place of the standard vinyl-covered “pivoting” sun visors.

CREW RESTRAINT SYSTEM

DIRECTIONAL AIR VENT AND LIGHT




Passenger Cabin

The passenger cabin measures approximately 194.7" from theforward cabin dividers to the aft pressure bulkhead, 59.2" fromsidewall to sidewall, and 56.0" from the lowest point of the floorto the ceiling. Passenger cabin configurations vary accordingto seating arrangement and installation of standard or optional

furnishings and equipment.

AFT PASSENGER CABIN

Standard configurations typi-cally include seating for sixpassengers, a refreshmentcenter, and a non-flushing

toilet. Optional configurationsinclude seating arrangementsfor up to ten passengers, aflush toilet, an executive writ-ing table, and various storagecabinets and/or refreshment

centers. Forward and aftdividers separate the passen-

ger area from the flight com-partment and aft baggagecompartment respectively.

REFRESHMENT CENTER





EXECUTIVE WRITING TABLE

The standard passenger seats, numbered 5 through 10 forpayload computation purposes, are installed on independentrail assemblies attached to longitudinal floor beams integral tothe fuselage structure. These seats may be high back, or lowback with an adjustable headrest. Seats 5 and 6, located

midcabin, may be installed facing forward or aft. Seats 7 and 8,located immediately forward of the aft cabin dividers, are in-

stalled forward facing only. Seats 9 and 10, located immedi-ately aft of the forward cabin dividers, are installed aft facingonly.

Each seat moves fore and aft along the rails on roller and guideassemblies attached to its frame. Stop pins engage the rails tosecure the seat in position. Fore and aft adjustment is accom-

plished by lifting the handle located below the forward center ofthe seat to disengage the stop pins from the rails, and slidingthe seat to the desired position. When located “over spar,”seats 7 and 8 are secured directly to the rails and are not ad-

justable fore and aft once installed.

Seat back angle adjustment from vertical to a reclining positionis provided by spring-loaded “hydrolock” actuators attached to

each seat frame. Pressing the button on the inboard side of thearmrest releases air pressure within the actuators allowing theseat back to be tilted to the desired position. When the button isreleased, air pressure trapped within the actuators holds theseat back in the selected position.




Passenger seats equipped for lateral adjustment may be posi-

tioned inboard away from the sidewalls to provide greaterheadroom. This is accomplished by lifting the control handlelocated below the inboard side of the seat and moving the seatto the desired position.

Note: Lateral adjustment seats should be locked in the out-board position during takeoff and landing.

Armrests are installed on each inboard seat frame. When not inuse, armrests are stowed flush with the bottom seat cushions.For use, each armrest is pulled upward until a spring-loadedlatching mechanism is engaged. Lifting the lever on the forwardend of each armrest disengages the latching mechanismpermitting the armrest to be stowed. Each passenger seat isequipped with a restraint system consisting of an adjustable

lap belt and inertia reel shoulder harness (550-0550 and after)or an adjustable lap belt only (550-0505 and earlier).

Optional passenger seatingarrangements may includetwo additional standard typeseats located at midcabin andnumbered 3 and 4 for payloadcomputation purposes, an aftportable seat, forward lounge

seats, and a two or three-

place forward facing divan(couch). The optional flushtoilet may be certified for useas a passenger seat whenlocated in the aft baggagecompartment (550-0550 and after) or when located in theforward passenger cabin (550-0505 and earlier). Non-flushingtoilets are generally not certified for use as a passenger seat.

Aft located toilets are also equipped with a relief tube incorpo-rating an electrically-heated drain/vent which is supplied with28 VDC right main bus power through the 7.5-amp TOILETcircuit breaker (not accessible from the flight compartment).Flush toilets are also powered by this circuit. Privacy for the afttoilet area may be provided by curtains, sliding doors (550-0550 and after), or a folding door (550-0505 and earlier) whichextend(s) between the left and right aft cabin dividers.

AFT FLUSH TOILET





1. Passenger Oxygen Masks

2. Reading Light

3. Reading Light Switch

4. Ventilation Air Outlet

Individually controlled readinglights, ventilation air outlets,and oxygen outlets are lo-

cated on the ceiling aboveeach passenger station.Airflow for heating is providedthrough registers located justabove floor level on each side ofthe cabin.

1

2

3

4




The aft baggage compartment is located between the aft cabin

dividers and the aft pressure bulkhead, and is accessibleduring flight. A cargo net and tie-down straps are provided tosecure the contents of the baggage compartment. The net isheld in place by attachment fittings that engage anchor plateslocated on the floor, aft pressure bulkhead, and sidewalls.

Though limits vary with passenger cabin configuration, themaximum volume and load capacity of the aft baggage com-partment are from 34 to 43 cubic feet and 220 to 600 pounds

respectively.

Note: Refer to the appropriate AFM and airframe placards forweight and balance limitations specific to your airplane.

Cabin Door and Stair Assembly

The cabin door and stair assembly are located in the forwardleft side of the center section. The door is a single-sectionassembly of aluminum alloy construction which houses theforward left passenger cabin window and, when secured with

all locking pins engaged, forms an integral part of the pressurevessel. The door swings forward to the open position on a

vertical hinge attached to the forward door frame structure andprovides an opening 50.7” high, 23.5” wide at the bottom, and19.7” wide at the top. The stair assembly is attached at twohinge points on the lower door frame structure and incorpo-rates two fold-out steps.





The cabin door is held securely closed by twelve locking pins

which engage sockets in the door frame structure when theinterior or exterior handle is rotated to the “LOCK” position.When either handle is rotated to the “OPEN” position, the lock-ing pins are disengaged. The overcenter locking design of thedoor handle linkage combined with an interior handle latching

mechanism function to prevent inadvertent opening of the doorparticularly from inside the airplane.

The overcenter locking designrequires that either handle berotated fully to the “LOCK” or“OPEN” position before beingstowed. The latching mecha-

nism secures the interiorhandle in the “STOW” position.Before rotating the interiorhandle to the “LOCK” or

“OPEN” position, the latchmust be disengaged by simul-taneously pressing a releasebutton and squeezing a trig-ger located on the top and backside of the handle respectively.

UPPER LOCKING PINS LOWER LOCKING PINS

INTERIOR HANDLE RELEASE




The cabin door is opened

from outside the airplane byunstowing the flush-mountedexterior handle, rotating itclockwise to the “OPEN”position, and pulling the door

outward and forward. A secu-rity lock prevents the exteriorhandle from being unstowed

without the appropriate key.From inside the airplane, thedoor is opened by disengag-ing the interior handle latchingmechanism as previouslydescribed, rotating the handlecounterclockwise to the

“OPEN” position, and pushingthe door outward and forward.

When fully open, a spring-loaded door stop/catch as-sembly, integral to the hinge,holds the door in position. Thecatch is disengaged by push-ing a release lever locatedimmediately forward of thedoor frame on the passenger

cabin sidewall.

EXTERIOR CABIN DOOR HANDLE

CATCH RELEASE LEVER

INTERIOR CABIN DOOR HANDLE




1. Upper Locking Pin

Indicator Windows

2. Lower Locking Pin

Indicator Windows

3. Overcenter Linkage

Indicator Window

Five position indicator win-

dows are provided on theinterior door trim panel tovisually confirm cabin doorsecurity. The upper two andlower two indicator windows permit inspection of locking pinengagement. The center indicator window permits inspection of

the overcenter locking position of the door handle linkage. Tofacilitate inspection, the visible portion of the upper two and

lower two “square” locking pins are green with white dots. Withthe cabin door properly secured, the white dot on each ofthese locking pins should be visible in its respective window.Through the center window, the alignment of two horizontalmarkings above the words DOOR CLOSED should be visible.

1

2

3





1. Cabin Door Microswitch

2. Inflatable Rubber Door Seal

Additional indication of cabin door security is provided by the

amber [DOOR NOT LOCKED] annunciator. The annunciator iscontrolled by a microswitch, installed on the door frame struc-ture, which is actuated by the lower forward locking pin. Withelectrical power applied to the airplane and the door unse-cured, the door warning circuit is complete and the annunciatorilluminates. When the door is properly secured, the door warn-

ing circuit is interrupted and the annunciator is extinguished.

An inflatable rubber seal is installed in a retainer around theperiphery of the door frame to enhance fuselage-to-door seal-ing. The seal is inflated by 23 ± 1 PSIG regulated engine bleed

air by way of a valve which is actuated open or closed by thelower forward locking pin. With either or both engines operatingand the door properly secured, the valve is actuated open and

the seal is inflated. A check valve prevents the seal from deflat-ing should loss of bleed air pressure occur. When the lockingpin is disengaged, the valve is actuated closed and the seal isdeflated to facilitate opening and closing of the cabin door.Additional protection against cabin door leakage is provided bya weather seal affixed to the inboard periphery of the doorexcept in the area of the hinge which is protected by a sepa-

rate rain seal.

1

2




Cabin/Cargo Door

The optional cabin/cargo door may be installed in place of thestandard cabin door in the forward left side of the center sec-tion. The door is a two-section assembly of aluminum alloyconstruction which, when secured with all locking mechanismsengaged, forms an integral part of the pressure vessel. The

upper section of the door swings upward to the open positionon a full length, piano-type hinge attached to the upper doorframe structure and houses the forward left passenger cabin

window. The lower (airstair) section of the door swings down-ward to the open position on a full length, piano-type hingeattached to the lower door frame structure and incorporatesthree fold-out steps. When fully open, the cabin/cargo doorprovides an opening 50.7” high and 35.12” wide.

The upper and lower sections of the door are each held se-

curely closed by six cable-operated latch fittings which engageeccentric latch posts on the door frame structure. Though each

door section is operated independently of the other, two lock-ing pins integral to the upper door handle linkage secure bothsections together. The overcenter locking design of the upperdoor handle linkage functions to prevent inadvertent opening ofthe door, particularly from inside the airplane by requiring thatthe interior or exterior handle be rotated fully to the “LOCK” or“OPEN” position before being stowed.

The upper section is opened from outside the airplane by

unstowing the flush-mounted exterior handle, rotating it clock-wise to the “OPEN” position to disengage the latch fittings andlocking pins, and pulling the door outward. A security lockprevents the exterior handle from being unstowed without theappropriate key. From inside the airplane, the upper section isopened by unstowing the interior handle, rotating it counter-clockwise to the “OPEN” position, and pushing the door out-ward. A pair of gas-operated extenders assist in opening the

upper section and hold it in position when fully open. A lockingmechanism on the forward extender stabilizes the upper sec-tion when fully open in windy conditions. The upper sectionmust be opened before the lower section is extended.





With the upper section open, lifting the lower handle disen-

gages the latch fittings allowing the lower section to be ex-tended. The steps are cable-operated and fold out automati-cally during extension. When fully extended, the lower sectionis supported by two cable assemblies attached to fittings onthe fore and aft door frame structure. A spring-loaded reel

assembly automatically retracts the support cables when thelower section is closed. A gas-operated snubber dampens theextension rate of the lower section. Cable assembly tension

should be inspected regularly to confirm that each carries anequal load with weight on the steps and that no load is carriedby the snubber.

Note: Uneven cable assembly tension or load on the snubbershould be corrected to prevent damaging the lower sectionattachment or support points on the door frame structure.

The cabin/cargo door is closed by lifting the lower section to

position it in the frame and returning the lower handle to theclosed position to engage the latch fittings. A T-handle on theaft support cable facilitates closing the lower section frominside the airplane. As the lower section is closed, the cable-operated stairs are automatically stowed against the interiortrim panel. With the lower section secured, the upper section ispulled down and positioned in the frame, and the upper handleis rotated (exterior-counterclockwise, interior-clockwise) to the

“CLOSE” position to engage the latch fittings as well as the

locking pins which secure both sections together.

Note: After closing the upper section from outside or inside theairplane, the handle used must be returned to the stowedposition.




Position indicator windows are provided on the upper and

lower section interior trim panels to visually confirm positiveengagement of each latch fitting. To facilitate inspection, thevisible portion of a stop arm on each latch fitting shaft is greenin color. With the cabin/cargo door properly secured, each stoparm should appear as a green vertical bar within its respective

window. An additional indicator window on the upper sectioninterior trim panel permits inspection of the overcenter lockingposition of the upper door handle linkage. Through this window,

the alignment of two vertical markings between the wordsDOOR CLOSED should be visible.

Additional indication of cabin door security is provided by theamber [DOOR NOT LOCKED] annunciator. The annunciator iscontrolled by five microswitches: two actuated by the uppersection latch fittings linkage, one actuated by a lower section

latch fitting linkage, and two actuated by the locking pins. Withelectrical power applied to the airplane and the upper or lower

section of the door unsecured, the door warning circuit iscomplete and the annunciator illuminates. When both sectionsare properly secured, the door warning circuit is interruptedand the annunciator is extinguished.

An inflatable rubber seal is installed in a retainer around theperiphery of the door frame to enhance fuselage-to-door seal-ing. The seal is inflated by 23 ± 1 PSIG regulated engine bleed

air by way of a valve which is actuated open or closed by the

upper door handle linkage. With either or both engines operat-ing and the upper section of the door properly secured, thevalve is actuated open and the seal is inflated. A check valveprevents the seal from deflating should loss of bleed air pres-sure occur. When the upper section is unsecured, the valve isactuated closed and the seal is deflated to facilitate openingand closing of the cabin/cargo door. A “cross seal,” installedon the lower door section, seals the gap between the upper

and lower section.





Emergency Exit Door

EMERGENCY EXIT EXTERIOR EMERGENCY EXIT INTERIOR

The emergency exit door is located in the forward right side of

the center section. The door is a single-section assembly ofaluminum alloy construction which houses the forward rightpassenger cabin window and, when properly secured, formsan integral part of the pressure vessel. The plug-type door isinstalled from inside the airplane and secured in place by twofixed retainers and a latch pin which engage the door framestructure. The latch pin is operated by an interior or exteriorhandle.

To prevent inadvertent open-ing of the door, particularly

from inside the airplane, thelatching mechanism is spring-loaded to the closed position,and the interior handle isguarded by a plastic cover.Ground security is provided

by a locking pin which pre-vents the latching mechanismfrom being operated wheninserted. The locking pin

incorporates a REMOVEBEFORE FLIGHT streamer.

EMERGENCY EXIT INTERIOR HANDLE




The emergency exit door is

opened from inside the air-plane by removing the lockingpin and the plastic cover (ifinstalled), then unstowing theinterior handle and rotating it

clockwise to the “OPEN”position. From outside theairplane, the flush-mounted

exterior handle must beunstowed and rotated coun-terclockwise to the “OPEN”position. With the latch pindisengaged, the top of the door is lowered until clear of theupper frame and the bottom of the door is lifted until the retain-ers are clear of the lower frame, thereby allowing the door to be

removed. Closing the door is accomplished from inside theairplane by carefully positioning the retainers on the bottom of

the door in the lower frame, pushing the top of the door intoposition in the upper frame, and rotating the interior handlefully-counterclockwise to engage the latch pin. A one-piece,self-inflating rubber seal is affixed to the inboard periphery ofthe door to enhance fuselage-to-door sealing. The seal isinflated by cabin pressure entering the seal through a series ofholes.

Note: After closing the cabin door from outside or inside the

airplane, the door handle used must be returned to the stowedposition.

Hand-Held Fire Extinguishers

Typically, two hand-held fireextinguishers are provided:one is secured to a bracket onthe copilot’s seat frame, the

other is located in the passen-ger cabin. Refer to the AFMemergency procedures andthe instructions printed on thecylinder for proper operation.

EMERGENCY EXIT EXTERIOR HANDLE

HAND-HELD FIRE EXTINGUISHER





Aft Fuselage Section

The unpressurized aft fuselage section houses control cablesand linkage for the tail control surfaces and the engines, andsuch major components as the autopilot servos, the battery,electrical power junction “J” boxes, the external power recep-tacle, the environmental air cycle machine (ACM), and theoxygen storage cylinder (550-0255 and after, or earlier air-planes in compliance with SB550-35-2). When optionally in-

stalled, the drag chute, components of the Freon air condition-ing system, and the tailcone baggage compartment are alsohoused within the aft fuselage.

Access to these componentsand to the tailcone baggagecompartment is through adoor attached to the lower leftside of the aft fuselage struc-

ture by a piano-type hinge.

The door swings downward tothe open position and is se-cured by two hook-typelatches and a key lock.

TAILCONE ACCESS DOOR




The tailcone baggage compartment consists of floor, side, and

ceiling panels attached to a support structure. A hinged ac-cess panel facilitates inspection and servicing of componentslocated forward of the baggage compartment. To reduce therisk of a baggage compartment fire, the structural elements ofthe baggage compartment and the inner surface of the door

are treated with a flame control coating.

A cargo net and tie-down straps are provided to secure the

contents of the baggage compartment. The net is held in placeby attachment fittings which engage anchor plates located onthe floor and aft panel of the baggage compartment. Thoughlimits vary with equipment installation, the maximum volumeand load capacity of the tailcone baggage compartment rangefrom 13 to 25 cubic feet and 200 to 500 pounds respectively.

Note: Refer to the appropriate AFM and airframe placards forweight and balance limitations specific to your airplane.

Indication of tailcone access door security is provided by theamber [DOOR NOT LOCKED] annunciator. The annunciator iscontrolled by a microswitch integral to the forward latch assem-bly. With electrical power applied to the airplane and the doorunsecured, the door warning circuit is complete and the annun-ciator illuminates. When the door is properly secured, the doorwarning circuit is interrupted and the annunciator is extin-

guished.

Limitations

Refer to the applicable airplane manufacturer’s FAA approvedflight manual or approved manual material, markings andplacards, or any combination thereof for all limitations.

Emergency Procedures

Refer to the applicable airplane manufacturer’s FAA approvedflight manual or approved manual material (supplementarychecklist) as revised, for procedural information.





14.80’

51.7’

17.59’

8.0’

18.83’

47.25’

18.35’

Citation II Airplane Dimensions




Chapter 2Engines


Table of Contents

Overview ....................................................2-1

Engine Installation ...................................... 2-1

Operational Theory andPower Ratings ............................................ 2-3

Engine Description...................................... 2-4

Low Compressor Section....................... 2-4

High Compressor Section ...................... 2-5

Combustion Section .............................. 2-7

High Turbine Section ............................. 2-8

Low Turbine Section.............................. 2-8

Accessory Gearbox ................................ 2-9

Engine Fuel System..................................2-10

Engine Driven Fuel Pump .................... 2-11

Fuel Control Unit ................................. 2-12

Oil-to-Fuel Heat Exchanger ................... 2-15

Flow Divider ........................................ 2-16

Fuel Manifold and Nozzles .................. 2-18

Fuel Drains ......................................... 2-18

Ignition System ........................................ 2-19

Engine Oil System .................................... 2-22

Oil Tank ............................................. 2-22

Oil Pressure System ........................... 2-23

Scavenge System ............................... 2-25

Breather System ................................. 2-25

Engine Controls ........................................ 2-26

Throttle Levers .................................... 2-26

Thrust Reverser Levers........................ 2-27

Engine SynchronizerSelector Switch ................................... 2-28

Automatic Fuel Shutoff Control ............ 2-28




Engines, continued

Table of Contents

Engine Indicating System .......................... 2-29

Fan Tachometer ..................................2-29

Inter-Turbine Temperature Gage ........... 2-30

Fuel Flow Gage ................................... 2-32

Turbine Tachometer............................. 2-32

Engine Oil Temperature Gage .............. 2-33

Engine Oil Pressure Gage.................... 2-33

Engine Starting System............................. 2-35

Engine Synchronizer System...................... 2-38

Thrust Reverser System............................ 2-40

Thrust Reverser Hydraulics........................ 2-40

Thrust Reverser Control Valves ............ 2-41

Thrust Reverser Isolation Valves .......... 2-41

Check Valves and Restrictors .............. 2-42

Thrust Reverser Levers........................ 2-42

Thrust Reverser Deployment ................ 2-43

Thrust Reverser Stowage..................... 2-45

Thrust ReverserEmergency Stowage ............................ 2-46

Thrust ReverserEmergency Stowage Test..................... 2-46

Engine Fire Detection andExtinguishing System ................................ 2-47

Thermal Detectors............................... 2-47

Detector Control Unit .......................... 2-47

Explosive Cartridges ............................ 2-47

Extinguisher Bottles ............................ 2-48

Annunciator Switches .......................... 2-48

System Operation ............................... 2-48

System Testing ................................... 2-51

Limitations ............................................... 2-51

Emergency Procedures.............................. 2-51




ENGINES 12/99 FOR TRAINING PURPOSES ONLY 2-1

Overview

This chapter will discuss the engines of the Cessna Citation II,their major components, controls, operation and indication. Theairplane is powered by two JT15D-4 lightweight, medium-bypass, axial-flow, turbofan engines manufactured by Pratt &Whitney Canada, Inc. The engines generate thrust which pro-

pels the airplane and rotational torque which powers all engine-driven accessories. Indication of engine operating parametersis displayed by gages on the upper center instrument panel.Other indications are displayed on the annunciator panel.

Engine Installation

The engines are attached to left and right pylons formed by

carry-thru beams which pass laterally through the aft fuselage.Each engine is secured to its associated pylon by two forwardmounts and one aft “steady” mount. The mounts incorporateisolators which function to reduce the transmission of engine

vibration through the fuselage structure. Each forward mount issecured by four bolts to the engine and one bolt to the forwardcarry-thru beam. Each steady mount is secured by four bolts tothe aft carry-thru beam and one bolt to the engine.




The outboard rib of each pylon is constructed of stainless steel

and sealed to form a firewall and vapor barrier through whichcontrol cables, wire bundles, and various air and fluid lines arerouted to and from the engines.

1. Engine Air Inlet Lip 4. Aft Cowling

2. Forward Cowling 5. Aerodynamic Tangs

3. Upper Cowling 6. Lower Cowling

The engines are enclosed by cowlings which provide aerody-namic smoothness. The forward cowling houses the engine airinlet lip and starter/generator cooling air inlet scoop and duct.

The aft cowling surrounds the outer exhaust nozzle and incor-

porates aerodynamic “tangs” that enclose the thrust reverseractuators when installed.

The upper and lower cowlings are attached to the forward andaft cowlings by quick-disconnect fasteners to facilitate removalfor inspection or servicing. The upper cowling incorporates the

oil filler access door. The lower cowling incorporates thestarter/generator cooling air outlet and an aerodynamic fairingwhich encloses various engine drain lines that extend throughits lower surface. To reduce the transmission of engine noise tothe cabin, the forward and aft cowlings are attached directly tothe engine such that they make no contact with the pylon.

2 1

5

4

3

6





Operational Theory and Engine Power Ratings

The power output of a jet engine, expressed in pounds ofthrust, is determined by the velocity to which it is capable ofaccelerating a given mass of air. As a rule, a smaller volume ofair must be accelerated to a higher velocity than that required

to generate the same amount of thrust from a larger volume ofair. In a turbojet engine, thrust is generated by compressingand combusting the entire volume of inlet air, and discharging

it from the exhaust nozzle at high velocity. In a turbofan engine,a portion of the thrust generated by compressing and combust-ing a smaller volume of air (in the same manner as a turbojetengine) is utilized to drive a fan which generates thrust from alarger volume of air at a lower velocity. This is accomplished bydividing inlet air drawn into the engine by the fan into primary(inner) and secondary (outer) paths.

The primary path directs a smaller volume of air through the

engine core where it is compressed, combusted, and acceler-ated to a higher velocity. In the JT15D-4, approximately 42% ofthis “core thrust” is used drive the compressor, fan, and acces-sory gearbox while the remaining 58% is discharged from theinner exhaust nozzle. The secondary path directs a largervolume of non-combusted air through a concentric bypass ductsurrounding the engine core and discharges it from the outerexhaust nozzle.

Of the 2500 LBS thrust produced by each engine at sea level,approximately 66% is “bypass thrust” generated by the fanwhile 34% is generated by the core. However, because fanefficiency decreases as altitude increases, the ratio of bypassthrust to core thrust (thrust ratio) progressively reverses be-tween sea level and approximately 40,000’ MSL.

The relationship between mass airflow through the bypass duct

and mass airflow through the engine core determines theengine’s “bypass ratio.” Of the total 77.8 LBS/SEC mass air flow(Wa) developed by the fan, 56.6 LBS/SEC is directed throughthe bypass duct while 21.2 LBS/SEC is directed through thecore thereby producing the engine’s 2.7:1 bypass ratio. Unlikethrust ratio, bypass ratio remains constant regardless of factorswhich influence air density such as altitude, temperature, andairspeed.




Power ratings are also based on the engine’s sea level staticoutput in ISA conditions. The power ratings and leading par-ticulars given for the JT15D-4 in the table below are for trainingpurposes only and are not DOT approved.

Operating Net Thrust Specific Fuel

Condition LBS LBS/HR/LBS THRUST

Takeoff 2500** 0.562

Max. Continuous 2375* 0.556Max. Climb 2375 0.556

Max. Cruise 2345 0.555

* Available to 15.0°C (59.0°F)** Available to 15.0°C (59.0°F) time limited to 5-minutes

Engine Description

This section is intended to provide a basic description of theJT15D-4 engine. Each engine is functionally divided into inde-pendent high and low pressure sections which rotate on con-centric, bearing supported shafts.

The outer, high pressure (N2) shaft supports the high turbine

and high compressor, associated primarily with the generationof core thrust. The inner, low pressure (N1) shaft supports thelow turbines, low compressor (fan), and booster stage, associ-ated primarily with the generation of bypass thrust. Both shafts

rotate in a clockwise direction at different speeds. In that nomechanical link exists between the shafts, the engine is classi-fied as a free-turbine.

The engine is further divided into six principle sections as

follows:

Low Compressor Section

The low compressor section consists of the fan, booster stage,and associated stator assemblies. The fan is installed on theforward end of the N1 shaft and incorporates 28 blades that

function to induct and compress all inlet air entering the en-gine. The induction of inlet air is optimized by the aerodynamic

shape of the nose cone attached to the front of the fan. Imme-diately aft of the fan, compressed inlet air is divided into con-centric primary (inner) and secondary (outer) paths.





The primary path directs a smaller volume of air through a rowof 38 core inlet guide vanes (stators) to the booster stage. Thebooster stage, installed aft of the fan on the N1 shaft, is asingle-stage, 43-blade, axial compressor that functions toincrease core inlet pressure. After passing through the boosterstage, core inlet air is directed through a row of 37 stators to

the high compressor.

The secondary path directs a larger volume of air to the bypass

duct through two staggered rows of 66 stators each. This airpasses through the bypass duct without undergoing combus-tion, after which is it discharged from the outer exhaust nozzleas bypass thrust.

A T2 thermocouple probe senses inlet air temperature prior toinduction by the fan. A T2.6 thermocouple probe senses by-

pass air temperature within the outer exhaust nozzle. In combi-nation, these probes constitute the T1 signal producing ele-

ments of the inter-turbine temperature (ITT) sensing system.

Protection against engine air inlet icing is provided by bleed airheating of the nose cone, core inlet stators, and T2 thermo-couple. Refer to Chapter 10 for a complete description ofengine ice protection systems.

High Compressor Section

The high compressor section consists of a single-stage cen-

trifugal impeller, stator assembly, impeller shroud, and an arrayof 24 diffuser pipes and deflector vanes. The centrifugal impel-ler is installed on the forward end of the N2 shaft and incorpo-rates 32 blades (16 full/16 splitter) that function to acceleratecore inlet air. This air is directed to the impeller through a row of23 stator vanes.

The impeller shroud, diffuser pipes, and deflector vanes are

installed within the gas generator case. The impeller shroudcontains the accelerated core inlet air and directs it radially tothe diffuser pipes. The diffuser pipes function primarily to de-celerate this air thereby maximizing its pressure. Their second-ary function is to restore axial flow from radial flow by redirect-ing the air through 90°. Exiting the diffuser pipes, compressordischarge air (P3) passes through the deflector vanes. Theangular orientation of the deflector vanes optimizes P3 air flow

for delivery to the combustion chamber.




JT15D-4 General Arrangement

N2 TACHGENERATOR PAD OIL PUMP

ASSEMBLY

P3 TOANTI-ICING

VALVEBEARING

SCAVENGELINE

ACCESSORYGEARBOX

OIL SCAVENGE LINERETURN TO TANK

STARTERGENERATOR

PAD

N1 TACHGENERATOR

PAD

ANTI-ICINGVALVE

OIL FILLER NECKAND DIPSTICK

ACCESSORYGEARBOX

DRAIN

NOSECONE

LOWCOMPRESSOR

FAN

OIL-TO-FUELHEAT EXCHANGER

OIL FILTERBYPASSVALVE

OIL PRESSURELINE

P3 CABIN BLEED

OIL PRESSURE LINETO NO 3 1/2 AND 4 BEARINGS

P3 LINES FORFCU AND T1 PROBE

ANTI-ICING

T2 PROBE

OIL PRESSUREREGULATING VALVE

MAIN OILPRESSURE BOSS

FUEL FILTERHOUSING

OIL FILTERHOUSING

Left Front View

Right Front View





Combustion Section

The combustion section consists of the combustion chamber,fuel nozzles, and igniters. The annular, reverse-flow combustionchamber, installed within the gas generator case, provides anarea for the mixture and ignition of air and fuel, and directscombustion gas flow to the turbines. Concentric outer and inner

liners extend forward from the aft (domed) end of the combus-tion chamber. Twelve adapters are positioned around thedomed end for insertion of the fuel nozzles. Two adapters are

positioned at 5 and 7 o’clock on the outer liner for insertion ofthe spark igniters. At the forward end of the combustion cham-ber, a large exit duct and small exit duct join the outer andinner liners respectively.

P3 air enters the interior of the combustion chamber through aseries of perforations in the inner and outer liners. Metered fuel

from the fuel control unit enters the interior of the combustionchamber through the fuel nozzles. The air/fuel mixture is initially

ignited by the spark igniters during engine start, after whichcombustion is self sustaining under normal engine operatingconditions. The shape, size, and location of the perforations, aswell as the location of the fuel nozzles and spark igniters,provide the best air/fuel ratio for engine starting and sustainedcombustion. The exit ducts redirect the combustion gas flowinward then aft through 180° to the turbine inlet, thus the term“reverse-flow.”

Air not used in the combustion process, referred to as second-ary air, is used by the engine for ice protection, hot sectioncooling, bearing compartment sealing, and fuel control. Toensure even temperature distribution and prevent flame contactwith the interior walls of the combustion chamber, cooling ringsdirect a layer of P3 air over these surfaces during combustion.To prevent exposing the high compressor to excessive com-bustion gas flow temperatures, the large and small exit ducts

incorporate heat shields through which P3 air is also directedfor cooling purposes.




High Turbine Section

The high turbine section consists of the single-stage highturbine, segmented shroud assembly, and turbine inlet guidevane assembly. The high turbine is installed on the aft end ofthe N2 shaft and incorporates 71 blades. The segmentedshroud assembly provides an ideal blade tip running clearance

to maximize turbine efficiency.

The turbine inlet guide vane assembly incorporates 14 stator

vanes that direct combustion gas flow against the blades at anoptimal angle and speed. The stator vanes feature cored pas-sages through which P3 air is directed for cooling purposes. Airenters through inlet ports at each vane’s root and is evacuatedinto the gas flow path through outlet ports at each vane’s trail-ing edge.

The high turbine extracts energy from the combustion gas flowto generate the rotational torque that drives the N2 shaft, high

compressor, and engine-driven accessories. The energy notabsorbed by the high turbine is directed to the low turbinesection.

Note: Engines in compliance with P&WC SB7293 feature “D.S.”high turbine blades produced using a directionally-solidified

casting process.

Low Turbine Section

The low turbine section consists of two low turbines and theirassociated turbine inlet stator vane assemblies. The low tur-bines (referred to as 2nd and 3rd stage turbines) are installedin tandem on the aft end of the N1 shaft. The 2nd stage turbineincorporates 61 blades. The 3rd stage turbine incorporates 55blades. Both feature shrouded blade tips to maximize turbineefficiency.

Each turbine inlet stator vane assembly incorporates 43 statorvanes that direct combustion gas flow against the turbineblades at an optimal angle and speed. The low turbines extractenergy from the combustion gas flow to generate the rotationaltorque that drives the N1 shaft, fan, and booster stage. Afterdriving the turbines, the combustion gas flow is dischargedfrom the inner exhaust nozzle as core thrust.





Six thermocouple probes sense combustion gas flow tempera-ture within the inner exhaust nozzle. In combination, theseprobes constitute the T6 signal producing elements of the ITTsensing system.

Accessory Gearbox

The accessory gearbox (AGB), located on the lower intermedi-ate case of the engine, houses the gearing and provides sup-port for all engine-driven accessories except the N1 tachom-

eter generator. The AGB consists of a cast magnesium alloyhousing and rear cover. The housing provides bearing sup-ported gear shafts which drive the centrifugal air/oil separatorand all externally-mounted engine accessories. External mount-ing pads are provided for the starter/generator, engine-drivenfuel pump, hydraulic pump, oil pump assembly, and N2 ta-chometer generator.

The AGB main shaft engages the starter/generator directly, and

is linked to the N2 shaft by a vertical tower shaft. The towershaft is splined at each end to engage upper and lower bevelgears. The upper bevel gear meshes with a bevel gear fitted tothe N2 shaft. The lower bevel gear meshes with a bevel gearfitted to the AGB main shaft.

During engine start, the tower shaft transmits starter/generatorrotation to the N2 shaft to turn the engine. During engine opera-

tion, the tower shaft transmits N2 shaft rotation to the AGB main

shaft to turn the starter/generator and other engine-drivenaccessories. A second vertical tower shaft engages the N1tachometer generator at its upper end. A bevel gear fitted tothe lower end of this shaft meshes with a bevel gear fitted tothe N1 shaft.




Engine Fuel System

The engine fuel system functions to regulate fuel flow to theengines. Major components for each engine include an engine-driven fuel pump, fuel control unit, oil-to-fuel heat exchanger,flow divider, fuel manifold, and 12 dual-orifice type nozzles.

Engine-Driven Fuel Pump

One single-stage, positive-

displacement, gear-type,engine-driven fuel pump ismounted on and driven by theaccessory gearbox of eachengine. Each pump functionsto supply clean fuel underhigh pressure to the fuel

control unit (FCU) of its asso-ciated engine, and motive flow

pressure to its associatedprimary ejector pump. Typicalpump capacity is 3935 PPHand 580 PSI at 100% N2. Fuelentering the pump passes through a 74 micron, wire mesh inletfilter before entering the pump chamber. The inlet filter is selfrelieving at 9 to 12 PSID should it become obstructed. Exitingthe pump chamber, high pressure fuel passes through a 10

micron, nonmetallic, disposable outlet filter en route to the FCU.

A spring-loaded, ball-type, bypass valve, preset to open at 40to 60 PSID, allows high pressure fuel to bypass the outlet filterelement should it become obstructed. Pump chamber inletpressure is maintained by a jet pump nozzle located upstreamof the inlet filter. When pump pressure exceeds metered pres-sure, a portion of FCU bypass fuel is returned to the pumpchamber inlet through this nozzle.

Fuel Control Unit (FCU)

The fuel control unit (FCU) is mounted on and driven by theengine-driven fuel pump through an integral splined couplingshaft. The FCU determines the correct fuel schedule to pro-duce desired engine power in response to THROTTLE levermovement. Major components include the fuel metering sec-tion, computing section, and N2 governing section. Otherassociated components include a T2 temperature compensator

and step modulator.

1. Fuel Pump Body

2. Fuel Filter Housing (inlet)

3. Fuel Filter Housing (outlet)

2

1

3





Fuel Metering Section

The fuel metering section includes a pressure relief valve,metering valve, and bypass valve. Unmetered fuel pressure(P1) is supplied to the FCU from the engine-driven fuel pump.The pressure relief valve functions to prevent systemoverpressurization by opening when P1 exceeds 1260 PSID,

thereby returning excess fuel pressure (P0) to the pump inlet.The metering valve is essentially a tapered needle valve posi-tioned by the FCU computing section to regulate fuel flow to the

engine. The bypass valve functions to maintain a constant 15 to24 PSI differential (delta P) between P1 and metered fuel pres-sure (P2) across the metering valve orifice. Because bypassvalve position and fuel flow are functions of metering valveposition, the bypass valve responds to increased fuel flowthrough the metering valve by decreasing the amount of P0returned to the pump inlet. When fuel flow through the metering

valve is decreased, the bypass valve responds by increasingthe amount of P0 returned to the pump inlet.

1. Fuel Control Unit 3. Throttle Lever Input2. Fuel Outlet (P2) 4. Fuel Bypass Line

Minimum flow through the metering valve is factory adjusted toapproximately 155 to 160 pounds per hour (LBS/HR) to providethe correct amount of fuel flow for engine starting and to pre-

vent engine failure during rapid deceleration.

3

4

1

2




Computing Section

The computing section includes a pneumatic governor bellowsand an acceleration bellows connected to a common torquetube linkage that positions the metering valve. In combination,these components control acceleration, deceleration, andsteady engine operating states in response to pneumatic sig-

nals supplied by the N2 governing section.

The body of the computing section is divided into two cham-

bers separated by the governor bellows. The upper chamber issupplied with Py (governing) pressure; the lower chamber issupplied with Px (enrichment) pressure. As such, the exterior ofthe governor bellows is exposed to Py pressure and the interiorof the governor bellows is exposed to Px pressure. The accel-eration bellows is sealed at absolute pressure and locatedwithin the lower chamber. The force of Px pressure acting on

the acceleration bellows is cancelled by the force of Px pres-sure acting on the same area of the interior of the governor

bellows. This “area of cancellation” ensures that any change inPy pressure will have a greater effect on metering valve posi-tion than an equal change in Px pressure.

Because Px and Py are derived from P3 (compressor dis-charge air), and because P3 is proportional to N2 and airdensity, the acceleration bellows provides an absolute pressurereference to compensate for reduced air density at higher

altitudes. An air filter is installed in-line to prevent foreign mate-

rial present in the P3 air from entering the FCU.

N2 Governing Section

The N2 governing section is associated primarily with enginespeed setting and speed control. Major components include aspeed scheduling cam, governor flyweights, feedback springs,a governor lever, enrichment lever, enrichment valve, andbackup valve.

The speed scheduling cam sets feedback spring resistance inresponse to THROTTLE lever position. The smaller of the twofeedback springs resists enrichment lever movement; thelarger spring resists governor lever movement. The enrichmentvalve is spring-loaded open and regulates the introduction ofP3 air into the body of the N2 governing section in response toenrichment lever movement. Movement of the governor lever

regulates the bleeding of Py air to the atmosphere throughvents in the body of the N2 governing section.





The flyweights are driven by the FCU coupling shaft such that

their lifting force varies with engine speed. Above approxi-mately 85% N2, when flyweight lifting force is sufficient toovercome the resistance of the smaller feedback spring, theenrichment lever moves the enrichment valve toward closedand less P3 air enters the FCU. As flyweight lifting force in-

creases, the enrichment lever makes contact with the governorlever pivoting about the same point. When flyweight lifting forceis sufficient to overcome the resistance of the larger feedback

spring, the enrichment lever moves the governor lever towardopen and more Py pressure is bled to the atmosphere. Thecorresponding decrease in Py pressure at the computingsection moves the metering valve toward closed and reducesfuel flow to the engine. Conversely, an increase in Py pressureat the computing section moves the metering valve towardopen and increases fuel flow to the engine. When Px and Py

are simultaneously decreased during decceleration or in-creased during acceleration, Py pressure will have a greater

effect on metering valve position than Px pressure.

Following a change in power setting, the corresponding “lag” inPx and Py pressure change at the computing section regulatesthe transition from the previous fuel flow rate through the meter-ing valve to that which will produce the desired engine power.As such, an excessively lean condition or excessively richcondition are prevented. When the sum of all forces acting on

the position of the metering valve are in equilibrium, fuel flow

and N2 remain essentially constant.

The backup (overspeed) valve is spring-loaded closed andnormally seals a secondary governing air pressure (Py) orificeplumbed to the computing section. When N2 exceeds theselected speed by approximately 9%, the valve is forced openby the governor lever and more Py pressure is bled to theatmosphere. The corresponding rapid decrease in Py pressure

at the computing section moves the metering valve towardclosed and reduces fuel flow to the engine, thereby preventingN2 overspeed.




1. T2 Temperature Compensator

The T2 temperature compensator functions to maintain consis-tent acceleration response by modifying the engine’s fuel flowschedule for variations in outside air temperature. To accom-plish this, the unit incorporates a bimetallic disk temperatureprobe that senses ambient air temperature, and a metering pin

that regulates the bleeding of Px pressure to the atmospherethrough vents in the compensator housing. The metering pin isspring-loaded open and moves toward closed when the bime-

tallic disks expand under increasing temperature. As such, Pypressure at the computing section is increased and decreasedinversely proportional to bypass air temperature.

2. Step Modulator

The step modulator functions to maintain consistent accelera-

tion response by maintaining a constant P3 supply pressure tothe FCU. To accomplish this, the unit incorporates an electri-cally-actuated restrictor orifice that is energized open primarilywhen the engine ice protection system is activated. Whenopen, more P3 air than normal is supplied to the FCU to com-pensate for that consumed in the bleed air heating of inductionair inlet components. The restrictor orifice is also energizedopen when the engine ignition system is activated. Whenclosed, the supply of P3 air to the FCU returns to normal.

12





Oil-to-Fuel Heat Exchanger

The oil-to-fuel heat exchanger functions to reduce engine oiltemperature and prevent ice formation in the fuel.

1. Fuel Inlet 3. Oil Inlet Manifold

2. Fuel Outlet 4. Oil Outlet Manifold

The unit consists primarily of a cylindrical outer “shell” and aninner “core.” The core is formed by 85 transfer tubes which runaxially through the interior of the shell. Fuel from the FCU entersthe core at the forward end of the unit and exits at the aft end ofthe unit en route to the flow divider. Oil from the pressure pumpenters the shell through an external inlet port and manifold that

directs the oil to the aft end of the unit. As the oil flows forwardthrough the interior of the shell, its heat is transferred to thelower temperature fuel flowing in the opposite direction throughthe core. An array of baffles, alternately positioned on oppositesides of the shell interior, repeatedly diverts the oil laterally overthe core to maximize heat transfer. Upon reaching the forwardend of the unit, the oil exits the shell through an external mani-

fold and outlet port en route to the flow divider.

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4

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Flow Divider

The flow divider integrates a minimum pressurizing and flowdivider valve, fuel cut-off valve, spill valve, and dump valve in asingle unit. The minimum pressurizing and flow divider valvemaintains a sufficient minimum system pressure and dividesmetered fuel flow between the primary and secondary fuel

manifolds. This plunger-type valve is spring-loaded closed andopened by fuel pressure. When fuel pressure reaches approxi-mately 75 PSID, the valve opens to an internal passage that

supplies the primary manifold. When engine speed increasesabove approximately 60% N2, increasing fuel pressure furtheropens the valve to an internal passage that supplies the sec-ondary manifold also.

1. Flow Divider 5. Primary Outlet

2. Fuel Inlet (P1) 6. Drain Outlet

3. Bypass Outlet 7. Secondary Outlet

4. P3 Line 8. Throttle Linkage

The rotary-type fuel cut-off valve is positioned by theTHROTTLE lever and functions to control fuel flow between theFCU and the flow divider. When the THROTTLE lever is posi-tioned to “OFF,” the valve is fully-closed and restricts fuel flow.When positioned to “IDLE” the valve is fully-open.

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6

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Engine Fuel System Schematic

During engine start, when the THROTTLE lever is initially posi-tioned to “IDLE,” the fuel pressure supplied to the primarymanifold is regulated by the spill valve as a function of P3pressure. When P3 pressure is low, the spill valve is fully-openand fuel in excess of that required for engine starting is re-turned to the pump inlet. As engine speed increases to idle, the

corresponding increase in P3 pressure moves the spill valvetoward closed and progressively less fuel is returned to thepump inlet. When P3 pressure reaches approximately 30 PSI,the spill valve is fully-closed.

During engine shutdown, when the THROTTLE lever is posi-tioned to “OFF” and the cut-off valve is fully-closed, the mini-mum pressurizing and flow divider valve returns to its spring-

loaded closed position. As fuel pressure falls below approxi-mately 5 PSI, the dump valve opens and residual fuel isdrained from the primary and secondary manifolds to an EPA

canister mounted below the engine. During engine operation,the EPA canister is charged with air tapped from the bypassduct and the residual fuel is returned to the tank. Each returnline incorporates a check valve to prevent backflow from thefuel tank to the canister. Residual fuel in excess of the canister’s15-ounce capacity is dumped overboard through a drain tube

that extends through the lower engine cowling.

© PCW




Fuel Manifold and Nozzles

The fuel manifold consists of primary and secondary transfertubes, and twelve manifold adaptors which carry metered fuelfrom the flow divider to the fuel nozzles. Each manifold adapterindependently supplies primary and secondary fuel flow to itsassociated nozzle. The dual-orifice type nozzles are enclosed

in sheaths which extend into the combustion chamber. Atom-ized fuel is introduced to the combustion chamber through theprimary orifice when the flow divider is supplying the primary

manifold, or through both the primary and secondary orificeswhen the flow divider valve is supplying the primary and sec-ondary manifolds. Each sheath is slotted to permit the passageof secondary air for nozzle cooling and improved atomization.

Fuel Drains

Two fuel drains, installed in the

6 o’clock position of the gasgenerator case below the

combustion chamber, ensurethat all residual fuel whichaccumulates in this area isdrained overboard after en-gine shut down. Each drain isfitted with a transfer tube thatcarries residual fuel to a com-mon drain valve installed on

the outer bypass duct. The

drain valve is spring-loaded tothe open position and heldclosed during engine opera-tion by P3 air pressure. When open, residual fuel is routed to anoverboard breather tube that extends through the lower enginecowling. Residual fuel is also drained from the inner exhaustnozzle through an overboard drain line that extends through thelower engine cowling.

1. Combustion Drain Line

2. Overboard Breather Tube3. EPA Canister

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32





Ignition System

1. Spark Igniter

2. Ignition Exciter Box

3. High Tension Wire

The ignition system functions primarily to supply the electricalspark necessary to initiate combustion during engine start.Each engine’s ignition system operates independently of the

other and consists of two engine mounted ignition exciters, two-high tension cables, and two spark igniters. Nominal 28 VDCpower is supplied to the system; however, the system is oper-

able between 9 and 30 VDC.

Note: On airplanes 550-0470 and earlier not incorporatingSB550-74-01 (P&WC SB 7178), each engine is equipped with asingle ignition exciter.

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The ignition exciters are sealed boxes which house the compo-nents that transform DC voltage into high-energy electricalpulses that are supplied to the spark igniters. The pulses aresupplied to the spark igniters through high tension cableswhich are encased in flexible metal braiding for protection fromheat. On engines with dual exciters, each is a single-output

type that supplies its associated igniter only. On engines with asingle exciter, each is a dual-output type that supplies bothignitors. In either installation, the system is designed such that

one igniter will remain operable should failure of the oppositeigniter occur.

The spark igniters are located at 5 and 7 o’clock positions onthe gas generator case and extend inward to the interior of thecombustion chamber through its outer liner. Each igniter iscomprised of a threaded outer casing and central electrode

separated by a semiconducting material. When the ignitionexciter is energized, a capacitor is progressively charged until

sufficient voltage is produced to ionize the gap between thepositive electrode and the negative casing. When this occurs,capacitor voltage is discharged across the gap in the form of ahigh-energy spark. To extend service life and reduce the risk offailure, the igniters are cooled by secondary air.

Each engine’s ignition systemis independently controlled by

a corresponding two-position

(ON/NORM) LH or RH IGNI-TION switch on the lower leftinstrument panel. Under mostoperating conditions, theseswitches should remain in the“NORM” (off) position. In thisposition, ignition system acti-vation occurs automatically

during engine start and en-gine ice protection systemoperation.

Independent indication of left (LH) or right (RH) ignition systemoperation is provided by green lights located above eachIGNITION switch. When power is being supplied to eithersystem, the corresponding light will be illuminated regardless of

operating condition.

IGNITION SWITCHES





Ignition System Schematic

During engine start, ignition system activation is carried out bya series of relays in the start control circuit which supply 28VDC hot battery bus power to the appropriate exciter(s)

through a 7.5-amp IGNITION circuit breaker located in the aft

fuselage electrical power junction “J” box. When the ENGINEANTI-ICE switches or IGNITION switches are set to the upper(on) position, each exciter receives 28 VDC power through itsassociated 15-amp LH IGN or RH IGN circuit breaker on theleft CB panel, correspondingly supplied by the left or right mainbus.

Note: The position of the LH IGN and RH IGN circuit breakershas no effect on ignition system operation during the enginestarting sequence.

Note: The IGNITION switches should be positioned to “ON”during takeoff, approach and landing, and turbulent air pen-etration.

© PCW




Engine Oil System

The engine oil system consists of pressure, scavenge, andbreather sub-systems. The pressure system pumps oil from thetank to lubricate and cool all internal moving engine compo-nents by way of cored passages and transfer tubes. The scav-

enge system returns oil to the tank by way of gravity drains,transfer tubes, cored passages, and scavenge pumps. Thebreather system functions essentially to vent air pressure from

the scavenge oil system. An oil-to-fuel heat exchanger is pro-vided to regulate engine oil temperature. Engine oil tempera-ture and pressure indication is provided by the independentOIL °C and OIL PSI gages on the upper center instrumentpanel. Total oil capacity is 2.08 US gallons, of which 1.25 USgallons are usable.

Oil TankThe oil tank is integral to the

intermediate case. The tank isserviced through a filler necklocated on the outboard inter-mediate case, and is acces-sible through a hinged door onthe upper cowling. A dipstick,integral to the filler cap, isprovided for oil level inspec-

tion. The dipstick is calibrated

to indicate the approximatetank quantity in US quartswhen the engine is hot, therefore, oil level should be inspectedwithin ten minutes of engine shutdown.

Note: The engine should be serviced with approved syntheticoils listed in the most current revision of P&WC SB 7001.

OIL FILLER NECK





Oil Pressure System

Engine oil pressure is devel-oped by a gear-type pressurepump that is mounted on andshaft-driven by the accessorygear box (AGB). The pressurepump is integral to an oilpump assembly that alsohouses the scavenge pumps.From the tank, oil is drawn

through a filter screen to theinlet side of the pump. Engineoil pressure developed by thepump is routed through an external transfer tube to an assem-bly that houses a pressure regulating valve, check valves, oilfilter, and filter bypass valve. The spring-loaded, piston-type

pressure regulating valve prevents operating pressure fromexceeding 73 ± 6 PSI by diverting excess oil pressure back to

the pump inlet. Regulated oil pressure is directed through theoil-to-fuel heat exchanger en route to the oil filter.

OIL PUMP ASSEMBLY

Engine Oil System Schematic

© PCW




The oil filter utilizes a 10 micron, cartridge-type element whichprevents particles of foreign matter present in the oil fromcirculating through the engine. Should the filter element be-come partially restricted or blocked, an integral spring-loadedbypass valve permits continued though unfiltered oil flow to theengine. The bypass valve is opened when the differential be-

tween filter inlet and outlet pressure reaches approximately 15to 24 PSID.

1. Oil Filter Housing

2. Fuel Filter Housing

3. Oil Pressure Regulating Valve

4. Oil-to-Fuel Heat Exchanger

The check valves open at 5 PSI when the engine is operating to

permit oil flow from the pressure pump to the engine, and fromthe pressure regulating valve to the pump inlet. When theengine is not operating, the check valves function to preventgravity draining of oil from the tank to the bearing cavities(primarily after engine shutdown), and allow the oil filter to be

removed for inspection without draining the tank.

From the filter, pressure oil flow is divided into three paths. Thefirst path directly supplies the AGB through an internal transfertube. The second path supplies the N1 shaft front (#1) bearing,the N2 shaft front and rear (#2 and #3) bearings, and the tower

shaft bevel gears through an external transfer tube and internalpassages. The third path supplies the N1 shaft intermediateand rear (#3½ and #4) bearings through an external transfer

tube and internal passages. Labyrinth air seals are used toconfine pressure oil to the bearing compartments.

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Scavenge System

The scavenge system incorporates a pair of gear-type pumpsthat are integral to the oil pump assembly and shaft-driven bythe AGB. The gearshaft that drives the oil pressure and scav-enge pumps also drives the N2 tachometer generator.

Scavenge oil gravity drains from the #1, 2, 3, and #3½ bearingsto the AGB assisted by secondary air pressure from the bear-ing compartment labyrinth seals. Scavenge oil from the #4

bearing is drawn through an external transfer tube by thesmaller of the two scavenge pumps. Scavenge oil from theAGB is drawn through a filter screen by the larger scavengepump. From these pumps, scavenge oil is returned to the tankthrough an external transfer tube and internal passage.

Breather System

The breather system incorporates a centrifugal impeller drivenby the AGB main shaft which also drives the starter/generator.

The impeller, also referred to as the centrifugal breather, sepa-rates air from the AGB scavenge oil by centrifugal force. Inoperation, air is drawn radially inward while oil is thrown radiallyoutward. Once separated, the relatively oil-free air is vented tothe atmosphere through an overboard breather tube that ex-tends through the lower engine cowling.




Engine Controls

Primary engine control is provided by left and right THROTTLElevers located on the center pedestal between the pilot’s andcopilot’s seats. When thrust reversers are optionally installed,thrust reverser control levers are mounted “piggyback” on the

THROTTLE levers. When the engine synchronizer system isoptionally installed, an ENGINE SYNC selector switch is alsolocated on the center pedestal.

1. Throttle Levers 3. Thrust Reverser Lever

2. Release Triggers 4. Engine Sync Switch

Throttle Levers

The THROTTLE levers function to set engine speed and shutdown the engines. To accomplish this, THROTTLE lever move-ment is transmitted to its associated FCU by “controlex” cablesand bellcrank assemblies. The FCU, in turn, is mechanically-linked to the flow divider cut-off valve by an interconnect rod.

When positioned forward of IDLE, the THROTTLE lever sets theFCU to maintain the selected engine speed. When positionedto “IDLE,” the FCU is set to provide approximately 49% N2minimum. The IDLE position also provides the correct fuelschedule during engine start. When positioned to “OFF,” theflow divider cut-off valve restricts fuel flow to the engine.

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A detent prevents inadvertent THROTTLE lever movement from“IDLE” to “OFF.” To clear this detent, spring-loaded triggers onthe outboard sides of the levers must be lifted.

Throttle lever friction is adjusted using an INCREASE FRICTIONknob located on the right side of the control pedestal. Rotating

this knob clockwise increases throttle lever friction; counter-clockwise rotation decreases friction.

Thrust Reverser Levers

The thrust reverser (TR) levers are mounted on the THROTTLElevers and function to control thrust reverser operation. Toaccomplish this, independent left and right, deploy and stowmicroswitches, located within the center pedestal, are actuatedby TR lever movement. The deploy switches are actuated whenthe TR levers are pulled aft; the stow switches are actuated

when the TR levers are pushed forward. When these switchesare actuated, corresponding deploy or stow solenoids integral

to the thrust reverser control valves are energized, therebypositioning these valves to permit thrust reverser operation asselected.

To prevent inadvertent thrust reverser operation, theTHROTTLE levers must be set to “IDLE” before the TR leverscan be moved. During thrust reverser transition, the throttlefeedback system holds the throttle linkage at idle until the

reverser doors are fully-deployed or fully-stowed. A pair of

locking solenoids prevents aft movement of the TR levers toincrease reverse thrust until the reverser doors are fully-de-ployed. TR lever stops limit reverse thrust N1 to approximately92% in sea level ISA conditions. Refer to the thrust reversersection of this chapter for functional detail.

Warning: The THROTTLE levers should not be restrained manu-ally or by lever friction during takeoff as this will prevent the

throttle feedback system from automatically returning thethrottle linkage and THROTTLE levers to “IDLE” should inad-vertent thrust reverser deployment occur.




Engine Synchronizer Selector Switch

The ENGINE SYNC selector switch is located aft of the rightTHROTTLE lever on the center pedestal and functions to con-trol engine synchronizer system operation. When positioned to“FAN,” the system synchronizes left (master) and right (slave)fan RPM (N1). When positioned to “TURB,” the system synchro-

nizes turbine RPM (N2). When positioned to “OFF,” the systemis deactivated. Refer to the engine synchronizer section of thischapter for functional detail.

Automatic Fuel Shutoff Control

The automatic fuel shutoff control is a mechanically-operatedsystem that functions to shutdown the engine in the event of #4bearing failure or N2 shaft decoupling. To accomplish this, thesystem incorporates an actuator rod that is linked by bellcrankto a piston-type valve. Should 0.070” rearward, axial displace-

ment of the N2 shaft occur, N2 shaft contact with the rod willactuate the valve thereby cutting off the supply of fuel to the

nozzles.





Engine Indicating System

The engine indicating system provides visual indication ofcritical engine operating parameters and provides annunciationwhere these parameters are exceeded. Each engine’s indicat-ing system operates independently of the other and includes

an N1 (fan) tachometer generator, inter-turbine temperature(ITT) sensing system, fuel flow transmitter, N2 (turbine) tachom-eter generator, oil pressure transmitter, and oil temperature

sensor which supply input signals to the corresponding left orright channel of their associated gages. Other engine indica-tions are displayed on the annunciator panel.

The primary engine gages are arranged in a horizontal row onthe upper center instrument panel. From left to right these areFAN % RPM, ITT, FUEL FLOW, TURBINE % RPM, OIL TEMP,

and OIL PRESS. With the exception of the TURBINE % RPMdigital readout, indication is provided by the position of inde-

pendent left (L) and right (R) vertical tape bars relative to agraduated, vertical instrument scale. The FAN % RPM gagealso features a digital readout. Where appropriate, coloredmarkings on the instrument scale denote operating ranges andlimitations corresponding to the indicated parameter.

Fan Tachometer

The FAN tachometer provides indica-

tion of N1 shaft rotational speed in %

RPM, with 100% equaling approxi-mately 15,904 RPM. The instrumentscale is graduated in 10% incre-ments between 20 and 90%, and 2%increments between 90 and 110%.Numerical values are marked at each10% increment between 20 and110%. Green markings between 25

and 104% denote the normal indicat-ing range. Red warning lines at104% denote the maximum N1limitation (16,540 RPM). Input signalsare supplied to each channel of thegage by its associated left or rightN1 tachometer generator located onthe upper intermediate case.




The digital readout on the lower instrument face providessupplemental indication of FAN % RPM. The readout displaysin tenths between 00.0 and 99.9%. A mechanical counter,mounted above the gage, may be set to reference a targetpower setting.

Each channel of the gage receives 28 VDC power through itsassociated 2-amp LH FAN SPEED or RH FAN SPEED circuitbreaker, correspondingly supplied by the left or right main bus.

When power is removed from either channel, the affecteddigital readout will be extinguished. The vertical tape indicatingportion of the gage is essentially “self-generating” and willremain operable above 50 N1 even when power is removedfrom the instrument. No OFF flag is provided.

Inter-Turbine Temperature Gage

The ITT gage provides indication ofinter-turbine temperature in °C. The

instrument scale is graduated in100° increments between 200 and800°C, and 10° increments be-tween 500 and 800°C. Numericalvalues are marked at each 100°increment between 200 and 800°C.Green markings between 150 and680°C denote the normal indicating

range. Yellow markings between

680 and 700°C denote the cautionrange. Red warning lines at 700°Cdenote the maximum ITT limitationduring engine start, which is timelimited to 2-seconds.

Note: During normal engine start, ITT indication should notexceed 500°C.

Note: ITT indication exceeding 500°C during normal enginestart should be investigated for cause.

Each channel of the gage receives 28 VDC power through itsassociated 2-amp LH ITT SPEED or RH ITT circuit breaker,correspondingly supplied by the left or right main bus. A redOFF warning flag will appear at the top of the vertical scale

whenever power is removed from the corresponding channel ofthe gage.





Input signals are supplied to each channel of the gage by itsassociated left or right ITT sensing system which computes a“simulated” inter-turbine temperature (T5) from signals pro-duced by eight thermocouple probes. One thermocoupleprobe senses inlet air temperature (T2) prior to induction by thefan. Another senses bypass air temperature (T2.6) within the

outer exhaust nozzle. In combination, the T2 and T2.6 probesconstitute the T1 signal producing elements of the system. Sixprobes sensing combustion gas flow temperature within the

inner exhaust nozzle constitute the T6 signal producing ele-ments of the system.

The bimetallic (alumel/chromel) thermocouples generate a mildelectrical current which varies in response to the temperaturesensed. The T2 and T2.6 probes contain three thermocoupleseach, wired in series. The T6 probes contain one thermocouple

each, wired in parallel to compensate for uneven heat distribu-tion and to obtain an average temperature reading. Each probe

is connected to its associated T1 or T6 wiring harness. Thewiring harnesses are also bimetallic, having an alumel negativecircuit and chromel positive circuit.

In operation, the difference between T2.6 and T2 representsthe temperature rise in non-combusted airflow as it passesthrough the bypass duct. Because the three thermocouples ineach of these probes are wired in series, the temperature rise is

multiplied by three to produce the T1 signal. The T1 signal and

composite T6 signal are corrected for sampling errors by avariable resistor. The variable resistor is set to a required trimvalue during final engine acceptance checks and is sealed atthat setting. The sum of the corrected signals, computed fromthe formula T1 + T6 = T5, is displayed by the ITT gage.




Fuel Flow Gage

The FUEL FLOW gage providesindication of engine fuel flow in LBS/ HR. The instrument scale is gradu-ated in 100 pound per hour incre-ments between 0 and 2000 LBS/HR.

Numerical values are marked at each200 pound per hour increment be-tween 0 and 2000 LBS/HR. Input

signals are supplied to each channelof the gage by its associated left orright fuel flow transmitter.

Each channel of the gage receives 28 VDC power through itsassociated 2-amp LH FUEL FLOW or RH FUEL FLOW circuitbreaker, correspondingly supplied by the left or right main bus.

A red OFF warning flag will appear at the top of the verticalscale whenever power is removed from the correspondingchannel of the gage. Refer to Chapter 3 for a complete descrip-tion of the fuel flow indicating system.

Turbine Tachometer

The TURBINE tachometer providesdigital indication of high pressure(N2) shaft rotational speed in %RPM, with 100% equaling approxi-

mately 32,760 RPM. The readoutdisplays in tenths between 00.0 and

99.9%, and single-digits above100%. A red light is located below each digital readout. WhenN2 reaches the maximum 96% limitation (31,450 RPM), thecorresponding light will illuminate and readout will flash. Inputsignals are supplied to each channel of the gage by its associ-ated left or right N2 tachometer generator installed on theforward AGB.

Each channel of the gage receives 28 VDC power through its

associated 2-amp LH TURB SPEED or RH TURB SPEED circuitbreaker, correspondingly supplied by the left or right main bus.When power is removed from either channel, the affecteddigital readout will be extinguished.





Engine Oil Temperature Gage

The OIL TEMP gage provides indica-tion of oil temperature in °C. Theinstrument scale is graduated in 10°increments between 0 and 130°C.

Numerical values are marked at each20° increment between 0 and 120°C.Green markings denote the normal

indicating range. Red warning linesat 121°C denote the maximum oiltemperature limitation. Input signalsare supplied to each channel of thegage by its associated left or right oiltemperature sensor installed in the oil filter housing.

Each channel of the gage receives 28 VDC power through itsassociated 2-amp LH OIL TEMP or RH OIL TEMP circuit

breaker, correspondingly supplied by the left or right main bus.A red OFF warning flag will appear at the top of the verticalscale whenever power is removed from the correspondingchannel of the gage.

Engine Oil Pressure Gage

The OIL PRESS gage provides indi-cation of oil pressure in PSI. The

instrument scale is graduated in 10

PSI increments between 0 and 100PSI. Numerical values are marked ateach 20 PSI increment between 0and 100 PSI. Green markings denotethe normal indicating range. Yellowmarkings denote the low pressurecaution range. Red warning lines at35 PSI denote the minimum oil pres-

sure limitation. Input signals aresupplied to each channel of the gageby its associated left or right oil pressure transmitter installed inthe oil filter housing.

Each channel of the gage receives 28 VDC power through itsassociated 2-amp LH OIL PRESS or RH OIL PRESS circuitbreaker, correspondingly supplied by the left or right main bus.

A red OFF warning flag will appear at the top of the verticalscale whenever power is removed from the correspondingchannel of the gage.




Should oil pressure fall below approximately 35 PSI, a red [LH]

or [RH] [OIL PRESS WARN] annunciator (550-0550 an after) or[L OIL PRESS LO] or [R OIL PRESS LO] annunciator (550-0505and earlier) will illuminate. Input signals are supplied by an oilpressure transmitter and low pressure switch installed on theAGB.

Note: The normal oil pressure indicating range applies to en-gine speeds above 60% N2.

Caution: Engine operation at oil pressures below 70 PSI is un-desirable. Should oil pressure fall below 70 PSI in-flight, flightmay be completed at oil pressures between 35 and 70 PSIwhen engine power is reduced.

Warning: Engine operation at oil pressures below 35 PSI is un-

safe. Should oil pressure fall below 35 PSI in-flight, the airplaneshould be landed as soon as possible using the minimum en-

gine power required to sustain flight.





Engine Starting System

Engine starting is accomplished semi-automatically by inde-pendent left and right start control circuits which activate anddeactivate the starter/generator, ignition system, and boostpump through a series of relays. The engines may be started

using battery power or external power. With one engine operat-ing, the opposite engine may also be started using generatorpower. LH and RH momentary-on, push-button ENGINE START

switches, located on the lower left instrument panel, controlstart sequence activation. A STARTER DISENGAGE switch,located between the ENGINE START switches, permits manualtermination of the start sequence. Either engine may be startedfirst, however, the left engine is typically started first due to itsproximity to the battery.

Of the numerous relays in each start control circuit, only threeare associated primarily with engine starting, these are theauxiliary (aux) start relay, start relay, and start control relay. Aleft engine start sequence is activated by pressing and releas-ing the LH ENGINE START switch. The resulting momentaryflow of current energizes the aux start relay, after which it re-mains energized through a latching circuit. Current flowing

across the closed contacts of the aux start relay simultaneouslyenergizes the start relay and start control relay and illuminatesthe engine instrument floodlights.




With the start relay energized, the LH ENGINE START switchilluminates and the starter-generator begins turning the engine.Current flowing across the closed contacts of the start controlrelay energizes the left boost pump and illuminates the amber[LH] [FUEL BOOST ON] annunciator (550-0550 and after) or [LFUEL BOOST ON] annunciator. When N2 reaches 8 to 10%,

lifting the spring-loaded trigger and advancing the leftTHROTTLE lever to “IDLE” opens the flow divider cut-off valveto supply fuel to the engine. Simultaneously, a throttle position

switch permits current flow across the closed contacts of thestart control relay to the ignition system and illuminates thegreen light above the LH IGNITION switch.

From this point on, normal initiation of combustion should beindicated by a steady rise in N1, N2, and ITT. If no indication ofN1 exists when N2 reaches 20 to 25%, or if ITT approaches

700°C or fails to rise within 10-seconds, the start sequenceshould be manually terminated by pressing the STARTER

DISENGAGE switch. If indications are normal, the start controlcircuit will automatically terminate the start sequence bydeenergizing the aux start relay, thereby deenergizing the startrelay and start control relay, under the following conditions:

On airplanes 550-0406 and after, and earlier airplanes incorpo-rating SB550-28-1, the starting sequence is terminated by theengine speed sensing circuit of the generator control unit

(GCU) when N2 reaches approximately 40%. On airplanes 550-

0405 and earlier not incorporating SB550-28-1, the startingsequence is terminated by the motive flow pressure switchwhen the engine-driven fuel pump is supplying at least 180 ± 5PSI motive flow to the primary ejector pump. This pressure istypically achieved when N2 reaches approximately 30%.Should the motive flow switch fail to open, the engine speedsensing circuit of the GCU will terminate the starting sequencewhen N2 reaches approximately 40%.

Note: During normal engine start, ITT indication should notexceed 500°C.

Note: ITT indication exceeding 500°C during normal enginestart should be investigated for cause.





The following precautions apply primarily during engine startingand ramp area ground operations.

Warning: An engine running at higher power is capable ofgenerating sufficient suction to pull a person into the intakeduct; even the suction generated by an idling engine is ca-

pable of pulling small objects through the intake duct sufficientto cause injury to persons and/or engine damage. As a rule,personnel and equipment should not be positioned within an

arc extending a minimum of 25 feet forward of and 90° eitherside of the intake duct when starting or running the engine.

Warning: An engine running at higher power is capable ofgenerating an exhaust wake sufficient in temperature andvelocity to cause injury to persons and/or damage to property.During engine start, it is possible for accumulated fuel within

the exhaust duct to be ejected from the engine as long streamsof flame. Exhaust gases can cause respiratory and/or eye

irritation. At high engine speeds, the jet wake may propel loosedirt, sand, stones, and/or other debris for considerable dis-tances. As a rule, personnel, equipment, structures, and flam-mable material should be clear of an area extending aft fromthe exhaust nozzle to a minimum width of 30 feet and a mini-mum distance of 160 feet when starting or running an engine.




SPEED SETTING ACTUATOR

Engine Synchronizer System

The engine synchronizer system automatically matches the fanor turbine speed of the right (slave) engine to the left (master)engine. Major components of the system include a control box,motorized speed setting actuator, trimmer assembly, flexible

drive shaft, selector switch, and indicator light.

The control box monitors RPM

signals supplied by the N1(fan speed) and N2 (turbinespeed) tachometer generatorsof each engine and providesoperating commands to thespeed setting actuator. Thespeed setting actuator, lo-

cated within the right enginepylon, operates the trimmer

assembly via the flexible driveshaft. The trimmer assemblyadjusts the right (slave)engine’s fuel control unit to maintain N1 or N2 within 1.5% of theleft (master) engine. Based on the RPM signals received andthe mode selected, the control box commands the speedsetting actuator to adjust slave N1 or N2

The system is controlled by

the ENGINE SYNC selectorswitch located aft of the rightTHROTTLE lever on the centerpedestal. When positioned to“FAN,” the system synchro-nizes N1. When positioned to“TURB,” the system synchro-nizes N2. When positioned to

“OFF,” the system is deacti-vated. The amber indicatorlight, located adjacent to theselector switch, illuminateswhen the engine synchronizer system is operating.

ENGINE SYNCHRONIZER SWITCH





The “FAN” position should be selected when passenger com-fort is a primary consideration, because fan speed mismatch ismore noticeable in the passenger cabin. The “TURB” positionshould be selected when crew comfort is a primary consider-ation, because turbine speed mismatch is more noticeable inthe flight compartment.

On airplanes 550-0165 and earlier, indication of synchronizersystem performance may be monitored using an optional

synchroscope mounted to the right of the engine instrumentson the upper center instrument panel.

The engine synchronizer system is powered by 28 VDC fromthe left main bus (550-0550 and after) or right main bus (550-0505 and earlier). Circuit protection is provided by a 5-amp(550-0626 and earlier) or 2-amp (550-0627 and after) ENGINE

SYNC circuit breaker located on the left CB panel.




Optional external target-type thrust reversers may be installedallowing engine thrust to be used to help decelerate the air-plane during landing rollout. The thrust reversers consist of twohydraulically-actuated doors per engine. Each door is attachedto the outer exhaust nozzle by means of an aluminum supportcasting and four interconnecting links. The interconnectinglinks are attached to sliding carriage mechanisms that are

driven by the hyrdraulic actuators. When deployed, the re-verser doors are positioned behind the exhaust nozzles. The

upper doors direct thrust upward and forward; the lower doorsdirect thrust downward and forward. When stowed, the reverserdoors are flush contoured to form the aft portion of the enginenacelle. The reverser doors are held in the stowed position bythe overcenter locking design of the linkage.

Thrust Reverser Hydraulics

Hydraulic pressure for thrust reverser operation is supplied by

the same system that supplies the and landing gear and speedbrakes. This section will describe the various valves andswitches that control thrust reverser operation. Refer to Chapter8 for a complete description of the hydraulic system.

Thrust Reverser System





Thrust Reverser Control Valves

The solenoid-operated thrust reverser control valves function todirect hydraulic pressure to, and return flow from, the deploy orstow ports of their corresponding actuators. To accomplish this,each control valve contains an internal selector spool that isspring-loaded to a neutral position and operated by indepen-

dent deploy and stow solenoids. In the neutral position, whenboth solenoids are deenergized, the deploy and stow ports areboth connected to the return line so that hydraulic pressure will

not be trapped in the lines between each control valve and itsassociated actuators. When the deploy solenoid is energized,the selector spool is positioned to direct hydraulic pressure tothe deploy ports of the actuators, and direct return flow fromthe stow ports of the actuators to the reservoir. Conversely,when the stow solenoid is energized, the selector spool ispositioned to direct hydraulic pressure to the stow ports, and

direct return flow from the deploy ports to the reservoir.

The solenoids are energized and deenergized primarily by thethrust reverser (TR) levers through corresponding deploy andstow microswitches. The deploy solenoid is energized throughthe deploy microswitch when its associated TR lever is pulledaft to the “idle deploy” position. The stow solenoid is energizedthrough the stow microswitch when its associated TR lever ispushed forward to the “stow” position.

To prevent overheating the deploy solenoid when energized,

holding voltage is gradually reduced from 28 VDC to 6 VDC bya transistorized circuit while its associated thrust reverser doorsare deployed. To prevent overheating the stow solenoid, limitswitches deenergize the solenoid when its associated thrustreverser doors are fully-stowed.

Thrust Reverser Isolation Valves

The thrust reverser isolation valves, installed upstream of each

control valve pressure inlet, function to isolate the hydrauliccomponents of the corresponding thrust reverser system fromthe main hydraulic system when the thrust reversers are not inuse. The isolation valves are normally spring-loaded closed,and energized open primarily by the TR levers through thedeploy and stow microswitches. Each isolation valve remainsopen when its associated TR lever is in the “idle deploy” posi-tion, and closes when its associated thrust reverser doors are

fully-stowed.




Check Valves and Restrictors

Check valves are installed in the return line to prevent reversefluid flow through the control valve and isolation valve. A re-strictor is installed between the control valve and the actuatorstow ports to prevent the rate of return flow from the actuatorsfrom exceeding the rate of pressure flow to the actuators when

the thrust reversers are deployed at ground speeds above 120knots.

Thrust Reverser Levers

Thrust reverser operation is controlled by the TR leversmounted “piggyback” on the THROTTLE levers. To preventinadvertent thrust reverser operation, the THROTTLE leversmust be set to “IDLE” before the TR levers can be moved.Additionally, thrust reverser deployment is inhibited by eithermain gear safety switch when the airplane is in flight.

During thrust reverser transition, the throttle feedback system

holds the throttle linkage at idle until the reverser doors arefully-deployed or fully-stowed. A pair of locking solenoidsprevents aft movement of the TR levers to increase reversethrust until the reverser doors are fully-deployed.

Warning: The THROTTLE levers should not be restrained manu-ally or by lever friction during takeoff as this will prevent thethrottle feedback system from automatically returning the

throttle linkage and THROTTLE levers to “IDLE” should inad-

vertent thrust reverser deployment occur.

Note: The thrust reversers should be deployed only duringlanding rollout following touchdown and only after all threewheels are on the ground. Thrust reverser deployment prior tonose wheel touchdown can generate sufficient pitch up move-ment to cause aft fuselage contact with the ground.

Note: The thrust reversers should not be deployed duringtouch-and-go landings due to increased takeoff distance result-ing from the time required to restow the reverser doors prior totakeoff power being set.





Thrust Reverser Deployment

Thrust reverser deployment is initiated by pulling the TR leversaft to the “idle deploy” position. In this position, the control valvedeploy solenoid, isolation valve, and hydraulic system centervalve are energized to permit thrust reverser deployment.

1. Stow Position

2. Idle Deploy Position

3. Full Reverse Position

When the center valve isenergized and hydraulicpressure is being supplied to

the thrust reversers, the amber[HYD PRESS ON] annunciator

will be illuminated. Whenpressure downstream of eachisolation valve reaches ap-proximately 200 PSI, a pressure switch causes the correspond-ing amber [ARM] annunciator on the glareshield panel to illumi-nate. As the actuators begin driving the sliding carriages for-ward along guides rods to extend the interconnecting links and

deploy the doors, a stow limit switch is actuated causing thecorresponding amber [UNLOCK] annunciator on the glare-shield panel to illuminate. When the doors have fully-deployed,a deploy limit switch is actuated causing the corresponding

white [DEPLOY] annunciator on the glareshield panel to illumi-nate.

Following thrust reverser deployment, engine thrust is in-

creased by pulling the TR levers further aft. Full reverse thrustis obtained when the TR levers are positioned fully-aft againsttheir stops. The TR lever stops limit reverse thrust N1 to ap-proximately 92% in sea level ISA conditions.

1 2

3




Thrust Reverser Deployment Schematic

Thrust Reverser Stowed Schematic

© PCW

© PCW





Thrust Reverser Stowage

Thrust reverser stowage is initiated by pushing the TR leversforward to the “stow” position. In this position, the control valvedeploy solenoid, isolation valve, and hydraulic system centervalve are energized to permit thrust reverser stowage.

As the actuators begin driving the sliding carriages aft, thecorresponding [DEPLOY] annunciator extinguishes. When thedoors are fully-stowed, the corresponding [UNLOCKED] annun

ciator extinguishes and the control valve deploy solenoid,isolation valve, and hydraulic system center valve will aredeenergized. In this condition, the [HYD PRESS ON] and[ARM] annunciators will also be extinguished.

1. Hydraulic Actuator 4. Stow Limit Switch

2. Sliding Carriage 5. Deploy Limit Switch

3. Interconnect Links 6. Throttle Feedback Linkage

Electrical components of the left and right thrust reverser sys-

tems are supplied with 28 VDC power from the left main busand right main bus respectively, through corresponding 7.5-amp LH and RH THRUST REVERSER (550-0550 and after) or

LH and RH THRU REV (550-0505 and earlier) circuit breakerson the left CB panel. The control valve and isolation valvefunction with an input power of 18 to 30 VDC. When electricalpower is removed from the system, the thrust reversers will failto the stowed position providing the linkage is not in itsovercenter position.

2

4

6

1

5

3




Thrust Reverser

Emergency Stowage

Should either reverser fail tostow in response to TR levermovement due to failure of thedeploy or stow switches or

due to loss of electrical powerto the control circuit, theaffected reverser(s) can be

stowed by positioning thecorresponding emergencystow switch (STOW SW) to the“EMER” position. The STOWSW for each thrust reverser iscorrespondingly located onthe left and right side of the

glareshield panel. Each emergency stow circuit receives elec-trical power through the opposite thrust reverser circuit breaker.

Thrust Reverser Emergency Stowage Test

Emergency stow switch function can be verified on the groundby deploying the reversers normally and then positioning eachswitch to “EMER.” In this condition, the reversers should stownormally, sequentially extinguishing the [DEPLOY] and [UN-LOCK] annunciators, while the [ARM] and [HYD PRESS ON]annunciators remain illuminated. After testing, each TR lever

should be returned to the “stow” position, and the correspond-

ing emergency stow switch should be positioned to “NOR-MAL.”

1. Emergency Stow Switch

2. Thrust ReverserAnnunciators

1

2




Extinguisher Bottles

The extinguisher bottles,located within the aft fuselage,contain Halon-1301 fire extin-guishing agent pressurizedwith dry nitrogen. Each bottle

is equipped with two dis-charge valves and outlets, acombination safety outlet/fill

port, and a pressure gage. Onairplanes equipped with thrustreversers, the bottles arecylindrical with a capacity of 125 cubic inches. On airplaneswithout thrust reversers, the bottles are spherical with a capac-ity of 86 cubic inches.

Annunciator/SwitchesThe fire extinguishing system

is controlled using the left andright red [ENG FIRE PUSH]and white [BOTTLE ARMEDPUSH] annunciator/switcheson the glareshield panel. TheTo prevent inadvertant actua-tion, the [ENG FIRE PUSH]annunciator switches are

guarded by hinged, spring-

loaded, transparent covers.

System Operation

If an overheat condition is detected, the appropriate [ENG FIREPUSH] annunciator/switch will illuminate. Depressing the [ENGFIRE PUSH] annunciator/switch closes the corresponding fueland hydraulic firewall shutoff valves, takes the correspondingstarter/generator off-line, and arms both extinguisher bottles.

When the appropriate [BOTTLE ARMED PUSH] annunciator/ switch is depressed, the corresponding explosive cartridge isdetonated and extinguishing agent is routed through tubingand discharged within the forward nacelle. Airplanes equippedwith thrust reversers also discharge extinguishing agent intothe area between the upper thrust reverser door and the engineexhaust duct assembly through holes in the front flange of theupper thrust reverser door.





© PCW

© PCW

Engine Fire Detection System Schematic

Engine Fire Extinguishing System Schematic




The fire extinguishing system is plumbed such that either orboth extinguishing bottles can be discharged into either engineby pressing the appropriate [ENG FIRE PUSH] and [BOTTLEARMED PUSH] annunciator/switches.

When each extinguisher bottle has been discharged, the corre-

sponding red [ENG FIRE PUSH] annunciator/switch should nolonger be illuminated, indicating that the fire in the associatedengine has been extinguished. If the [ENG FIRE PUSH] annun-

ciator/switch remains illuminated, the remaining extinguishingagent can be discharged into the affected engine by pressingthe appropriate [BOTTLE ARMED PUSH] annunciator/switch.

A gage indicating extinguishing agent supply pressure is lo-cated on each extinguisher bottle within the aft fuselage com-partment. Each gage provides the only indication that extin-

guishing agent may be leaking from its respective bottle. Atable of acceptable bottle pressures per ambient temperature

is placarded adjacent to each pressure gage and/or printed inthe Operating Limitations section of the Operating Manual.

On airplanes 550-0550 and after, 28 VDC electrical power issupplied to the left fire detection system and the left firewallshutoff valves/extinguishing system by the right main bus, andthe right fire detection and the right firewall shutoff valves/ extinguishing system by the left main bus, through correspond-

ing 2-amp LH and RH FIRE DET and 7.5-amp LH and RH F/W

SHUTOFF circuit breakers on the left CB panel. On airplanes550-0505 and earlier, the left and right fire detection systemand the left and right firewall shutoff valves/extinguishing sys-tems are supplied with 28 VDC power from the left main busand right main bus respectively, through corresponding 2-ampLH and RH FIRE DETECT circuit breakers on the left CB panel.





System Testing

A test function is provided totest the fire detection systemannunciator/actuators and tocheck continuity of their asso-ciated sensors and detector

control units. When the rotaryTEST selector switch on thelower left instrument panel is

positioned to “FIRE WARN”,both red [ENG FIRE PUSH]annunciator/actuators shouldilluminate. Pressing either[ENG FIRE PUSH] annunciator/actuator will then illuminate both[BOTTLE ARMED PUSH] annunciator/actuators. The fire extin-guishing system may be tested on the ground or in flight.

Limitations

Refer to the applicable aircraft manufacturers FAA approvedflight manual or approved manual material, markings andplacards, or any combination thereof for all limitations.


Refer to the applicable aircraft manufacturers FAA approved

flight manual or approved manual material (supplementary

checklist) as revised, for procedural information.

TEST SELECTOR SWITCH




Chapter 3Fuel System

Table of Contents

Overview ...............................................3-1

Fuel Storage System .............................3-1

Fuel Tanks .......................................3-1

Servicing...........................................3-2

Fuel Venting System .............................3-5

Fuel Drains ...........................................3-6

Distribution System ...............................3-8

Motive Flow Ejector Pumps ................3-8

Motive Flow Pressure Switches .......... 3-9

Motive Flow Shutoff Valves ................3-9

Boost Pumps ....................................3-9

Fuel Filters .....................................3-11

Maintenance Shutoff Valves.............3-11

Fuel Firewall Shutoff Valves .............3-12

Crossfeed Valves ............................ 3-13




Fuel System, continued

Table of Contents

Fuel System Indication ........................3-14

Fuel Quantity Indication ................... 3-14

Low Fuel Level Warning ...................3-15

Low Fuel Pressure Warning ..............3-15

Fuel Flow Indication.........................3-16

Fuel Remaining/

Consumed Indication .....................3-17

Operational Summaries .......................3-18

Engine Starting ...............................3-18Normal Operation ............................3-19

Low Fuel Pressure ..........................3-19

Crossfeeding...................................3-20

Limitations..........................................3-26





FUEL SYSTEM 12/99 FOR TRAINING PURPOSES ONLY 3-1

Overview

This chapter will describe the Cessna Citation II fuel system, its

major components, controls, operation, and indication. Theairplane fuel supply is contained in two independent wingsystems connected by crossfeed lines. Fuel transfer and deliv-

ery of fuel to the engines is carried out by a distribution systemof motive flow ejector pumps, electrically-powered boostpumps, engine-driven fuel pumps, and control valves. Thedistribution system supplies fuel to the engines in excess ofthat required for all operations. Indication of fuel quantity andflow are displayed by gages on the upper center instrument

panel. Other system indications are displayed on the annuncia-tor panel. Emergency fuel shutoff systems are provided.

Fuel Storage System

The fuel storage system includes the fuel tanks, filler caps,venting system, and fuel drains.

Fuel Tanks

The airplane fuel supply is stored in one integral tank per wing.Each tank occupies roughly the entire inner area of the wingforward of the rear spar, excluding the main wheel well, and isformed by sealing all structural joints between the extremeinboard and outboard ribs. The front spars and all interior ribsincorporate holes that permit fuel migration within the tank.

Outboard interior ribs incorporate baffle plates to prevent rapidfuel load shift when the airplane transitions to and from a wing-

low attitude. The fuel storage area is chemically-treated andfinished with an epoxy primer for corrosion resistance. Duringnormal operation, each engine is supplied by its associatedtank. During crossfeeding, fuel may be supplied from one tankto both engines or, in the event of engine failure, from eithertank to the operative engine. Refer to the fuel system opera-tional summaries in this chapter for a complete description ofthese conditions.




Each fuel tank is divided into a main area and sump area. The

sump area is located within the main area, inboard of the mainwheel well, and functions as a reservoir to which fuel is eithergravity-fed or transferred, and from which fuel is supplied to theengines. Each sump area is enclosed by a cover assembly thathouses the boost pump, the primary motive flow ejector pump,

and a pair of flapper-type check valves. The flapper valvespermit the gravity feeding of fuel from the main area to thesump area while preventing back flow from the sump area to

the main area. This arrangement ensures that sufficient fuel iscontained within each sump area to supply the engines duringall normal maneuvering, and a minimum of five seconds fuelsupply during negative gravity maneuvers not exceeding -0.5G. Aside from gravity feeding, all other transferring of fuel iscarried out by the distribution system.

ServicingEach tank is serviced through a single, flush-mounted filler cap

located on the outboard upper surface of the wing. The loca-tion of the filler caps ensures that sufficient fuel expansionspace will exist within the tanks when topped off. Filler capsecurity should be checked during preflight. A fuel nozzlegrounding point is located on the lower surface of each wingtip.

The fuel tanks have an approximate usable capacity of 371 US

gallons each, and 742 US gallons (5000 LBS) total. The fueltanks should be kept full between flights (providing weight andbalance considerations permit) to reduce explosive vapors andcondensation. Maintaining fuel load symmetry during servicingis unnecessary; however, the maximum permissible asymmetryis 200 LBS during normal flight operations and 600 LBS in anemergency.





Approved fuels are JET A, JET A-1, JET B, JP-4, JP-5, or JP-8,

all with 0.15% PFA55MB anti-icing additive in solution. Whenpreblended fuel is not available, anti-icing additives conformingto MIL-I-27686E (Ethylene Glycol Monomethyl Ether (EGME)) orMIL-I-85470 (Diethylene Glycol Monomethyl Ether (DIEGME))specifications such as “Prist” may be introduced directly into

the nozzle fuel stream during servicing. Concentrations of lessthan 0.06% (20 fluid ounces of additive per 260 gallons of fuelor more) may be insufficient to prevent fuel system icing or

microbiological contamination. Conversely, concentrations ofmore than 0.15% (20 fluid ounces of additive per 104 gallons offuel or less) could cause damage to internal components of thefuel system or erroneous fuel quantity indications.

Caution: EGME and DIEGME are aggressive chemicals andshould not exceed 0.15% of fuel volume. Improperly handled,

these materials will damage the epoxy primer and sealantsused in the fuel tanks, O-ring seals, and any part of the

airplane’s exterior finish with which it comes in contact.

Warning: Anti-icing additives containing EGME or DIEGME areharmful if inhaled, swallowed, or absorbed through the skin,and will cause eye irritation. Refer to all instructions and warn-ings regarding toxicity and flammability before using thesematerials.

All grades of aviation gasoline (AVGAS) conforming to MIL-G-

5572 specifications are approved for use under emergencycircumstances only, providing the airplane is operated in ac-cordance with related procedures and limitations specified inthe AFM. Use of AVGAS is limited to no more than 3500 USgallons or 50 hours of engine operation during any periodbetween engine overhaul. For record keeping purposes, 1 hourof engine operation may be considered equivalent to 70 USgallons.

Note: For a complete listing of approved fuels and additives,refer to the latest revision of Pratt & Whitney Service Bulletin7144.

Warning: Do not allow open flame or smoking in the vicinity ofthe airplane during fuel system servicing.

Warning: Do not operate avionics, communications, or otherelectrical equipment on the airplane during fuel system

servicing.




Warning: Appropriate fire fighting equipment must be availableduring fuel system servicing.

Access panels on the lower surface of each wing are providedto facilitate inspection and maintenance of the fuel tanks aswell as components of the distribution and indicating systems.These panels incorporate liquid-tight perimeter seals. Prior to

flight, the lower surface of each wing should be inspected forevidence of fuel leakage. If observed, the source and cause ofleakage should be determined by maintenance personnel andevaluated against classification criteria and repair action re-quirements. Generally, light to moderate seepage does not limitflight operations, unless observed in proximity to the heatedleading edge panels. Conversely, heavy seepage, runningleaks, or any fuel leakage observed in proximity to the heated

leading edge panels require immediate repair before resuming

flight operations.

Caution: Any fuel leakage caused by structural failure such ascracks, or failure of components such as fuel lines, must berepaired before resuming flight operations.




Fuel Drains

Fuel drain valves, used to check for contamination and toremove water and sediment from the tanks, are located on thelower surface of each wing. Four (550-0060 and after) or six(550-0059 and earlier) “quick-drain” type valves are providedfor each fuel tank. One drain valve is also provided for eachfuel filter assembly. All drain valves are actuated by pushing upthe inner portion (poppet) of the valve assembly.

Note: Rotating the poppet of the quick-drain valve assemblieswhile draining fuel will lock the valve in its open position. This isaccomplished using a phillips head screwdriver (550-0060 andafter) or flat-bladed screwdriver (550-0059 and earlier).





Airplanes 550-0409 and after incorporate semi-flush mounted

fuel tank drain valves which permit poppet O-ring replacementwithout defueling the airplane or valve assembly removal.

Normal preflight fuel draining procedures will generally removemost excess water from the fuel tanks. However, small amounts

of water will remain in solution within the fuel. This residualwater will facilitate microbe and bacterial growth in settlementareas of the fuel system which can lead to fuel flow obstruction

and/or corrosion. To minimize the effects of fuel contamination,the pilot should ensure that the fuel does not contain unap-proved additives and has been properly handled by thesource. Before every flight, a sample should be taken fromeach fuel tank drain and inspected for contamination. If con-tamination is detected, draining from that point should continueuntil contamination is no longer present. If after continued

draining contamination still exists, the airplane should not beflown.

Note: Operators not acquainted with a particular airfield shouldconfirm that the fuel supply there is routinely checked for con-tamination, contains approved additives, and is properly fil-tered before allowing the fuel system to be serviced.

Note: At least 30 minutes should elapse between fuel systemservicing and sample taking.

Warning: Under no circumstances should the airplane be flownwith contaminated or unapproved fuel.




Distribution System

The distribution system functions to maintain a continuoussupply of fuel to the engines in excess of that required for alloperations. Major components of the system include primaryand transfer motive flow ejector pumps, electrically-powered

boost pumps, and engine-driven fuel pumps. Other compo-nents, located in the “dry bay” area between the front and rearspars within each lower fuselage to stub wing fairing, include

maintenance shutoff valves, firewall shutoff valves, crossfeedvalves, motive flow check valves, fuel filters, motive flow pres-sure switches (550-0405 and earlier not incorporating SB550-28-1), and motive flow valves.

Motive Flow Ejector Pumps

Motive flow ejector pumps are operated by fuel pressure sup-

plied primarily by the engine-driven fuel pumps. During enginestart and crossfeeding, motive flow pressure is supplied by the

boost pumps. One primary ejector pump and two transferejector pumps are located in each fuel tank. The transferpumps transfer fuel from the main area of the fuel tank to thesump area. The primary pump supplies fuel from the sumparea to the engine-driven fuel pump. Each primary ejectorpump operates on bypass fuel from its associated engine-driven fuel pump when the engine is operating and the engine-driven fuel pump is developing sufficient pressure to maintain

motive flow. Each pair of transfer ejector pumps operates on

bypass fuel from its associated primary ejector pump.

As motive flow fuel enters each ejector pump through its pres-sure inlet, pressure is reduced by venturi effect thereby draw-ing fuel through its suction inlet. The suction inlet incorporatesa wire mesh fuel strainer to prevent solid particles from enteringthe pump. A spring-loaded, ball-type check valve is installed ineach primary pump outlet to prevent reverse flow through the

pump when its associated boost pump is in operation. Onairplanes 550-0405 and earlier not incorporating SB550-28-1, amotive flow pressure switch is located in each primary ejectorpump supply line. Refer to the motive flow pressure switchdescription and engine starting operational summary in thischapter for functional details.





Motive Flow Pressure Switches

On airplanes 550-0405 and earlier not incorporatingSB550-28-1, a motive flow pressure switch is installed in eachprimary ejector pump supply line. Each switch is normally-closed permitting current flow through the start control circuit toits respective starter motor, ignition system, and boost pump

during the engine starting sequence. After light-off has oc-curred and the engine-driven fuel pump is developing at least180 ± 5 PSI motive flow, the switch is opened interrupting

current flow to the starter motor, ignition system, and boostpump, thereby terminating the starting sequence. During en-gine shutdown, as motive flow pressure decreases belowapproximately 120 PSI, each switch returns to its normally-closed position. The motive flow pressure switches should notbe confused with the fuel pressure switches also described inthis chapter.

On airplanes 550-0406 and after, and earlier airplanes incorpo-

rating SB550-28-1, termination of the starting sequence occursas a function of engine speed. Refer to the engine startingoperational summary in this chapter for functional details.

Motive Flow Shutoff Valves

An electrically-operated motive flow shutoff valve is installed ineach primary ejector pump supply line. Each shutoff valve isnormally-open permitting motive flow to its corresponding

primary ejector pump. During crossfeeding, the shutoff valve

for the system not supplying fuel is energized closed by 28VDC power supplied through its associated 15-amp LH BOOSTor RH BOOST circuit breaker on the left CB panel. Refer to thecrossfeeding operational summary in this chapter for functionaldetails.

Boost Pumps

One electrically-powered, centrifugal-type boost pump is sub-

merged in the sump area of each fuel tank. Each boost pumpprovides fuel pressure to its respective engine during enginestart and opposite engine during crossfeeding. During condi-tions of low fuel pressure or low fuel level, the boost pumpsmay be activated to ensure uninterrupted fuel flow to the en-gines. A check valve is installed in each pump outlet to preventreverse flow through the pump when it is not in operation.




Each boost pump is indepen-

dently controlled by a three-position (ON/OFF/NORM)switch labeled FUEL BOOST,located on the lower left instru-ment panel. Under most

operating conditions, theseswitches should remain in the“NORM” (normal) position. In

this position, boost pumpactivation occurs automati-cally during engine start,crossfeeding, and conditionsof low fuel pressure.

During engine start, automatic boost pump activation is carried

out by a series of relays in the start control circuit that supply28 VDC power to the appropriate pump. During crossfeeding

and conditions of low fuel pressure, automatic boost pumpactivation is carried out by various relays and switches thatsupply 28 VDC power to the appropriate pump(s). Refer to thecorresponding operational summaries in this chapter for com-plete descriptions of these conditions.

With the FUEL BOOST switches set to “ON,” 28 VDC power issupplied directly to each boost pump through its associated

15-amp LH BOOST or RH BOOST circuit breaker on the left CB

panel. In this condition, on airplanes 550-0550 and after, theleft boost pump is supplied by the right main bus and the rightboost pump is supplied by the left main bus. On airplanes 550-0505 and earlier, the left boost pump and right boost pump arecorrespondingly supplied by the left main bus and right mainbus. When set to “OFF,” automatic activation of the boostpumps will occur only during engine start or crossfeeding aspreviously described.

Note: The position of the LH BOOST and RH BOOST circuitbreakers on the left CB panel has no effect on boost pumpoperation during the engine starting sequence.

BOOST PUMP SWITCHES





Indication of boost pump operation is provided by an amber,

three lens [LH] [RH] [FUEL BOOST ON] annunciator (550-0550and after) or by independent, amber [L FUEL BOOST ON] and[R FUEL BOOST ON] annunciators (550-0505 and earlier).When power is being supplied to either boost pump, the corre-sponding annunciator will be illuminated regardless of operat-

ing condition.

Fuel Filters

Each fuel filter incorporates a disposable paper element thatfunctions to trap solid particle contaminants present in the fuel.A differential pressure sensing switch and bypass valve areintegral to each filter head assembly. Should the differentialbetween filter inlet and outlet pressure exceed approximately3.75 PSID, the corresponding amber [LH] or [RH] [FUEL FLTRBYPASS] annunciator (550-0550 and after), or the amber [FUEL

FILT BYPASS] annunciator (550-0505 and earlier) will illuminateindicating filter element obstruction and an impending bypass

condition. Should this differential exceed approximately 4.75PSID, the bypass valve will open providing continued thoughunfiltered fuel flow to the engine.

A spring-loaded drain valve is installed in the base of each filterbowl and extends through the lower surface of the stub wing.Before flight, a sample should be drained from each of thesevalves and inspected for contamination. A manually-operated

shutoff valve in the outlet port of each filter head permits re-

placement of the paper element without fuel drainage from theengine supply line.

Maintenance Shutoff Valves

The maintenance shutoff valves function to isolate componentsof the distribution system downstream of the fuel tanks duringmaintenance operations. Each ball-type valve is spring-loadedto the open position and manually closed by rotating its handle

to a detent in the valve body.





Crossfeed Valves

The crossfeed valvesfunction to permit sup-plying fuel from onetank to both engines or,in the event of engine

failure, from either tankto the operative engine.Each ball-type valve ismotor-operated andelectrically-controlledby a three-position LHTANK/OFF/RH TANKCROSSFEED selector

switch on the lower leftinstrument panel. When positioned to “LH TANK” or “RH

TANK,” both valves are simultaneously energized open by 28VDC power supplied through corresponding 15-amp LHBOOST and RH BOOST circuit breakers on the left CB panel.When open, either or both engines are supplied with fuel fromthe selected tank. When positioned to “OFF,” both valves aresimultaneously deenergized closed and each engine is sup-plied by its associated tank. Refer to the crossfeeding opera-

tional summary in this chapter for functional details.

Note: During crossfeeding, the motive flow shutoff valve for thesystem not supplying fuel is energized closed by 28 VDC

power supplied through the same circuit that opens thecrossfeed valves.




Fuel System Indication

Indication of fuel quantity and flow are displayed by indepen-dent gages located on the upper center instrument panel. Fuelused and fuel remaining are displayed on optional gages,typically located on the lower copilot’s instrument panel.

Fuel Quantity Indication

The fuel quantity indicating system consists of capacitance-type probes, compensator modules, and a dual-channel, verti-

cal-scale FUEL QTY gage.

Five probes measure the level of fuelin each tank. The combined capaci-tance signal generated by theprobes is transmitted to the corre-

sponding channel of the gage wherethey are displayed as a measure-ment of fuel quantity in pounds

(LBS). The compensator module,located in the sump area of eachtank, modifies these signals to cor-rect for changes in fuel temperature.The instrument scale is graduated in100 pound increments between 0 to3000 LBS. The position of indepen-dent left (L) and right (R) white verti-

cal tape bars against the instrumentscale, indicates the fuel quantity ofthe corresponding tank.

28 VDC left main bus power is supplied to the left fuel quantityindicating system through the 2-amp LH FUEL QTY circuitbreaker. 28 VDC right main bus power is supplied to the rightfuel quantity indicating system from the right main bus throughthe 2-amp RH FUEL QTY circuit breaker. On airplanes 550-

0550 and after, the LH FUEL QTY and RH FUEL QTY circuitbreakers are located on the left and right CB panels respec-tively. On airplanes 550-0505 and earlier, the LH FUEL QTYand RH FUEL QTY circuit breakers are both located on theright CB panel.





Low Fuel Level Warning

A normally-open, float-type switch monitors the level of fuel inthe sump area of each tank. When the fuel quantity in eithertank decreases to between 169 and 219 pounds, the affectedfloat switch closes thereby illuminating the correspondingamber [LH] or [RH] [FUEL LOW LEVEL] annunciator (550-0550

and after) or amber [L FUEL LEVEL LO] or [R FUEL LEVEL LO]annunciator (airplanes 550-0505 and earlier). The low fuel levelwarning system operates independently of the fuel quantity andoptional fuel remaining indicating systems.

Low Fuel Pressure Warning

A normally-open, pressure switch monitors motive flow fuelpressure between each primary ejector pump and its associ-ated engine-driven fuel pump. When this pressure falls belowapproximately 5 PSI, the affected fuel pressure switch closes

thereby illuminating the corresponding amber [LH] or [RH][FUEL LOW PRESS] annunciator (550-0550 and after) or amber[L FUEL PRESS LO] and [R FUEL PRESS LO] annunciator (550-0505 and earlier), and activating the corresponding boostpump. The low fuel pressure warning system operates inde-pendently of all other fuel indicating systems. Refer to the fuelsystem operational summaries in this chapter for a completedescription of low fuel pressure conditions.




Fuel Flow Indication

The fuel flow indicating system consists of fuel flow transmitters

and a dual-channel, vertical-scale FUEL FLOW gage.

The fuel flow transmitters are posi-tioned in-line between the FCU and

the oil-to-fuel heat exchanger of eachengine. Each transmitter generatesan electrical signal proportional tothe rate of fuel flow to its associated

engine. These signals are suppliedto the corresponding channel of thegage where they are displayed as ameasurement of fuel flow in poundsper hour (LBS/HR). The instrumentscale is graduated in 100 pound per

hour increments between 0 to 2000LBS/HR. The position of independentleft (L) and right (R) white vertical tape bars against the instru-

ment scale, indicates the rate of fuel flow to the correspondingengine. Although the indicating range of the gage is 0 to 2000LBS/HR, the operating range of each transmitter is approxi-mately 145 to 1800 LBS/HR. Typical fuel flow rates at cruisepower settings vary between approximately 450 and 650 LBS/ HR per engine depending on operating conditions.

28 VDC left main bus power is supplied to the left fuel flow

indicating system through the 2-amp LH FUEL FLOW circuitbreaker. 28 VDC right main bus power is supplied to the rightfuel flow indicating system through the 2-amp RH FUEL FLOWcircuit breaker. On airplanes 550-0550 and after, the LH FUELFLOW and RH FUEL FLOW circuit breakers are located on theleft and right CB panels respectively. On airplanes 550-0505and earlier, the LH FUEL FLOW and RH FUEL FLOW circuitbreakers are both located on the right CB panel. To prevent

erratic indication at low engine power settings, each fuel flowindicating system channel is disabled by a correspondingthrottle cutoff switch when its associated THROTTLE lever ispositioned below approximately 10% N2.





Fuel Remaining/Consumed Indication

Optional fuel remaining or fuel consumed indicators may be

located on the lower copilot’s instrument panel if installed.

On airplanes 550-0550 and after, a four-digit LCD readoutdisplays the total FUEL LBS remaining from signals supplied by

the left and right fuel quantity indicating channels. The indicatorhas a range of 0 to 6000 LBS.

On airplanes 550-0505 and earlier, a four-digit electromechani-

cal counter displays the total FUEL REMAINING in LBS. Onairplanes 550-0062 through -0505, an additional four-digitelectromechanical counter displays the total FUEL CONSUMEDin LBS. Each of these indicators receives signals from the leftand right fuel flow indicating channels and has a range of 0 to9999 LBS. Before takeoff, the FUEL REMAINING indicator must

be set to the known pounds of fuel on board the airplane. TheFUEL CONSUMED indicator must be set to 0000. This is ac-complished using the PRESET knob on the indicator face plate.

Rotating this knob clockwise or counterclockwise correspond-ingly increases or decreases the displayed value. The normalrate of change is approximately 4 pounds per second in bothdirections. When the knob is pressed inward, the rate ofchange increases to approximately 20 pounds per second. Theknob may be locked in this “fast” position when pressed fully-inward, and returned to the “slow” position when pulled out-ward. When released, the knob returns to its spring-loaded

center position.





Normal Operation

With the engine operating and the engine-driven fuel pump

developing sufficient pressure to maintain motive flow, theprimary ejector pump supplies the engine-driven fuel pumpand transfer ejector pumps. En route to the engine-drivenpump, fuel passes through the maintenance shutoff valve, fuel

filter, and firewall shutoff valve. From the engine-driven pump,fuel is directed through the FCU, fuel flow transmitter, oil-to-fuelheat exchanger, and flow divider, to the fuel nozzles.

Low Fuel Pressure

To ensure uninterrupted fuel flow to the engine-driven fuelpumps, boost pump activation occurs automatically as a func-tion of fuel pressure. When the output of either primary ejectorpump falls below approximately 5 PSI, the corresponding fuelpressure switch closes thereby energizing its associated pres-

sure switch relay and boost pump relay. With the boost pumprelay closed, 28 VDC is supplied to the boost pump which thensupplies fuel pressure to the engine-driven fuel pump.

Closure of the fuel pressure switch is indicated by illuminationof the corresponding amber [LH] or [RH] [FUEL LOW PRESS]annunciator (550-0550 and after) or amber [L FUEL PRESS LO]and [R FUEL PRESS LO] annunciator (550-0505 and earlier).Boost pump operation is indicated by illumination of the corre-sponding amber [LH] or [RH] [FUEL BOOST ON] annunciator(550-0550 and after) or amber [L FUEL BOOST ON] and [R

FUEL BOOST ON] annunciator (550-0505 and earlier).

The boost pump relay is initially energized through the pressureswitch relay, but remains energized through an integral latchingcircuit. In this condition, the relay will remain energized closedand the boost pump will continue operating as long as theassociated FUEL BOOST switch remains in the “NORM” posi-tion, regardless of fuel pressure. Should indication of boostpump operation exist without corresponding indication of low

fuel pressure, the boost pump relay circuit should be reset bymoving the switch to “ON” and returning it to “NORM.” Shouldindication of low fuel pressure accompany indication of boostpump operation, the associated FUEL BOOST switch shouldremain in the “NORM” position with the pump operating. Shouldindication of low fuel pressure exist without indication of boostpump operation, the associated FUEL BOOST switch should beset to the “ON” position.




Automatic boost pump activation will occur during conditions oflow pressure only when the FUEL BOOST switches are in the

“NORM” position. When activated during conditions of lowpressure, each boost pump receives 28 VDC power through itsassociated 15-amp LH BOOST or RH BOOST circuit breaker,located in the aft fuselage electrical power “J” box, corre-

spondingly supplied by the left or right main bus. Each pres-sure switch relay and boost pump relay is energized closed by28 VDC power through its associated 15-amp LH BOOST or RHBOOST circuit breaker, located on the left CB panel in the flight

compartment.

To prevent low fuel pressure boost pump activation at lowengine power settings, each pressure switch relay is disabledby a corresponding throttle cutoff switch when its associatedTHROTTLE lever is positioned below approximately 10% N2.

Crossfeeding

Crossfeeding permits fuel to be supplied from one tank to bothengines or, in the event of engine failure, from either tank to theoperative engine. Under normal operating conditions,crossfeeding for the purpose of maintaining fuel load symmetryis seldom necessary unless asymmetry exceeds 200 LBS.When crossfeeding is necessary, the crossfeed selector switchshould be positioned to the tank indicating the higher fuelquantity until fuel load symmetry is achieved.

When either tank is selected, the corresponding boost pump isactivated and both crossfeed valves are simultaneously ener-gized open. Approximately one second is required for thecrossfeed valves to fully open. During this time, a green[INTRANSIT] annunciator, located above the crossfeed selectorswitch, should be illuminated. Approximately three secondsafter the crossfeed valves have fully opened, the non-selectedtank motive flow shutoff valve is energized closed through a

time delay relay. This time delay provides sufficient time for thecrossfeed valves to fully open before motive flow is interrupted,and prevents non-selected tank boost pump activation due tolow motive flow pressure. Should activation of both boostpumps occur, crossfeeding would be prevented due to equalpressure at each crossfeed valve. Therefore, indication of boostpump operation should be closely monitored whencrossfeeding is initiated.





If both boost pumps are activated, the non-selected tank boost

pump should be deenergized by moving the associated FUELBOOST switch to “ON” and returning it to “NORM.” Whencrossfeeding, indication of boost pump operation should occurimmediately, correspond to the selected tank only, and becontinuous until crossfeeding is terminated.

When both crossfeed valves are open, either or both enginesare supplied with fuel under boost pump pressure from the

selected tank. Regardless of which tank is supplying theengine(s), a portion of crossfed fuel is returned to the non-selected tank through the transfer ejector pumps at a rate ofapproximately 600 LBS/HR. As a result, the indicated quantityof the non-selected tank will increase as the indicated quantityof the selected decreases.

When fuel load symmetry is achieved, the crossfeed selectorswitch should be positioned to “OFF.” In this position, the non-

selected tank motive flow shutoff valve is deenergized open.Approximately three seconds later, both crossfeed valves andthe corresponding boost pump are deenergized. Approxi-mately one second is required for the crossfeed valves to fullyclose. During this time, the green [INTRANSIT] annunciatorshould be illuminated. When both crossfeed valves are fully-closed, each engine is supplied by its associated tank.

Automatic boost pump activation will occur during

crossfeeding with the FUEL BOOST switches in the “NORM” or“OFF” position. When activated during crossfeeding, the leftboost pump, crossfeed valve, and motive flow valve are sup-plied with 28 VDC power from the left main bus (550-0550 andafter) or right main bus (550-0505 and earlier) through the 15-amp LH BOOST circuit breaker located on the left CB panel.The right boost pump, crossfeed valve, and motive flow valveare supplied with 28 VDC power from the right main bus (550-

0550 and after) or left main bus (550-0505 and earlier) throughthe 15-amp RH BOOST circuit breaker also located on the leftCB panel.



F U E L S Y S T E M

1 2 / 9 9

F O R T R A I N I N G P U R P O S E S O N L Y

3 - 2 3

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C I T A T I O N I I C H A P T E R 3

1 2 / 9 9

3 - 2 4

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F U E L S Y S T E M

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3 - 2 5

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Limitations

Refer to the applicable aircraft manufacturers FAA approvedflight manual or approved manual material, markings andplacards, or any combination thereof for all limitations.


Refer to the applicable aircraft manufacturers FAA approvedflight manual or approved manual material (supplementary





Chapter 5Flight Controls

Table of Contents

Overview ...............................................5-1

Control Wheels .....................................5-2

Ailerons ................................................5-3

Aileron Trim ......................................5-4

Elevators ..............................................5-6

Elevator Trim ....................................5-7

Electric Elevator Trim.........................5-9

Rudder Pedals ....................................5-10

Rudder ...............................................5-11

Rudder Trim....................................5-12

Control Lock .......................................5-14

Nosewheel Steering.............................5-15

Wing Flaps .........................................5-18

Flap Actuation System.....................5-19

Flap Control ....................................5-19Flap Position Indication ................... 5-20




Flight Controls, continued

Table of Contents

Speed Brakes .....................................5-22

Speed Brake Hydraulics...................5-23

Speed Brake Control Valve ..............5-23

Speed Brake Safety Valve ...............5-23

Speed Brake Thermal Relief Valve ...5-24

Speed Brake Switch ........................5-24

Stall Warning ......................................5-28

Stick Shaker ...................................5-28

Stick Shaker Self Test .................... 5-28

Limitations..........................................5-29





FLIGHT CONTROLS 12/99 FOR TRAINING PURPOSES ONLY 5-1

Overview

The flight controls of the Cessna Citation II consist of staticallymass-balanced ailerons, elevators, rudder, and associated trimsystems for each. The control surfaces are bearing supportedand operated through conventional cable systems and me-

chanical linkage. Trim tabs are attached to the trailing edge ofthe left aileron, right elevator, and rudder. The elevator trim tabis positioned manually or by an electrically operated servo-

actuator. The rudder trim tab and aileron trim tab are positionedmanually only. The manual trim controls and their associatedmechanical position indicators are located on the center ped-estal. This chapter also includes coverage of the wing flaps,speed brakes, nosewheel steering system, stick shaker, andstall warning system.




Control Wheels

Dual control wheels for aileron and elevator operation arepositioned on columns in front of the pilot’s and copilot’s seats.Control wheel rotation is transmitted to a drum around whichthe aileron control cables are attached. The control wheel and

drum are installed on opposite ends of a common shaft that isbearing supported within a cover assembly attached to the topof each column. The control cables are guided by pulleysthrough the interior of the column to the aileron operating link-age. The control wheels are interconnected by crossover andsynchronizing cables such that both rotate simultaneously.

Each control column is bearing supported at its base and

pivots about this point in response to fore and aft control wheelmovement. The control columns are interconnected by a torquetube such that both move simultaneously. Control columnmovement is transmitted to the elevator operating linkage by apush-pull rod attached to the torque tube.





Ailerons

The ailerons are of aluminum alloy, semi-monocoque construc-tion and are attached at two hinge points on the rear spar ofeach wing, outboard of the flaps. On airplanes 550-0049 andafter, and earlier airplanes incorporating SB550-57-2, a small

fence installed on the inboard edge of each aileron functions tomaintain aileron effectiveness by reducing airflow spillagewhen the flaps are extended.

Control wheel rotation is transmitted through cables and pulleysto an aileron sector assembly located below the passengercabin floor. Rotation of the sector is transmitted through cablesand pulleys to an aileron actuator assembly located within eachwing. Each actuator assembly consists of a quadrant, yoke,

and pivot/stop plate. The quadrant rotates on a sealed bearingat its center axis and provides attachment points for the aileroncontrol cables. An additional sealed bearing pressed into thearm of the quadrant provides an off-center pivoting attachment

point for the yoke which serves as the mechanical link betweenthe quadrant and the aileron. As the quadrant rotates, theeccentric motion of the yoke positions the aileron up or downaccordingly.




To prevent overtravel, aileron deflection between approximately

19° up and 15° down is limited by the quadrant arm’s range ofmotion between up and down travel limit stop bolts fitted to thepivot/stop plate.

During autopilot operation, the aileron autopilot servo-actuator

rotates the sector assembly through separate cables andpulleys, and sector rotation is transmitted back to the controlwheels. In the event of malfunction, the servo-actuator may be

overridden with control wheel pressure.

Aileron Trim

Aileron trim is provided by atrim tab attached to the in-board trailing edge of the leftaileron by a full-length, piano-

type hinge. The trim tab isactuated by a pair of push-pull

rods attached to a dual jackscrew type actuator installedwithin the left wing, forward ofthe aileron. The actuator isdriven through a chain/cableassembly that is operated by the aileron trim control knob. Dueto the positioning of the actuator, the trim tab moves in theopposite direction of aileron movement, thereby functioning as

a servo-type trim tab and reducing the control forces required

to position the aileron during flight.

To prevent overtravel, aileron trim tab deflection between ap-proximately 20° up and 20° down is limited by travel stopblocks fitted to the chain/cable assembly.





Ailerons and Aileron Trim

PILOT’SCONTROL

WHEEL

COPILOT’SCONTROLWHEEL

RIGHT CONTROLCOLUMN CABLE

ASSEMBLY

INTERCONNECTCABLES ANDTURNBUCKLE

PULLEYS

CABLE ANDPULLEY

ASSEMBLY

AILERONCABLES

AILERON

QUADRANT

AILERON

TAB ACTUATOR

AILERONCABLES

ACTUATORCHAINS

CHAINGUARD

INBOARDADJUSTABLE

PUSHROD

CABLE ANDPULLEY

ASSEMBLY

TRIMKNOB

SYNCHRONIZINGCABLES

TURNBUCKLES

AILERON SECTOR

ASSEMBLY

FUSELAGECABLES

WING

CABLES

TRIM TABHORN




Elevators

The elevators are of aluminum alloy, semi-monocoque con-struction and are attached at two hinge points each on the rearspar of the horizontal stabilizer. The left and right elevator areinterconnected by torque tubes and an elevator horn therebyforming a single moving elevator assembly.

Fore and aft control wheel movement is mechanically transmit-

ted by a push-pull rod to an elevator sector assembly locatedbelow the flight compartment floor. Rotation of the sector istransmitted through cables and pulleys to an elevator bellcrankassembly located within the aft fuselage. Bellcrank rotation ismechanically transmitted to the elevator assembly by a pair ofpush-pull rods. An elevator bob weight (attached to the controlcolumn interconnect tube) and an elevator down spring (at-tached to the bellcrank) function to improve stability and eleva-

tor balance during flight.

To prevent overtravel, elevator deflection between approxi-mately 20° up and 15° down is limited by the sector’s range ofmotion between up and down travel limit stop bolts, and by thebellcrank’s range of motion between up and down limit stopblocks fitted to their corresponding support brackets.




Elevators and Elevator Trim

ELEVATOR BELLCRANK

ASSEMBLY

ELEVATOR SECTOR

ASSEMBLY

TRIM CONTROL WHEEL

SPROCKET

CABLE/CHAINASSEMBLY

STOP BLOCK

CABLE

BEARING

ELEVATOR TRIM TAB

AND PUSHROD

ASSEMBLY

TAB HORN

TAB PUSHRODSELEVATOR HORN

TORQUE TUBE

ELEVATORBELLCRANK

ELEVATORPUSHRODS

ELEVATOR

TRIM SECTOR

TURNBUCKLES

PULLEYS

ELEVATORTAB ACTUATOR




Rudder Pedals

Pilot and copilot rudder pedals are provided for rudder opera-tion, nosewheel steering, and airplane braking. Each rudder

pedal is fitted to an arm assembly that is suspended from apair of concentric, bearing supported torque tubes. The pilot’sand copilot’s left rudder pedals are connected to the innertorque tube; the pilot’s and copilot’s right rudder pedals areconnected to the outer torque tube. The outer torque tube is

comprised of two sections interconnected by a bridge assem-bly which provides travel clearance for the copilot’s left rudderpedal. Each torque tube incorporates link arms for the attach-

ment of its associated rudder and nosewheel steering controlcables. The inner and outer torque tubes are linked by aninterconnect cable such that each pair of pedals moves simul-taneously.

Each rudder pedal is independently adjustable fore and aft bymeans of a lever located on the pedal arm. Pressing on the

lower end of this lever disengages a spring-loaded locking pinfrom a hub that is fitted to the torque tube. With the locking pin

disengaged, the pedal may be moved to one of three positionsprovided by the hub. Reengaging the locking pin secures thepedal arm in the selected position.





Rudder

The rudder is of aluminum alloy, semi-monocoque constructionand is attached at three hinge points on the rear spar of thevertical stabilizer. Rudder pedal operation is transmittedthrough cables and pulleys to a bellcrank in the aft fuselage.The bellcrank is attached directly to a torque tube extendingfrom the base of rudder. In addition to the three hinge points,

the bellcrank pivots on a bearing/stop plate.

To prevent overtravel, rudder deflection up to approximately22° either side of center is limited by the bellcrank’s range ofmotion between left and right travel limit stop bolts fitted to thebearing/stop plate.

During autopilot operation, the rudder autopilot servo-actuatorrotates the bellcrank through separate cables and pulleys

attached by clevis fittings to the rudder control cables withinthe aft fuselage. Bellcrank rotation is transmitted back to therudder pedals and in the event of malfunction, the servo-actua-tor may be overridden with rudder pedal pressure.




Rudder Trim

Rudder trim is provided by a trim tab attached to the trailingedge of the rudder by a full-length, piano-type hinge. The trimtab is actuated by a pair of push-pull rods attached to a dualjack screw type actuator installed within the vertical stabilizer,forward of the rudder. The actuator is driven through a chain/

cable assembly that is operated by the rudder trim controlknob. Due to the positioning of the actuator, the trim tab movesin the opposite direction of rudder movement, thereby function-

ing as a servo-type trim tab and reducing the control forcesrequired to position the rudder during flight.

To prevent overtravel, rudder trim tab deflection up to approxi-mately 10° either side of center is limited by travel stop blocksfitted to the chain/cable assembly.





Rudder and Rudder Trim

TURNBUCKLE

RUDDER PEDAL

ASSEMBLY

LEFT FORWARDCABLE

RIGHT FORWARDCABLE

OUTER TUBE ASSEMBLY

INTERCONNECT CABLE

PEDAL ASSEMBLY

TRIMCONTROL

WHEEL SPROCKET

CABLE/CHAINASSEMBLY

RUDDER TORQUE TUBE

RUDDERBELLCRANK

TRIM TABHORN

ADJUSTABLEPUSHROD

LEFT CABLEPULLEY

RUDDER TRIM

ACTUATOR

AUTOPILOTSERVO CABLE

AUTOPILOTSERVO CABLE

BEARINGSUPPORTBRACKET




Control Lock

The control lock (also calledthe gust lock) prevents move-ment of and possible damageto the ailerons, elevator, and

rudder by locking these sur-faces in a neutral position andlocking the THROTTLE leversin the "OFF" position. With the

control lock engaged, nose-wheel deflection is limited to60°. The control lock systemconsists of a control handle,cable assemblies, push rods, slide plate, and bellcranks.

Operationally, pulling the control lock handle moves the slideplate, which pulls the aileron, elevator, and rudder cables bycapturing balls swaged onto the cables. The cables also move

the bellcrank, which rotates the throttle locking cams into thelocked position. When control surfaces reach the neutral posi-tion and the throttle cams reach the locked position, the controllock handle should reach its locking detent. Before pulling thecontrol lock handle, the nosewheel should be centered, theaileron control wheels should be level, and the THROTTLElevers should be in the "OFF" position.

The control lock is released by rotating the control lock handle45° and lowering it into the release position.

Warning: The control lock should be released before startingthe engine.

Caution: The control lock should be released before towing theairplane.





Nosewheel Steering

The nosewheelsteering system isoperated by therudder pedals and

allows airplanedirectional controlduring groundoperations. Rudder

pedal operation istransmitted bycables to abellcrank locatedwithin the nosewheel well.

Bellcrank movementis transmitted by aspring-loaded

steering rod to asteering arm thatoperates a steeringgear mechanism mounted atop the trunnion. The steering armand gear mechanism are interconnected by a universal jointthat automatically centers the nosewheel during retraction. Thespring-loaded steering rod (bungee) allows the nosewheel tobe positioned beyond rudder pedal travel limits when using

differential braking or power, or when the airplane is beingtowed.

The nosewheel steering gear mechanism is attached usingshear bolts that are designed to protect nose gear componentsby breaking at a torque load exceeding 15,000 inch pounds,which corresponds to a left or right tow bar excursion of morethan 60° left or right with the control lock engaged or more than95° under any condition. Should the shear bolts break, the

nose gear strut becomes free wheeling. Differential brakingmust then be used for steering control.

1. Steering Bellcrank

2. Steering Rod

3. Steering Arm

3

2

1




Steering is accomplished by allowing the aircraft to roll whiledisplacing the appropriate rudder pedal. The nosewheel can

be steered up to 20° left or right of center during taxi.

When taxiing, the minimum wing tip turning radius using differ-ential braking and partial power is 69.3 feet. This method

causes excessive wear on the tires, and should be employedsparingly. If the airplane is parked with the nose wheel cockedto one side, initial taxiing should proceed with caution.

The nosewheel steering gear mechanism is attached usingshear bolts that are designed to protect nose gear componentsby breaking at a torque load exceeding 15,000 inch pounds,which corresponds to a left or right tow bar excursion of morethan 95°.

The airplane may be towed if the parking brake is not engaged.If the control locks are engaged, turning angle during tow islimited to 60° to avoid control lock damage. If the control locks

are not engaged, turning angle during tow is limited to 95° toavoid breaking the steering gear shear bolts. Should the shearbolts break, the nose gear strut becomes free wheeling. Differ-ential braking must then be used for steering control.

Caution: Nose gear forced beyond the towing stop (95° limit)will shear bolts attaching steering gear assembly to cylinder.





Minimum Turning Radius (wing tip)

1 7 . 6

’

AIRPLANES 550-0627 AND AFTER = 35’

AIRPLANES 550-0626 AND EARLIER = 35’2"

2 0 . 2

’

P I V O T

P O I N T

AIRPLANES 550-0627 AND AFTER = 69’4"

AIRPLANES 550-0626 AND EARLIER = 69’2 ‰ "

TURNING RADIUS:AIRPLANES 550-0627 AND AFTER = 34’8"

AIRPLANES 550-0626 AND EARLIER = 34’7 ‰ "




Wing Flaps

Each wing is fitted with a single-piece, Fowler-type flap posi-tioned by an electromechanical drive system and “track-and-roller” type operating linkage. When extended, the flaps moverearward and downward, effectively modifying wing camberand increasing wing area to reduce the stalling speed of theairplane. When retracted, the flaps form the trailing edge of thewing.

Each flap is of aluminum alloy, semi-monocoque constructionand is attached to the wing structure at three positions by itsoperating linkage. A pair of rollers is attached to the inboardand outboard ends, and lower middle surface of each flap.

Each pair of rollers engages a corresponding inboard, out-board, and center track which extend aft from their attachmentpoints on the wing structure. Three bracket assemblies formattachment points for inboard, center, and outboard bellcrankassemblies, each of which is linked to the flap by a push-pull

rod.





Flap Actuation System

The flap operating linkage iselectromechanically driven by

two 28 VDC motors, reductiongearing, and chain/cableassemblies that actuate the

bellcranks in each wing. Thechain/cable assemblies matewith sprockets fitted to driveshafts that each engages itsassociated reduction gearbox.Each reduction gearbox is

driven by its associated motor.The drive shafts are linked byan interconnect chain suchthat they operate simulta-

neously. Normally, both motors and drive shafts operate to-gether to position the flaps. Should one motor or drive shaft fail,the functioning motor or drive shaft should permit continuedflap operation.

Flap Control

Flap position is controlledusing the FLAP lever locatedon the center pedestal to theright of the THROTTLE levers.The FLAP lever can be set to

any flap position between“FLAPS UP” (0°) and “LAND”(40°); the FLAP lever incorpo-

rates a mechanical detent atthe “T.O. & APPR” (15°) posi-tion. Full flap extension isselected by pushing the FLAP lever fully down past the “T.O. &APPR” detent to the “LAND” position.

1

3

2

1. Bellcrank

2. Flap Track

3. Roller Pivot Point




Flap Position Indication

A flap position indicator is located on the center pedestal to theleft of the FLAP lever. The indicator is mechanically connectedto the flaps by a cable-operated sector assembly; the indicatortherefore moves with the flaps and provides verification that theflaps have assumed the selected position. Down (extend) and

up (retract) position switches are attached to the position indi-cator sector assembly. A cam on the FLAP lever actuates theseswitches. When the FLAP lever is positioned to extend the

flaps, the cam contacts the down position switch, therebyenergizing the flap actuator motors such that they extend theflaps. When the FLAP lever is positioned to retract the flaps, thecam contacts the up position switch, thereby energizing theflap actuator motors such that they retract the flaps. As the flapindicator moves to correspond with FLAP lever position, thesector assembly carries the respective position switch out of

contact with the cam, thereby deenergizing the flap motors. Upand down flap limit switches function as backups to the posi-

tion switch, and deenergize the flap motors when the flapsreach the fully-retracted or fully-extended position.

The landing gear warning horn sounds if the FLAP lever is setbelow the T.O. & APPR position and the gear is not down andlocked, regardless of airspeed or THROTTLE lever position.The horn is energized by one or more downlock switches withinthe landing gear actuators in conjunction with a flap approach

switch incorporated within the FLAP lever assembly. The flap

approach switch is actuated closed by the FLAP lever when setbelow the T.O. & APPR position (approximately 15°).

The flap control circuit and actuator motors receive 28 VDC leftmain bus power through the 5-amp FLAP CONTROL and 15-amp FLAP MOTOR circuit breakers respectively, each locatedon the left CB panel.





Flap Motor and Actuator Assembly

OUTBOARD FLAP BELLCRANK

ACTUATION PULLEY ASSEMBLY

ACTUATION CABLE

FLAP DRIVEGEARBOX

FLAP DRIVEMOTOR

LH RETURN

CABLE ASSEMBLY

TO INBD

BELLCRANK

ACTUATOR

TO FLAPINTERCONNECT

ASSEMBLY

BRACKET AND

PULLEY ASSEMBLY

RIGHT FLAP

INTERCONNECTCABLES

RIGHT FLAP

ACTUATIONCABLES




Speed Brakes

1. Operating Linkage

2. Limit Switch

3. Hydraulic Actuator

Hydraulically-operated, electri-cally-controlled speed brakesare located on the upper and

lower surfaces of the wings,forward of the flaps. Whenextended, the speed brakesincrease drag sufficiently toallow increased airplane rate-

of-descent without exceeding VMO /MMO. The speed brakes mayalso be extended during landing rollout, to spoil lift and provideaerodynamic braking.

The speed brakes are of aluminum-reinforced magnesium alloy

construction and are attached to the rear wing spar at fivehinge points each. Operating linkage for each pair of speedbrakes consists of a bellcrank, push-pull rods, and a hydraulicactuator. The push-pull rods link the upper and lower speed

brakes to the bellcrank such that they operate simultaneously.The cylinder end of the hydraulic actuator is attached to the

wing structure; the rod end is attached to the bellcrank. Normaloperation is initiated by the SPEED BRAKE switch on the centerpedestal.

1

2

3





Speed Brake Hydraulics

Hydraulic pressure for speed brake operation is supplied bythe same system that supplies the landing gear and thrustreversers. This section will describe the various valves andswitches that control speed brake operation. Refer to Chapter 8

for a complete description of the hydraulic system.

Speed Brake Control Valve

The solenoid-operated speed brake control valve functions todirect hydraulic pressure to, and return flow from, the extend orretract ports of the actuators. To accomplish this, the controlvalve contains an internal selector spool that is spring-loadedto a neutral position and operated by independent extend andretract solenoids. In the neutral position, when both solenoidsare deenergized, the extend and retract ports are blocked,

trapping hydraulic pressure in the lines between the controlvalve and the actuators. When the extend solenoid is ener-

gized, the selector spool is positioned to direct hydraulic pres-sure to the extend ports of the actuators, and direct return flowfrom the retract ports of the actuators to the reservoir. Con-versely, when the retract solenoid is energized, the selectorspool is positioned to direct hydraulic pressure to the retractports, and direct return flow from the extend ports to thereservoir.

The solenoids are energized and deenergized primarily by the

SPEED BRAKE switch through up (extend) and down (retract)limit switches. The limit switches are mechanically-actuatedduring speed brake extension and retraction. On airplanes 550-0015 and after, and earlier airplanes incorporating SB550-27-2,the limit switches are actuated by the speed brake bellcranksand lower speed brakes. On airplanes 550-0014 and earlier notincorporating SB550-27-2, the limit switches are integral to thespeed brake actuators.

Speed Brake Safety Valve

The speed brake safety valve, installed between the controlvalve extend port and the return line, functions to inhibit speedbrake extension by relieving hydraulic pressure from the extendlines when engine speed is set above approximately 85% N2.The safety valve is normally energized closed, spring-loadedopen, and controlled by throttle position switches. A check

valve, installed downstream of the safety valve, preventsbackflow from the return line to the extend lines.




Speed Brake Thermal Relief Valve

The speed brake thermal relief valve, installed between thecontrol valve retract port and the return line, functions to relievehydraulic pressure in excess of approximately 1500 PSI fromthe retract lines to prevent system overpressure caused prima-rily by thermal expansion. A check valve, installed downstream

of the relief valve, prevents backflow from the return line to theretract lines.

Speed Brake Switch

Speed brake operation isinitiated by the SPEED BRAKEswitch on the center pedestal.The switch is solenoid-oper-ated, spring-loaded to theupper (RETRACT) positionand requires electrical power

to remain in the lower (EX-TEND) position. The circuitthat supplies electrical powerto the solenoid is completedthrough the throttle position switches when the speed of bothengines is set below approximately 85% N2. When the speedof either or both engines is set above approximately 85% N2,the circuit is interrupted, the switch returns to the "RETRACT"position, and speed brake extension is inhibited.

Note: On airplanes 550-0231 and earlier not incorporating SB-550-27-4, the throttle position switch circuit may be bypassedand the speed brakes extended by holding the SPEED BRAKEswitch in the "EXTEND" position.

Electrical components of the speed brake system are suppliedwith 28 VDC power from the left main bus (550-0550 and after)

or right main bus (550-0505 and earlier) through the 5-ampSPEED BRAKE circuit breaker on the left CB panel. The controlvalve and safety valve function with an input power of 18 to 30VDC. When electrical power is removed from the system, the

speed brakes fail to the retracted position.




Speed Brake Extension Schematic

© PCW

Speed brake extension is initiated by setting the SPEED BRAKEswitch to "EXTEND." In this position, the control valve extendsolenoid, safety valve, and hydraulic system center valve are

energized to permit speed brake extension. During extension,the extend solenoid and center valve are energized through theup limit switches. When the speed brakes are fully-extendedand the up limit switches are open, the extend solenoid andcenter valve are deenergized and the white [SPEED BRAKEEXTEND] annunciator (550-0550 and after) or [SPD BRAKEEXTENDED] annunciator (550-0505 and earlier) is illuminated.In this condition, the hydraulic system returns to "open center"mode and the control valve selector spool returns to its neutral

position, thereby trapping hydraulic pressure in the extendlines and holding the speed brakes in the selected position.

Note: Setting the speed of either or both engines above ap-proximately 85% N2 with the speed brakes extended will causethem to retract.

Note: When the center valve is energized and hydraulic pres-sure is being supplied to the speed brakes, the amber [HYD

PRESS ON] annunciator will be illuminated.





Speed brake retraction is normally initiated by setting theSPEED BRAKE switch to "RETRACT." In this position, the con-trol valve retract solenoid, safety valve, and hydraulic systemcenter valve are energized to permit speed brake retraction.During retraction, the retract solenoid and center valve areenergized through the down limit switches. When the speedbrakes are fully-retracted and the down limit switches are open,the retract solenoid and center valve are deenergized and thewhite [SPEED BRAKE EXTEND] or [SPD BRAKE EXTENDED]

annunciator is extinguished. In this condition, the hydraulicsystem returns to "open center" mode and the control valveselector spool returns to its neutral position, thereby trappinghydraulic pressure in the retract lines and holding the speedbrakes in the selected position.

On airplanes 550-0014 and earlier not incorporating SB550-27-2, a self-locking device integral to each actuator secures its

associated speed brake in the retracted position. On airplanes550-0015 and after and earlier airplanes incorporating SB550-27-2, each lower speed brake incorporates two spring-loadedretainers that function to prevent droop when the actuators aredepressurized following retraction.

Speed Brake Retraction Schematic

© PCW




Stall Warning

Stall warning is provided by stall strips installed on the inboardsection of each wing deice boot. These strips disrupt airflow,thereby providing aerodynamic warning of an impending stallby inducing prestall buffet. Aerodynamic prestall warning buffet

commences at an airspeed of approximately VS1 +10 in theclean configuration and VSO +5 in the landing configuration.

Stick Shaker

The optional stick shaker functions to provide warning of animpending stall by imparting a low-frequency vibration to thepilot’s control column when the AOA system senses an im-pending stall. The stick shaker is mounted on the forward sideof the pilot’s control column and consists of an electric motor,rotating weights, stick shaker relay, resistor, and test switch.

In-flight activation of the stick shaker is initiated by a signalfrom the AOA system that closes an angle-of-attack stall warn-

ing switch in the AOA indicator and energizes the stick shakerrelay. With the relay energized, 28 VDC left main bus power issupplied to the motor through the 5-amp ANG OF ATTACKcircuit breaker. The resistor regulates power to maintain a stickshaker frequency of between 23 and 24 Hz.

Activation of the stick shaker during ground operations is inhib-ited by the left main gear safety switch except during system

testing.

Stick Shaker Self Test

Operation of the stick shaker may be verified by rotating theTEST selector switch on the lower left instrument panel to the“STICK SHAKER” position. If the system is functioning normally,the AOA indicator should flag and drive to zero. The indicatorflag should then pull from view and the indicator needle shoulddrive to 1.0. As the needle passes 0.75, the stick shaker should

activate for several seconds. The cycle should repeat until theTEST selector switch is repositioned.





Limitations

Refer to the applicable aircraft manufacturer’s FAA approvedflight manual or approved manual material, markings andplacards, or any combination thereof for all limitations.


Refer to the applicable aircraft manufacturer’s FAA approvedflight manual or approved manual material (supplementary





Chapter 6Electrical System

Table of Contents

Overview ...............................................6-1

DC Power Sources ................................6-2

Battery .............................................6-2

Battery Overheat Warning .................. 6-3

Starter/Generators ............................6-4

Generator Control Units .....................6-5

Starter/Generator Ground Cooling ......6-7

External Power System ..........................6-8

External Power Requirements .............6-8

Overvoltage/Overcurrent Protection ....................... 6-9

DC Power Distribution .........................6-10

Hot Battery Bus ..............................6-10

Battery Bus ....................................6-10

Left and Right Main Busses ............6-11

Emergency Bus ...............................6-11

Battery Switch.................................6-12

Starter/Generator Switch ................. 6-13

DC System Indication ..........................6-14

Voltmeter........................................6-14

Ammeters.......................................6-14

DC Circuit Protection ........................... 6-15

Circuit Breakers ..............................6-15

Current Limiters ..............................6-15




Electrical System, continued

Table of Contents

Bus Distribution Tables .......................6-34

Alternating Current(AC) Power System..............................6-44

Dual Split Bus AC System ...................6-44

Inverters .........................................6-44

AC Control ......................................6-44

Inverter Testing ...............................6-45

“Tied” Split Bus AC System ................ 6-46

Inverters .........................................6-46

AC Control ......................................6-46


Single Bus System..............................6-48

Inverters .........................................6-48

AC Control ......................................6-48


AC Circuit Protection ...........................6-51

Lighting ..............................................6-52

Exterior Lighting ..................................6-52

Taxi/Landing Lights .........................6-54

Wing Recognition Lights...................6-54

Rotating/Flashing Beacon ................ 6-55

Wing Inspection Light ...................... 6-55

Tail Floodlights ................................6-55




Electrical System, continued

Table of Contents

Interior Lighting ...................................6-56

Control and Instrument Lighting ....... 6-56

Instrument Panel Floodlights ............6-56

Vertical Scale EngineInstrument Light..............................6-56

Counter Light ..................................6-57

Cathode Tube Lights .......................6-57

Map Lights .....................................6-57

Indirect Cabin Lights .......................6-58

Overhead Console Sign ................... 6-58

Entrance Lights ...............................6-58

Emergency Exit Sign ........................6-59

PSU Light .......................................6-59

Passenger Reading Lights................6-59

Aft Fuselage Interior Light ................6-60

Baggage Compartment Lights .......... 6-60

Limitations..........................................6-61





ELECTRICAL SYSTEM 12/99 FOR TRAINING PURPOSES ONLY 6-1

Overview

This chapter describes systems that supply and control air-plane electrical power. Interior and exterior airplane illuminationwill also be discussed.

The Cessna Citation II electrical system is powered by two28 Volt Direct Current (VDC), 400-ampere (amp), negative-ground, engine-driven starter/generators. A 24 VDC nickel-cadmium battery provides current for engine starting and

serves as an emergency source of power. An external powerreceptacle located below the left engine pylon allows an exter-nal power source to energize the airplane for ground operationsor engine start. Electrical power from these sources is distrib-uted to the airplane’s systems through a multiple bus arrange-ment designed to provide continued operation in the event of

an electrical source failure. Items in the system requiring alter-nating current (AC) for their operation are powered by two

inverters rated at 300 to 600 volt-amp (VA).

PRIMARY ELECTRICAL SYSTEM CONTROLS AND INDICATORS




DC Power Sources

The section describes the sources of airplane DC power, in-cluding the battery, starter/generators, and external powersystem.

Battery

The 20-cell 44 amp-hour (550-0550 andafter) or 19-cell 40

amp-hour (550-0505and earlier) nickel-cadmium (NiCad)battery is secured bya hold-down clampto a battery tray

mounted within theaft fuselage. The

amp-hour ratingindicates that whennew and fullycharged, the batteryis capable of deliver-ing one amp ofcurrent for 44 hours(in the case of 44amp-hr battery), before reaching a fully discharged condition.

The following formula: amp-hr rating ÷ amp load = hours avail-able, may be used to approximate hours of battery power avail-able in the event of a dual generator failure. In the case of a 44amp-hr battery, this is accomplished as shown in the followingexample: 44 amp-hr ÷ 88-amp load = 0.5 hours.

Note: A nickel-cadmium battery will maintain a constant outputvoltage during approximately 90 percent of its discharge cycle,

after which available power will rapidly deteriorate. The electro-lyte in a NiCad battery serves only as a conductor and does notreact with the battery plates. Because of these characteristics,the condition of a NiCad battery cannot be reliably determinedby voltage checks or specific gravity readings. However, theplates within a NiCad battery absorb electrolyte as the batterydischarges, so battery condition can be approximated byobserving electrolyte level.

1

3

2

1. Battery Vent Tube

2. Battery Hold Down

3. Battery Tray





Electrolyte level in the NiCad battery should be 1/8 inch abovethe visible insert or plates two to four hours following a full

charge, or 1/4 inch above the visible insert or plates immedi-ately after full charge.

To prevent the accumulation of fluid and vapors, the battery box

is equipped with independent drain and vent tubes that extendthrough the lower surface of the aft fuselage.

It is recommended that the battery electrolyte level be checked

every 100 flight hours or every 14 days, whichever occurs first.Distilled water should be used when servicing is required.

Battery Overheat Warning

Indication of excessive batterytemperature is provided by the

battery overheat warning system.It consists of a battery temperature

sensor, temperature module, andbattery overheat annunciator. Thetemperature sensor is installedbetween the cells near the centerof the battery and provides inputto the temperature module. Bat-tery temperature should remainbelow 145°F. Should battery temperature exceed 145 to 160°F,the temperature module will cause illumination of the red [BATT

O’TEMP] (airplanes 550-0550 and after) or [BATT O’HEAT](550-0505 and earlier) annunciator. If battery temperature ex-ceeds 160°F, the [BATT O’TEMP] or [BATT O’HEAT] annuncia-tor will flash. A battery temperature gage may be optionallyinstalled on airplanes 550-0626 and earlier and is standard onairplanes 550-0627 and after. The battery overheat warningsystem receives 28 VDC left main bus power through the 2-ampBATT TEMP circuit breaker on the left CB panel.

Note: The battery should be serviced if battery temperaturesexceeding 145°F are indicated.

Battery overheat warning system operation may be verified byrotating the TEST selector switch, located on the lower leftinstrument panel, to the “BATT TEMP” position. In this position,a test mode is activated that simulates a temperature exceeding

160°F, thereby causing the [BATT O’HEAT] or [BATT O’TEMP]annunciator to flash.

BATTERY TEMPERATURE GAGE




Starter/Generators

The starter/generators supply 28 VDC electrical power to all

distribution busses, provide charging current to the battery, andare also used as engine starting motors. The starter/generatorsare each capable of producing a continuous current of 400amps.

The starter/generators can each sustain an overload of up to600 amps for five minutes. On airplanes 550-0550 and after,and earlier airplanes incorporating SB550-54-4, sustainedgenerator load is limited to 325 amps above 35,000 feet. Onairplanes 550-0505 and earlier not incorporating SB550-54-4,sustained generator load is limited to 250 amps above 25,000feet.

Each generator is capable of powering the entire airplane

electrical system with the exception of the optional Freon airconditioning system, which should be turned off in the event ofgenerator failure.





Generator Control Units

Once on-line and supplying power, each generator’s output is

controlled by an independent generator control unit (GCU). TheGCU provides voltage regulation, load sharing, ground faultprotection, overvoltage and overexcitation protection, as well asdifferential voltage and reverse current protection. The GCU

also regulates field excitation and starter cut-off functions dur-ing engine start.

Voltage Regulation

Voltage regulation is accomplished by an integrated circuitwhich essentially compares actual generator output to aregulated reference voltage (28.5 VDC ± 1 VDC). Whensensing a differential between these two voltages, the inte-grated circuit regulates field excitation thereby regulatinggenerator output. The circuit also incorporates a field relay

which operates in conjunction with its respective GCU’scontrol relay. Input to each GCU relative to the output of its

respective generator is routed through the 10-amp LH GENSENSE and RH GEN SENSE circuit breakers (not accessiblefrom the flight compartment.)

Load Sharing

Load sharing is accomplished by an equalizer connectionbetween the left and right GCUs and an integrated equalizercircuit resident to each GCU. This circuit essentially “regu-lates” voltage regulator output to maintain load sharing within

± 40 amps (under normal operating conditions). The circuitincorporates an equalizer relay which operates in conjunc-tion with its respective GCU control relay. Essentially, when-ever the control relay is deenergized, the equalizer relay isdeenergized thereby interrupting the equalizer circuit andisolating a “tripped” generator should a ground fault occur.




Reverse Current Protection

Differential voltage and reverse current protection are pro-vided by each GCU through independent sensing circuits.Each differential voltage sensing circuit enables closure ofits associated power relay (permitting current flow from thegenerator to its associated main bus) when the output volt-

age of the generator is within .30 volts of its associated mainbus. Input to each GCU relative to the voltage of its corre-sponding main bus is through the 2-amp LH BUS SENSEand RH BUS SENSE circuit breakers (not accessible fromthe flight compartment). Once the power relay is closed,reverse current protection is enabled thereby preventing afailing generator from imposing a load on the other. In thiscondition, when generator output falls 10% or more below itsrated output, the generator is taken off-line until output isrestored to a level which will ensure forward current flow toits corresponding main bus.

Overvoltage/Overexcitation Protection

Overvoltage and overexcitation protection is provided byeach GCU in conjunction with its voltage regulation and loadsharing equalizer circuits. Should either GCU’s voltageregulation circuit fail, generator output will increase to 35VDC and an overvoltage integrator will trip the associatedfield relay after a predetermined period of time thereby takingthe generator off-line. When the generators are paralleledand sufficiently loaded, overvoltage may not occur; however,

a malfunctioning voltage regulation circuit can result in itsassociated generator assuming a greater percentage of theload. When this occurs, a “deexcitation” signal is providedby the equalizer circuit to the voltage regulation circuit andthe overvoltage integrator thereby taking the generatoroff-line.

Field Weakening

The GCU field weakening feature regulates field excitation toassure that the starter/generator operates as a starter anddoes not generate power during engine start. This function is

accomplished by controlling field excitation so as to main-tain starter/generator interpole winding current below thelevel needed to initiate power generation. The field weaken-ing circuitry initiates current regulation when starter/genera-tor interpole winding current drops below the field weaken-ing threshold value, and continues until starter speedreaches the cut-off value (approximately 40% N2). To avoidnuisance trips, all other GCU protection functions are dis-abled during engine start.





Ground Fault Protection

Ground fault protection is provided by a “feeder to ground

short” sensing circuit integral to each GCU. After engine startand during initial generator “build-up” (between opening ofthe start relay and closing of the control relay), if the genera-tor begins to assume a load equal to or greater than its rated

output before the control relay has closed, a ground short willbe sensed. In this condition, closure of the control relay andcontinuous excitation of the generator are prevented until“tripping” (opening) of the start relay occurs. After the control

relay has closed and the generator is on-line, occurrence ofa ground fault condition will cause its associated field relayto trip open thereby taking the generator off-line. Additionalprotection against generator build-up with an open fieldrelay is a function of the voltage regulation reset circuit. Thiscircuit operates in conjunction with the RESET position of the

generator switches. Essentially, generator build-up with anopen field relay cannot occur until the switch is momentarily

positioned to “RESET.” With the field relay reset, the resetcircuit is isolated such that build-up cannot occur if thegenerator is reset into a ground fault condition.

Starter Cut-Off

A starter cut-off circuit, integral to each GCU, functions toterminate the start sequence as a function of engine speed(approximately 40% N2). Speed sensing is provided by atach drive integral to the starter/generator. The field weaken-

ing circuitry of each GCU, functions to regulate field excita-tion during the engine start sequence until starter cut-offoccurs. Should the starter cut-off circuit malfunction belowcut-off speed, the power relay will be closed, the equalizerrelay will be open, and the starter will continue motoring theengine until manually disengaged.

Starter/Generator Ground Cooling

When operating on the ground, each starter/generator is cooledby an internal fan connected to the generator shaft. The fandraws cooling air through an inlet scoop and duct, located inthe lower forward engine cowling, after which it is exhaustedthrough an outlet in the lower cowling. In flight, starter/generatorcooling is primarily accomplished by ambient air enteringthrough the inlet scoop.




External Power System

The external power systemconsists of an external powerreceptacle located below theleft engine pylon and an

external power relay located inthe aft fuselage electricalpower junction “J” box. Thereceptacle is a three-pin type

design with permanent mark-ings identifying the positiveand negative pins. The relay isenergized closed upon appli-cation of external power,thereby permitting electrical

flow to the hot battery bus.

With the battery switch set to “BATT” and the battery relayclosed, external power is made available to the distributionsystem and charging current is supplied to the battery. Duringengine start using external power, a battery disconnect relayopens, thereby isolating the battery from the distribution system.With 28 VDC external power applied, battery power is con-served during engine starting, or when testing electrical equip-ment on the ground.

External Power RequirementsMaximum external power source output should be at least 28VDC and no more than 1,000 amps.

Caution: If an external power source without reverse currentprotection is turned off while connected to the airplane, rapidbattery discharge and battery damage can result. If the externalpower source is turned off, it should be disconnected from the

airplane.

EXTERNAL POWER RECEPTACLE





Note: External power engine starts may be accomplished withthe generator switches positioned to either “ON” or “OFF;”

however, it is recommended that the generator switches bepositioned to “OFF” during external power engine start. If thegenerator switch is in the “ON” position, the generator will comeon-line automatically upon completion of the start sequence. If

the generator switch is in the “OFF” position, the generator mustbe brought on-line by positioning the generator switch to “ON.”When the generator begins to supply power to the DC bus, anexternal power disable relay automatically disconnects externa

power. Therefore, the generator switch for the operating enginemust be positioned to “OFF” to start the second engine usingexternal power.

Overvoltage/Overcurrent Protection

The overvoltage/overcurrent protection system prevents dam-

age to the starter/generators, avionics equipment, lights, and/orother electrical equipment if external power source voltage and/

or current exceeds limits. The system also prevents batterycurrent from augmenting external power source current, whichwould cause an electrical overload.

The system includes an overvoltage/overcurrent monitor and acurrent sensor. The monitor disconnects external power fromthe airplane electrical system if external power voltage remainsbetween 32 and 33 VDC for more than 200 milliseconds. Themonitor also operates in conjunction with the current sensor to

protect against overcurrent. The current sensor comparesexternal input voltage to a reference voltage supplied by themonitor. When the current sensor detects external power currentexceeding 1100 to 1300 amps for 1.7 to 2.3 seconds, the moni-tor disconnects the airplane electrical system from externalpower.




DC Power Distribution

Electrical power is distributed to the airplane’s systems throughinterconnected busses normally supplied by the battery or thestarter/generators. The busses may also be supplied by anexternal power source.

Hot Battery Bus

The hot battery bus is connected directly to the battery, and isconnected to the battery bus and the emergency bus through

their associated relays. External power is also supplied directlyto the hot battery bus.

Battery Bus

Battery power is routed from the hot battery bus to the batterybus through the battery relay, which is controlled by the battery

switch. When the battery switch is in the “BATT” position andbattery voltage is at least 17 volts, the battery relay is energized

closed and battery power is supplied to the battery bus. Whenset to “OFF” the battery is isolated from all but the hot batterybus. When set to “EMER” the battery is isolated from all but thehot battery bus and the emergency battery bus.

With both engines operating and both generators on-line, thebattery bus is supplied with 28 VDC power from the left andright main busses. With the starter/generators or an externalpower source supplying 28 volts and the battery supplying 24

volts or less, current flow reverses, thereby charging the battery.

Note: If there are no indications of battery power availability tothe system with the battery switch in the “BATT” or “EMER”positions, battery service may be required.

Note: Charging current from the starter/generator will be avail-able to the battery only when the battery switch is set to “BATT.”





Left and Right Main Busses

Electrical power may be supplied to the left and right main

busses by the battery, the starter/generators, or an externalpower source. Battery power or external power are supplied tothe main busses through the battery bus. When the starter/ generators are on-line, 28 VDC power is routed directly to the

main busses through their corresponding power relays. The leftand right main busses are tied together by the battery busthrough corresponding 225-amp current limiters. Each mainbus supplies its associated circuit breaker panel through three

80-amp current limiters and three 75-amp circuit breakers. Themajority of the airplane’s electrical components receive theirpower from these busses.

Emergency Bus

The emergency bus is powered by the battery when the battery

switch is set to “BATT” or “EMER”. When in the “EMER” positionthe battery relay is opened, disconnecting the main DC busses

and the battery bus from the emergency bus. Use of the emer-gency bus enables critical airplane components to be poweredby the battery but electrically isolated from malfunctioningstarter/generator(s) or other components. The emergency buscircuit is protected by a 20-amp EMER POWER circuit breaker(not accessible from the flight compartment).




Battery Switch

The battery switch primarily

controls the battery relaythrough which battery power issupplied to the battery bus orcharging current is supplied to

the battery. The battery switchalso controls the emergencyrelay through which power issupplied to the emergency

bus.

When set to the upper “BATT” position, the battery relay andemergency relay are simultaneously energized closed therebyconnecting the battery to the battery bus and the emergencybus to the hot battery bus (550-0626 and earlier) or to the bat-

tery bus (550-0627 and after). When set to the center “OFF”position, the battery relay and emergency relay are

deenergized open thereby isolating the battery from all but thehot battery bus. When set to the lower “EMER” position, thebattery relay is deenergized open while the emergency relayremains energized closed thereby isolating the battery from allbut the hot battery bus and the emergency bus. In this condi-tion, no more than 30 minutes of battery power is available tothe emergency bus.





Starter/Generator Switch

Each starter/generator switch,

located on the lower left instru-ment panel, controls the oper-ating state of its associatedstarter/generator. When set to

the upper (on) position eachgenerator switch supplies an“on-line” signal to its associ-ated GCU thereby activating

its generator mode functions.

When either switch is set to the center “OFF” position, the on-line signal is interrupted. The momentary “RESET” position ofeach generator switch functions to restore the signal suppliedto the GCU.

It is sometimes necessary to set the generator switch to “RE-

SET” following a windmilling airstart of an engine.




DC System Indication

Indication of electrical system operation is provided by a volt-meter and two ammeters located on the lower left instrumentpanel, and by annunciators located on the upper center instru-ment panel.

Voltmeter

The voltmeter provides a means of monitoring starter/generatoroutput voltage or battery bus voltage as determined by theposition of a selector switch adjacent to the meter. When set to“LH GEN” (left starter/generator), or “RH GEN” (right starter/

generator), the voltage of the selected source is indicated.When set to “BATT,” electrical system voltage monitored at thebattery bus is indicated.

The voltmeter scale is graduated in 1-volt increments between10 and 40 D.C. Volts, with numerical values marked at each 10-volt increment.

Ammeters

Independent left and right ammeters indicate the load carriedby each generator. The ammeter scale is graduated in 50-ampincrements from 0 to 400 amps, with numerical values markedat each 100-amp increment.





DC Circuit Protection

Individual electrical system circuits and components are pro-tected against overload by circuit breakers and current limiters.

Circuit Breakers

Most circuit breakers are located on left and right CB panels onthe flight compartment sidewalls so as to be readily accessibleto the flight crew. The applicable amperage is marked on eachcircuit breaker. Labels above each breaker identifies the circuit

protected. These push-to-reset type circuit breakers will popout, or “trip,” when heat is generated by an electrical overload.Should an overload occur, a tripped circuit breaker may bereset after a cooling period of approximately three minutes bypushing it back in. If the circuit breaker trips a second time, ashort circuit is indicated and it should not be reset, as this could

cause system damage. Additional circuit breakers are installedin various electrical power junction “J” boxes located within the

aft fuselage.

Current Limiters

Primary bus tie circuit protection is provided by a 225-ampcurrent limiter in-line between the battery bus and the left andright main DC busses. Three 80-amp bus feeder current limitersare wired in parallel between each main bus and its associatedcircuit breaker panel bus.

Note: The flight crew should ensure that all circuit breakers areengaged and serviceable fuses are installed before all flights.The airplane should never be operated with any disengagedcircuit breakers or open current limiters without a thoroughknowledge of the consequences.




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Hot Battery Bus

Aft & Nose Baggage Compartment Lights AFT/FWD COMP LT 3Emergency Lights EMER LT 5Engine Ignition (during start) IGNITION 7.5Emergency Power EMERGENCY POWER 20

Battery Bus

Battery Voltmeter BATT VOLTAGE 2

Left Isolation Bus

Left Generator Ammeter LH AMMETER (2) 2Left Generator Sense LH GEN SENSE 10Left Engine Start Light LH START LT 2Left Generator Voltmeter LH VOLTMETER 2

Right Isolation Bus

Right Generator Ammeter RH AMMETER (2) 2

Right Generator Sense RH GEN SENSE 10Right Engine Start Light RH START LT 2Right Generator Voltmeter RH VOLTMETER 2

Left Main Bus

Left Bus Sense LH BUS SENSE 2Left Fuel Boost Pump LH BOOST 15Left Generator Off Light LH GEN OFF 2Left Landing Light LH LDG LT 15Passenger Advisory Lights OXY/SEAT BELT 5Left Recognition Light LH RECOG LT 5Indirect Cabin Lights INDIRECT LT 7.5Entertainment Center ENT CTR 5Tail Flood Lights TAIL LIGHTS 5

Right Main Bus

Right Bus Sense RH BUS SENSE 2Right Fuel Boost Pump RH BOOST 15Right Generator Off Light RH GEN OFF 2Right Landing Light RH LDG LT 15Right Recognition Light RH RECOG LT 5Cabin Lights CABIN LT 7.5Toilet/Shaver Outlet TOILET 7.5

DC Bus Distribution Table (550-0550 and after)

Electrical Power Junction Box

Power Source Circuit

and Equipment Breaker Amperage





Left CB Panel

Left Engine Fan Speed Digital Indicator LH FAN SPEED 2Left Engine Turbine Speed Digital Indicator LH TURB SPEED 2Left 5V Panel Lights LH PANEL 5AC Inverter 1 AC INVERTER NO 1 15Anti-Collision Light ANTICOLL 7.5AOA System Control ANG OF ATTACK 5AOA System Heater AOA HTR 5Battery Overtemp System BATT TEMP 2Cabin Fan CABIN FAN 20Cockpit Voice Recorder VOICE RECORDER 5Left Digital Clock LH CLOCK 2Electroluminescent Panel Lights EL PANEL 1Left Engine Control LH ENG 7.5Engine Synchronization System ENGINE SYNC 5Right Engine Fire Detection RH FIRE DET 2Right Engine Firewall Shutoff RH FW SHUTOFF 7.5

Flap Control FLAP CONTROL 5Flight Data Recorder FLIGHT RECORDER 5Right Fuel Boost Pump RH BOOST 15Left Engine Fuel Flow Indicator LH FUEL FLOW 2Left Fuel Quantity Indicator LH FUEL QTY 2Right Engine Ignition System RH IGN 7.5Left Engine ITT Indicator LH ITT 2Landing Gear Control GEAR CONTROL 5Landing Gear Warning LDG GEAR 2Left Engine Start Control LH START 7.5Wing Inspection Light WING INSP 5Nose Wheel Spinup System NOSE WHL RPM 2Outside Air Temperature OAT 2Left Engine Oil Pressure Indicator LH OIL PRESS 2Left Engine Oil Temp Indicator LH OIL TEMP 2Pitch Trim Control PITCH TRIM 5

Left Pitot/Static Heater LH PITOT STATIC 7.5Power Brakes and Anti-Skid Control SKID CONTROL 20Normal Pressurization System NORM PRESS 5Right Circuit Breaker Panel RH CB PANEL 35Rotating Beacon ROTATING BEACON 5Speed Brake Control SPEED BRAKE 5Standby Gyro STBY GYRO 5True Airspeed Probe Heater (Sperry) TAS HTR 5Cabin Temperature Control TEMP 5Left Engine Thrust Reverser Control LH THRUST REVERSER 7.5Warning Lights 1 WARN LTS 1 2Windshield Bleed Air Control W/S BLEED AIR TEMP 15Windshield Bleed Air Power W/S BLEED AIR 5Wing Navigation Lights NAV 5Cockpit Voice Recorder VOICE RECORDER 5Flap Motors FLAP MOTOR 15


Left Circuit Breaker Panel






Left Main Bus

Left CB Panel Bus LH BUS NO 1 75Left CB Panel Bus LH BUS NO 2 75Left CB Panel Bus LH BUS NO 3 75

Right Crossover Bus

Center 5V Panel Lights CENTER PANEL 5Right 5V Panel Lights RH PANEL 5Windshield Alcohol Pump W/S ALCOHOL 5Right Altimeter Vibrator RH ALT 2Right Digital Clock RH CLOCK 2Emergency Pressurization System EMER PRESS 5Right Engine Control RH ENG 7.5Left Engine Fire Detection LH FIRE DET 2Left Engine Firewall Shutoff LH FW SHUTOFF 7.5Flight Hour Meter FLT/HR 2Left Engine Ignition LH IGN 7.5

Overspeed Warning OVERSPEED 2Right Pitot/Static Heater RH PITOT STATIC 7.5Surface Deice Boots SURFACE DEICE 5Right Thrust Reverser Control RH THRUST REVERSER 7.5Warning Lights 2 WARN LTS 2 5Left Fuel Boost Pump LH BOOST 15









Avionics Bus 1

Automatic Direction Finder 1 ADF 1 3Audio Control 1 AUDIO 1 3Autopilot Warning WARN 3DME 1 (550-0627 and after) DME 1 3EADI 1 (550-0627 and after) EADI 1 5EFIS 1 (550-0627 and after) EFIS 1 5EFIS 1 Control (550-0627 and after) EFIS 1 CONT 1EHSI 1 (550-0627 and after) EHSI 1 5Flight Director 1 FD 1 3Navigation 1 NAV 1 3Radio Altimeter RAD ALT 2Radio Magnetic Indicator 1 RMI 1 2Transponder 1 XPDR 1 3B&D True Airspeed System TAS 2Communication 2 COMM 2 7.5Directional Gyro 1 DG 1 5

Avionics Bus 2

EFIS 2 (550-0627 and after) EFIS 2 5EFIS 2 Control (550-0627 and after) EFIS 2 CONT 1EHSI 2 (550-0627 and after) EHSI 2 5Flight Director 2 FD 2 3Transponder 2 XPDR 2 3Communication 2 COMM 2 7.5DME 2 (550-0627 and after) DME 2 3EADI 2 (550-0627 and after) EADI 2 5True Air Speed Heater TAS HTR 2Audio Control 2 AUDIO 2 3Automatic Direction Finder 2 ADF2 2

Avionics Bus 3

AFIS AFIS 7.5

Autopilot Servo AP 7.5Communication 3 COMM 3 5Flight Management System FMS 5VLF Navigation VLF 5

Avionics Bus 4

Flitefone PHONE 5Multifunction Display MFD DISP 5Multifunction Symbol Generator MFD SYM GEN 7.5Weather Radar RADAR 7.5Radio Magnetic Indicator 2 RMI 2 2

Emergency Bus

Directional Gyro 2 DG 2 3Cockpit Flood Lights FLOOD 7.5Radio Magnetic Indicator 2 RMI 2 2

Navigation 2 NAV 2 3Attitude Director Indicator 2 ADI 2 2Audio Control 1 AUDIO 1 3Communication 1 COMM 1 7.5


Right Circuit Breaker Panel






Right CB Panel

Right Engine Fan Speed Digital Indicator RH FAN SPEED 2Right Engine Turbine Speed Digital Indicator RH TURB SPEED 2Right Engine ITT Indicator RH ITT 2Right Fuel Flow Indicator RH FUEL FLOW 2Right Fuel Quantity Indicator RH FUEL QTY 2Right Oil Pressure Indicator RH OIL PRESS 2Right Oil Temperature Indicator RH OIL TEMP 2Right Engine Start Control RH START 7.5AC Inverter 2 AC INVERTER 2 25Left CB Panel LH CB PANEL 35

Right Main Bus

Right CB Panel Bus RH BUS NO 1 75Right CB Panel Bus RH BUS NO 2 75Right CB Panel Bus RH BUS NO 3 75





Dual Split AC Bus Distribution Table (550-0550 and after)




115 VAC BUS 1

Autopilot Control AP 1Flight Director 1 FD 1 1Sperry True Airspeed System AIR DATA 2Vertical Gyro 1 VG 1 1Weather Radar RADAR 1

115 VAC BUS 2

Flight Director 2 FD 2 1Vertical Gyro 2 VG 2 1

26 VAC BUS 1

EFIS 1 (550-0627 & after) EFIS 1 2Navigation 1 NAV 1 3Radio Magnetic Indicator 1 RMI 1 2

26 VAC BUS 2

Navigation 2 NAV 2 3Radio Magnetic Indicator 2 RMI 2 2Horizontal Situation Indicator 2 HSI 2 2EFIS EFIS 5

Right Sub CB Panel

115VAC 1 115VAC 1 5

115VAC 2 115VAC 2 526VAC 1 26VAC 1 1026VAC 2 26VAC 2 10





Hot Battery Bus

Aft/Forward Cargo Compartment Lights AFT/FWD COMP LT 3Emergency Flood Lights EMER LT 5Engine Ignition IGNITION 7.5Emergency Power EMERGENCY POWER 20

Battery Bus

Battery Voltmeter BATT VOLTAGE 2

Left Isolation Bus

Left Generator Ammeter LH AMMETER (2) 2Left Generator Sense LH GEN SENSE 10Left Engine Start Light LH START LT 2Left Generator Voltmeter LH VOLTMETER 2

Right Isolation Bus

Right Generator Ammeter RH AMMETER (2) 2

Right Generator Sense RH GEN SENSE 10Right Engine Start Light RH START LT 2Right Generator Voltmeter RH VOLTMETER 2

Left Main Bus

Left Bus Sense LH BUS SENSE 2Left Fuel Boost Pump LH BOOST 15Left Generator Off Light LH GEN OFF 2Left Landing Light LH LDG LT 15Passenger Advisory Lights OXY/SEAT BELT 5Left Recognition Light LH RECOG LT 5Indirect Lighting System INDIRECT LT 7.5Entertainment Center ENT CTR 5

Right Main Bus

Right Bus Sense RH BUS SENSE 2

Right Fuel Boost Pump RH BOOST 15Right Generator Off Light RH GEN OFF 2Right Landing Light RH LDG LT 15Right Recognition Light RH RECOG LT 5Cabin Lights CABIN LT 7.5Toilet/Shaver Outlet TOILET 7.5

DC Bus Distribution Table (550-0505 and earlier)

Electrical Power Junction Box






Left CB Panel

Left Engine Fan Speed Digital Indicator LH FAN SPEED 2Left Engine Turbine Speed Digital Indicator LH TURB SPEED 2Left 5V Panel Lights LH PANEL 5AC Inverter 1 AC INVERTER NO 1 15Anti-Collision Light ANTICOLL 7.5AOA System Control ANG OF ATTACK 5AOA System Heater AOA HTR 5Battery Overtemp System BATT TEMP 2Cabin Fan CABIN FAN 20Cockpit Voice Recorder VOICE RECORDER 5Left Digital Clock LH CLOCK 2Electroluminescent Panel Lights EL PANEL 1Left Engine Control LH ENG 7.5Engine Synchronization System ENGINE SYNC 5Left Engine Fire Detection LH FIRE DET 2Left Engine Firewall Shutoff LH FW SHUTOFF 7.5

Flap Control FLAP CONTROL 5Flight Data Recorder FLIGHT RECORDER 5Right Fuel Boost Pump RH BOOST 15Left Engine Fuel Flow Indicator LH FUEL FLOW 2Left Fuel Quantity Indicator LH FUEL QTY 2Right Engine Ignition System RH IGN 7.5Left ITT Indicator LH ITT 2Landing Gear Control GEAR CONTROL 5Landing Gear Warning LDG GEAR 2Left Engine Start Control LH START 7.5Left Wing Inspection Light LH WING INSP 5Nose Wheel Spinup System NOSE WHL RPM 2Outside Air Temperature OAT 2Left Engine Oil Pressure Indicator LH OIL PRESS 2Left Engine Oil Temp Indicator LH OIL TEMP 2Pitch Trim Control PITCH TRIM 5

Left Pitot/Static Heater LH PITOT STATIC 7.5Power Brakes and Anti-Skid Control SKID CONTROL 20Normal Pressurization System NORM PRESS 5Right Circuit Breaker Panel RH CB PANEL 35Rotating Beacon ROTATING BEACON 5Speed Brake Control SPEED BRAKE 5Standby Gyro STBY GYRO 5True Air Speed Probe Heater (Sperry) TAS HTR 5Cabin Temperature Control TEMP 5Left Engine Thrust Reverser Control LH THRUST REVERSER 7.5Warning Lights 1 WARN LTS 1 2Windshield Bleed Air Control W/S BLEED AIR TEMP 5Windshield Bleed Air Power W/S BLEED AIR 5Wing Navigation Light NAV 5Cockpit Voice Recorder Voice Recorder 5Flap Motors FLAP MOTOR 15

Left Main Bus

Left CB Panel Bus LH BUS NO 1 75Left CB Panel Bus LH BUS NO 2 75Left CB Panel Bus LH BUS NO 3 75









Right Crossover Bus

Center 5V Panel Lights CENTER PANEL 5Right 5V Panel Lights RH PANEL 5Winshield Alcohol Pump W/S ALCOHOL 5Right Altimeter Vibrator RH ALT 2Right Digital Clock RH CLOCK 2Emergency Pressurization System EMER PRESS 5Right Engine Control RH ENG 7.5Right Engine Fire Detection RH FIRE DET 2Right Engine Firewall Shutoff RH FW SHUTOFf 7.5Flight Hour Meter FL/ HR 2Left Engine Ignition LH IGN 7.5Overspeed Warning OVERSPEED 2Right Pitot/Static Heater RH PITOT STATIC 7.5Surface Deice Boots SURFACE DEICE 5Right Thrust Reverser Control RH THRUST REVERSER 7.5Warning Lights WARN LTS 5

Left Fuel Boost Pump LH BOOST 15Equipment Cool Equipt Cool 7.5

Emergency Bus

Directional Gyro 2 DG 2 3Cockpit Flood Lights FLOOD 7.5Navigation 2 NAV 2 3Communication 1 COMM 1 7.5








Right CB Panel

Right Engine Fan Speed Digital Indicator RH FAN SPEED 2Right Engine Turbine Speed Digital Indicator RH TURB SPEED 2AC Inverter 2 AC INVERTER NO 2 15Right Fuel Flow Indicator RH FUEL FLOW 2Right Fuel Quantity Indicator RH FUEL QTY 2Right Engine ITT Indicator RH ITT 2Left Circuit Breaker Panel LH CB PANEL 35Right Engine Oil Pressure Indicator RH OIL PRESS 2Right Oil Temperature Indicator RH OIL TEMP 2Right Engine Start Control RH START 7.5Communication 2 COMM 2 7.5Distance Measuring Equipment 2 DME 2 3Transponder 2 XPDR 2 3Automatic Direction Finder 2 ADF 2 2Communication 3 COMM 3 5Audio Control 2 AUDIO 2 3

AC Switch AC SWITCH 3Autopilot Warning WARN 3Area Navigation AREA NAV 2Ground Proximity Warning GROUND PROX 1True Air Speed Heater HTR TAS 15VLF Navigation VLF NAV 7.5Nav Data Bank NAV DATA BANK 5Flight Management Systems FMS 7.5Weather Radar RADAR 7.5

Right Main Bus

Right DC Bus Number 1 RH BUS NO 1 75Right DC Bus Number 2 RH BUS NO 2 75Right DC Bus Number 3 RH BUS NO 3 75

Left Crossover Bus

Navigation 1 NAV 1 3Automatic Direction Finder 1 ADF 1 3Audio Control 1 AUDIO 1 3Distance Measuring Equipment 1 DME 1 3Attitude Director Indication 1 ADI 1 5EFIS Disp EFIS Disp 1EFIS EFIS 5EHSI 1 EHSI 1 5Flight Director 1 FD 1 3Radio Altimeter RAD ALT 2Radio Magnetic Indicator 1 RMI 1 2Transponder 1 XPDR 1 3Communication 2 COMM 2 7.5Directional Gyro 1 DG 1 5Autopilot Servo AP 7.5Flitefone PHONE 5

Horizontal Situation Indicator 1 HSI 1 5Cockpit Voice Advisory VOICE ADV 5Flight Management System FMS 5

Emergency Bus

Navigation 2 NAV 2 2Communication 1 COMM 1 7.5Directional Gyro 2 DG 2 3Cockpit Flood Lights FLOOD 5









115 VAC BUS 1

Autopilot Control AP 1Flight Director 1 FD 1 1Sperry True Airspeed System AIR DATA 2Vertical Gyro 1 VG 1 1Weather Radar RADAR 1

115 VAC BUS 2

Flight Director 2 FD 2 1Vertical Gyro 2 VG 2 1

26 VAC BUS 1

Navigation 1 NAV 1 3RMI /ADF 1 RMI /ADF 1 2Attitude Director Indicator 1 ADI 1 1

26 VAC BUS 2

Navigation 2 NAV 2 3Radio Magnetic Indicator 2 RMI 2 2Horizontal Situation Indicator 2 HSI 2 2EFIS EFIS 2RMI /ADF 2 RMI /ADF 2 2Attitude Director Indicator 2 ADI 2 2

Right Sub CB Panel

115VAC 115VAC 526VAC 26VAC 10

Split AC Bus Distribution Table (550-0505 and earlier)




115 VAC BUS 1

Flight Director 1 FD 1 1Flight Director 2 FD 2 1Vertical Gyro 1 VG 1 1Vertical Gyro 2 VG 2 1AC Monitor AC MONITOR 3Sperry True Airspeed System AIR DATA 2Autopilot Control AP 1Weather Radar RADAR 1

26 VAC BUS 1

Navigation 1 NAV 1 3Radio Magnetic Indicator 1 RMI 1 2

Radio Magnetic Indicator 2 RMI 2 1Automatic Direction Finder 1 ADF 1 2Automatic Direction Finder 2 ADF 2 2Attitude Director Indicator 1 ADI 1 2Horizontal Situation Indicator 1 HSI 1 1

Single AC Bus Distribution Table (550-0505 and earlier)







Alternating Current (AC) Power System

Various airplane components requiring AC power for theiroperation are supplied by AC inverters through a multiple busdistribution system which varies depending on airplane con-figuration and unit number.

Dual Split Bus AC System (550-0550 and after)

Inverters

During normal operation, each AC bus is supplied by its asso-ciated inverter. Should failure of either inverter occur, the oppo-site inverter supplies power to all AC busses through an auto-matic switching circuit. When an inverter has failed, the red [ACFAIL] annunciator and the corresponding [1] or [2] [INV FAIL]annunciator will illuminate, and the red [MASTER WARNING]

light/switch will flash. Resetting the [MASTER WARNING] light/ switch will extinguish the [AC FAIL] annunciator but will not

reset the failed inverter.

Should an AC Bus circuit breaker trip, either or both [INV FAIL]annunciators, the [AC FAIL] annunciator and the red [MASTERWARNING] light/switch will illuminate. The tripped circuitbreaker may be reset by pushing it back in, and the annuncia-tors extinguished by pressing the [MASTER WARNING] light/ switch which may return either or both inverters to operationproviding the fault has cleared. If the A/C Bus circuit breaker

cannot be reset, the bus isolated by the associated circuitbreaker is no longer energized and all systems powered by itbecome inoperative.

AC Control

Positioning the AC switch tothe upper (on) position acti-vates inverter 1 and 2, momen-

tarily illuminating the [ACFAIL] annunciator until bothinverters are on-line and inphase. A synchronizationcircuit between the inverters isused as the reference forphase relationship.





Inverter Testing

Each inverter may be tested by positioning the AC switch to the

upper (on) position and holding the test switch to the “INV 1” or“INV 2” position and observing illumination of the correspond-ing [1] or [2] [INV FAIL] annunciator. When the test switch isreleased the annunciator should extinguish.

Each 115 VAC and 26 VACdual split bus circuit breaker,located on the right flight

compartment sidewall, func-tions to isolate its associatedAC bus when overloaded andto illuminate the [AC FAIL]annunciator.

AC SYSTEM CIRCUIT BREAKERS





Inverter Testing

Each inverter may be tested by positioning the AC switch to the

upper (on) position and holding the test switch to the “INV 1” or“INV 2” position and observing illumination of the correspond-ing [1] or [2] [INV FAIL] annunciator. When the test switch isreleased the annunciator should extinguish.

Each 26 VAC bus and 115VAC bus are tied togetherthrough corresponding 26V

and 115V AC POWER BUSTIE circuit breakers located onthe right flight compartmentsidewall.

AC SYSTEM CIRCUIT BREAKERS





Dual Split Bus AC Power System (550-0550 and after)

AC Power Indication (550-0550 and after)




Lighting

The airplane is equipped with a variety of exterior lights tofacilitate takeoff, landing, in-flight recognition, and wing inspec-tion, plus interior lights that provide varying degrees of cockpit,cabin, and baggage compartment illumination.

Exterior Lighting

Exterior lighting consists of navigation lights, anti-collisionlights, taxi/landing lights, a rotating or flashing beacon, tail floodlights, a wing inspection light, and recognition lights. Exteriorlighting switches are located on the lower left instrument paneland the pilot’s lower instrument panel.

LOWER LEFT INSTRUMENT PANEL PILOT’S LOWER INSTRUMENT PANEL





1. Navigation Lights

Navigation lights are installedon each wing tip and on thetailcone stinger. The lights arecontrolled by the NAV switchand supplied with 28 VDC

power from the right main bus(550-0550 and after) or leftmain bus (550-0505 and

earlier) through the NAV circuitbreaker on the left CB panel.

When darkness permits, operation of the navigation lights canbe confirmed during preflight by observing their reflection onthe ground and/or other surrounding objects. During daylighthour preflight, the operation of each light should be confirmed

from outside of the airplane.

2. Anti-Collision Strobe Lights

High-intensity anti-collision strobe lights are installed on eachwing tip adjacent to the navigation lights. The strobe lights areenergized by bus voltage boosted through independent powersupplies installed within the wing tip. The lights are controlledby the ANTI COLL switch and supplied with 28 VDC left mainbus power through the ANTI COLL circuit breaker on the left CBpanel.

Note: Strobe lights should not be operated in clouds, fog, orhaze as their reflection on water droplets in the atmosphere caninduce disorientation or vertigo.

Note: To avoid interfering with the vision of other pilots, strobelights should not be operated when taxiing in the vicinity ofother aircraft.

1

2




Taxi/Landing Lights

A sealed beam taxi/landinglight is installed on the forwarddrag brace of each main gearassembly (airplanes not incor-porating SB550-33-5), or on

each main gear door (air-planes incorporating SB550-33-5). The lights are exposed

and operable only when themain gear is extended. Eachlight is independently con-trolled by its associated LH or RH LANDING switch. 28 VDCpower is supplied to each from its corresponding left or rightmain bus, through circuit breakers located in the aft fuselageelectrical power junction “J” box.

Wing Recognition Lights

Optional recognition lights,used to provide additionalexterior lighting and increaseairplane visibility, are installedon the leading edge of eachwing tip. Glareshields areinstalled slightly inboard of thelights to reduce glare within

the cockpit. The recognition

lights are controlled by theRECOG switch. 28 VDC poweris supplied to each from its corresponding left or right main bus,through circuit breakers located in the “J” box.

Caution: Recognition lights use a pressurized, halogen cycle-type lamp which produces extremely high intensity light which

could cause eye damage if viewed directly. The lamp shouldbe protected from abrasions, scratches, impact, and contactwith liquids. Handling the lamp should be avoided. Allow thelamp to cool, and wear protective clothing and dark glasses ifcontact is necessary.

Note: Recognition lights should be turned on shortly beforetakeoff and during descent, and extinguished during climb,

cruise, and after landing as soon as the airplane is clear of therunway.





Rotating/Flashing Beacon

Rotating or flashing beacons, used to increase airplane visibil-ity, may be installed on the top of the rudder. Any of three differ-ent types of beacons may be installed on individual airplanes,two of which use flashing lamps and one of which uses a rotat-ing light assembly that employs a stationary lamp and electri-

cally-driven rotating reflector. The beacon is controlled by theBEACON switch and supplied with 28 VDC left main bus powerthrough the ROTATING BEACON circuit breaker on the left CB

panel.

Wing Inspection Light

The wing inspection light,used to check for ice accumu-lation on the leading edge ofthe wing during night opera-

tions, is flush-mounted on theleft side of the fuselage for-

ward of the wing. The light iscontrolled by the WING INSPswitch and supplied with 28VDC power from the left mainbus (550-0550 and after) orright main bus (550-0505 and earlier) through the WING INSPcircuit breaker on the left CB panel.

Note: Operation of the wing inspection light is mandatory for

flight in icing conditions as defined by the FAA.

Tail Floodlights

Optional floodlights may be installed on the upper left and rightsurfaces of the horizontal stabilizer to illuminate the verticalstabilizer. These lights may also be referred to as identificationlights, logo lights, or tail lights. The floodlights are controlled by

the BEACON switch (550-0038 and after) or the NAV switch(550-0037 and earlier), and supplied with 28 VDC left main buspower through the TAIL LIGHTS circuit breaker located in the“J” box.




Interior Lighting

Airplane interior lighting systems provide illumination, advisoryand warning within the flight compartment and passengercabin, plus baggage compartment illumination.

Control and Instrument LightingElectroluminescent “backlit” panels provide primary illumina-tion of labels on various switch, control, and circuit breakerpanels as well as the center pedestal. The electroluminescentpanels are powered by 40 to 60 VAC, 400 Hz inverters. Addi-tional panel and instrument illumination is provided by integral5 VDC powered lights. Panels and instruments that are notinternally lighted are illuminated by 5 VDC or 28 VDC poweredpost lights. Panel and instrument illumination is controlled byON/OFF switches and rheostats located on the pilot’s lowerinstrument panel.

PANEL AND INSTRUMENT LIGHT CONTROLS

Instrument Panel Floodlights

The instrument panel floodlights, located on the aft overheadconsole behind a blue tinted lens, are normally used duringthunderstorms to provide supplemental instrument panel illumi-nation to compensate for lightning-induced night vision loss.These lights are controlled by the FLOOD LTS rheostat andsupplied with 28 VDC emergency bus power through theFLOOD circuit breaker on the right CB panel.

Vertical Scale Engine Instrument Light

The vertical scale engine instrument light, located on the lowersurface of the glareshield panel fire tray, illuminates the verticalscale engine instruments on the upper center instrument panel.





This light is automatically illuminated during the engine start

sequence and extinguished upon start sequence completion.During engine start, the light is supplied with 28 VDC emer-gency bus power through the EMER LT circuit breaker locatedin the “J” box. During normal operation, the light is suppliedwith 28 VDC emergency bus power through the FLOOD circuit

breaker.

In the event of electrical system failure the light is supplied with

28 VDC power from the standby gyro battery pack (if installed).The light is also supplied with 28 VDC power from an emer-gency lighting battery pack, located above the cabin headliner,through an inertial switch which closes when exposed to anacceleration force of 5Gs or more.

Counter Light

The post-type counter light illuminates the mechanical counterlocated above the FAN tachometer on the upper center instru-

ment panel. This light is controlled by the CENTER panel rheo-stat and supplied with 28 VDC left main bus power through theCENTER PANEL circuit breaker on the left CB panel

Cathode Tube Lights

Two cathode tube lights are installed under the glareshield toprovide supplemental instrument panel lighting. The lights arepowered by a high-voltage inverter located within the left side

console. This light and the inverter are controlled by the EL

panel rheostat and supplied with 28 VDC left main bus powerthrough the EL PANEL circuit breaker on the left CB panel.

Map Lights

Map lights, providing direc-tional flight compartmentillumination, are located on the

overhead console. Each lightis independently controlled byan associated rheostat, lo-cated on the forward end ofeach side console, and sup-plied with 28 VDC right mainbus power through the 5-ampRH PANEL circuit breaker on

the left CB panel.




Indirect Cabin Lights

Primary cabin illumination isprovided by an indirect light-ing system, consisting oftwelve overhead fluorescentlights powered by two invert-

ers located above the cabinheadliner. The lights and theinverters are controlled by a

three-position (OFF/BRIGHT/ DIM) switch, located on thecabin sidewall forward of theentrance door, and suppliedwith 28 VDC left main bus power through the INDIRECT LTcircuit breaker located in the “J” box. The indirect lightingsystem is optional on airplanes 550-0550 and after, and stan-

dard on airplanes 550-0505 and earlier.

Overhead Console Sign

The lighted overhead consolesign, mounted just aft of theforward divider, displaysuniversal “no smoking” and“fasten belt” symbols. Thislight is controlled by a PASSSAFETY switch and supplied

with 28 VDC left main power

through the OXY/SEAT BELTSIGN circuit breaker located inthe “J” box.

Entrance Lights

Entrance lights are provided to illuminate the passenger doorentrance, emergency exit door, and aft baggage compartment.

Each entrance light may be illuminated by an integral switch, orby the PASS SAFETY switch. 28 VDC hot battery bus power issupplied to these lights through the CABIN LIGHTS circuitbreaker located in the “J” box.

In the event of electrical system failure these lights may besupplied with 28 VDC power from the emergency lightingbattery pack and are automatically illuminated through the

inertial switch with an acceleration force of 5Gs or more.

OFF/BRIGHT/DIM SWITCH





Emergency Exit Sign

A lighted EMERGENCY EXIT sign is located over the emer-gency exit door to clearly identify its location. The light is con-trolled by an illuminated switch located on the cabin sidewallforward of the entrance door, or by the PASS SAFETY switch. 28VDC left main bus power is supplied to the light through the

OXY/SEAT BELT circuit breaker located in the “J” box.

In the event of electrical system failure this light may be sup-

plied with 28 VDC power from the emergency lighting batterypack and is automatically illuminated through the inertial switchwith an acceleration force of 5Gs or more.

PSU Light

On airplanes 550-0550 and after, an optional fluorescent light isavailable to provide passenger service unit (refreshment center

or vanity) area illumination. The light is controlled by the three-position (OFF/BRIGHT/DIM) switch, located on the cabin

sidewall forward of the entrance door, and supplied with 28VDC left main bus power through the INDIRECT LT circuitbreaker located in the “J” box.

Passenger Reading Lights

Reading lights, providingdirectional illumination, arelocated above each passen-

ger station. Each light is inde-

pendently controlled by anintegral switch and suppliedwith 28 VDC right main buspower through the CABINLIGHTS circuit breaker on theleft CB panel.





Limitations

Refer to the applicable airplane manufacturers FAA approvedflight manual or approved manual material, markings andplacards, or any combination thereof for all limitations.


Refer to the applicable airplane manufacturers FAA approved

flight manual or approved manual material (supplementarychecklist) as revised, for procedural information.




Chapter 7Flight Instrumentation

Table of Contents

Overview ...............................................7-1

Flight Environment Data System.............7-2

Pitot-Static System ................................7-2

Pitot Tubes......................................7-2

Static Ports .....................................7-3

Alternate Static Air Source................7-3

Airspeed/Mach Indicators ...................... 7-4

Airspeed Pointer...............................7-4

Mach Sub-Dial .................................7-4Airspeed Mach Indicator Markings.....7-5

Index Marker ...................................7-5

Overspeed Warning System ................... 7-6

True Airspeed System (optional).............7-7

Sperry TAS System ..........................7-7

B&D TAS System.............................7-8

Barometric Altimeters ....................... 7-9

Pilot’s Altimeter .............................7-10

Copilot’s Altimeter..........................7-11

Altitude Alerting and

Reporting System................................7-12

Altitude Alerting..............................7-12

Altitude Reporting...........................7-13

Vertical Speed Indicators .....................7-15




Flight Instrumentation, continued

Table of Contents

Radio Altimeter System (optional) ........ 7-16

Transceiver and Antennas...............7-16

Conventional Indicator(excluding RAD/BAR) ...................... 7-17

Conventional Indicator (RAD/BAR) ... 7-18

Mechanical Flight DirectorRadio Altitude Indication .................7-19

EFIS Radio Altitude Indication ......... 7-20

Attitude and Direction System..............7-21

Conventional Attitude Indicator ........7-21Air Driven Gyro...............................7-22

Gyro Pressure Gage .......................7-23

Electrically-Driven Gyro .................... 7-23

Turn-and-Bank Indicator ...................7-24

Mechanical Flight Directors .................. 7-25

Attitude Directional Indicator ........... 7-25

Horizontal Situation Indicator .......... 7-25

Gyro Slaving ..................................7-26

ADI Display Features ...................... 7-27

HSI Display Features ......................7-31

Electronic Flight

Instrumentation System .......................7-33

Autopilot/Flight Director System ........... 7-33

Limitations..........................................7-33




FLIGHT INSTRUMENTATION 12/99 FOR TRAINING PURPOSES ONLY


7-1

Overview

PILOT’S FLIGHT INSTRUMENTATION

The primary flight instrumentsof the Cessna Citation II arepositioned on panels directlyin front of the pilot and copilot.Pitot-static instrumentationconsists of airspeed/mach

indicators, barometric altim-eters, and vertical speedindicators. Gyroscopic instru-mentation consists of attitudedirectional indicators (ADIs)and horizontal situation indica-tors (HSIs) of mechanical orelectronic (EFIS) type depending on installation. Turn coordina-tion information is provided either by independent mechanicalindicators or by a rate-of-turn indicator and conventional incli-nometer integral to each ADI/EADI. A standby attitude indicator

may be installed as an emergency backup to the ADI(s) orEADI(s). Additional navigational guidance is provided by radiomagnetic indicators (RMIs) displaying both VOR and ADFmagnetic bearing information, an optional radio altimeter, and amagnetic compass mounted on the windshield center postabove the glareshield. An outside air temperature (OAT) indica-tor, clock, and optional angle-of-attack (AOA) indicator arelocated on the upper left instrument panel. An optional trueairspeed (TAS) system may also be installed.

COPILOT’S FLIGHT INSTRUMENTATION




For organizational purposes, this chapter is divided into sepa-rate flight environment data and attitude and direction systemsaccording to the conditions or forces utilized in their operation.The flight environment data system includes equipment andinstruments which are sensitive to environmental conditionswhich influence navigation. The attitude and direction system

includes equipment and instruments which are sensitive togyroscopic, inertial, and magnetic forces which influence navi-

gation.

Flight Environment Data System

The flight environment data system includes the pitot-staticsystem and associated flight instruments, the overspeed warn-ing system, altitude alerting and reporting system, as well asthe optional true airspeed (TAS), radio altimeter, and AOA

systems when installed.

Pitot-Static System

The pilot’s and copilot’s pitot-static flight instruments are sup-plied by independent pitot-static systems consisting of onepitot tube and two static ports each.

Pitot Tubes

The pitot tubes are located onthe lower left and right surface

of the nose section. The leftpitot tube supplies ram pres-sure to the pilot’s airspeed/ mach indicator, the air datacomputer, and airspeed/machwarning switch. The right pitot

tube supplies ram pressure tothe copilot’s airspeed indica-tor, the landing gear warningairspeed switch (550-0627and after), B&D TAS pressure transducer (if installed), andairspeed/mach warning switch (Canadian (CAA) certified air-planes).

PITOT TUBE





7-3

Static Ports

The static ports are located onthe left and right sides of thefuselage. The upper right andlower left static port are inter-connected and supply ambi-

ent pressure to the pilot’sairspeed/mach indicator, IVSI,

the air data computer, andairspeed/mach warningswitch. The upper left andlower right static port areinterconnected and supplyambient pressure to the copilot’s airspeed/mach indicator,barometric altimeter, IVSI, the landing gear warning airspeedswitch (550-0627 and after), B&D TAS pressure transducer (if

installed), and airspeed/mach warning switch (CAA certificatedairplanes). Ambient pressure is also provided to the cabindifferential pressure indicator through the copilot’s static ports.The interconnection and location of the static ports on oppositesides of the fuselage minimizes system pressure errors causedby uncoordinated flight.

The pitot tubes and static ports are protected against icing byintegral, electrically-powered heating elements. Refer to Chap-

ter 10 for a complete description of pitot-static ice protection.

Note: The pitot tubes and static ports must be clear and free ofobstructions for proper operation.

Alternate Static Air Source (if installed)

Should restriction of the static ports occur, as evidenced byerratic indication of the pitot-static flight instruments, the

copilot’s instruments may be provided with ambient pressurefrom within the nose section through an alternate static airsource. A manually-operated control valve, located below thecopilot’s instrument panel, is used to select the normal or alter-nate static air source. Though optional on most Citations, thealternate static air source system is standard on French(DGAC) certificated airplanes.

STATIC PORTS




Airspeed/Mach Indicators

The airspeed/mach indicatorsprovide visual indication of theairplane’s speed in relation tothe ambient air and the speed

of sound. Each unit consists ofa stationary indicator dial,

airspeed pointer, rotatingmach sub-dial, moveableindex marker, and an airtightinstrument case which housesa pair of airtight diaphragmslinked to drive mechanismsthat operate the pointer andsub-dial. The instrument case

is supplied with static (ambi-ent) air pressure through itsassociated static ports. The airspeed diaphragm is suppliedwith pitot (ram) pressure through its associated pitot tube whilethe mach diaphragm is sealed at standard sea level atmo-spheric pressure (29.92 inHg/1013.2 mb (reference pressure)).

Airspeed Pointer

As airspeed increases or decreases, the differential between

static pressure and pitot pressure causes the airspeed dia-phragm to expand or contract. As it does, its movement is

transmitted by the drive mechanism to position the pointer atthe corresponding KIAS value on the indicator dial.

Mach Sub-Dial

The inner (rotating) mach sub-dial is visible through a windowon the face of the instrument between 140 and 320 knots on the

outer (stationary) KIAS indicator dial. Unlike the airspeedpointer which is positioned relative to speed, the mach sub-dialis positioned relative to altitude. As altitude increases or de-creases, the differential between static pressure and referencepressure causes the mach diaphragm to expand or contract.As it does, its movement is transmitted by the drive mechanismto rotate the sub-dial such that the relationship between itsposition and that of the pointer will correspond to the approxi-

mate mach number (M).

1. Airspeed Pointer

2. Mach Sub-Dial

3. Index Marker Knob

3

2 1





7-5

Airspeed/Mach Indicator Markings

The airspeed indicator dial is calibrated in knots per hour(KIAS) and incorporates two radial slots (9500 LB ZFW) or oneradial slot (11,000 LB ZFW) in positions corresponding to theairplane’s Maximum Operating Speed (VMO) limitation(s) listedbelow. Rotation of the sub-dial, occurring with changes in

altitude, causes the color red to become visible through theslot(s) within the appropriate altitude range.

Zero Fuel Weight Altitude Range VMO

9500 LB SL ~ 14,000' 262 KIAS

9500 LB 14,000' ~ 28,000' 277 KIAS

11,000 LB SL ~ 30,500' 262 KIAS

A red radial line on the mach sub-dial denotes the airplane’sMaximum Mach Operating (MMO) limitation of 0.705 M at alti-

tudes above 28,000 feet (9500 LB ZFW) or 30,500 feet (11,000LB ZFW). Essentially, when operating above these altitudes,alignment of the airspeed pointer with the MMO radial line willproduce a true airspeed of 0.705 M regardless of indicated

speed. To avoid exceeding this limitation, the airspeed pointermust never be permitted to rotate beyond the MMO radial line.

Index Marker

The index marker is controlled by a knob located on the lowerleft corner of the instrument face and may be moved to anyposition around the KIAS scale to reference a desired airspeed.




Overspeed Warning System

The overspeed warning system consists primarily of an air-speed/mach warning switch located on the forward pressurebulkhead, and a warning horn located in the overhead flightcompartment outboard of the pilot’s station. The warning switch

senses airspeed and altitude via the pilot’s pitot-static system,or copilot’s pitot-static system on Canadian (CAA) certified

airplanes, and causes the warning horn to sound when VMO /

MMO is reached or exceeded.

On 9500 LB ZFW airplanes, the warning switch integrates twoairspeed switches, an altitude switch, and a mach numberswitch. One airspeed switch (S1) operates in conjunction withthe altitude switch (S2) to sound the horn when airspeedreaches 262 KIAS at altitudes below 14,000 feet. The other

airspeed switch (S3) and the mach number switch (S4) functionto sound the horn at 277 KIAS and 0.705 M respectively, re-gardless of altitude.

On 11,000 LB ZFW airplanes, the warning switch utilizes air-speed switch S1 and the mach number switch S4 to sound thehorn at 262 KIAS and 0.705 M respectively, regardless of alti-tude (switches S2 and S3 are not required to be operational inthis installation).

28 VDC power is normally

supplied to the warning hornby the right main bus throughthe 2-amp OVERSPEED circuitbreaker. Operation of theoverspeed warning systemmay be verified by rotating the

TEST selector switch, locatedon the lower left instrumentpanel, to the “OVERSPEED”position. In this position,power supplied through the 2-amp WARN LTS 1 circuitbreaker will cause the horn to sound if the system is functional.






7-7

True Airspeed (TAS) System (optional)

True airspeed systems incorporate equipment which sensesand measures pitot-static pressures and ambient air tempera-ture to compute a correction for compressibility and ram rise.Once computed, a signal corresponding to true airspeed is

supplied to the TAS indicator and/or flight guidance systemswhich utilize this data in their operation. One of two systems

manufactured by Sperry or B&D may be installed.

Sperry TAS System

Components of the Sperry TAS system include the air datacomputer (ADC), located in the nose avionics bay; a tempera-ture probe, located on the lower right surface of the nose sec-tion; and an indicator, located on the center instrument panel.

AC power is supplied to the ADC, temperature probe, andindicator through the 2-amp AIR DATA circuit breaker. The ADCis supplied by the number one 115 VAC bus (split bus ACconfiguration), or by the 115 VAC bus (single bus AC configura-tion) and incorporates a transformer which steps 115 VACdown to 26 VAC to supply the temperature probe and indicator.

The temperature probe is protected against icing by an inte-gral, electrically-powered heating element controlled by the

PITOT & STATIC switch on the lower left instrument panel. Theheating element is supplied with 28 VDC power through the 15-

amp TAS HTR circuit breaker.




B&D TAS System

Components of the B&D system include a true airspeed (TAS)computer, located below the flight compartment floor; a tem-perature probe, located on the lower right surface of the nosesection; and a pressure transducer located on the forward sideof the lower right instrument panel. The TAS computer is in-

stalled in addition to, and functions independently of, the ADCaddressed in the Sperry TAS system description.

The TAS computer processes signals supplied by the copilot’spitot-static system and the temperature probe. Like the Sperrysystem, the temperature sensed by the probe is corrected forMach effect to obtain a measurement of static air temperature(SAT) which is computed with pitot-static pressures to producean electrical signal proportional to true airspeed (TAS). Unlikethe Sperry system, however, this signal is supplied only to the

flight guidance systems which require TAS data for their opera-tion. The B&D system does not feature a digital indicator ortemperature probe ice protection.

Should malfunction of thissystem occur, as evidencedby erratic indication of thecopilot’s pitot-static flightinstruments, the copilot’s pitot-

static system may be isolatedfrom the TAS computer by

closing a pair of valves lo-cated on the lower right instru-ment panel. The valves arelabeled TAS COMP - STATIC -PITOT and protected by redguard covers which must be

lifted when actuation is required. In the normal “OPEN” position,TAS signals are supplied to the flight guidance systems requir-ing this data for their operation. In the “CLOSED” position, TASdata is no longer supplied by the computer, but may still besupplied to flight guidance equipment which provides formanual entry of this data.

The system is supplied with 28 VDC power through the 2-amp

TAS circuit breaker. When closing the TAS COMP valves isrequired, this circuit breaker should be pulled to ensure thatflight guidance equipment does not receive erroneous airspeeddata.

TAS COMPUTER PITOT-STATIC VALVES





7-9

Barometric Altimeters

The barometric altimeters provide visual indication of airplanealtitude in relation to mean sea level (MSL) atmospheric pres-sure. Each unit incorporates an indicating pointer, drum-typecounter, and barometric scale. The pointer dial is calibrated inincrements of 20 and 100 feet, while the counter displays alti-

tude in increments of 20, 100, 1000, and 10,000 feet. A blackand white crosshatched area appearing in the left most digit of

the counter signifies that indicated altitude is below 10,000 feet.The barometric scale, calibrated in inches of mercury (inHg)and millibars (mb), displays the current correction setting fornon-standard atmospheric conditions.

At flight altitudes below 18,000 feet, the barometric scale mustbe set to the current altimeter setting provided by en routereporting stations within 100 nautical miles of the airplane’s

position. This setting reflects a computed correction, for non-standard conditions, of the barometric pressure measured nearground level in the vicinity of the reporting station and providesindication of true altitude above mean sea level (MSL). At flightaltitudes 18,000 feet (FL 180) and above, the barometric scalemust be set to standard sea level pressure (29.92 inHg/1013.2mb). This setting provides indication of pressure altitude abovethe standard datum plane, a theoretical level where atmo-spheric pressure is equal to standard sea level pressure. Since

the assigned altitudes of all aircraft operating at FL 180 andabove are referenced to this setting, collision avoidance and

vertical separation are assured.

To ensure adequate separation from aircraft operating below18,000 feet and compliance with minimum altitude rules, localreported pressure must be monitored and cruise altitudes in thevicinity of FL 180 proportionally increased by 500 feet for every

half inch that current altimeter setting falls below 29.92 inHg/ 1013.2 mb as specified in FAR 91 and the table below:

Altimeter Lowest Usable Adjustment

Setting Flight Level Factor

29.92 or higher FL 180 0'

29.91 ~ 29.42 FL 185 500'

29.41 ~ 28.92 FL 190 1000'

28.91 ~ 28.42 FL 195 1500'

28.41 ~ 27.92 FL 200 2000'

27.91 ~ 27.42 FL 205 2500'

27.41 ~ 26.92 FL 210 3000'




Pilot’s Altimeter

The pilot’s altimeter is a servo-type indicator which is electri-cally-driven by the air datacomputer (ADC) located in thenose avionics bay. The ADC

processes pneumatic signalssupplied by the pilot’s pitot-

static system and computesan electrical signal propor-tional to the airplane’s altitude.This signal is transmitted tothe altimeter drive mechanism such that the correspondingaltitude is displayed by the pointer and counter. Correction fornon-standard atmospheric conditions is accomplished byrotating the barometric setting knob, located on the lower left

instrument bezel, until the desired pressure appears on theinHg or mb scale. This setting transmits an electrical signal tothe ADC which, in turn, transmits a barometrically-correctedsignal to the altimeter drive mechanism. Should an error be-tween these signals occur, a failure warning flag will extendacross the counter indicating that altitude readout is not reli-able.

AC power is supplied to the ADC and the pilot’s altimeter

through the 2-amp AIR DATA circuit breaker. The ADC is sup-plied directly by the 115 VAC bus (single bus AC configuration)

or the number one 115 VAC bus (split bus AC configuration)and incorporates a transformer which steps 115 VAC down to26 VAC to power the altimeter. The failure warning flag will alsoextend across the counter whenever electrical power is re-moved from the instrument.

An optional radio/barometric (RAD/BAR) altimeter may beinstalled in place of the standard pilot’s altimeter. Operation andsetting of the barometric portion of the instrument are consis-tent with the standard altimeter. Operation of the radio altimeterportion of the instrument is addressed in the description ofradio altimeter systems.





7-11

Copilot’s Altimeter

The copilot’s conventionalaltimeter consists of an airtightinstrument case which housesan aneroid diaphragm assem-bly linked to a drive mecha-

nism that operates the pointerand counter. The instrument

case is supplied with static(ambient) air pressure throughthe static ports, while thediaphragm assembly is sealedat standard sea level atmospheric pressure (29.92 inHg/1013.2mb (reference pressure). As altitude increases or decreases,the differential between static pressure and reference pressurecauses the diaphragm assembly to expand or contract. As it

does, its movement is mechanically transmitted by the drivemechanism such that the corresponding altitude is displayedby the pointer and counter.

Correction for non-standard atmospheric conditions is accom-plished by rotating the barometric setting knob, located on thelower left instrument bezel, until the desired pressure appearson the inHg or mb scale. This setting rotates the indicator drivemechanism to produce the necessary altimeter correction.

The copilot’s altimeter incorporates a vibrator which functions to

optimize indicator response. 28 VDC power is supplied to thevibrator by the right main bus through the 2-amp RH ALT circuitbreaker.

When the airplane is configured for dual-altitude reportingcapability, the standard copilot’s altimeter is replaced by an

optional encoding altimeter. Operation and setting of the baro-metric portion of the instrument are consistent with the standardaltimeter. Operation of the encoder portion of the instrument isaddressed in the description of altitude alerting and reportingsystems.




Altitude Alerting and Reporting System

The altitude alerting and reporting system incorporates thosecomponents which provide visual and aural alerting ofpreselected reference altitudes, and those which supply en-coded altitude data to the transponder(s).

Altitude Alerting

Altitude alerting components include the air data computer(ADC), located in the nose avionics bay; vertical navigationcomputer/controller (VNCC), located on the center instrumentpanel; altitude alert light(s) and a warning horn. One altitudealert light is located on the upper right instrument bezel of thestandard pilot’s altimeter. Others are located on the upper leftbezel of the optional pilot’s RAD/BAR altimeter and optionalcopilot’s encoding altimeter when these instruments are in-

stalled. The altitude alert warning horn is located within the leftflight compartment sidewall.

The VNCC references altitudesignals transmitted to thepilot’s altimeter from the ADC,as previously described, andprovides data input and outputfor altitude alerting and vari-

ous vertical navigation (VNAV)modes. Mode selection and

data input are controlled by arotary switch and concentricsetting (SET) knob, and an-nunciated by a three-digitincandescent display. A dim-mer (DIM) knob permits adjustment of display intensity. When

“ALT” (altitude mode) is selected, rotating the SET knob slewsthe display in increments of 100 feet between 000(00) and500(00) feet to the desired altitude preselect. Slew rate is pro-portional to the speed at which the SET knob is rotated.

VNAV COMPUTER/CONTROLLER





7-13

As the airplane climbs or descends to within 1000 feet of thepreselected altitude, the warning horn will sound for one sec-ond and the altitude alert light(s) will illuminate and remainilluminated until the airplane is within 250 feet of thepreselected altitude. Should the airplane deviate from thepreselected altitude by 250 feet or more, the warning horn will

sound for one second and the altitude alert light(s) will illumi-nate and remain illuminated until the airplane returns to within

250 feet of the preselected altitude or until a new altitudepreselect is set.

AC power is supplied to the VNCC by the number one 115 VACbus (split bus AC configuration) or the 115 VAC bus (single busAC configuration) through the 1-amp FD 1 circuit breaker. 28VDC power is supplied to the warning horn and altitude alertlight(s) by the corresponding DC AVIONICS bus through the 3-

amp FD 1 circuit breaker.

Altitude Reporting

Altitude reporting components include the air data computer(ADC), transponder(s), and the optional copilot’s encodingaltimeter when installed. These components function to gener-ate a logic code corresponding to the airplane’s pressure alti-tude which is transmitted to the Air Traffic Control Radar Bea-con System (ATCRBS) through the transponder in response to

Mode C (altitude reporting) interrogations. Barometric pressuresetting of the altimeter(s) has no effect on this function since

encoder output is always referenced to standard sea levelatmospheric pressure (29.92 inHg/1013.2 mb).

In standard configuration, a single transponder (transponder 1)receives encoded altitude signals from the ADC. One optionalconfiguration adds a second transponder (transponder 2)

which also receives encoded altitude signals from the ADC. Inthis installation, either transponder may be assigned to altitudereporting by means of a select switch, co-located with thetransponder controls, while the other remains in standby condi-tion. Another option configures the system for dual-altitudereporting. In this installation, the selected transponder mayreceive encoded altitude signals from the ADC (primary source)or the optional copilot’s encoding altimeter (secondary source)

by means of a combination annunciator/switch (550-0258 andafter) or crossover switch and relay (550-0257 and earlier).





7-15

Vertical Speed Indicators

The vertical speed indicatorsare instantaneous type (IVSI)and provide visual indicationof the airplane’s rate of climb

or descent in feet per minute(FT/MIN X 1000). Each unit

consists of an indicator dial, apointer, and an airtight instru-ment case which houses anairtight diaphragm and accel-erometer linked to a drivemechanism that operates thepointer. Static (ambient) air pressure from the static ports issupplied to the diaphragm directly, and to the instrument case

through a calibrated restrictor orifice. As the airplane climbs ordescends, the static air pressure supplied to the instrumentdecreases or increases accordingly. Due to the restrictor ori-fice, however, the rate of pressure change within the instrumentcase occurs more slowly than within the diaphragm. This pro-duces a pressure differential which causes the diaphragm toexpand or contract in proportion to the rate of altitude change.As it does, its movement is transmitted by the drive mechanismto position the pointer at the corresponding value on the indica-

tor dial.

During a climb, the pressure within the diaphragm decreases ata faster rate than the pressure within the instrument case. Theresulting differential causes the diaphragm to contract and thepointer to indicate a rate of climb. During descent, the pressurewithin the diaphragm increases at a faster rate than the pres-sure within the instrument case. The resulting differential

causes the diaphragm to expand and the pointer to indicate arate of descent. As the airplane resumes level flight, the pres-sure within the instrument case and the diaphragm becomeequalized and pointer indication returns to zero.

The time required to stabilize the pressure differential whichcauses pointer deflection can result in a delay of up to nineseconds before vertical speed indication becomes reliable. To

compensate for this, the accelerometer’s sensitivity to verticalG-loading provides instantaneous indication of vertical speedby displacing the pointer prior to the instrument’s response tochanges in pressure. During level flight or steady rates of climbor descent, the IVSI function as a conventional VSI.




Radio Altimeter System (optional)

The radio altimeter system measures absolute altitude aboveground level (AGL) and supplies this data to flight guidancesystems, ground proximity warning systems (GPWS), and/orindicators when installed. Though models and manufacturers of

these systems differ, all share a basic configuration whichincludes an indicator, a transceiver, and corresponding transmit

and receive antennas. Operation is continuous when electricalpower is supplied to the system, however, radio altitude readoutis limited by the indicator’s usable range.

Transceiver and Antennas

The transceiver incorporates solid-state circuitry which makesinstantaneous comparisons between the frequency of a fre-quency-modulated microwave signal that is beamed down from

the transmit antenna to a return signal that is reflected back tothe receive antenna from the terrain. Because the differencebetween these frequencies is proportional to the transmitsignal’s “round-trip” time to the terrain and back, the frequencydifference is processed to generate an electrical signal propor-tional to absolute altitude. Once generated, this signal is sup-plied to the indicator(s) and/or flight guidance systems whichutilize radio altitude data in their operation. Depending oninstallation, the transceiver may be located within the right flight

compartment sidewall or below the cabin floor. The antennasare located on the lower surface of the fuselage.





7-17

Conventional Indicator

(excluding RAD/BAR)

The indicator may be locatedon the pilot’s or copilot’s instru-ment panel and incorporatesan altitude scale and pointer,

decision height (DH) indexmarker and setting knob, DH

alert light, and a test switch.The altitude scale is calibratedin 10 foot increments below500 feet, 100 foot incrementsabove 500 feet, and identifies the indicator’s usable range. Thepointer is visible at the corresponding AGL value within theusable range only. Outside this range, or if the airplane issteeply banked, the pointer will be positioned behind a mask in

the upper left corner of the indicator dial.

The DH setting knob, located on the lower right instrumentbezel, is rotated to preselect a desired decision height. The DHalert light, located on the upper right instrument bezel, illumi-nates in conjunction with the sounding of an alert tone whendecision height is reached or when the system is tested. Thetest switch is located on the lower left instrument bezel. The DHalert horn is located in the overhead flight compartment out-

board of the pilot’s station.

28 VDC power is supplied to the system by the emergencyavionics bus through the 5-amp RAD ALT circuit breaker. Afailure warning flag will appear across the upper instrumentface whenever electrical power is removed from the instrumentor when radio altitude indication becomes invalid. Refer to theappropriate Operating Manual or Airplane Flight Manual for test

procedures and information regarding specific capabilities ofthe system installed in your airplane.




Conventional Indicator (RAD/BAR)

The RAD/BAR indicator is located on the pilot’s instrumentpanel as previously described and incorporates a four-digitincandescent radio altitude display, decision height (DH) set-ting knob and alert light, and a test switch. The radio altitudedisplay functions independently of the barometric portion of the

instrument. During initial climb-out, radio altitude is displayed in10 foot increments between 0 and 990 feet AGL. During de-

scent, radio altitude is displayed in 100 foot increments be-tween 2500 and 1000 feet AGL, and 10 foot increments be-tween 1000 and 0 feet AGL. Above 2500 feet AGL the displayis blank. The DH setting knob, located on the upper right instru-ment bezel, is rotated to preselect a desired decision height.The ALT alert light, located on the upper left instrument bezel,illuminates in conjunction with the sounding of an alert tone orGPWS vocal alert when decision height is reached or when the

system is tested. The test switch is located on the lower leftinstrument bezel. The ALT alert light, also functions as a baro-metric altitude alert light by way of the VNCC as previouslydescribed.

The RAD/BAR system also incorporates a converter, located inthe nose avionics bay, which processes radio altitude signalssupplied by the transceiver and glideslope signals supplied bythe flight guidance system to support GPWS vocal alert func-

tions. 28 VDC power is supplied to the system by the emer-gency avionics bus through the 5-amp RAD ALT circuit breaker.

Should radio altitude data become invalid, a failure warning flagwill appear adjacent to the RAD/ALT display. The failure warn-ing flag in the upper center of the instrument face appears onlywhen power is removed from the barometric portion of theinstrument. Refer to the appropriate Operating Manual or Air-plane Flight Manual for test procedures and information regard-

ing specific capabilities of the system installed in your airplane.





7-19

Mechanical Flight Director Radio Altitude Indication

Radio altitude and decision height may also be displayed onthe ADI(s) in mechanical flight director installations. Thoughconfigurations vary, reference to absolute altitude is generallyprovided by a bar with chevron markings or a rising runwaysymbol which appears in the lower center of the attitude sphere

as the airplane descends below 200 feet AGL and movestoward the airplane symbol in relation to ground proximity until

contact between the two occurs at the point of touchdown.

Decision height is generally annunciated by a DH alert light,located on the upper right instrument bezel, which illuminates inconjunction with the sounding of an alert tone when decisionheight is reached or when the radio altitude system is tested.With the exception of the Sperry AD-650, the DH alert lightdoes not illuminate when the ADI is tested.

In addition to the featuresdescribed above, radio alti-tude and decision height arealso digitally displayed on theSperry AD-650 (pictured)when installed. The four-digitincandescent RAD ALT dis-play, located on the lower right

instrument face, is calibratedin 5 foot increments between 0

and 200 feet AGL, and 10 footincrements between 200 and2500 feet AGL. Above 2500feet AGL the RAD ALT displayis blank. Reference to groundproximity is provided by a

rising runway symbol as previ-ously described. Should radio altitude data become invalid,four dashes will appear in the RAD ALT display.

The three-digit decision height (DH) display, located in thelower left corner of the ADI, is calibrated in 10 foot incrementsbetween 0 and 990 feet AGL. The inner decision height (DHSET) setting knob, located on the lower right instrument bezel,

is rotated to preselect a desired decision height between 0 and990 feet AGL. The surrounding DIM ring permits adjustment ofRAD ALT and DH display intensity. Should decision height databecome invalid, three dashes will appear in the DH display.

1. RAD ALT Display

2. DH Display

3. DH SET/DIM Knob

4. RA Test Button

12

4 3




In dual AD-650 flight director installations, the alert tone isactivated through the DH setting of the pilot’s ADI only. The DHsetting of the copilot’s ADI and/or conventional radio altitudeindicator, if installed, have no effect on alert tone activation.Refer to the appropriate Operating Manual or Airplane FlightManual for test procedures and information regarding specific

capabilities of the system installed in your airplane.

EFIS Radio Altitude Indication

Radio altitude and decision height may also be digitally dis-played on the EADI(s) in EFIS installations. The four-digit radioaltitude (RA) display, located in the lower right corner of theEADI, is calibrated in 5 foot increments between 0 and 200 feetAGL, and 10 foot increments between 200 and 2500 feet AGL.Above 2500 feet AGL the RA display is blank. Additional refer-ence to absolute altitude is provided by a rising runway symbol

which appears in the lower center of the attitude sphere as theairplane descends below 200 feet AGL and moves toward theairplane symbol in relation to ground proximity until contactbetween the two occurs at the point of touchdown. Should radioaltitude data become invalid, four amber dashes will appear inthe RA display and the rising runway will not be visible in theattitude sphere.

The three-digit decision height (DH) display, located in the

lower left corner of the EADI, is calibrated in 5 foot incrementsbetween 0 and 200 feet AGL, and 10 foot increments between

200 and 990 feet AGL. The decision height (DH/TST) settingknob, located on the EFIS display controller, is rotated topreselect a desired decision height between 0 and 990 feetAGL. Rotating this knob fully counterclockwise removes the DHdisplay from the EADI. As the airplane descends to within 100feet of decision height, a white box will appear above and left of

the radio altitude (RA) display. An amber DH will appear withinthis box in conjunction with the sounding of an alert tone whendecision height is reached or when the EFIS system is tested.Should decision height data become invalid, three amberdashes will appear in the DH display.

In dual EFIS installations, the alert tone is activated through theDH setting of the pilot’s EADI only. The DH setting of the

copilot’s EADI and/or conventional radio altitude indicator, ifinstalled, have no effect on alert tone activation. Refer to theappropriate Operating Manual or Airplane Flight Manual for testprocedures and information regarding specific capabilities ofthe system installed in your airplane.





7-21

Attitude and Direction System

The attitude and direction system consists of all gyroscopicflight instruments (including mechanical and EFIS flight direc-tors), turn-and-bank indicators, the RMIs, and the magneticcompass.

Conventional Attitude Indicator

The attitude indicator provides visual indication of pitch and rollin relation to the actual horizon. The unit consists of a gyrostabilized in the horizontal plane, an attitude sphere, airplanesymbol, and roll index pointer. The gyro is air-driven on air-planes 550-0626 and earlier, or electrically-driven on airplanes550-0627 and after, and responds to pitch and roll movementsof the airplane. An electrically driven attitude indicator may alsobe installed on airplanes 550-0550 through 0626 as optional

equipment.

The attitude sphere is divided into sky and ground hemispheresby a horizon bar which provides visual reference to the actualhorizon. The airplane symbol is secured to the instrument faceand provides visual reference of the airplane’s attitude relativeto the horizon bar. The roll index pointer is located at the top ofthe instrument face and provides visual reference of theairplane’s bank angle relative to the actual horizon.

An inclinometer is installed on the lower instrument bezel of the

electrically-driven attitude indicator to provide visual indicationof turn coordination. The inclinometer is comprised of a ballcontained in a sealed, silicone liquid filled, glass tube andresponds to gravitational and centrifugal forces acting on theairplane. The tube is curved and mounted such that the ball willrest in the center lowest position when the airplane is in coordi-

nated flight. In uncoordinated flight, the ball will move from thecenter to the outside of a turn (indicating a skid) or the inside ofa turn (indicating a slip).




Air-Driven Gyro

On airplanes 550-0626 and earlier, the gyro is air-driven byregulated engine bleed air. Bleed air used by the system istapped from the supply tubes between the compressor sectionof each engine and the environmental flow control/shutoffvalves within the fuselage tailcone. On airplanes 550-0484,

0483, 0481 and earlier, the bleed air is routed through aprecooler within each engine nacelle prior to entering the fuse-

lage tailcone. On airplanes 550-0482, 0485 and after, availabil-ity of bleed air to the system is continuous when either or bothengines are operating and is not influenced by the position ofthe PRESS SOURCE selector. On airplanes 550-0484, 0483,0481 and earlier, however, setting the PRESS SOURCE selectorto “LH” or “RH” correspondingly results in bleed air being madeavailable to the system from the left engine or right engine only.

The bleed air tapped from each source is routed through inde-

pendent supply tubes to a common cross fitting within thefuselage tailcone. A check valve in each of these tubes pre-vents the backflow of bleed air to either engine when the oppo-site engine has failed or is operating at a sufficiently lower RPM.

From the cross fitting, bleed air is routed to the windshield anti-ice/rain removal system, cabin pressurization control systemejector, and pneumatic distribution pressure regulator. Fromthis regulator, 23.0 +/- 1.0 PSIG bleed air is routed to the pneu-matic surface deice system, cabin temperature manual control

system, inflatable cabin door seal, and the instrument air sys-

tem. The instrument air system consists of a water separator/ filter, instrument pressure regulators, the air-driven attitudeindicator, and gyro pressure gage.

The water separator/filter is installed on the forward pressurebulkhead within the nose section and functions to filter and

extract moisture from the bleed air prior to being routed to theinstrument pressure regulators. Extracted moisture collects inthe lower filter bowl and is eliminated through an orificed draineither to be vented overboard through an instrument air vent/ drain line which extends through the lower surface of the nosesection (550-0173 ~ 0626), or evaporated within the nosesection (550-0172 and earlier).





7-23

The instrument pressure regulators are also installed on theforward pressure bulkhead within the nose section and functionin series to reduce and stabilize bleed air pressure prior tosupplying the instruments. The “first-stage” regulator reducesthe pressure to approximately 6.0 PSIG; the “second-stage”regulator further reduces the pressure to approximately 2.5

PSIG. Regulated bleed air entering the attitude indicator case isdirected against “buckets” machined into the rim of the gyro

causing it to spin at a high rate of speed. After driving the gyro,bleed air is exhausted from the instrument case through theoverboard instrument air vent/drain line previously described.

Gyro Pressure Gage

On airplanes 550-0626 and earlier, the gyro pressure gageprovides visual indication of the bleed air pressure driving thegyro. The normal indicating range is denoted by a green arc

between 2.0 and 3.0 PSIG. Pressure indications which areerratic and/or outside the normal range indicate that a malfunc-tion may exist in the pneumatic system. From the gyro pressuregage, bleed air is also exhausted through the overboard instru-ment air vent/drain line.

Electrically-Driven Gyro

On airplanes 550-0627 and after, the gyro is electrically drivenby 28 VDC power through the 2-amp ADI 2 circuit breaker.

Power is supplied to the gyro from the emergency avionics buswhenever the battery switch is in the “BATT” (on) position and

the DC avionics power switch is also in the “ON” position.Power is also supplied to the gyro when the battery switch is inthe “EMER” (emergency) position, regardless of DC avionicspower switch position. A red GYRO failure warning flag will bevisible in the upper left instrument face when power is removedfrom the instrument. A spring-loaded caging knob is located on

the lower right instrument bezel.




Turn-and-Bank Indicator

In single mechanical flight director systems, a turn-and-bank in-dicator is installed on the pilot’s and copilot’s instrument panel. Indual mechanical flight director installations, a turn-and-bank indi-cator is typically installed on the copilot’s instrument panel only.In dual 5” mechanical flight director installations, an air-driven

“standby” attitude gyro replaces the copilot’s turn-and-bank indi-cator.

The turn pointer is attached to a DC electrically-driven gyro whichindicates the airplane’s turning rate in degrees per second. If in-stalled, the copilot’s turn indicator is vacuum-driven. Operation ofthe turn indicator can be checked by initiating a standard rateturn and cross checking the turn rate with the heading indicator.An indicated standard rate turn should show a turning rate of 3°per second on the heading indicator.

Visual indication of turn coordination is provided by an inclinom-

eter on the lower instrument face. The inclinometer is comprisedof a ball contained in a sealed, silicone liquid filled, glass tubeand responds to gravitational and centrifugal forces acting onthe airplane. The tube is curved and mounted such that the ballwill rest in the center lowest position when the airplane is in coor-dinated flight. In uncoordinated flight, the ball will move from thecenter to the outside of a turn (indicating a skid) or the inside of aturn (indicating a slip).





7-25

Mechanical Flight Directors

Mechanical flight director system configurations are generally clas-sified as 4” or 5”, single or dual flight director. In dual flight direc-tor systems, the pilot’s and copilot’s ADI and HSI may be identicaor one of various combinations of 4” or 5” instruments depending

on installation.

The combination of instruments in each installation varies suffi-

ciently between airplanes as to make specific descriptions of everypossible system configuration impractical. The following descrip-tion, therefore, primarily addresses the functional characteristicsand operational features which are common among the mechani-cal ADIs and HSIs, as well as any relevant technical differencesthat exist between them.

Attitude Directional IndicatorThe ADI or flight director indi-

cator (FDI) functions as a con-ventional attitude indicator anddisplays command informationprovided by the autopilot/flightdirector computer. Dependingon avionics equipment installedand modes available, the ADImay be utilized to intercept and

maintain a desired heading, al-

titude, VOR radial, or localizercourse and glideslope. Refer to the appropriate operating manuafor specific capabilities of the system installed in your airplane.

Horizontal Situation Indicator

The HSI functions essentially

as a slaved heading indicatorand (depending on avionicsequipment installed andmodes available) providesvisual indication of airplaneposition relative to VOR radi-als, RNAV courses, localizercourses, and glideslope

beams. Refer to the appropri-ate operating manual forspecific capabilities of the system installed in your airplane.




To function as a slaved heading indicator, the HSI incorporates

components which electromechanically produce a controlled pre-cession of the gyro which corrects the compass card to agreewith the correct magnetic heading. This installation eliminates theneed for periodic resetting of the gyro due to precession drift.

Gyro Slaving

Gyro slaving is provided by a remotely-mounted magnetic fluxdetector and slaving accessory. The flux detector senses the di-

rection of the earth’s magnetic field and transmits these signalsto the slaving accessory where they are compared with gyro ref-erence signals corresponding to the position of the compass card.The resulting error signal is amplified and transmitted to compo-nents which electromechanically produce a controlled preces-sion of the gyro that corrects indicated heading to agree with themagnetic heading.

The gyro slaving system is con-

trolled by switches located onthe lower left instrument paneland lower copilot’s instrumentpanel. A slaving meter, used tomonitor displacement errorswhich may exist between indi-cated heading and magneticheading, may also be installed.

Indicator needle deflection in

either direction from center cor-responds to the polarity of theerror. 1. Mode Selection Switch

2. Manual Slaving Switch

1 2

The gyro slaving switches permit selection of automatic slaving(slaved gyro) when positioned to “AUTO,” or manual slaving (freegyro) modes when positioned to “MAN.” In the slaved gyro mode,heading displacement errors are corrected automatically throughthe gyro slaving circuit. In the free gyro mode, heading displace-ment errors are corrected using the corresponding LH/RH switch

to rotate the compass card left or right to agree with the magneticheading, and return the slaving meter indicating needle to center(if installed). The rate of manual compass card rotation is ap-proximately 30° per minute. Refer to the appropriate operatingmanual for specific capabilities of the system installed in yourairplane.





7-27

ADI Display Features

The following summary describes the functional characteristicsand operational features which are common to mechanical ADIstypically installed in the Cessna Citation II. Refer to the appropri-ate operating manual for specific capabilities of the system in-stalled in your airplane.

Attitude Sphere

The attitude sphere is divided into sky (blue) and ground

(brown) hemispheres by a horizon line which provides visuareference to the actual horizon.

Eyelid Display

The eyelid display surrounds the attitude sphere and providesvisual reference of the relative position of the sky (blue) andground (brown), independent of attitude sphere position, to

facilitate recovery from unusual flight attitudes.

Airplane Symbol

The airplane symbol is located in the center of the instrumentface and provides visual reference of the airplane’s attituderelative to the horizon line. Depending on ADI installation, thesymbol may be conventional or reference delta type.

Pitch Attitude Indication

Pitch attitude is indicated by the relative position of the air-

plane symbol’s nose above or below the horizon line. A pitch

scale on the attitude sphere references deviation above orbelow the horizon line in increments of 5°.

Roll Attitude Indication

Roll attitude is indicated by the relative position of the airplanesymbol’s wings in relation to the horizon line. A roll attitudepointer and scale on the upper center instrument face refer-

ences deviation from wings level attitude. The scale is markedat 10, 20, 30, 45, 60, and 90° to the left and right of the triangu-lar 0° index mark at its center. To facilitate roll attitude recogni-tion, the 30 and 60° marks are longer and heavier, while each45° mark is displayed as a dot or a triangle.




Depending on ADI installation, the pointer may be a moveable

“sky pointer” or fixed “roll pointer.” In sky pointer configuration,the moveable pointer references the relative position of the skyin relation to the fixed scale index mark which references theairplane’s vertical axis. Roll recovery, therefore, is made in thedirection of the pointer. In roll pointer configuration, the fixed

pointer references the airplane’s vertical axis in relation to themoveable scale index mark which references the relative posi-tion of the sky. Roll recovery, therefore, is made in the direction

of the scale index mark.

Flight Director Command Bar(s)

Depending on ADI installation, the command bar(s) may bedouble or single-cue. In double-cue flight directors, computedcommands are displayed by independent pitch (horizontal) andsteering (vertical) command bars as a conventional VOR/ILS

indicator. Movement of the bars indicates pitch and steeringcontrol inputs required to satisfy computed commands of the

selected flight director operating mode. In single-cue flightdirectors, computed pitch and steering commands are dis-played by a single (delta) command bar. To satisfy computedpitch and steering commands of the selected flight directoroperating mode, the airplane symbol is “flown” to align withthe command bar. Should loss of pitch or steering commandsignals from the flight director computer occur, the single-cuecommand bar or the affected double-cue command bar will

retract from view. The non-affected double-cue command bar

will continue to display normally.

Glideslope Indication

Glideslope deviation is indicated by the relationship betweena moveable pointer and fixed vertical deviation scale commonlylocated on the right side of the instrument face. When a validglideslope signal is being received, deviation above the beam

centerline is indicated by displacement of the pointer belowscale center. Deviation below the beam centerline is indicatedby displacement of the pointer above scale center. Each dotabove and below scale center represents approximately 0.4°deviation from the beam centerline. A green area on the scaledenotes the category II (CAT II) approach window. Pointer dis-placement in this area indicates glideslope deviation withinapproximately 0.2° of the beam centerline.





7-29

Expanded Localizer Indication

Expanded localizer deviation is indicated by the relationshipbetween a moveable “runway” pointer and fixed lateral devia-tion scale located on the lower center instrument face. When avalid localizer signal is being received, deviation left of the beamcenterline is indicated by displacement of the pointer to the

right of scale center. Deviation right of the beam centerline isindicated by displacement of the pointer to the left of scalecenter. Amplification of localizer data from the navigation re-

ceiver permits increased pointer sensitivity within the CAT Iapproach window. Because this increase in sensitivity makestracking the localizer more difficult throughout the entire ap-proach, the expanded localizer pointer should be referencedfor position assessment only, until established on final approachDuring final approach within the CAT II window, pointer dis-placement within the lateral deviation scale indicates localizer

deviation within approximately 0.25° of the beam centerline, owithin 33 feet of the runway centerline.

Radio Altitude Indication

Refer to the description of radio altimeter systems.

Test Switch

The attitude (ATT) test switch, located on the lower left instru-ment bezel, initiates an attitude self-test function. Whenpressed, the attitude sphere should be positioned to indicate

approximately 20° right bank and 10° positive pitch, and the

ATT warning flag should appear across the instrument faceRefer to the appropriate Operating Manual or Airplane FlightManual for test procedures specific to the system installed inyour airplane.

GA Light

The go-around light illuminates when the go-around mode has

been selected

DH Light

The decision height light illuminates when the airplane de-scends below the selected decision height as set on the radioaltitude indicator.

Mode Annunciators (5”)

Ten/twelve annunciators indicate which vertical and horizontamodes are engaged with the flight director.




Turn and Bank Indication

Rate-of-turn is indicated by the relationship between a move-able pointer and fixed scale located on the lower center instru-ment face of 5" ADIs. Pointer displacement over the left or rightscale markings indicates a standard rate turn (2-minute/3° persecond) in that direction.

A conventional inclinometer is installed on the lower instrumentbezel of ADIs to provide visual indication of turn coordination.

The inclinometer is comprised of a ball contained in a sealed,silicone liquid filled, glass tube and responds to gravitationaland centrifugal forces acting on the airplane. The tube is curvedand mounted such that the ball will rest in the center lowestposition when the airplane is in coordinated flight. In uncoordi-nated flight, the ball will move from the center to the outside ofa turn (indicating a skid) or the inside of a turn (indicating a

slip).

Angle-of-Attack Indication (5” ADIs)

Angle-of-attack (AOA) is indicated on 5” ADIs by the relation-ship between a moveable “speed command” pointer and fixedFAST/SLOW scale commonly located on the left side of theinstrument face. The circular speed command pointer is posi-tioned by signals received from the AOA transmitter. The pointeris calibrated such that its position relative to the scale corre-sponds with the AOA indicator on the upper left instrument

panel. The FAST and SLOW scale markings correspond to the

.4 and .8 AOA indicator markings respectively. The scale cen-ter marking corresponds to the .6 AOA indicator marking andrepresents the optimum landing approach speed (1.3 timesstalling speed) for the current airplane configuration. Refer tothe description of AOA systems for complete detail.

Failure Warning Flags

Depending on ADI installation, various failure warning flagswill appear across the instrument face whenever electricalpower is removed from the instrument or when the correspond-ing indication becomes invalid. Typically, 4” ADIs provide warn-ing flags for attitude (ATT) and flight director (FD) indicationfailure only, while most 5” ADIs also provide warning flags forlocalizer (LOC), glideslope (GS), rate-of-turn (RT), and angle-of-attack (SPD) indication failure.





7-31

HSI Display Features

The following summary describes the functional characteristicsand operational features which are common to mechanical HSIstypically installed in the Cessna Citation II. Refer to the appropri-ate operating manual for specific capabilities of the system in-stalled in your airplane.

Compass Card and Heading Index

A 360° rotating white compass card indicates airplane head-

ing when referenced to the white triangular heading index inthe upper center of the display. The compass scale is dividedinto 5° increments with 10° division markings being twice aslong as the 5° markings. Fixed 45° index markings are posi-tioned adjacent to the scale. Compass heading is referencedto magnetic north.

Airplane SymbolThe airplane symbol, located in the center of the instrument

face, provides visual indication of the airplane’s position in re-lation to the course deviation bar. Alignment of the airplanesymbol with the course deviation bar simulates alignment ofthe airplane’s flight path to the centerline of the selected navi-gation course or localizer.

Course Deviation Indication

Lateral deviation from the centerline of a selected navigation

course or localizer is indicated by the relationship between a

moveable bar and fixed deviation scale located in the centerof the instrument. The deviation scale consists of two filled whitecircles evenly spaced on each side of the airplane symbol.The outer circles reference full scale deviation while the innercircles reference half scale deviation. Alignment of the coursedeviation bar with the airplane symbol represents alignment ofthe airplane’s flight path with the centerline of the selected navi-

gation course or localizer.

Heading Bug

An orange heading bug is manually rotated about the com-pass card by the heading (HDG) select knob on the lower rightcorner of the instrument bezel. Once set, the heading bug ro-tates with the compass card. The heading bug functions toindicate desired heading and provides selected heading ref-

erence for autopilot steering.




Course Pointer

The course pointer is manually rotated about the compass cardby the course (CRS) select knob on the lower left corner of theinstrument bezel. Once set, the course pointer rotates with thecompass card and provides indication of the desired naviga-tion course to be flown. Depending on HSI installation, the se-

lected course may be indicated in the upper left corner of thedisplay.

To/From Indicator

A filled white triangle, pointing either toward the head (to) orthe tail (from) of the course pointer, functions to provide visualindication that the selected course will take the airplane to orfrom the selected navaid or waypoint. The to/from indicator isnot displayed during ILS operation or when an invalid navaidor waypoint signal is received.

Distance to Station (DME) Indication

Depending on HSI installation, DME information in nautical milesmay be indicated in the upper right corner of the display.

Glideslope Indication

Glideslope deviation is indicated by the relationship betweena moveable pointer and fixed vertical deviation scale commonlylocated on the right side of the instrument face. The deviationscale consists of two filled white circles evenly spaced above

and below a filled white diamond. The outer circles reference

full-scale deviation while the inner circles reference half-scaledeviation. Alignment of the green pointer with the center dia-mond represents alignment of the airplane’s glide path withthe glideslope centerline.

Failure Warning Flags

Depending on HSI installation, various failure warning flags will

appear across the instrument face whenever electrical poweris removed from the instrument or when the corresponding in-dication becomes invalid. Typically, warning flags are providedfor navigation (NAV) and heading (HDG) indication failure, andvertical gyro (VERT) failure.




Chapter 8Hydraulics and Landing Gear

Table of Contents

Overview ...............................................8-1

Landing Gear System Description........... 8-1

Nose Gear ............................................8-2

Main Gear ............................................8-6

Hydraulic Actuators...........................8-8

Uplock Sequence Actuators ..............8-9

Left Main Gear Safety Switch .............. 8-10

Landing Gear Control and

Position Indicator ................................ 8-12Landing Gear Control Handle .......... 8-12

Landing Gear Warning Horn ................. 8-15

Landing Gear Indicator Lightand Warning Horn Test .................. 8-16

Landing Gear Hydraulics ...................... 8-17

Landing Gear HydraulicPressure Source ............................8-17

Hydraulic PowerSystem Components ...................... 8-17

Hydraulic Fluid Reservoir.................8-17Engine-Driven Hydraulic Pumps ........8-21

Hydraulic Filters .............................8-21

Hydraulic Firewall Shutoff Valves ..... 8-22

Hydraulic System Indication ................. 8-23

Hydraulic Flow Annunciators ............8-23

Hydraulic Pressure Annunciator ....... 8-24

Landing Gear Control Valve.............8-25




Hydraulics and Landing Gear, continued

Table of Contents

Landing Gear Operation .......................8-26

Landing Gear Retraction ................. 8-26

Landing Gear Extension ..................8-28

Static (Open Center) Condition........8-28

Landing Gear Extension/RetractionSpeeds and Cycle Times ................8-28

Auxiliary/Emergency Landing Gear

Extension System ...............................8-30

Auxiliary/Emergency AirStorage Bottle ...............................8-30

Auxiliary/Emergency Gear ExtensionSystem Operation ..........................8-31

Emergency Hydraulic Dump Valve ....8-33

Brakes ...............................................8-34

Power Brake System ...................... 8-34

Touchdown Protection .....................8-37

Anti-Skid System ............................ 8-38

Anti-Skid System Test .................... 8-39

Locked Wheel CrossoverProtection ......................................8-40

Auxiliary/Emergency Braking System..............................8-40

Auxiliary/Emergency Braking System Usage ................... 8-41

Parking Brake System .................... 8-41

Limitations..........................................8-42




HYDRAULICS AND LANDING GEAR 12/99 FOR TRAINING PURPOSES ONLY


8-1

Overview

The Cessna Citation II is equipped with retractable tricyclelanding gear that is electrically-controlled and hydraulically-operated. The main gear assemblies are located in each wing.The nose gear assembly is located in the nose section of the

fuselage. Normal gear extension and retraction is activated bythe landing gear control handle. Emergency gear extension is

provided by a mechanical uplock release “free-fall” system anda pneumatic “blowdown” system. The airplane has an indepen-dent hydraulic system for the main gear wheel brakes. Emer-gency braking, anti-skid and other braking systems are pro-vided.

Landing Gear System Description

Normal extension and retraction is accomplished by directinghydraulic fluid under engine-driven pump pressure to an actua-tor at each gear assembly. All three gear assemblies are heldin the extended position by mechanical downlock latchesinternal to each actuator, and held in the retracted position byuplock hooks. Hydraulic pressure releases the downlock

latches during gear retraction, and uplock hooks during gearextension.

Six microswitches are incorporated; three actuated by theuplock hooks and three actuated by the downlock latches.

These microswitches operate in conjunction with the landinggear control, position indication, and warning circuitry.




The nose gear is of conventional design incorporating a lowerpiston/fork assembly, upper trunnion assembly, shock strutassembly, torque links, and a hydraulic shimmy damper. Thepiston/fork assembly provides attachment points and runningclearance for the nosewheel. Attachment points for installationof the nose gear to the nose wheel well structure are providedby the trunnion assembly which also houses the shock strut

assembly. Upper and lower torque links connect the piston/forkassembly to the trunnion assembly and maintain alignment of

the nosewheel.

Nose Gear





8-3

The nose gear incorporates an air-oil type shock strut assemblywhich functions to support the weight of the airplane and ab-sorb shock loads encountered during ground operation. Theshock strut assembly consists primarily of an inner cylinder re-tained within a hydraulic fluid filled outer cylinder. The inner cyl-inder is integral to the piston/fork assembly while the outer cyl-

inder is housed within the trunnion assembly. A floating isola-tion piston divides the interior of the inner cylinder into a hy-

draulic fluid filled upper chamber and nitrogen gas or dry aircharged lower chamber. A fixed orifice separates the upperchamber of the inner cylinder from the outer cylinder. A taperedmetering pin regulates the flow of hydraulic fluid through thisorifice in relation to increasing or decreasing load.

Under increasing load, hydraulic fluid flows from the outer cylin-der to the upper chamber of the inner cylinder. As the strut

compresses, the metering pin progressively restricts the orifice.When the rate of hydraulic fluid flow through the orifice is insuf-ficient to absorb compression shocks, the isolation piston isforced downward against gas/air pressure to assume the addi-tional load. Under decreasing load, the isolation piston isforced upward by gas/air pressure and hydraulic fluid flowsfrom the upper chamber of the inner cylinder to the outer cylin-der until the pressure on each side of the orifice is equal.

The shimmy damper consistsof a hydraulic fluid filled outer

cylinder assembly and aninternal piston assembly,attached to the nose gearsuch that the piston effectivelymoves within the cylinder asthe nosewheel is turned.

Movement of the piston,caused by lateral oscillation ofthe nosewheel, is dampenedby the hydraulic fluid whichmust be forced through ori-fices in the piston as it moves

within the cylinder. Theshimmy damper also incorporates a compensating chamberwhich houses a spring-loaded valve that relieves thermal ex-pansion of the hydraulic fluid.

1. Shimmy Damper

2. Compensating Chamber

1

2




The nose gear retractsforward into the nose wheel

well and, when fully re-tracted, is enclosed bythree doors. The two for-ward doors are linked by

push-pull rods to a com-mon torque tube andbellcrank assembly suchthat they operate simulta-

neously. The torque tubeand bellcrank assembly islinked by a single push-pull rod to the trunnionassembly such that theforward doors mechani-

cally open during exten-sion and retraction, and

close following extension or retraction. The aft door is alsolinked by a single push-pull rod to the trunnion assembly suchthat it mechanically opens during extension and closes duringretraction. A universal joint straightening mechanism centersthe nose gear during retraction.

1. Torque Tube

2. Push-Pull Rods

3. Trunnion Assembly

4. Forward Door Hinges

2

34

2

1





8-5

Landing Gear Assemblies

HYDRAULIC OIL CHAMBERFILL PLUG

BRAKE LINES

TRUNNION

ASSEMBLY

SWIVELFITTING

DRAG

BRACE

MAIN GEAR

DOOR LINKAGE

AXLE

BRAKE

ASSEMBLY

TORQUE

LINKS

SQUATSWITCH

SHOCK

STRUT

AIR CHAMBER

AIR VALVE

ACTUATOR

STEEL

UNIVERSAL

JOINT

STEERING

GEARS

SHIMMY

DAMPER

FORWARDDOOR

LINKAGE

TRUNNION

AFT

DOOR

LINKAGE

TORQUE

LINKS

HYDRAULIC

ACTUATOR

DOWNLOCK

SWITCH

Main Gear

Nose Gear





8-7

Each main gear assembly incorporates an air-oil type shockstrut assembly which functions to support the weight of the air-plane and absorb shock loads encountered during ground op-eration. The shock strut assembly consists primarily of an innercylinder retained within a hydraulic fluid filled outer cylinder.The inner cylinder is integral to the axle/piston assembly while

the outer cylinder is housed within the trunnion assembly. Afloating isolation piston divides the interior of the inner cylinder

into a hydraulic fluid filled upper chamber and nitrogen gas ordry air charged lower chamber. A variable orifice separates theupper chamber of the inner cylinder from the outer cylinder andregulates the flow of hydraulic fluid between these areas in rela-tion to increasing or decreasing load.

Under increasing load, hydraulic fluid flows from the outer cylin-der to the upper chamber of the inner cylinder. As the strut

compresses, the variable orifice is progressively restricted.When the rate of hydraulic fluid flow through the orifice is insuf-ficient to absorb compression shocks, the isolation piston isforced downward against gas/air pressure to assume the addi-tional load. Under decreasing load, the isolation piston isforced upward by gas/air pressure and hydraulic fluid flowsfrom the upper chamber of the inner cylinder to the outer cylin-der until the pressure on each side of the orifice is equal.

The main gear assemblies retract inward into the wing wheelwells and, when fully retracted, each is partially enclosed by a

hinged door which opens during extension and closes duringretraction. Push-pull rods mechanically link the doors to their re-spective trunnions such that they operate simultaneously.




Hydraulic Actuators

Each hydraulic actuator is comprised of a cylinder body with aninternal piston and rod assembly. The cylinder body of eachactuator is attached to the airframe structure. Each rod end isattached to its associated gear trunnion assembly. Retract and

extend ports on each cylinder are plumbed to the landing gear

hydraulic system. Separate pneumatic extend ports areplumbed to the pneumatic blowdown system.

1. Cylinder Body 4. Hydraulic Retract Port

2. Piston Rod 5. Hydraulic Extend Port

3. Downlock Indicator 6. Pneumatic Retract Port

The nose gear actuator piston “pulls” the nose gear to theextended position and “pushes” it to the retracted position. The

main gear actuator piston “pushes” the main gear to the ex-tended position and “pulls” it to the retracted position. The nosegear actuator incorporates a shuttle valve that is normally

spring-loaded open to the hydraulic extend port. During auxil-iary/emergency gear extension, the shuttle valve is repositionedopen to the pneumatic extend port by gas/air pressure. Eachmain gear actuator incorporates a separate, concentric, pneu-matic extension chamber that is always open to the pneumaticextend port. All three gear actuators incorporate integral “ring

and groove” type downlock latches which hold the gear assem-blies in the fully-extended position when hydraulic pressure is

removed. Approximately 300 PSI is required to release theselatches.

3

2

4

6

5

1





8-9

Uplock Sequence

Actuators

Independent uplocksequence actuators areprovided for each gearassembly to hydrauli-

cally release the uplockhooks during normal

gear extension. Eachconsists of an actuatorbody, a spring-loadedinternal piston and rodassembly, and a checkvalve. The actuatorbody is attached to theairframe structure. The

uplock hook is linked tothe rod end. Threeports are located oneach actuator body: the first (pressure inlet) is plumbed to theextend circuit of the landing gear hydraulic system, the second(pressure outlet) is plumbed to the extend port of its associatedgear assembly’s hydraulic actuator, the third is plumbed to theretract port of its associated gear assembly’s hydraulic actua-tor. The check valve is positioned between the pressure inlet

port and pressure outlet port such that hydraulic fluid flow tothe extend port of the hydraulic actuator is restricted until the

uplock hook is released. During normal gear extension, hydrau-lic pressure applied to the internal piston “pulls” the rod inwardthereby releasing the uplock hook. Following uplock hookrelease, continued inward movement of rod unseats the checkvalve allowing hydraulic fluid flow to the extend port of thehydraulic actuator. During gear retraction, return hydraulic fluid

flow from the extend port of the hydraulic actuator unseats thecheck valve. During emergency gear extension, the uplockhooks are mechanically-released.

1. Actuator Body

2. Uplock Hook

3. Release Cable4. Uplock Switch

1

4

2

3




Left Main Gear Safety Switch

A safety switch, located on theleft main gear, is installedprimarily to prevent inadvert-ent retraction of the gear

during ground operation whenthe left main gear strut iscompressed. Numerous addi-tional functions are listed

below.

On Ground, Strut Compressed, Safety Switch Open

Enables:1. Generator-assisted engine start

2. Automatic ground cabin depressurization

3. Wheel brake anti-skid

4. Locked wheel crossover protection (550-0437 and after)

5. Thrust reverser deployment

6. Sperry TAS heater probe low heat level (550-0505 and earlier)

7. Tailcone pressurization

Disables:

1. Landing gear handle movement2. Emergency pressurization control valve

3. Touchdown protection (550-0437 and after)

4. Air Data Computer (550-0324 and after)

5. Stick shaker

6. Optional approach indexer (550-0627 and after)

7. Air data warning horn (550-0505 and earlier)

8. Ground Proximity Warning System (550-0376 and after)

9. Flight data recorder (550-0550 and after)

10. Cockpit voice recorder (550-0550 and after)

11. Angle of attack probe heat

12. Sperry TAS heater probe high heat level (550-0505 and earlier)

13. Hobbs meter

14. Davtron digital clock flight time function

15. Panel light dimming (550-0689 ~ 0698; 550-0703 and after)

16. Angle of attack indexer dimming (550-0550 and after)





8-11

In Flight, Strut Extended, Safety Switch Closed

Enables:

1. Landing gear handle movement

2. Emergency pressurization control valve

3. Touchdown protection (550-0437 and after)

4. Air Data Computer (550-0324 and after)

5. Stick shaker

6. Optional approach indexer, if nose gear is down and locked

(550-0627 and after)

7. Altitude alert warning horn

8. Ground Proximity Warning System (550-0376 and after)

9. Flight data recorder (550-0550 and after)

10. Cockpit voice recorder (550-0550 and after)

11. Angle of attack probe heat

12. Sperry TAS heater probe high heat level (550-0505 and earlier)13. Hobbs meter

14. Davtron digital clock flight time function

15. Panel light dimming (550-0689 ~ 0698; 550-0703 and after)

16. Angle of attack indexer dimming (550-0550 and after)

Disables:

1. Generator-assisted engine start

2. Automatic ground cabin depressurization

3. Wheel brake anti-skid

4. Locked wheel crossover protection (550-0437 and after)5. Thrust reverser deployment

6. Sperry TAS heater probe low heat level (550-0505 and earlier)

7. Tailcone pressurization




Landing Gear Control and Position Indication

Landing Gear Control Handle

The landing gear hydraulic system is activated by a two-posi-tion, wheel-shaped control handle, labeled LDG GEAR - UP/ DOWN, located in the lower left corner of the center instrumentpanel. Three microswitches are actuated by the control handle:

a retract switch, an extend switch, and a selector switch. Theretract switch is actuated when the control handle is moved to

the “UP” position; the extend switch is actuated when the con-trol handle is moved to the “DOWN” position. When eitherswitch is actuated, a corresponding retract or extend solenoidintegral to the landing gear control valve is energized, therebypositioning the valve to permit gear retraction or extension asselected. The control handle is spring-loaded to the selected

position and must be pulled outward before it can be moved.

Protection against inadvertent retraction of the gear duringground operation is provided by a spring-loaded, solenoid-

operated handle lock. The solenoid is energized and deener-gized through the left main gear safety switch. When the leftmain gear strut is compressed, the solenoid is deenergizedand the spring-loaded lock prevents the control handle frombeing moved to the “UP” position. When the left main gear strut

is uncompressed, the solenoid is energized by 28 VDC leftmain bus power through the 2-amp LDG GEAR circuit breaker,

and the handle lock is released.





8-13

The selector switch is integral to the landing gear positionindication and warning system and has two positions (up and

down) that correspond with control handle position. When thecontrol handle is in the “UP” position, the selector switch estab-lishes a circuit to the GEAR UNLOCKED light module throughthe uplock and downlock microswitches. When the control

handle is in the “DOWN” position, the selector switch estab-lishes a circuit to the GEAR UNLOCKED light module throughdownlock microswitches only.




1. Gear Down and

Locked Light Module

A light module incorporatingthree green lenses individuallylabeled NOSE, LH, and RH,located to the right of the

landing gear control handle,provides indication that thelanding gear is in the downand locked position. The LH

and RH main gear positionindicators are illuminated byone bulb each. The NOSEgear position indicator is illuminated by two bulbs wired inparallel for continued operation should one bulb fail. The downand locked position indicators are individually illuminated

through the downlock microswitches as each gear assemblyreaches its fully-extended position and its associated downlock

latch is engaged.

2. Gear Unlocked Light Module

A light module incorporating a single red lens labeled GEARUNLOCKED, located below the gear down and locked lightmodule, provides indication that the landing gear is in transit orthat one or all three landing gear assemblies are not in thesame position as the landing gear control handle. This indicatoris illuminated by two bulbs wired in parallel for continued

operation should one bulb fail. During gear retraction, theindicator is illuminated through the selector switch, downlockmicroswitches, and uplock microswitches when the controlhandle is positioned to “UP,” the downlock latches are re-leased, and uplock hooks are not engaged. During gear exten-sion, the indicator is illuminated through the selector switch andthe downlock microswitches only, when the control handle ispositioned to “DOWN” and the downlock latches are not en-

gaged. During extension and retraction, the downlockmicroswitches provide the ground required for GEAR UN-LOCKED indicator illumination only when downlock latches arenot engaged. When all three gear assemblies have reachedtheir fully-retracted or fully-extended position the GEAR UN-LOCKED indicator should be extinguished. When all three gearassemblies are up and locked, all position indicators should beextinguished.

R-6/27/96

1

2





8-15

Landing Gear Warning Horn

The landing gear warning horn is located in the overhead flightcompartment, aft and outboard of the pilot’s station, and pro-vides audible indication that the gear is not down and lockedwhen the airplane is configured for landing. The horn is ener-

gized through a pair of throttle position switches, a flap positionswitch, and an airspeed switch (550-0627 and after), integral to

the landing gear position indication and warning circuit. Eachthrottle position switch is actuated closed when its associatedTHROTTLE lever is set below approximately 70% N2. The flapposition switch is actuated closed by the FLAP lever when setbelow the T.O. & APPR position (approximately 15°). The air-speed switch, incorporated on airplanes 550-0627 and afteronly, is closed below approximately 150 KIAS. In combination,these switches and the downlock microswitches will cause the

horn to sound when any one or all three gear assemblies arenot down and locked under the following conditions:

Airplanes 550-0626 and earlier

1. Either or both THROTTLE levers set below approximately 70% N2regardless of airspeed

2. FLAP lever set below T.O. & APPR position, regardless of airspeed orTHROTTLE position

Airplanes 550-0627 and after

1. Either or both THROTTLE levers set below approximately 70% N2 whenairspeed is below approximately 150 KIAS

2. FLAP lever set below T.O. & APPR position, regardless of airspeed orTHROTTLE lever position

A HORN SILENCE button,located right of the controlhandle, energizes a pair ofrelays associated with thethrottle position switches.When this button is pressed,

both relays are energizedthereby interrupting the throttleposition switch circuit andsilencing the warning horn.

HORN SILENCE BUTTON




Each relay remains energized through an integral latchingcircuit as long as its associated throttle position switch remains

closed and all three gear assemblies are not down and locked.If the horn has been energized through the throttle positionswitches only, it may also be silenced by advancing theTHROTTLE levers above approximately 70% N2, or by increas-

ing airspeed above 150 KIAS (550-0627 and after only). If thehorn has been energized through the flap position switch,pressing the HORN SILENCE button, advancing the THROTTLElevers, or increasing airspeed (550-0627 and after only), will not

silence the horn.

Landing Gear Indicator Light

and Warning Horn Test

Operation of the landing gearposition indication and warn-ing system may be verified byrotating the TEST selectorswitch, located on the lowerleft instrument panel, to the“LDG GEAR” position. This

energizes a test relay thatbypasses the downlockmicroswitches which normallyprovide the ground required toilluminate the position indicators and sound the warning horn.In this condition, if the system is functioning normally, the LH,

RH, NOSE, and GEAR UNLOCKED position indicators shouldbe illuminated and the warning horn should sound. Operation ofthe HORN SILENCE button may also be verified during this test.

If functioning normally, pressing the button should silence thehorn.

The position indicators and warning horn normally receive 28VDC power from the left main bus (550-0550 and after) or rightmain bus (550-0505 and earlier) bus through the 2-amp LDG

GEAR circuit breaker. During system testing, the test relay andGEAR UNLOCKED position indicator receive 28 VDC power

from the left main bus (550-0550 and after) or right main bus(550-0505 and earlier) through the 5-amp WARN LTS 1 circuitbreaker.






8-17

Landing Gear Hydraulics

Landing Gear Hydraulic Pressure Source

The hydraulic power system provides hydraulic pressure forlanding gear retraction and normal extension. The system alsosupplies hydraulic pressure for operation of the speed brakes

and optional thrust reversers when installed. A detailed de-scription of the hydraulic power system is provided below.

Hydraulic Power System Components

Major components of the hydraulic power system include ahydraulic fluid reservoir, engine-driven hydraulic pumps, sole-noid-operated control valve, and annunciators that indicatesystem operating conditions.

Hydraulic Fluid Reservoir

The hydraulic fluid reservoir, located within the aft fuselage,stores fluid and supplies the hydraulic pumps. The reservoir iscomprised of two cylindrical sections joined end-to-end. Thelarger cylinder houses the fluid reservoir; the smaller cylinder(neck) primarily houses a fluid level sight gage. A piston di-vides the larger cylinder into separate fluid-filled and air-filledchambers. On airplanes 550-0180 and after, and airplanes 550-0065 and earlier not incorporating SB550-29-1 or -2, the reser-voir is considered pressurized. On airplanes 550-0066 through0179, and airplanes 550-0065 and earlier incorporating SB550-

29-1 or -2, the reservoir is considered non-pressurized.

In pressurized reservoirs, the large reservoir piston is attachedto and positioned by a small piston located within the neck. Thesmall piston is positioned by hydraulic pressure. With 1350 to1500 PSI hydraulic pressure acting on the small piston, thelarge piston maintains approximately 15 PSI within the fluid-filled chamber. The large piston is spring-loaded to maintain

approximately 3 to 4 PSI within the fluid-filled chamber whenhydraulic pressure is not available.

In non-pressurized reservoirs, the piston is spring-loaded tomaintain approximately 3 to 4 PSI within the fluid-filled chamberThe air-filled chamber is vented to facilitate piston movement.




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8-19

Overpressurization of the fluid-filled chamber is prevented by a

relief valve that begins opening at 40 PSI and fully opens at 60PSI. When open, excess hydraulic fluid is routed through anoverboard relief line. On airplanes 550-0482, 0485 through0698 incorporating SB550-29-06, and airplanes 550-0699 andafter, this fluid drains into a plastic EPA bottle within the aft

fuselage. On airplanes 550-0484, 0483, 0481 and earlier, andairplanes 550-0482, 0485 through 0658 not incorporatingSB550-29-06, this fluid drains overboard through the hydraulic

service panel vent mast located on the lower right surface of theaft fuselage.

The fluid level sight gage consists of an indicator rod, visible

through a window on the reservoir neck, that is attached to andpositioned by the reservoir piston. The position of the rod inrelation to REFILL, FULL, and OVERFULL markings above thewindow, and corresponding 38, 125, and 150 IN3 markingsbelow the window, indicates the reservoir fluid level. The 38,125, and 150 IN3 markings correspond to 0.2, 0.5, and 0.6gallons respectively. A microswitch is attached to the neck suchthat it is held open by the indicator rod when it is positionedabove REFILL. When the rod is positioned below REFILL, the

microswitch closes illuminating the amber [HYD LOW LEVEL]annunciator (550-0550 and after) or [HYD LEVEL LO] annuncia-

tor (550-0505 and earlier).

© PCW




The reservoir is serviced through a pressure line coupling on

the hydraulic service panel. Reservoir capacity is approxi-mately 0.65 gallons with the gear down and locked. The hy-draulic system reservoir may be serviced with Skydrol 500A,500B, 500B-4, 500C, and LD-4; Hyjet, Hyjet W, III, or IVA; or anyequivalent phosphate ester based hydraulic fluid. Mixing hy-

draulic fluids should not impair system operation.

Caution: Phosphate ester based hydraulic fluid will attack a

wide range of materials, including rubber, copper, variousplastics, and paints.

Caution: If heated beyond 270°F, Skydrol decomposes intoacids and other products that can damage metal structures.





8-21

Engine-Driven Hydraulic Pumps

Hydraulic pressure is providedby constant-displacementhydraulic pumps that are gear-driven by and mounted on theaccessory gearbox of each

engine. When hydraulic pres-sure is not required to operateairplane systems, the hydraulic

system functions in an “opencenter” fashion, as the pumpsdraw hydraulic fluid from thereservoir, circulate it through thehydraulic system, and return it to the reservoir through the returnline. When hydraulic pressure is required to operate landinggear or other systems, a bypass valve is energized closed,

thereby allowing hydraulic pressure to increase. A check valveis installed in the return line to the reservoir to prevent reverse

flow from the reservoir. Check valves, installed downstream fromeach pump and its associated filter, prevent reverse fluid flow inthe event of opposite pump failure. These check valves alsoincorporate flow switches or flow detectors that illuminate corre-sponding annunciators to indicate low hydraulic flow/pressurefrom each hydraulic pump. A restrictor check valve is installedin the pressure line downstream of the right hydraulic pump toprevent reverse fluid flow during servicing.

Hydraulic FiltersA hydraulic filter is installed in each pump pressure line and inthe hydraulic reservoir return line to prevent foreign materialfrom contaminating the hydraulic fluid. The pump pressure linefilters have a 3 GPM nominal flow capacity, and incorporate a100 PSI differential bypass valve that permits continued flow inthe event of filter blockage. The reservoir return line filter has a

12 GPM nominal flow capacity and incorporates a 100 PSIdifferential bypass valve. These filters have a 5-micron nominalrating and a 15-micron absolute rating.




Hydraulic Firewall Shutoff Valves

A hydraulic firewall shutoff valve is installed within the aft fuse-

lage in each hydraulic pump suction line. The hydraulic firewallshutoff valves function to terminate hydraulic fluid flow in the

event of an engine fire. Each ball-type valve is motor-operatedand electrically-controlled by independent [LH ENG FIRE] and[RH ENG FIRE] annunciator/switches on the glare shield panelfire tray. A transparent, spring-loaded guard is installed overeach switch to protect against inadvertent actuation.

When either switch is pressed, the corresponding valve isclosed by 28 VDC power supplied through its associated 7.5-amp LH F/W SHUTOFF or RH F/W SHUTOFF circuit breaker onthe left CB panel. When fully closed, the corresponding amber[LH] or [RH] [HYD FLOW LO] annunciator (550-0550 and after)

or [L] or [R] [HYD PRESS LO] annunciator (550-0505 and ear-lier) should be illuminated. Additionally, both [BOTTLE ARMEDPUSH] annunciator/switches on the glare shield panel fire tray

should be illuminated. A closed valve may be reopened bypressing the corresponding switch to release it from its lockedposition.

Note: The fuel firewall shutoff valve is also closed when thehydraulic firewall shutoff valve is closed. Refer to Chapter 3 fora complete description of the fuel firewall shutoff valve. Refer to

Chapter 2 for a complete description of the engine fire protec-tion system.





8-23

Hydraulic System Indication

Indication of hydraulic system operation is provided by flow,pressure, and fluid level lights located on the annunciatorpanel.

Hydraulic Flow Annunciators

Independent amber [LH] and [RH] [HYD FLOW LOW] annun-ciators (550-0550 and after) or [L HYD PRESS LO] and [R HYDPRESS LO] annunciators (550-0505 and earlier) illuminate to

indicate low hydraulic flow from each engine-driven pump.

On airplanes 550-0050 through 0063, 0065 and after, the hy-draulic fluid flow annunciators are illuminated by independentleft and right flow switch/check valves. Each flow switch/checkvalve is comprised of a sliding magnet that is attached to and

moves with the check valve poppet. If the system is operatingnormally, each annunciator will illuminate when the airplane

electrical system is energized, and will remain illuminated untilits associated hydraulic pump develops sufficient pressure tounseat the check valve, thereby magnetically opening theswitch and extinguishing the corresponding annunciator.

On airplanes 550-0064, 0049 and earlier, the hydraulic fluidflow annunciators are illuminated by a combination checkvalve/flow detector that is comprised of a sliding magnet andtwo switches. One switch controls illumination of the amber

[L HYD PRESS LO] annunciator; the other switch controls illumi-nation of the amber [R HYD PRESS LO] annunciator. The slidingmagnet is spring-loaded to a neutral position. When an inletport pressure differential exists, the magnet is forced toward thearea of lowest pressure. If one pump is generating at least 25PSI more than the other, the magnet closes the switch corre-sponding to the pump that is generating the lower pressure,thereby illuminating the associated annunciator. If the system is

operating normally, the annunciator corresponding to the en-gine started last will illuminate following initial engine start, andwill remain illuminated until its associated hydraulic pumpdevelops sufficient pressure to move the flow detector magnetto the neutral position, thereby magnetically opening the switchand extinguishing the corresponding annunciator.




Hydraulic Pressure Annunciator

Pressurization of hydraulic components is indicated by an

amber [HYD PRESS ON] annunciator. This annunciator isilluminated when pressure exceeds approximately 155 PSI, bya hydraulic pressure switch, located upstream of the landinggear control valve.





8-25

Landing Gear Control Valve

The landing gear control valve functions to direct hydraulic fluidto the hydraulic actuator retract or extend ports. The hydrauliclanding gear control valve is operated by “retract” and “extend”solenoids. The control valve contains an internal selector spoolthat is spring-loaded to a neutral position when the solenoids

are not energized. In this position, both the retract and extendports are connected to the return port so that fluid will not betrapped under pressure within the hydraulic lines. The other

selector spool positions depend upon the position of the land-ing gear control handle. When the control handle is moved tothe “UP” position, the retract solenoid is energized, whichpositions the selector spool to connect the hydraulic inlet port tothe retract port and connect the extend port to the return port.When the control handle is moved to the “DOWN” position theextend switch is actuated, energizing the extend solenoid,

which positions the selector spool to connect the hydraulic inletport to the extend port and connect the retract port to the return

port. The control handle is spring-loaded to the selected posi-tion and must be pulled outward before it can be moved.

The landing gear control valve functions with an input power of18 to 30 VDC.

System pressure is regulated by a relief valve that begins toopen at approximately 1350 PSI, and prevents hydraulic system

pressure from exceeding 1500 PSI. The relief valve also pre-

vents hydraulic fluid flow rate from exceeding 6.6 GPM. Excessfluid released through the relief valve flows into the reservoirreturn line.

During landing gear extension or retraction, the [HYD PRESSON] annunciator should be illuminated as an indication ofadequate hydraulic pressure. Failure of this light to illuminate

indicates that hydraulic pressure is insufficient for system op-eration. Continued illumination after all three landing gearassemblies have reached the selected position indicatesmalfunction of a hydraulic system component.




Landing Gear Operation

The following paragraphs describe normal landing gear retrac-tion, as well as normal and emergency landing gear extension.The airplane has no emergency landing gear retractioncapability.

Landing Gear Retraction

Landing gear retraction is initiated by positioning the landing

gear control handle to “UP.” In this position, the control valveretract solenoid and hydraulic system center valve are ener-gized to permit landing gear retraction. During retraction, thecontrol valve selector spool to is positioned to direct hydraulicfluid to the retract port of each actuator and the center valvecloses, illuminating the amber [HYD PRESS ON] annunciator.Once hydraulic pressure releases the downlock latches, the

retract side of the gear actuator pistons is pressurized, therebyinitiating gear retraction. Release of the downlock latchesactuates downlock switches that illuminate the red GEAR UN-LOCKED light while the gear is in transit.

As each gear assembly reaches its fully-retracted position, it isheld in position by a gear uplock hook that engages an uplockroller on the gear trunnion. An uplock switch is actuated closedby each uplock hook. When the landing gear is fully retractedand all three microswitches have closed, the red GEAR UN-

LOCKED light is extinguished, the hydraulic system center

valve opens, and the retract solenoid is deenergized. In thiscondition, the control valve selector spool returns to its neutralposition and the hydraulic system returns to “open center”mode, extinguishing the amber [HYD PRESS ON] annunciator.

Note: When the center valve is energized and hydraulic pres-sure is being supplied to the landing gear, the amber [HYD

PRESS ON] annunciator will be illuminated.



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Landing Gear Extension

Landing gear extension is normally initiated by positioning thelanding gear control handle to “DOWN.” In this position, thecontrol valve extend solenoid and hydraulic system centervalve are energized to permit landing gear extension. Duringextension, the control valve selector spool to is positioned to

direct hydraulic fluid to the uplock sequence actuators and thecenter valve closes, illuminating the amber [HYD PRESS ON]annunciator. As sufficient hydraulic pressure builds within these

actuators, the uplock hooks are released. After each uplockhook is fully released, fluid is routed to the extend port of itsrespective landing gear actuator and the gear begins to extend.Individual microswitches, actuated open by the release of theuplock hooks, energize the red GEAR UNLOCKED light whilethe gear is in transit.

As each gear actuator reaches its fully-extended position,internal downlock latches mechanically engage to hold the

gear extended. Individual microswitches are actuated closedby the downlock latches, illuminating their respective greenNOSE, LH, and RH gear down indicators. When all three gearassemblies are fully down and locked and their respectivemicroswitches are closed, the red GEAR UNLOCKED light isextinguished, the hydraulic system center valve opens, and theextend solenoid is deenergized. In this condition, the controlvalve selector spool returns to its neutral position and the hy-

draulic system returns to “open center” mode, extinguishing the

amber [HYD PRESS ON] annunciator.

Note: When the center valve is energized and hydraulic pres-sure is being supplied to the landing gear, the amber [HYDPRESS ON] annunciator will be illuminated.

Static (Open Center) Condition

When normal extension or retraction is complete, the hydraulicfluid contained in both the retract and extend circuits remainsstatic while fluid circulation through the pumps, filters, checkvalves, and reservoir continues.

Landing Gear Extension/Retraction Speeds and Cycle Times

On airplanes 550-0626 and earlier, maximum landing gearextended speed is 250 KIAS. On airplanes 550-0627 and after,maximum landing gear extended speed is 262 KIAS. Maximumlanding gear retraction speed is 200 KIAS. Normal cycle time

for landing gear operation is approximately six seconds.




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Auxiliary/Emergency Landing Gear Extension System

An auxiliary/emergency landing gear extension system isinstalled for use in the event of hydraulic system failure. Auxil-iary/emergency gear extension is provided by a mechanicaluplock release “free-fall” system and a pneumatic “blowdown”

system. The free-fall system allows mechanical gear releaseusing a T-handle connected by cables to each uplock hook.Pulling the T-handle releases the uplock hooks mechanically.The pneumatic blowdown system is used to assure that the

gear is fully extended and locked, and is actuated using ablowdown knob located behind the T-handle. Major compo-nents of the system include an auxiliary/emergency air storage“blowdown” bottle and discharge valve, the auxiliary/emer-gency gear release T-handle, the blowdown knob, and associ-ated cables and plumbing.

Auxiliary/Emergency

Air Storage Bottle

The auxiliary/emergency airstorage “blowdown” bottle islocated behind the aft dividerwithin the right nose baggagecompartment and is pressur-ized with dry nitrogen or clean,dry compressed air. The bottlesupplies pneumatic pressure

to operate both the auxiliary/ emergency landing gearblowdown system and theauxiliary/emergencybraking system.

A bottle pressure gage is visible through an inspection windowpositioned on an access panel. The access panel is hinged to

facilitate servicing and is accessible through the right nosebaggage door. The gage is marked with a green arc, denotingthe normal system indicating range, from 1800 to 2050 PSIG.The bottle should be serviced when indicated bottle pressure isless than 1800 PSIG. The bottle holds from 75 to 100 cubicinches of air.

1. Pressure Gage

2. Air Storage Bottle

1

2





8-31

The blowdown bottle is equipped with a thermal relief valve thatfunctions to release excessive pressure overboard through a

vent line. The relief valve is designed to rupture at approxi-mately 3750 to 4250 PSIG. The thermal relief valve is not reus-able; if it ruptures, the valve or the complete bottle assemblymust be replaced.

Warning: Thermal relief valve rupture will render the auxiliary/ emergency landing gear extension and braking systems inop-erable.

Note: The blowdown bottle requires hydrostatic testing everyfive years.

Auxiliary/Emergency Gear

Extension System Operation

Should the landing gear fail toextend hydraulically, requiring

operation of the auxiliary/ emergency extension system,the red AUX GEAR CONTROLT-handle should be pulledfully out to release the landinggear uplocks. The T-handleshould then be rotated 45°clockwise to unlock. Thisaction allows the landing gear

to free-fall, and also unlocksthe red, collar-type pneumaticblowdown knob. To assure that the landing gear is fully downand locked, the blowdown knob should be pulled to actuate avalve that discharges the blowdown bottle.

The blowdown knob is connected by cable to the blowdownbottle discharge valve. Pulling the blowdown knob opens the

discharge valve, and routes pneumatic pressure directly to thenose and main gear actuator pneumatic extend ports, therebyforcing the gear into the down and locked position. Hydraulicfluid within the gear actuators is returned directly to the hydrau-lic fluid reservoir by means of a dump valve.

Note: Before operating the AUX GEAR CONTROL T-handle, theflight crew should assure that the landing gear control handle is

in the “DOWN” position.

1. T-Handle

2. Blowdown Knob

2

1




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8-33

Note: When the blowdown knob is pulled, the blowdown bottledischarge valve latches in the open position and cannot be

reset by the flight crew. Therefore, the blowdown system cannotbe reused until the discharge valve is reset and the storagebottle is refilled by service personnel. When the blowdownbottle discharge valve is reset to the closed position during

servicing, it connects the landing gear actuators to a vent line,allowing gas/air trapped within the gear actuators to be re-leased overboard when the gear is next operated hydraulically.

Note: To ensure the highest probability of full extension, theauxiliary/emergency gear extension system should be operatedat an airspeed of approximately 150 KIAS with flaps retracted.

The landing gear may not fully extend if free-fall landing gearextension is attempted at airspeeds above 200 KIAS.

As in normal extension, respective gear down indicators shouldilluminate when each gear assembly reaches its fully-extended

position.

Emergency Hydraulic Dump Valve

Following auxiliary/emergency (pneumatic) gear extension,hydraulic fluid within the retract side of the landing gear actua-tors is returned directly to the hydraulic fluid reservoir by meansof a dump valve that is connected to the hydraulic gear retractline, the hydraulic system return line, and the pneumatic gearextension line. When pneumatic pressure within the landing

gear system exceeds 200 PSI, the dump valve opens, allowinghydraulic fluid to flow through the dump valve into the return lineto the reservoir.




Brakes

The left and right wheel brakesare hydraulically operated byindependent master cylindersattached to the pilot’s and

copilot’s rudder pedals. Thebrake system is pressurizedwhen either pilot depressesthe toe pedals. Interconnect

assemblies allow either pilot tooperate the brakes with equalauthority. The brake system ispower-assisted and providestouchdown protection capability. An anti-skid system providesskid and locked wheel crossover protection. A backup pneu-

matic system can be used for auxiliary/emergency braking inthe event of hydraulic brake system failure; on airplanes 550-

0460 and earlier not incorporating SB550-32-12, backupmanual braking is available as well. Parking brake capability isprovided by locking the normal brakes.

Power Brake System

The power brake system is composed of a brake hydraulic fluidreservoir, an electrically-driven hydraulic pump and filter as-sembly, one or two accumulators, and an anti-skid servo valve.

Hydraulic fluid for brake sys-tem operation is supplied froma reservoir installed within thenose compartment on the rightside of the forward pressurebulkhead. On airplanes 550-0281 and after, airplanes 550-0039 and earlier, and air-

planes incorporating SB550-32-8, the brake reservoir ispressurized by cabin pres-sure, which enters the reservoirthrough an assembly thatincorporates a check valve toprevent hydraulic fluid orfumes from entering the cabin,

and a filter to protect the brake fluid from contamination. Thereservoir is vented by an overboard vent line connected to arelief valve that incorporates a restrictor orifice.

1. Fluid Reservoir

2. AccumulatorPressure Gage

2

1




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The restrictor orifice allows pressurization to accumulate withinthe reservoir, but is never fully closed, thereby allowing continu-

ous venting of fumes. On airplanes 550-0040 through 0280 andairplanes not incorporating SB550-32-8, the reservoir is notpressurized, and is vented by an overboard line.

Pressure for power brake system operation is provided bypressurized hydraulic fluid supplied by an electrically-drivenhydraulic pump installed within the left side of the nose com-partment. The pump is controlled by a pressure switch that

opens when the pressure approaches 1300 PSI and closeswhen system pressure approaches 900 PSI. The pump sup-plies hydraulic fluid to the power brake system and to one ortwo accumulators under the left nose baggage compartmentfloor. Each accumulator is divided by a floating piston into anair compartment and a fluid compartment. The air compartment

is pneumatically pressurized to a nominal pressure of 650 to700 PSIG, and acts to pressurize the contents of the fluid com-

partment.

The accumulator(s) functions to maintain hydraulic pressurewithout the need to continuously operate the hydraulic pump,thereby assuring that hydraulic power-assist is immediatelyavailable to the brake actuators. Fluid expelled from the accu-mulator during brake operation is returned to the accumulator

by the pump. On airplanes 550-0453 and after, the main gearpower brake system includes one 50 cubic inch accumulator.

On airplanes 550-0437 through 0452, the main gear powerbrake system includes two 25 cubic inch accumulators. Onairplanes 550-0436 and earlier, the main gear power brakesystem includes one 25 cubic inch accumulator.





8-37

Applying pressure to the brake pedals actuates a piston as-sembly in each master cylinder. Master brake cylinder piston

displacement transfers hydraulic pressure to the anti-skid servovalve. Braking is power-assisted by hydraulic pressure sup-plied from the accumulator(s). Hydraulic fluid from theaccumulator(s) is released by the anti-skid servo valve through

the parking brake valve below the flight compartment floor tothe main gear brake assemblies. The additional fluid pressurehydraulically amplifies the effect of brake pedal pressure.

A color band gage is provided for checking brake hydraulicaccumulator pressure. The pressure gage is visible through aninspection window positioned on an access panel. The accesspanel is hinged to facilitate servicing and is accessible throughthe right nose baggage door. The gage is marked with a red

arc between 0 and 650 PSIG denoting underpressure, a green

arc between 650 to 700 PSIG denoting normal operating pres-sure range, a yellow arc between 700 to 900 PSIG denoting

caution range, a green crosshatched arc between 900 to 1350PSIG denoting precharge range, and a red arc between 1350 to1500 PSIG denoting overpressure.

Note: The accumulator pressure gage denotes pressureranges only; it is not marked with numerals denoting specificpressures.

The power brake system receives 28 VDC left main bus power

through a 20-amp SKID CONTROL circuit breaker on the left CBpanel.

Touchdown Protection

On airplanes 550-0437 and after, the touchdown protectionsystem prevents landing with pressure applied to the brakes.

The touchdown protection system energizes the power brakeanti-skid control valve open at airplane touchdown through theleft main gear safety switch. The valve therefore remains closed

until the airplane is on the ground, preventing hydraulic fluidfrom entering the brake actuators and thereby assuring that thebrakes are not applied at airplane touchdown.




Anti-Skid System

The anti-skid system is integral to the power brake system and

provides maximum braking capability on wet or icy runways.The system consists of two wheel speed transducers, an elec-tronic control box, an anti-skid servo valve, pressure switches,mode switch, circuit breakers, and indicator lights.

The wheel speed transducers are installed in the main gear

axles. Each transducer generates electrical signals proportionalto wheel rotational speed; these signals are then transmitted tothe electronic control box. The control box also receives signalsfrom the left main gear safety switch and the brake hydraulicpump pressure switch. The control box averages signals fromleft and right transducers to generate a composite referencevoltage and compares this voltage to left and right transducer

voltage. The control box then generates signals that actuate the

anti-skid servo valve within the appropriate disc brake assem-bly to reduce braking pressure as required to prevent wheel

skidding.

The system detects incipient skids by using a wheel speedtransducer to measure the deceleration of each landing wheel,and then prevents skids by reducing the brake pressure inproportion to the deviation of each wheel from normal braking

deceleration. The system also modulates brake pressure tomaximize braking efficiency.

The anti-skid system is acti-vated by positioning the ANTI-SKID switch on the LDG GEARcontrol panel to “ON.” If theANTI-SKID switch is set to“OFF,” power braking shouldbe available without the anti-skid function, and the amber

[ANTISKID INOP] annunciatorshould illuminate. If the brakesystem hydraulic pressuredecreases to 900 PSI, thebrake hydraulic pump pres-sure switch closes, causing the control box to illuminate theamber [POWER BRAKE LOW PRESS] annunciator (550-0550and after) or [POWER BRK PRESS LO] annunciator (550-0505

and earlier).

ANTI-SKID SWITCH





8-39

The anti-skid brake system automatically disengages whenground speed falls below approximately 12 knots.

The anti-skid system receives 28 VDC left main bus powerthrough the 20-amp SKID CONTROL circuit breaker on the leftCB panel.

Warning: On airplanes 550-0461 and after and earlier airplanesincorporating SB550-32-12, power brake system failure rendershydraulic braking completely inoperable. If the power brake

system fails, auxiliary/emergency pneumatic braking must beused.

Note: When the anti-skid system is operating, the pilot shouldapply maximum braking pressure throughout the braking run. Ifthe pilot attempts to modulate brake pressure while the system

is releasing applied brake pressure to avoid a skid, the appliedbrake signal can be interrupted, resulting in a considerable

increase in braking distance.

Anti-Skid System Test

Anti-skid system ground self-test may be activated byrotating the test selector switchon the lower left instrumentpanel to the “ANTISKID”position, then returning it to the

“OFF” position. The anti-skidtest circuit monitors anti-skidsystem electrical function. Ifthe anti-skid system is func-tioning normally, the amber[ANTISKID INOP] annunciatorwill illuminate, then extinguish 3 to 4 seconds after the TESTswitch is returned to “OFF.” If an anti-skid system fault is de-

tected, the [ANTISKID INOP] annunciator will remain illumi-nated. While airborne, the anti-skid test circuit is automaticallyactivated when the landing gear control handle is positioned to“DOWN” (if the ANTI-SKID switch is positioned to “ON”). If anti-skid system test detects a fault, the [ANTISKID INOP] annuncia-tor will remain illuminated.

Note: If an anti-skid system fault is detected, the ANTI-SKID

switch can be positioned to OFF. If the SKID CONTROL circuitbreaker is engaged, normal power-assisted hydraulic brakingis available without anti-skid protection.






8-41

Note: The capacity of a properly serviced air storage bottle issufficient for at least ten individual brake applications if the

landing gear has not been extended pneumatically. Bottlecapacity is adequate to provide auxiliary/emergency brakingfor most conditions even if the landing gear has been extendedpneumatically. After use of auxiliary/emergency braking, it is

recommended that the engines be shut down and the airplanebe towed to the ramp, as there is no gage, light, or other warn-ing device in the cockpit to alert the flight crew when the pneu-matic bottle is depleted.

Auxiliary/Emergency Braking System Usage

On airplanes 550-0460 and earlier incorporating SB550-32-12and airplanes 550-0461 and after, the auxiliary/emergencybraking system must be used if the power brake system fails.On airplanes 550-0460 and earlier not incorporating SB550-32-

12, the brakes can be applied without power assist and/or theauxiliary/emergency braking system can be used if the power

braking system fails.

Parking Brake System

The parking brake handle,located below the lower leftinstrument panel, operates aparking brake control valveinstalled in-line downstream ofthe master cylinders and

upstream of the brake assem-blies. The parking brake is setby pressing the rudder toepedals until sufficient hydrau-lic pressure has developed inthe lines, then pulling the parking brake handle out to close thecontrol valve. With the control valve closed, hydraulic pressureis retained in the lines, thereby holding the brakes in the ap-

plied position. Pushing the parking brake handle in opens thecontrol valve, thereby releasing the brakes.

Note: The parking brake should not be set if the flight crewsuspects that the brakes may be unusually hot. Setting thebrake increases cool down time by impeding airflow, and

therefore may allow sufficient heat transfer to open the parkingbrake thermal relief valves and/or melt the thermal relief plugs

in the wheel, causing tire deflation.




Limitations



Refer to the applicable airplane manufacturer’s FAA approvedflight manual or approved manual material (supplementary





Chapter 9Environmental System

Table of Contents

Overview ...............................................9-1

Pressurization Air Source and Selection ..9-2

Source Selection..............................9-2

Air Cycle Machine and TemperatureControl System .....................................9-6

Air Cycle Machine ............................ 9-6

Temperature Control System.............9-8

Temperature Control SystemIndication and Warning ..................... 9-9

ACM Overpressure Warning andProtection ........................................9-9

ACM Overheat Warning andProtection ...................................... 9-10

Pressurization Air Source andSelection ............................................9-13

Nacelle Precooler System ............... 9-14

Source Selection............................9-15

Air Cycle Machine and TemperatureControl System ...................................9-18

Air Cycle Machine .......................... 9-18ACM Overpressure Protection..........9-22

ACM Overheat Warning andProtection ...................................... 9-22

Temperature Control System................9-24

Conditioned Air TemperatureWarning and Protection .................. 9-25




Environmental System, continued

Table of Contents

Air Distribution System ........................ 9-27

Distribution Control ........................ 9-27

Blowers .........................................9-29

Distribution Ducting andAir Outlets ..................................... 9-30

Cabin Pressurization Control System .... 9-31

Cabin Outflow Valves......................9-31

Cabin Pressurization Controller ........9-34

Pneumatic Relay ............................9-37

Cabin Altitude Limit Valves ............. 9-38Pressurization System Indication ..... 9-39

Cabin Altitude and DifferentialPressure Indicator .......................... 9-39

Cabin Rate-of-Change Indicator ........9-39

Cabin Altitude WarningAnnunciator ...................................9-39

Emergency Dump ...........................9-40

Ambient Air Sources............................9-41

Vapor Cycle Air Conditioning System .... 9-43

Refrigerant Circulation System ........ 9-43

Vapor Cycle Air ConditioningSystem Protection ..........................9-45

Vapor Cycle Air Conditioning SystemControls and Indicators .................. 9-46

Vapor Cycle Air ConditioningSystem Protection ..........................9-46




Environmental System, continued

Table of Contents

Emergency Oxygen System .................. 9-47

Oxygen Outlets...............................9-48

Crew Oxygen Masks .......................9-49

Passenger Oxygen Masks ...............9-50

Oxygen System Controls, MalfunctionWarning, and Indication .......................9-51

Limitations..........................................9-53





ENVIRONMENTAL SYSTEM 12/99 FOR TRAINING PURPOSES ONLY 9-1

Overview

The environmental system of the Cessna Citation II functions to

control cabin pressure, temperature, and ventilation to ensurethe comfort and safety of the flight crew and passengers.Engine bleed air is the primary source of air for cabin pressur-

ization and ventilation. An air cycle machine conditions thebleed air for delivery to the cabin. A cabin pressurization con-trol system regulates cabin pressure. During unpressurizedoperation, ambient air may be used for cabin ventilation. Anoptional flood cooling system may be installed to enhance

ambient air ventilation. An optional vapor cycle air conditioningsystem may be installed to provide supplemental cabin coolingprimarily during ground operations.

In this chapter, the environmental system is divided into the

following major groupings: pressurization air source and selec-tion, the air cycle machine and temperature control system, theair distribution system (including cabin ventilation and wind-shield defogging), cabin pressurization control and indication,

ambient air sources (including flood cooling and tailcone pres-surization), the vapor cycle air conditioning system, and theemergency oxygen system.

Depending on airplane unit number, two basic environmentalsystem configurations exist, differing primarily in relation topressurization air source and selection as well as the air cycle

machine (ACM) and temperature control system. For organiza-tional purposes and clarity, these primary differences aredescribed separately according to unit number range.




Pressurization Air Source and Selection

(550-0482, 0485 and after)

Bleed air used for cabin pressurization and ventilation is ex-

tracted from the compressor section of each engine throughports located at 4 and 8 o’clock positions on the gas generator

case. From these ports, bleed air is primarily routed throughindependent supply tubes to the ACM located within the aftfuselage. An environmental flow control/shutoff valve in each ofthese tubes controls the flow of bleed air from each source tothe ACM and functions as a check valve to prevent thebackflow of bleed air to the opposite source when either engine

has failed or is operating at a sufficiently lower RPM. A groundshutoff valve bypasses the right environmental flow control/ shutoff valve permitting increased right engine bleed air flowthrough the ACM to enhance cabin ventilation during ground

operation. During normal operation, all bleed air flows throughthe ACM en route to the cabin. In an emergency, left enginebleed air may be supplied directly to the cabin through anemergency supply tube and pressurization valve.

The environmental flow control/shutoff valves are normally-open, electrically-actuated closed, and have a nominal flowrate of approximately 6 pounds per minute (PPM) each. Theground shutoff valve is motor-operated, electrically-actuatedopen and closed, and has a nominal flow rate of approximately18 PPM. The emergency pressurization valve is normally-

closed and electrically-actuated open.

Source Selection

All four valves are controlledprimarily by the PRESSSOURCE selector switch onthe environmental “tilt” panel.The valves are also controlledby various switches that sense

bleed air pressure and tem-perature. The effects of rotat-ing the selector switch to eachof its six positions are de-scribed in the following para-graphs.





NORM Position

During normal operation with both engines operating andthe PRESS SOURCE selector switch set to “NORMAL,” both

environmental flow control/shutoff valves will be open,permitting bleed air flow through the ACM and into thecabin at a rate of approximately 12 PPM.

LH and RH Positions

When set to “LH” or “RH,” bleed air is correspondinglysupplied by the left engine or right engine only at a rate ofapproximately 6 PPM. In this condition, the environmental

flow control/shutoff valve for the non-selected source isenergized closed by 28 VDC left main bus power throughthe 5-amp NORM PRESS circuit breaker on the left CBpanel.

OFF PositionWhen set to “OFF,” both environmental flow control/shutoffvalves are energized closed through the NORM PRESScircuit breaker.

GND Position

To enhance cabin ventilation during ground operation,primarily when the right engine is operating only, the PRESSSOURCE selector switch should be set to “GND.” In thisposition, both environmental flow control/shutoff valves areenergized closed, the ground shutoff valve is energized

open, and the amber [BLD AIR GND] annunciator is illumi-nated through the NORM PRESS circuit breaker. With theground shutoff valve open, right engine bleed air flowsthrough the ACM and into the cabin at a rate of approxi-mately 18 PPM. Selection of this source is inhibited by theleft main gear safety switch when the airplane is in flight.




EMER Position

Should bleed air flow through the ACM be insufficient tomaintain selected cabin altitude, the PRESS SOURCE selec-tor switch should be set to “EMER.” In this position, theamber [EMERG PRESS ON] annunciator is illuminated, bothenvironmental flow control/shutoff valves are energized

closed, and the emergency pressurization valve is ener-gized open. 28 VDC right main bus power is supplied to thevalves and the annunciator through the 5-amp EMER

PRESS circuit breaker on the left CB panel. Selection of thissource is inhibited by the left main gear safety switch whenthe airplane is on the ground.

With the emergency pressurization valve open, uncondi-tioned left engine bleed air is supplied directly to the cabinthrough the emergency supply tube. The emergency supply

tube terminates within a mixing tube below the aft passen-ger cabin floor where it forms an ejector nozzle. The ejector

nozzle produces a suction force that opens a check valvethrough which cabin air is drawn into the mixing tube. Thecabin air mixes with and reduces the temperature of thebleed air prior to entering the distribution system. Additionalbleed air temperature reduction is provided by the emer-gency supply tube itself, which features a “beaded” or“spiral” exterior that increases surface area to maximizeheat transfer. A check valve installed in the aft pressure

bulkhead prevents cabin pressure backflow through the

emergency supply tube during normal pressurized opera-tion.

Note: When emergency pressurization is selected, the ACMand temperature control system are disabled. Limited control ofcabin temperature may be accomplished using the leftTHROTTLE lever to regulate bleed air flow; however, excessive

engine power reduction can cause an increase in cabin alti-tude.

Note: Emergency pressurization is automatically activatedwhen the temperature of bleed air flow through the ACM ex-ceeds approximately 435°F. Refer to the ACM Overheat Warn-ing and Protection section of this chapter for a complete de-scription of ACM overheat protection.



E N V I R O N M E N T A L S Y S

T E M 1 2 / 9 9


9 - 5

© P C W




Air Cycle Machine and Temperature Control System

(550-0482, 550-0485 and after)

Bleed air is cooled, mixed with uncooled bleed air, and dehu-

midified to provide conditioned air with the desired temperatureto the air distribution subsystem. Major components of the

system include an air cycle machine (ACM) containing aprecooler, primary and secondary heat exchangers and acooling turbine, a water separator, a bypass modulating valve,a water ejector nozzle, a fan, and necessary ducting.

Air Cycle Machine

From the environmental control/shutoff valves or the groundshutoff valve, bleed air is supplied to the ACM, directed to thebypass modulating valve, and passes through the precooler.From the precooler, bleed air passed through the primary heat

exchanger and is cooled by heat transfer. After passingthrough the primary heat exchanger, the bleed air is suppliedto the cooling turbine. The cooling turbine essentially consistsof an impeller-type compressor and a turbine, mounted on thesame shaft. The shaft rotates at approximately 80,000 RPM and

its bearings are lubricated by oil drawn by wicks from a sumpmounted on the turbine housing. A fan, external to the coolingturbine and used to circulate ambient air for cooling, is alsomounted on the shaft.

© PCW





In the cooling turbine, bleed air first enters the compressor,where its pressure and temperature are increased. The air isthen directed to the secondary heat exchanger, where it is

cooled again by heat transfer. Water, drawn by the water ejec-tor nozzle from the water separator, is sprayed over the sec-ondary heat exchanger to provide additional cooling. A small

amount of high velocity air from the secondary heat exchangeroutput line is used to create the suction required to draw thewater from the water separator.

From the secondary heat exchanger, the compressed air is

directed to the turbine where its temperature and pressure arerapidly reduced by expansion. From the turbine outlet, thissuper-cooled air is passed through a mixing tube where it ismixed with hot bleed air supplied through the bypass modulat-ing valve. The electrically controlled and operated bypass

modulating valve is located in a bypass duct connected be-tween the bleed air inlet and the mixing tube at the outlet sideof the cooling turbine. The valve functions to control the tem-perature of the conditioned air by opening and closing to

modulate the flow of hot bleed air to the mixing tube.

From the mixing tube, the conditioned air passes through thewater separator, which collects moisture from the passing airand forms large droplets that are removed by centrifugal force.This removed moisture is drawn away by the water ejectornozzle previously described. An integral spring-loaded relief

valve allows air to bypass the unit should the water separatorbecome obstructed by ice or foreign material.

Cooling air for the precooler, primary, and secondary heatexchangers is drawn from within the aft fuselage by the fan thatis driven by the cooling turbine shaft. After passing over theheat exchangers, the cooling air is exhausted through anoverboard vent on the lower surface of the aft fuselage.




Temperature Control System

Temperature control is primarily achieved by varying theamount of hot bleed air that the bypass modulating valve mixeswith cold air from the ACM. Cabin temperature may be setmanually or controlled automatically as desired.

Automatic

Temperature

Selection

Automatic tempera-ture selection isaccomplished usingthe AUTOMATICtemperature controlknob on the environ-mental panel. Rotat-

ing the knob clock-wise for a higher cabin temperature or counterclockwise for a

lower cabin temperature operates a potentiometer that estab-lishes a reference voltage corresponding to the selected tem-perature. The reference voltage is supplied to the temperaturecontrol computer, where it is compared with signals from theduct temperature sensor (ACM conditioned air temperature)and the cabin temperature sensor (actual cabin temperature).The temperature control computer then generates a signal todrive the bypass modulating valve open or closed to maintain

the desired cabin temperature. The bypass modulating valve

receives power from the left main bus through the 5-amp TEMPcircuit breaker.

Manual Temperature Selection

The manual temperature selection mode is selected by rotatingthe AUTOMATIC temperature control knob fully counterclock-wise till it clicks into the “MANUAL” position. In this mode, the

three-position manual mode toggle switch controls the positionof the bypass modulating valve. The switch is spring-loaded tothe “OFF” position. When held in the “MANUAL HOT” position,the bypass valve moves toward open, allowing more hot air tomix with the cooled air. When released, it returns to the “OFF”position, but the bypass valve remains in the selected position.When held in the “MANUAL COLD” position, the bypass valvemoves towards closed. The manual mode toggle switch is only

usable when the AUTOMATIC temperature control knob is inthe “MANUAL” position.

1. Automatic Temperature Control Knob

2. Manual Mode Toggle Switch

12





Temperature Control System Indication and Warning

When conditioned air temperature at the water separator outlet(550-0550 and after) or inlet (550-0482 and 0485 ~ 0505)exceeds approximately 315°F, a temperature switch functionsto illuminate the amber [AIR DUCT O’HEAT] annunciator. Illumi-nation of this annunciator indicates that corrective action

should be taken to lower the cabin temperature in order toavoid duct damage.

Following [AIR DUCT O’HEAT] annunciator illumination, theTEMP circuit breaker should be pulled and reset, and manualtemperature control mode should be selected. The manualmode toggle switch should be held in the “MANUAL COLD”position until the [AIR DUCT O’HEAT] annunciator extinguishesAutomatic temperature control should then be reselected. If the[AIR DUCT O’HEAT] annunciator reilluminates, cabin tempera-

ture should be controlled manually for the remainder of theflight.

ACM Overpressure Warning and Protection

To protect the ACM from overpressurization, a primary andsecondary pressure switch are installed in the environmentalsupply tubing upstream and downstream of the ground shutoffvalve respectively.

During ground operations with the PRESS SOURCE selectorswitch in the “GND” position, when bleed air pressure reaches

approximately 38 PSI, the primary pressure switch functions toclose the ground shutoff valve and extinguish the [BLD AIRGND] annunciator. In this condition, retarding the rightTHROTTLE lever below approximately 72% N2 should reducebleed air pressure sufficiently to cause the valve to reopen andthe [BLD AIR GND] annunciator to illuminate.

Should the primary pressure switch fail, the secondary pres-sure switch will activate when bleed air pressure reachesapproximately 42 PSI, functioning to close the ground shutoffvalve, extinguish the [BLD AIR GND] annunciator, and illumi-nate the amber [ACM O’PRESS] annunciator. In this condition,the valve will remain closed and the [ACM O’PRESS] annuncia-tor will remain illuminated regardless of N2/bleed air pressure

reduction or PRESS SOURCE selector switch position until thepressure switch circuit is restored. This may be accomplishedby pulling and resetting the NORM PRESS circuit breaker;however, the cause of the malfunction should be identified andrepaired before resuming flight operations.




Note: Illumination of both the [ACM O’PRESS] and [BLD AIRGND] annunciators may indicate that the primary pressureswitch has closed, but the ground shutoff valve is stuck open.Should this occur, the cause of the malfunction should beidentified and repaired before resuming flight operations.

Should the ground shutoff valve malfunction and open duringflight, the [BLD AIR GND] and [ACM O’PRESS] annunciatorswill both be illuminated. In this condition, the PRESS SOURCEselector switch should be set to “RH” to close the left flowcontrol shutoff valve, and the right THROTTLE lever should beretarded below 80% N2 to reduce bleed air pressure.

Failure of the [ACM O’PRESS] annunciator to extinguish may

indicate that it was illuminated by the secondary pressureswitch because the primary pressure switch failed to close. In

this condition, the [ACM O’PRESS] annunciator will remainilluminated until the pressure switch circuit is restored. Duringflight this may be accomplished by first pulling the EMERPRESS circuit breaker to prevent inadvertent activation ofemergency pressurization, pulling and resetting the NORMPRESS circuit breaker, and then resetting the EMER PRESS

circuit breaker.

Note: If the [ACM O’PRESS] annunciator remains illuminated,the PRESS SOURCE selector switch should remain in the “RH”position, right engine N2 should remain below 80%, and the left

engine should be operated normally for the duration of theflight. After landing, the cause of the malfunction should beidentified and repaired before resuming flight operations.

ACM Overheat Warning and Protection

To protect the ACM from overheating, an overheat sensor isinstalled in the bleed air tube between the ACM compressoroutlet and secondary heat exchanger inlet.

During flight with the PRESS SOURCE selector switch in the“NORMAL” position, when bleed air temperature exceedsapproximately 435°F, the overheat sensor functions to close

both environmental flow control/shutoff valves, open the emer-gency pressurization valve, and illuminate the amber [EMERGPRESS ON] annunciator. Power is supplied to the valves andthe annunciator through the EMER PRESS circuit breaker. Inthis condition, bleed air flow through the ACM is interruptedand unconditioned left engine bleed air is supplied directly tothe cabin through the emergency pressurization valve.





If bleed air temperature falls below approximately 405°F within

12-seconds of [EMERG PRESS ON] annunciator illumination,normal system operation will be automatically restored. If nor-mal system operation is not automatically restored within thisperiod of time, an emergency pressurization lockout relay willbe energized through the EMER PRESS circuit breaker. With

this relay energized, both environmental flow control/shutoffvalves will remain closed, the emergency pressurization valvewill remain open, and the [EMERG PRESS ON] annunciator will

remain illuminated until the overheat circuit is reset. This maybe accomplished by rotating the PRESS SOURCE selectorswitch to the “EMER” position to deenergize the emergencylockout relay, waiting one minute, and then reselecting the“NORMAL” position to restore normal operation.

Note: If the [EMERG PRESS ON] annunciator remains illumi-

nated, the PRESS SOURCE selector switch should be set to the“EMER” position, the right engine should be operated normally,

and the left THROTTLE lever should be used to control cabintemperature for the duration of the flight. After landing, thecause of the malfunction should be identified and repairedbefore resuming flight operations.

During ground operations, the overheat sensor functions as itdoes in flight; however, the left main gear safety switch pre-vents the emergency pressurization valve from opening when

the airplane is on the ground. During ground operations with

the PRESS SOURCE selector switch in the “GND” position, theoverheat sensor functions to close the ground shutoff valve andextinguish the [BLD AIR GND] annunciator.

As in flight, system operation will be automatically restored ifbleed air temperature falls below approximately 405°F within12-seconds of [EMERG PRESS ON] annunciator illumination. If

system operation is not automatically restored within this periodof time, the overheat circuit must be reset by rotating thePRESS SOURCE selector switch to the “EMER” position, wait-ing one minute, and then reselecting the previous position torestore operation.

Note: If the [EMERG PRESS ON] annunciator remains illumi-nated, the cause of the malfunction should be identified and

repaired before resuming flight operations.




Note: Automatic activation of emergency pressurization may

occur under such conditions as low airspeed climbs at highaltitudes with a low cabin temperature selected. Should thisoccur, increasing airspeed and selecting a higher cabin tem-perature after restoring normal operation should prevent reacti-vation of emergency pressurization.

The conditioned and dehumidified air is routed to the distribu-tion system described later in this chapter.





Pressurization Air Source and Selection

(550-0484, 0483, 0481 and earlier)

Bleed air used for cabin pressurization and ventilation is ex-tracted from the compressor section of each engine throughports located at 4 and 8 o’clock positions on the gas generator

case. From these ports, bleed air is primarily routed throughindependent supply tubes to a precooler installed within eachengine nacelle. The precoolers reduce the temperature ofengine bleed air supplied to various airplane systems, includ-ing the ACM located within the aft fuselage. An environmental

flow control/shutoff valve in each of these tubes controls theflow of bleed air from each source to the ACM and functions asa check valve to prevent the backflow of bleed air to the oppo-site source when either engine has failed or is operating at asufficiently lower RPM. A ground shutoff/pressure regulating

valve bypasses the right environmental flow control/shutoffvalve, permitting increased right engine bleed air flow throughthe ACM to enhance cabin ventilation during ground operation.During normal operation, all bleed air flows through the ACM

en route to the cabin. In an emergency, left engine bleed airmay be supplied directly to the cabin through an emergencysupply tube and pressurization valve.

The environmental flow control/shutoff valves each have anominal flow rate of approximately 6 pounds per minute (PPM)and a maximum flow rate of approximately 9 PPM. Flow rate is

controlled by a primary solenoid and secondary solenoidintegral to each valve. Both solenoids are normallydeenergized; nominal flow rate occurs in this condition. Maxi-mum flow rate occurs when the primary solenoid isdeenergized and the secondary solenoid is energized.The ground shutoff/pressure regulating valve is motor-oper-ated, electrically-actuated open and closed, has a nominal flowrate of approximately 18 PPM, and incorporates an indepen-dent pressure relief valve.

The emergency pressurization valve is normally-closed andelectrically-actuated open.




Nacelle Precooler System

The nacelle precooler system functions to regulate the tem-perature of the bleed air used by airplane systems to approxi-mately 500°F. This is accomplished by routing bleed air fromthe engine through a heat exchanger, where it is cooled bybypass air. The primary components of the nacelle precooler

system are heat exchangers, temperature sensor valves, tem-perature control valves, and tubular plumbing.

The temperature sensor valves monitor engine bleed air tem-perature and send pneumatic signals to open and close thebleed air temperature control valves. When a temperaturesensor valve senses that bleed air temperature is more thanapproximately 500°F, it sends a signal to open the associatedtemperature control valve and allow more bypass air to passover the heat exchanger, lowering the temperature of bleed air

supplied to distribution tubes within the aft fuselage. Con-versely, when bleed air temperature is less than 500°F, the

signal from the temperature sensor valve allows the tempera-ture control valve to close, raising the temperature of the bleedair supplied to the environmental and other airplane systems.

Overheat switches are located in the bleed air tubes betweenthe nacelle precoolers and the environmental flow control/ shutoff valves. The switches will operate at a temperature ofapproximately 540°F and cause the amber [L PRECOOL FAIL]

or [R PRECOOL FAIL] annunciator to illuminate, indicating

excessively hot bleed air from the respective precooler.





Source Selection

All four valves are controlledprimarily by the PRESS

SOURCE selector switch onthe environmental “tilt” panel.The valves are also controlled

by various switches that sensebleed air pressure and tem-perature. The effects of rotat-ing the selector switch to eachof its seven positions are

described below:

NORMAL Position

During normal operation with both engines operating andthe PRESS SOURCE selector switch set to “NORMAL,” both

environmental flow control/shutoff valves will be open,permitting bleed air flow through the ACM and into thecabin at a rate of approximately 12 PPM.

LH and RH Positions

When set to “LH” or “RH,” bleed air is correspondinglysupplied by the left engine or right engine only at a rate ofapproximately 6 PPM. In this condition, the environmentalflow control/shutoff valve for the non-selected source isenergized closed by 28 VDC left main bus power throughthe 5-amp NORM PRESS circuit breaker on the left CB

panel

BOTH HI Position

When set to “BOTH HI”, both environmental flow control/ shutoff valves remain open and both secondary solenoidsare energized open. In this condition, approximately 18PPM of bleed air from both engines flows through the ACMinto the cabin and the amber [BLEED AIR GND/HI] annun-ciator is illuminated through the NORM PRESS circuit

breaker.

OFF Position

When set to “OFF,” both environmental flow control/shutoffvalves are energized closed through the NORM PRESScircuit breaker.




GND Position

To enhance cabin ventilation during ground operation,primarily when the right engine is operating only, the PRESSSOURCE selector switch should be set to “GND.” In thisposition, both environmental flow control/shutoff valves areenergized closed, the ground shutoff/pressure regulating

valve is energized open, and the amber [BLEED AIR GND/ HI] annunciator is illuminated through the NORM PRESScircuit breaker. With the ground shutoff/pressure regulating

valve open, right engine bleed air flows through the ACMand into the cabin at a rate of approximately 18 PPM.

Selection of this source is inhibited by the left main gearsafety switch when the airplane is in flight.

EMER Position

Should bleed air flow through the ACM be insufficient tomaintain selected cabin altitude, the PRESS SOURCE selec-

tor switch should be set to “EMER.” In this position, theamber [EMER PRESS ON] annunciator is illuminated, bothenvironmental flow control/shutoff valves are energizedclosed, and the emergency pressurization valve is ener-gized open. 28 VDC right main bus power is supplied to thevalves and the annunciator through the 5-amp EMERPRESS circuit breaker on the left CB panel. Selection of thissource is inhibited by the left main gear safety switch when

the airplane is on the ground.

With the emergency pressurization valve open, uncondi-tioned left engine bleed air is supplied directly to the cabinthrough the emergency supply tube. The emergency supplytube terminates within a mixing tube below the aft passen-ger cabin floor where it forms an ejector nozzle. The ejectornozzle produces a suction force that opens a check valve

through which cabin air is drawn into the mixing tube. Thecabin air mixes with and reduces the temperature of thebleed air prior to entering the distribution system. Additionalbleed air temperature reduction is provided by the emer-gency supply tube itself, which features a “beaded” or“spiral” exterior that increases surface area to maximizeheat transfer. A check valve installed in the aft pressurebulkhead prevents cabin pressure backflow through the

emergency supply tube during normal pressurized opera-tion.





Note: When emergency pressurization is selected, the ACMand temperature control system are disabled. Limited control ofcabin temperature may be accomplished using the left

THROTTLE lever to regulate bleed air flow; however, excessiveengine power reduction can cause an increase in cabin alti-tude.

Note: Emergency pressurization is automatically activatedwhen the temperature of bleed air flow through the ACM ex-ceeds approximately 435°F. Refer to the ACM Overheat Warn-ing and Protection section of this chapter for a complete de-

scription of ACM overheat protection.




Air Cycle Machine and Temperature Control System

(550-0484, 0483, 0481 and earlier)

Bleed air is cooled, mixed with uncooled bleed air, and dehu-midified to provide conditioned air with the desired temperatureto the air distribution subsystem. Major components of the

system include an air cycle machine (ACM) containing primaryand secondary heat exchangers and a cooling turbine, a waterseparator, a bypass modulating valve, a water ejector nozzle, a

bleed air ejector solenoid valve, a bleed air ejector, and neces-sary ducting.

Air Cycle Machine

From the environmental control/shutoff valves or the groundshutoff/pressure regulating valve, bleed air is supplied to theACM and directed to the bypass modulating valve and bleed

air ejector nozzle supply tube. Within the ACM, the bleed airpasses through the primary heat exchanger and is cooled by

heat transfer. After passing through the primary heat ex-changer, the bleed air is supplied to the cooling turbine. Thecooling turbine essentially consists of an impeller-type com-pressor and a turbine, mounted on the same shaft. The shaftrotates at approximately 80,000 RPM and uses airfoil-typebearings that require no lubrication.

© PCW





In the cooling turbine, bleed air first enters the compressor,where its pressure and temperature are increased. The air isthen directed to the secondary heat exchanger, where it is

cooled again by heat transfer. Water, drawn by the water ejec-tor nozzle from the water separator, is sprayed over the sec-ondary heat exchanger to provide additional cooling. A small

amount of high velocity air from the secondary heat exchangeroutlet line is used to create the suction required to draw thewater from the water separator.

From the secondary heat exchanger, the compressed air is

directed to the turbine, where its temperature and pressure arerapidly reduced by expansion. From the turbine, this super-cooled air is passed through a mixing tube where it is mixedwith hot bleed air supplied through the bypass modulatingvalve. The pneumatically controlled and operated bypass

modulating valve is located in a bypass duct connected be-tween the primary heat exchanger inlet and the mixing tube atthe outlet side of the cooling turbine. The valve functions tocontrol the temperature of the conditioned air by opening and

closing to modulate the flow of hot bleed air to the mixing tube.

From the mixing tube, the conditioned air passes through awater separator, which collects moisture from the passing airand forms large droplets that are removed by centrifugal force.This removed moisture is drawn away by the water ejectornozzle previously described. An integral spring-loaded relief

valve allows air to bypass the unit should the water separatorbecome obstructed by ice or foreign material.




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A duct routes cooling air to the

ACM primary and secondaryheat exchangers from aNACA-type scoop on thedorsal fairing (dorsal scoop).After passing over the heat

exchangers, the air is ex-hausted through an overboardvent on the lower surface of

the aft fuselage. In flight,sufficient air is available due toram effect. During groundoperations, ram effect is notavailable to move air over the heat exchangers. To compen-sate, a bleed air ejector nozzle is installed within the heatexchanger exhaust duct. Bleed air is admitted to the bleed air

ejector nozzle by the solenoid valve installed in the ejectornozzle supply tube.

The ejector nozzle directs bleed air toward the ambient airexhaust outlet. Bleed air flow exiting through the exhaust outletcreates a suction that draws ambient air through the heatexchangers. Bleed air and conditioned air are then exhaustedoverboard through an exhaust outlet below the right enginepylon.

The shutoff valve is enabled by the left main gear safety switch

when the airplane is on the ground, and is disabled by thesafety switch in flight. To maximize engine power during take-off, throttle position switches, brake switches, and a differentialpressure switch act together to close the ejector nozzle shutoffvalve and thereby disable the ACM bleed air ejector nozzleduring takeoff roll. The amber [ACM EJECTOR ON] annunciatorilluminates when the ejector nozzle shutoff valve is open.

DORSAL SCOOP




ACM Overpressure Protection

The ACM is protected from overpressurization by a relief valveincorporated into the ground shutoff/pressure regulating valve.The relief valve, which functions independently of the ground

shutoff/pressure regulating valve, opens at approximately 50PSIG, releasing excess pressure overboard through a pneu-

matic tube. The valve reseats at approximately 40 PSIG.

ACM Overheat Warning and Protection

To protect the ACM from overheating, an overheat sensor isinstalled in the bleed air duct between the ACM compressoroutlet and secondary heat exchanger inlet.

During flight with the PRESS SOURCE selector switch in the“NORMAL” position, when bleed air temperature exceedsapproximately 435°F, the overheat sensor functions to close

both environmental flow control/shutoff valves, open the emer-gency pressurization valve, and illuminate the amber [EMERPRESS ON] annunciator. Power is supplied to the valves andthe annunciator through the EMER PRESS circuit breaker. In

this condition, bleed air flow through the ACM is interruptedand unconditioned left engine bleed air is supplied directly tothe cabin through the emergency pressurization valve.

If bleed air temperature falls below approximately 405°F within12-seconds of [EMER PRESS ON] annunciator illumination,normal system operation will be automatically restored. If nor-

mal system operation is not automatically restored within thisperiod of time, an emergency pressurization lockout relay willbe energized through the EMER PRESS circuit breaker. Withthis relay energized, both environmental flow control/shutoffvalves will remain closed, the emergency pressurization valvewill remain open, and the [EMER PRESS ON] annunciator willremain illuminated until the overheat circuit is reset. This maybe accomplished by rotating the PRESS SOURCE selectorswitch to the “EMER” position to deenergize the emergency

lockout relay, and then reselecting the “NORMAL” position torestore normal operation.

Note: If the [EMER PRESS ON] annunciator remains illumi-nated, the PRESS SOURCE selector switch should be set to the“EMER” position, the right engine should be operated normally,and the left THROTTLE lever should be used to control cabintemperature for the duration of the flight. After landing, the

cause of the malfunction should be identified and repairedbefore resuming flight operations.





During ground operations, the overheat sensor functions as itdoes in flight; however, the left main gear safety switch pre-vents the emergency pressurization valve from opening when

the airplane is on the ground. During ground operations withthe PRESS SOURCE selector switch in the “GND” position, theoverheat sensor functions to close the ground shutoff/pressure

regulating valve.

As in flight, system operation will be automatically restored ifbleed air temperature falls below approximately 405°F within12-seconds of [EMER PRESS ON] annunciator illumination. If

system operation is not automatically restored within this periodof time, the overheat circuit must be reset by rotating thePRESS SOURCE selector switch to the "OFF" or “EMER” posi-tion and then reselecting the previous position to restore opera-tion.

Note: If the [EMER PRESS ON] annunciator remains illumi-nated, the cause of the malfunction should be identified andrepaired before resuming flight operations.

Note: Automatic activation of emergency pressurization mayoccur under such conditions as low airspeed climbs at highaltitudes with a low cabin temperature selected. Should thisoccur, increasing airspeed and selecting a higher cabin tem-perature after restoring normal operation should prevent reacti-vation of emergency pressurization.




Temperature Control System

Temperature controlis primarily achievedby varying theamount of hot bleed

air that the bypassmodulating valvemixes with cold airfrom the air cyclemachine. Cabintemperature may beset manually orcontrolled automati-cally. In either mode, pneumatic pressure is used to open the

normally closed bypass modulating valve and allow hot bleedair to mix with cold air from the ACM. Major components of thetemperature control system are a selector switch, the bypassmodulating valve, a solenoid shutoff valve, an automatic tem-perature control pressure regulator, automatic temperaturecontrol knob and manual temperature control knob, cabintemperature, supply temperature, and low limit sensors, and anair duct temperature switch.

Automatic Mode

Automatic temperature control mode selection is accomplishedby positioning the selector switch on the environmental panel to

“AUTOMATIC.” When automatic mode is selected, the solenoidshutoff valve energizes, supplying 15 PSI from the automaticcontrol pressure regulator, located on the right side of the aftfuselage, to the automatic cabin temperature selector, the low

limit sensor, and the cabin temperature sensor. These sensors,the supply duct temperature sensor, and the temperatureselector interact to develop a control pressure that causes thebypass modulating valve to open when sensed cabin tempera-ture is too low or too close when sensed cabin temperature istoo high. When duct temperature is less than approximately35°F, the low limit sensor causes the bypass valve to open,

raising duct temperature regardless of cabin temperature orselected temperature.

1. Manual Temperature Control Knob

2. Automatic Temperature Control Knob

3. Selector Switch

21

3





Note: The automatic temperature control operates on bleed airfrom the right engine only; the manual system operates onbleed air from both engines. Therefore, the automatic tempera-

ture control system cannot be used unless the right engine isoperating.

Manual Mode

When the cabin temperature selector switch is positioned to“MANUAL,” the temperature solenoid shutoff valve isdeenergized, routing 23 PSI manual control pneumatic pres-sure to the bypass modulating valve and manual temperature

selector. The manual temperature control knob operates aninternal spring-loaded poppet, which determines the amount ofcontrol pressure that is allowed to bypass the poppet and ventoverboard through the forward pressure bulkhead. This actionvaries the control pressure applied to the bypass modulating

valve, thereby increasing or decreasing cabin temperature aswith the automatic system. Therefore, in manual control mode,compensations for changes in cabin temperature must beachieved by rotating the manual temperature control knob. The

manual mode is available as a backup should the automaticsystem fail.

Note: When in manual mode, the low-limit sensor is inoperativeand ice formation in the water separator/ducting is possible. Iceformation is more likely if the outside relative humidity exceeds40%. When operating in manual mode, the ACM system should

be carefully monitored to detect overheating.

Note: The design of the temperature control system is such thatinsufficient bleed air pressure will cause the temperature con-trol system to operate fully cold.

Conditioned Air Temperature Warning and Protection

Warning of excessive temperature of conditioned air is pro-vided by an amber [AIR DUCT O’HEAT] annunciator that illumi-

nates to alert the flight crew of a conditioned air overheat. Anair duct overheat switch, installed in the air supply duct down-stream from the water separator outlet, closes at approximately315°F and causes the annunciator to illuminate.




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Air Distribution System

The air distribution system directs and controls the flow ofpressurization and ambient (fresh) air to the pressurized part ofthe fuselage. Ambient air is used to augment cabin airflowduring ground operations. Major components of the air distribu-

tion system include distribution control devices, blowers, distri-bution ducting, and air outlets.

Distribution Control

Pressurization and ambient air are introduced into the cabinthrough flapper-type check valves that function to prevent lossof cabin pressurization through the fresh air duct or duringoperation of the emergency pressurization system. After pass-ing through the check valves, air enters a ventilation junctionbox, located below the aft passenger cabin floor at the aftpressure bulkhead.

The ventilation junction box functions to control the source ofair directed to the overhead duct according to the temperatureof the conditioned pressurization air entering the cabin. Theventilation junction box contains a swing-type door,thermoswitch, temperature motor, two limit switches, and theoverhead blower. Operation of the overhead blower is de-scribed later in this chapter.

When pressurization air temperature exceeds approximately

100°F, the thermoswitch closes an electrical circuit, causingthe temperature motor to close the swing-type door. With thedoor closed, pressurization air is prevented from entering theoverhead duct and recirculated cabin air is admitted. Whenpressurization air temperature falls below 100°F, thethermoswitch deenergizes and causes the temperature motor

to run in the opposite direction, opening the door, and therebyreadmitting pressurization air into the overhead ducts. Doortravel is controlled by limit switches that turn off the tempera-ture motor when the door is fully open or fully closed. The motoreceives 28 VDC power from the left (550-482, 550-485 and

after) or right (550-484, 550-483, 550-0481 and earlier) mainbus through the 5-amp TEMP circuit breaker.




A flexible air duct, connected to the ventilation junction box

blower inlet, draws air from the aft baggage compartment to becirculated through the overhead duct network. This arrange-ment allows the aft divider door to be kept closed, yet stillprovides a means for the crew and passengers to smell smokeoriginating in the aft baggage compartment earlier than would

otherwise be the case.

From the ventilation junction box, conditioned air is passed

through the emergency pressurization mixing tube. Duringemergency pressurization, hot engine bleed air used to pres-surize the cabin is released into the mixing tube ejector nozzle.The ejector nozzle produces a suction force that opens acheck valve through which cabin air is drawn into the mixingtube. The cabin air mixes with and reduces the temperature ofthe bleed air prior to entering the distribution system.

From the mixing

tube, cabin airenters a flow divider.The flow divider,located below thecabin floor, containsa divider vane andmotor, and functionsto apportion air

between the flight

compartment andthe passengercabin. Proportions ofconditioned airdelivered to thecockpit and cabinare controlled using

CKPT/CABIN control, located on the environmental panel.Rotating this switch energizes the motor to move the flow di-vider vane. Clockwise rotation increases cabin airflow;counterclockwise rotation increases cockpit airflow.

CKPIT/CABIN CONTROL





Blowers

Two impeller-type blowers areused to enhance cabin air

circulation. Both blowersreceive 28 VDC power fromthe left main bus through the

20-amp CABIN FAN circuitbreaker.

The overhead blower is anintegral part of the ventilation

junction box and functions toforce pressurization air, recir-culated air, or ambient airthrough the overhead duct.The overhead blower is controlled by the three-position (HI/

OFF/LOW) FAN OVHD switch on the copilot’s lower instrumentpanel.

The defog blower is located in the underfloor ducting down-

stream from the flow divider and functions primarily to increasethe flow of conditioned air to the forward part of the cabin. Thedefog blower is controlled by the three position (HI/OFF/LOW)FAN DEFOG switch on the copilot’s lower instrument panel.

1. Overhead Blower Switch

2. Defog Blower Switch

1 2




Distribution Ducting and Air Outlets

Within the cabin interior, distribution ducting is located in theoverhead and at floor and armrest level. Additional distributionducting is located under the cabin floorboards.

Conditioned pressurization air

below 100°F, recirculated air,or ambient air is taken fromthe distribution junction box

and circulated through theoverhead duct to adjustableoutlets in the passenger cabinand flight compartment. Pas-senger air outlets are locatedover each seat position andare fully adjustable from no

flow to maximum flow. Air fromthe aft baggage compartment

is also supplied to the overhead duct to maximize the crew’sability to detect smoke in the baggage compartment. Becauseof this method of smoke detection, air outlets in the flight com-partment are configured so they can not be fully shut off. Air-planes with optional vapor cycle air conditioning systems usethe overhead duct to circulate air conditioned air; these air-planes are equipped with overhead outlets that have higherflow rates.

Air from the upper branch of the flow divider is routed to a mainand auxiliary plenum. The main plenum supplies air to thefootwarmer and armrest manifolds on the left side of the pas-senger cabin. The auxiliary plenum supplies the correspondingright side components. Passenger footwarmer manifolds arelocated along the outboard cabin walls at floor level. Thefootwarmer manifolds are assembled in segments, each seg-

ment including several outlet holes. The armrest manifolds arelocated along the outboard cabin walls at passenger seatarmrest level. Air outlet holes are located beneath the armrestcover assemblies. The armrest and footwarmer manifolds areconnected together by ducting at several points.

PASSENGER AIR OUTLET





Air from the lower branch of the flow divider is drawn throughthe defog fan and routed through underfloor ducting to theforward part of the pressurized cabin. In the flight compart-

ment, conditioned air is routed to cockpit footwarmer manifoldslocated on both sides of the cockpit, to the windshield and sidewindow defog vents, and to optional supplementary ducts. In

all airplanes, the underfloor ducting supplies air to thefootwarmer and armrest warmer associated with the passengerseat immediately forward of the main entrance door. On air-planes 550-0482 and 550-0485 and after, the underfloor duct-ing also supplies air to warm the cabin door seal. On these

airplanes and in airplanes 550-0010 through 550-0049 notincorporating SB550-21-1, the underfloor ducting is also con-nected to the left footwarmer/armrest manifold by a duct lo-cated aft of the main entrance door and to the right footwarmer/armrest manifold by another duct located immediately aft of the

flight compartment divider.

Cabin Pressurization Control System

The pressurization control system provides for passengercomfort by allowing the selection of a desired cabin altitudeand rate-of-change during ascent or descent. The cabin ispressurized using engine bleed air as previously described.Cabin pressurization is regulated using outflow valves thatopen to allow pressurized air to exit the cabin, raising cabinaltitude; and close to retain pressurized air in the cabin, lower-

ing cabin altitude. Major components of the system include twocabin outflow valves, a pneumatic relay, two cabin altitude limitvalves, a depressurization (dump) toggle valve, a pressuriza-tion source selection system, cabin altitude controls and indicators, and associated circuitry and plumbing.

Cabin Outflow Valves

Two cabin outflow valves, both mounted on the aft pressurebulkhead below the passenger cabin floor, vent pressurization

air overboard to maintain the selected cabin altitude or pres-sure differential in reference to the ambient air pressure. Themaximum pressure differential of 8.8 PSI is primarily deter-mined by the structural limitations of the airplane’s pressurizedcenter section. The system is designed to maintain a cabinaltitude of 8,000 feet at airplane altitudes of up to 43,000 feet.




Both cabin outflow valves function primarily to regulate theoutflow of pressurization air from the cabin based on pneumaticsignals received from the controller through the pneumaticrelay. These signals establish a reference pressure that isapplied to one side of an internal poppet valve. The other sideof the poppet valve is exposed to actual cabin pressure. The

poppet valve is spring-loaded closed and suspended by aflexible diaphragm between these two pressures such that theoutflow of pressurization air increases when cabin pressureexceeds reference pressure, and decreases when referencepressure exceeds cabin pressure. In this way, the outflow valvemodulates to maintain the selected cabin altitude or to effect adesired change in cabin altitude at a selected rate.

Positive pressure relief and negative pressure relief functionsare also provided by the outflow valves. Both functions override

the controller and the pneumatic relay.

Positive pressure relief is provided by a Schrader-type valvesuspended by a flexible diaphragm between the referencepressure control chamber and an ambient pressure chamberthat is vented to the atmosphere. The valve is spring-loaded

closed and factory preset to open when the differential be-tween reference pressure and ambient pressure exceeds thenominal differential of approximately 8.6 PSID. Should thisoccur, the release of control pressure to the atmospherethrough the open valve would allow the poppet valve to modu-

late toward open, increasing pressurization air outflow. In thiscondition, cabin pressure is maintained at the nominal differen-tial and cabin rate-of-change follows that of the airplane.

Protection against exceeding the airplane’s negative pressurestructural limits is provided by a flexible diaphragm exposed tocabin pressure on one side and ambient pressure on the other.Should ambient pressure exceed cabin pressure, as in duringrapid descent, this diaphragm would raise and lift the poppetvalve open allowing ambient pressure to enter the cabin untilboth pressures become approximately equal.



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Cabin Pressurization

Controller

The cabin pressurizationcontroller, located on theenvironmental panel, incorpo-rates two knobs: one for

selecting cabin altitude andthe other for selecting cabinrate-of-change. The cabin

altitude selector features anouter CABIN scale and aninner ACFT scale, both cali-brated in feet X 1000 and visible through a window at the topcenter of the selector face. Both scales rotate simultaneouslyas the selector is rotated. The selected altitude is indicated bythe alignment of each scale with the twelve o’clock positionrelative to the selector. The CABIN scale indicates the cabin

altitude the controller is set to maintain. The ACFT scale indi-cates the maximum altitude to which the airplane may ascendwithout causing the selected cabin altitude to be exceeded.The cabin rate selector is marked with an arrow for positionreference only.

Cabin Pressurization Controller

The pressurization controller establishes desired cabin altitude

and rate of climb by modulating reference air pressure to thepneumatic relay. The controller body is divided into three

chambers: cabin pressure, rate pressure, and reference pres-sure. Cabin air enters the cabin pressure chamber through afiltered orifice. The cabin pressure chamber houses an abso-lute bellows. Rotating the cabin altitude selector mechanicallycompresses or extends this bellows to a position that sets thecontroller to maintain the selected cabin altitude.

The rate pressure chamber houses a rate spring secured to theabsolute bellows on one side and a rate diaphragm on theother. The rate diaphragm separates the rate chamber from thereference chamber. Airflow passage between these two cham-

bers is regulated by a needle valve that sets the controller toprovide the desired cabin rate-of-change according to theposition of the CABIN RATE selector.





Clockwise rotation of the CABIN RATE selector decreasesairflow through the needle valve and increases cabin rate-of-change; counterclockwise rotation increases airflow through

the needle valve and decreases cabin rate-of-change. In thetwelve o’clock position, indicated cabin rate-of-change shouldbe approximately 500 FPM.

The rate chamber is ported to a small tank, installed below theairplane floor, which supplies auxiliary volume to the rate cham-ber to provide greater accuracy in cabin rate-of-change con-trol. Should rate pressure exceed cabin pressure, a check

valve will permit airflow from the rate chamber to the cabinpressure chamber.

The reference pressure chamber houses a metering valve andfollower spring linked to the rate diaphragm. The chamber is

ported to cabin pressure, the airplane suction supply, and thecabin outflow valve. Cabin pressure enters the reference cham-ber through a filtered orifice. The metering valve regulates theflow of cabin pressure to airplane suction to produce the refer-

ence pressure, which is then applied to the pneumatic relay.When the metering valve is modulating towards closed, the flowof cabin pressure to airplane suction is reduced and referencepressure is increased. Conversely, when the metering valve ismodulating towards open, the flow of cabin pressure to air-plane suction is increased and reference pressure is reduced.

With a desired cabin altitude and rate-of-change selected,changes in cabin pressure cause the absolute bellows toexpand or contract. As it does, the metering valve is reposi-tioned to maintain the correct reference pressure. Airflow be-tween the rate chamber and the reference chamber producesa pressure differential across the rate diaphragm, which furtherrepositions the metering valve to provide the correct cabin rate-of-change.

Increasing cabin altitude generates an increasing pressuredifferential between the cabin and reference pressure cham-bers, causing the rate diaphragm to move, which routes refer-ence air to the pneumatic relay.




Before takeoff, the cabin altitude selector should be set to 500

feet above the planned cruise altitude on the inner ACFT scaleor 500 feet above the destination field pressure altitude on theouter CABIN scale, whichever is greater. The CABIN RATEselector should be positioned to provide a rate-of-changewhich will result in the cabin reaching the altitude indicated on

the CABIN scale as the airplane reaches the correspondingaltitude indicated on the ACFT scale. During the takeoff roll,when the airplane is on the ground and the left main gear

safety switch is closed, throttle advancement beyond approxi-mately 85% N2 closes a solenoid valve that traps cabin airpressure within the auxiliary volume tank for reference by thecabin pressurization controller, closes a solenoid valve thatremoves suction to the outflow valves, and closes anothersolenoid valve that allows the cabin to prepressurize to 60 feetbelow field altitude at a fixed 500 FPM rate of change. After

liftoff, the safety switch functions to open this valve, therebyrestoring cabin rate of change control. As the airplane climbs,

the absolute bellows expands and contracts as minutechanges in cabin pressure are sensed. Expansion and contrac-tion of the bellows is resisted by the pressure differential acrossthe rate diaphragm such that the sum of these forces reposi-tions the metering valve to apply the correct reference pressureto the pneumatic relay.

During climb, increasing reference pressure causes the outflow

valves to be modulated toward the closed position such that

the selected cabin rate-of-change is maintained to the selectedaltitude. As the airplane reaches the planned cruise altitude atthe selected rate and levels off, the pressure differential acrossthe rate diaphragm equalizes and the flow of cabin pressure toairplane suction becomes steady. In this condition, referencepressure becomes essentially constant and cabin altitudestabilizes.

If required to ascend beyond the altitude indicated on theACFT scale, the controller should be reset to a higher altitudeto maximize passenger comfort and to prevent unscheduleddifferential pressure control by the outflow valve. If required todescend below the altitude indicated on the ACFT scale, reset-ting the controller is normally not required unless the descentwill result in airplane altitude being less than the selected cabin

altitude. In this case, the controller should be reset to a lowercabin altitude to maximize passenger comfort and to preventunpressurized operation.





Before descent to landing, the cabin altitude selector should beset to 500 feet above the destination field pressure altitude onthe inner ACFT scale and the CABIN RATE selector should be

positioned to provide a rate-of-change that will allow the cabinto reach the altitude selected on the CABIN scale as the air-plane reaches the corresponding altitude indicated on the

ACFT scale. During descent, decreasing reference pressurecauses the outflow valve to modulate toward open such thatthe cabin depressurizes at the selected cabin rate-of-changeuntil the airplane reaches the altitude indicated on the CABINscale. Below this altitude, the outflow valve will be fully open,

the cabin will be unpressurized, and cabin rate-of-change willfollow airplane rate-of-change until touch-down.

Note: The desired cabin altitude should be set as early aspractical to provide the lowest cabin rate-of-change. Rate-of-

change should be adjusted as necessary during ascent ordescent so that the cabin reaches the altitude indicated on theCABIN scale at approximately the same time that the airplanereaches the altitude indicated on the ACFT scale.

Note: To calculate the approximate field pressure altitude, add100 feet to the field elevation for each 0.10 inHg that the fieldaltimeter setting is over 29.92 inHg. Subtract 100 feet from thefield elevation for each 0.10 inHg that the field altimeter settingis below 29.92 inHg.

Pneumatic RelayThe pneumatic relay is the primary control device for the out-flow valves. The pneumatic relay amplifies reference pressurefrom the cabin pressurization controller by mixing it with suctionair flow generated by an ejector-type pump installed in the leftengine environmental supply tube. The pneumatic relay con-tains four chambers separated by two spring-loaded dia-phragms. The upper diaphragm is exposed on the upper sideto rate pressure and on the lower side to reference pressure.

The lower diaphragm is exposed on the upper side to refer-ence pressure and on the lower side to cabin pressure. Bothdiaphragms are connected to a metering valve that controlssuction air flow to create the amplified reference pressurewhich is then used to modulate (control) the outflow valves.




Cabin Altitude Limit Valves

Two cabin altitude limit valves, mounted on the aft pressurebulkhead to the right of the outflow valves, serve as backups tothe pressurization controller. Each altitude limit valve containsan evacuated bellows within a chamber that is open to cabinpressure through an inlet port, and a spring-loaded poppet

valve within a chamber that is connected to the cabin outflowvalve reference line. The poppet valve is normally held closedby pressure from a valve spring. If a malfunction causes out-

flow valve control suction to exceed normal limits, therebyopening the outflow valves excessively, the reduction in cabinpressure allows the altitude limit valve bellows to expand,unseating the poppet valve. With the poppet valve open, cabinpressure enters the outflow valve reference line, increasingoutflow valve reference pressure, thereby modulating the out-flow valves toward the closed position. The cabin altitude limit

valves function to prevent cabin altitude from exceeding ap-proximately 13,000 feet.





Pressurization System Indication

Pressurization system operation status is indicated on the cabinaltitude and differential pressure indicator, the cabin rate-of-change indicator, and by the cabin altitude warningannunciator.

1. Cabin Altitude Scale

2. Differential Pressure Scale

3. Cabin Rate of Change Indicator

Cabin Altitude and Differential Pressure Indicator

The cabin altitude and differential pressure indicator, locatedon the environmental panel, is a combination gage having anouter CABIN ALT scale denoting cabin altitude from 0 to45,000 feet in feet X 1000, and an inner DIFF PRESS scaledenoting 0 to 9 PSI differential pressure between the cabin andthe atmosphere. The DIFF PRESS scale features a green arc

between 0 and 8.7 PSI, denoting the differential pressure rangewithin the normal operating limitations of the system. A red lineat 8.8 PSI denotes system overpressure.

Cabin Rate-of-Change Indicator

The cabin rate-of-change indicator denotes the rate of cabinpressure change from 0 to 6,000 feet per minute in FT/MIN X1000.

Cabin Altitude Warning AnnunciatorA red [CAB ALT 10000 FEET] annunciator, activated by abarometric pressure switch, illuminates to indicate that thecabin altitude has exceeded 10,000 feet and the use ofsupplemental oxygen is required.

32

1




Emergency Dump

The cabin can berapidly depressur-ized by using theemergency dump(depressurization

toggle) valve. Theguarded operatinglever for this valve is

labeled EMERDUMP and is lo-cated on the envi-ronmental panel.Activation of theemergency dumpvalve applies suc-tion that opens the

pressurization outflow valves, thereby releasing cabin pressureand allowing cabin altitude to equalize with airplane altitude.

EMERGENCY DUMP VALVE





Ambient Air Sources

For the purposes of this chapter, ambient air is used for cabin

and flight compartment ventilation during ground operation, toprovide a positive pressure differential in the aft fuselage whilein flight, and to act as the heat exchange medium for air-to-air

heat exchangers. Other uses are discussed elsewhere in thismanual.

On airplanes 550-0484, 0483, 0482 and earlier and on air-planes 550-0627 and after, ambient ventilation air for the cabin

is routed from the NACA-type scoop on the dorsal fairing to thecabin fresh air check valve through ducting. On airplanes 550-0482,0485 through 0626, ambient ventilation air is taken from thetailcone. On some of these airplanes the cabin fresh air check

valve, attached to the aft pressure bulkhead, is mounted at theend of a duct. On other airplanes no duct is used.

An optional flood cooling

system is available on air-planes 550-0356 and after.This system is primarily usedon the ground, but may alsobe used at flight altitudesbelow 10,000 feet. The systemfunctions to supply a mixture

of ambient air and conditionedair directly to the passengercabin, bypassing the normalair distribution system. Floodcooling is activated by a two-position FLOOD COOLING switchon the tilt panel. Positioning the switch to “ON” activates thesystem by operating an electric actuator in a flow divider andan electric motor/fan. The divider diverts the flow of conditionedair to the axial fan, where it is mixed with ambient air before

entering the cabin through a grille at the top of the rear pres-sure bulkhead. The fan and actuator receive 28 VDC from theright main bus through the 20-amp FLOOD COOLING circuitbreaker located in the aft fuselage electrical power junction “J”box.

In flight, the aft fuselage (tailcone) is pressurized (relative tooutside pressure) to prevent ingestion of external fluids. This is

accomplished using ram effect air taken in to the aft fuselagethrough the dorsal scoop.

FLOOD COOLING SWITCH




On the ground, the tailcone pressurization system is disabled

to prevent pressure transients in the environmental systemduring the takeoff roll. On airplanes 550-0593 and after (andearlier airplanes incorporating SB550-21-23), the tailconepressurization system incorporates a valve in the inlet duct fromthe dorsal scoop that, during ground operation, is energized

closed to prevent air from entering the aft fuselage. On air-planes 550-0592 and earlier, a valve in the aft fuselage skin isenergized open to allow air to vent overboard until the airplane

is airborne. Both types of valve are actuated by thermal expan-sion of an enclosed fluid, so operating time for the valve willvary according to initial conditions. A 60-second time delaybefore valve actuation is incorporated on airplanes 550-0550and after. The pressurization valve receives power from the leftmain bus through a 5-amp circuit breaker located in the “J”box. The circuit breaker is labeled TAIL PRESSURIZATION or

TAIL BUMP PRESS, depending on service bulletin incorpora-tion.

On airplanes 550-0482, 0485 and after, ambient air from the aftfuselage is drawn by an ACM-driven fan that directs it throughducting, where the air absorbs heat from both ACM heat ex-changers and the precooler before it is dumped overboardthrough an outlet located on the lower side of the aft fuselagebelow the engine pylon. On airplanes 550-0484, 0483, 0481and earlier, ambient air to cool the ACM heat exchangers is

routed through ducting from the dorsal scoop to the ACM.

Moisture drains are located at the bottom of the ambient airinlet duct and on the bottom of the aft pressure bulkhead checkvalve duct attach connection. These moisture drains directaccumulated water into the lower aft fuselage area, where itexits through skin drain holes.





Vapor Cycle Air Conditioning System

The optional vapor cycle air conditioning system is electrically-controlled, thermostatically-regulated, and uses Freon (R-12)as a refrigerant. This system consists of one cockpit evapora-tor, two cabin evaporators, and either a nose baggage com-

partment or aft fuselage mounted condenser/compressor withassociated controls, wiring, and plumbing.

The optional vapor cycle air conditioning system provideseffective cockpit and cabin cooling. It is used primarily duringground operations, but may also be operated at flight altitudesup to 18,000 feet. The vapor cycle air conditioning system canbe used alone or in conjunction with the ACM.

Refrigerant Circulation System

The refrigerant circulation system functions to activate andcontrol the vapor cycle that reduces the temperature of cabinair. Major components of the system include a compressor,condenser, condenser blower, receiver-dryer, and three evapo-rator modules. A compressor/condenser unit is located in eitherthe nose baggage compartment or aft fuselage. The receiver-drier unit is installed within the compressor/condenser unit. Oneevaporator unit is located within the cockpit and two additional

evaporator units are located within the cabin.

The Freon (R-12) refrigerant used in this system is normally in a

gaseous state at standard atmospheric temperatures andpressures. Within specific ranges of temperature and pressure,however, the state of Freon may be transformed between liquidand gas. This characteristic of Freon is critical to understandingthe vapor cycle, because in the transformation from gas toliquid (condensation), heat is emitted; and in the transformation

from liquid to gas (evaporation), heat is absorbed.

Compressor

The compressor functions to provide the pressure and suctionthat circulates Freon through the condenser, the receiver-dryer,and the evaporator modules during air conditioning systemoperation. The compressor is belt-driven by means of a pulleyattached to an electric motor. The same motor also drives an

axial fan that provides airflow through the condenser. Thisairflow provides a cooling effect that condenses the hot gas-eous Freon from the compressor into a liquid.




When air conditioning is not required, the compressor is idle.

When air conditioning is required, the compressor initiates thevapor cycle by drawing low-pressure, low-temperature Freongas through a suction tube and delivering high-pressure, high-temperature Freon gas to the condenser.

Condenser

The condenser is mounted in proximity to the compressor andfunctions to transform the high-pressure, high-temperature

Freon gas into a high-pressure, low-temperature liquid. Con-densation occurs as heat energy in the Freon gas passingthrough the condenser coils is transferred to cooling fins whichare exposed to lower temperature airflow provided by a com-pressor-driven axial fan. The cooled, high-pressure, liquidFreon is then routed to the receiver-dryer.

Receiver-DryerThe receiver-dryer, installed in the high-pressure tube between

the condensers and the evaporator modules, functions toremove moisture from the liquid Freon when the air conditioningsystem is operating. Moisture removal is critical not only in theprevention of corrosion damage, but in the prevention of refrig-erant circulation blockage caused by thermal expansion valvefreeze-up. Normal operation of the air conditioning system forseveral minutes followed by the loss of cooling airflow mayindicate that freeze-up has occurred.

Evaporator ModulesOne evaporator module is installed within the cockpit, and isaccessed by removing floor panels located behind the pilot’sseat. Two additional evaporator modules are installed within theaft baggage compartment in proximity to the aft pressurebulkhead. Each module contains a thermal expansion valve,evaporator coil, drain tube, and evaporator blower. The evapo-

rator modules transform high-pressure liquid Freon into a low-pressure, low-temperature gas, completing the vapor cycle thatreduces the temperature of the cabin air.





High-pressure liquid Freon enters each evaporator modulethrough its respective thermal expansion valve that reducespressure. Each expansion valve incorporates a variable orifice

that is regulated automatically through a temperature-sensingbulb attached to the Freon suction tube near the evaporatoroutlet. When the temperature of the Freon gas leaving the

evaporator is too high, the orifice constricts to provide in-creased cooling. When this temperature becomes too low, theorifice opens to reduce cooling.

From the low-pressure side of the expansion valve, reduced

pressure liquid Freon is routed through the evaporator coilwhere it is transformed into a gas. In the transformation fromliquid to gas, heat is absorbed from the cabin air as it is drawnthrough each evaporator coil by its respective blower. Therefrigerated cabin air is then forced by the evaporator blowers

into the conditioned air distribution tubing. Cooled air from theforward (cockpit) evaporator is introduced into the cabinthrough armrest-level vents in the flight compartment. Air thathas been cooled by passing over the aft evaporators is intro-

duced into the passenger cabin and flight compartmentthrough the overhead duct.

As heat is absorbed from the cabin air, moisture accumulateson the evaporator coil and collects in the lower portion of eachevaporator module. Drain tubes carry this moisture to forwardand aft heated drain assemblies located below the cabin floor,

which automatically control drainage of moisture from theairplane. The condensate drain valves within these assembliesincorporate a two-stage orifice that provides maximum drain-age during ground operation, and reduced drainage duringpressurized flight to minimize loss of cabin pressure.

Vapor Cycle Air Conditioning System Protection

Vapor cycle air conditioner compressor protection is providedby low-pressure, high-pressure, and suction switches and a

150-amp current limiter.

The low and high-pressure switches are mounted on fittingsinstalled on the compressor housing, or on a fitting assemblythat is installed in proximity to the compressor.




Compressor power is routed through an anti-cycle relay. If

compressor pressure exceeds 350 PSI, the high-pressureswitch opens, thereby removing power from the anti-cyclerelay. The anti-cycle relay then interrupts current to the recy-cling timer, thereby shutting down the compressor.

A suction switch is installed into a fitting on the compressorhousing. If compressor suction drops below 9.5 PSI, the suc-tion switch opens, thereby removing power from the anti-cycle

relay. The anti-cycle relay then interrupts current to the recy-cling timer, thereby shutting down the compressor.

On airplanes 550-0505 and earlier incorporating SB550-21-16,the low-pressure switch is disabled.

On airplanes 550-0505 and earlier not incorporating

SB550-21-16, if compressor pressure drops below 32 PSI, thelow-pressure switch opens, thereby removing power from the

anti-cycle relay. The anti-cycle relay then interrupts current tothe recycling timer, thereby shutting down the compressor.

Vapor Cycle Air Conditioning System Controls and Indicators

Airplanes equipped with vapor cycle air conditioning have aFREON AIR CONDITIONER control panel mounted on thecopilot’s instrument panel. This air conditioning control panelincludes a rotary switch with four positions, labeled OFF,

FAN FWD, FAN ALL, and COMP and a FAN SPEED toggle

switch with HI and LO positions. The FAN FWD position ener-gizes only the flight compartment blower, the FAN ALL positionenergizes the all the blowers, and the COMP position energizesthe compressor and all the blowers. The selected blowers willrun at the speed selected by the FAN SPEED toggle switch.

Vapor Cycle Air Conditioning System Protection

The compressor motor will automatically shut down shouldmotor current exceed 350 amps, condenser Freon dischargepressure exceed 350 PSIG, or Freon condenser suction pres-sure fall below 10 PSIG.





Emergency Oxygen System

The oxygen system is designed to provide emergency breath-

ing oxygen for the crew and passengers in the event of apressurization system emergency occurring at flight altitudesabove 10,000 feet, or if the cabin fills with smoke. The standard

oxygen system includes a 22 cu-ft oxygen bottle that will pro-vide emergency oxygen for crew and six passengers for up to15 minutes; an optional system incorporates a 64 cu-ft bottlethat will provide emergency oxygen for crew and six passen-gers for up to 50 minutes.

Emergency oxygen system operation is initiated by an altitudepressure switch installed in the pressurized cabin area. At acabin altitude of 12,900 to 14,000 feet, the altitude pressureswitch energizes a solenoid valve that allows oxygen to flow to

the oxygen distribution system. Oxygen pressure of approxi-mately 70 PSI automatically opens passenger oxygen maskstorage compartment doors, thereby dropping the masks.

Both oxygen systems provide distribution plumbing for theflight compartment and passenger cabin, individual outlets foreach airplane occupant (location depending on seating con-figuration), an oxygen bottle, regulator, filler valve, pressuregage, and control switch. In all installations, the system regula-tor is assembled directly to the oxygen bottle and functions toreduce bottle pressure to a lower, constant supply pressure.

The regulator incorporates a mechanically-operated supplyshutoff valve and ports for a supply tube, filler tube, pressuregage tube, and overboard discharge indicator tube. The pres-sure regulator outlet ports in the 64 cu-ft system are installed inslightly different positions than the ports in the 22 cu-ft system.

On airplanes 550-0255 and after and earlier airplanes incorpo-rating SB550-35-2, the standard or optional oxygen bottle isinstalled on the aft fuselage compartment structure using

mounting brackets.

On airplanes 550-0254 and earlier not incorporatingSB550-35-2, the standard or optional oxygen bottle is installedbelow the right nose baggage compartment floor using mount-ing brackets.




All oxygen bottles require a U.S. Department of Transportation

(D.O.T.) designation that identifies bottle specification andservice pressure. The D.O.T. designation also identifies thebottle’s life limitation and hydrostatic testing requirements.

The 22 cu-ft bottle bears a DOT-3AA 1800 designation and

requires hydrostatic testing to 167% of its service pressureevery five years. There is no life limitation for this bottle unlessfailure occurs during hydrostatic testing. The 22 cu-ft bottle is

charged to 1800 PSI under standard atmospheric conditions.Two types of 64 cu-ft bottles are available. The lightweightbottle bears a DOT-3HT 1850 designation; the fiber-woundbottle bears a DOT-3FC1850 designation. Both types of bottlerequire hydrostatic testing to 167% of service pressure everythree years. Life limitation is twenty-four years from date ofmanufacture.

All oxygen bottles should be filled with breathing oxygen that

conforms to the requirements of MIL-0-27210, Type 1.

On airplanes and 550-0255 and after and earlier airplanesincorporating SB550-35-2, the oxygen filler valve is locatedinside the tailcone baggage compartment door. On airplanes550-0254 and earlier not incorporating SB550-35-2, the oxygenfiller valve is accessible through the right nose baggage com-partment door. The oxygen filler valve incorporates a filter and

a protective cap. A check valve, installed in-line between the

regulator and the filler valve, prevents the escape of bottlepressure from the filler tube or its connections.

Oxygen Outlets

Oxygen outlets for the pilot and copilot are located on thepilot’s and copilot’s side consoles in the flight compartment. Upto eight oxygen outlets may be installed in the overhead pas-

senger cabin. Due to differences in seating configurations, thelocation of the passenger cabin outlets will vary between air-planes. Each outlet incorporates a spring-loaded valve thatprevents oxygen flow from the outlet unless a hose assembly isconnected.





Warning: Permit no smoking when using oxygen. Materials thatwill not normally flash in the atmosphere will readily burn orexplode in the presence of concentrated oxygen. Oil, grease,

soap, lipstick, lip balm, and other fatty materials constitute aserious fire hazard when in contact with oxygen. Be sure handsand clothing are oil free before handling oxygen equipment.

Crew Oxygen Masks

Two types of crew oxygenmasks are available in theCitation II. The standard oxy-

gen mask is a diluter demandtype with integral oxygenregulator and microphone.Each oxygen regulator in-cludes a lever that allows

selection of diluter demand(NORMAL) or demand (100%OXY) modes. The demandmode should be selected to insure adequate supplemental

oxygen at altitudes above 20,000 feet. The standard maskqualifies as quick donning when it is worn with the head straparound the neck.

The optional crew oxygen mask is a quick-donning sweep-ontype with a regulator and microphone attachment. This mask isa diluter demand type with pressurized flow (100% oxygen)

selectable by placing the regulator in the “EMER” (demand)position. The EMER position should be selected to insure ad-equate supplemental oxygen at cabin altitudes above 20,000feet. To conserve oxygen, the regulator may be set to “NOR-MAL” if cabin altitude is below 20,000 feet. To qualify as quick-donning, the mask must be properly stowed in its retainer.

Either mask should be set to the 100% oxygen (“100% OXY” or“EMER”) position when it is used for smoke protection.




Oxygen mask microphones

are operated using a two-position toggle switch on thepilot’s and copilot’s side con-soles. Setting this switch to“MIC OXY MASK” energizes

the mask microphone; settingthe switch to “MIC HEAD SET”energizes the headset micro-

phone. The selected micro-phone may then be used fortransmission by depressingthe microphone button on the control wheel.

Passenger Oxygen Masks

Passengers are provided oro-

nasal type oxygen masks thatdeform to seal around the

nose and mouth area. Eachmask consists of a face plate,economizer bag, plastic sup-ply tube, and a lanyard cordwith pintle pin attached. Thepintle pins are installed toprevent oxygen loss fromunused masks. After oxygen

masks are deployed, the lanyard cord must be pulled to with-

draw the pintle pin and thereby initiate oxygen flow to eachmask. Passenger oxygen masks provide a constant flow rate of4.5 liters per minute.





Oxygen System Controls, Malfunction Warning,

and Indication

On airplanes 550-0501 andafter, the rotary, three-positionOXYGEN CONTROL VALVE

switch is located on the pilot’sside console. When this switchis set to the center “NORMAL”position, oxygen automaticallyflows to both the cockpit and

cabin. When the switch isrotated counterclockwise tothe “CREW ONLY” position,oxygen flow is limited to thecockpit. The switch is rotated clockwise to the “MANUALDROP” position to manually deploy passenger oxygen masks ifthe automatic mask deployment system should fail.

On airplanes 550-0500 andearlier, two-position toggleswitches labeled OXYGENPRIORITY VALVE and PASSOXY MASKS are located onthe pilot’s side console. Whenthe OXYGEN PRIORITY VALVEswitch is set to the upper

“NORMAL” position, oxygenautomatically flows to both thecockpit and cabin. When theswitch is set to the “CREWONLY” position, oxygen flow is limited to the cockpit. ThePASS OXY MASKS switch is repositioned from“NORMAL” to “MANUAL DROP” to manually deploy passengeroxygen masks if the automatic mask deployment systemshould fail.




The oxygen pressure gage is

located on the right instrumentpanel. The gage providesvisual indication of bottlepressure and is marked with ayellow arc from 0 to 400 PSI, a

green arc from 1600 to 1800PSI, and red line at 2000 PSI.When fully charged and stabi-

lized at approximately 70°F,indicated pressure should be1800 PSI for the 22 cu-ft bottleand 1850 PSI for the 64 cu-ft bottle; however, indicated pres-sure will vary with ambient temperature. Either bottle will requirerecharging if indicated pressure falls below 300 PSI.

Evidence of oxygen bottleoverpressure is provided by

an indicator disc (originallygreen in color).

Should overpressure occur, ahigh-pressure rupture fittingwithin the regulator releasesbottle pressure through theoverboard discharge indicator

tube. When oxygen bottle

pressure exceeds 2850 ± 150PSI (at 70°F/21°C), the disc is “blown out,” thereby providingvisual indication that oxygen was discharged overboard.

On airplanes with oxygen bottles located in the nose section,the indicator disc is located on the lower right surface of thenose. On airplanes with oxygen bottles located in the aft fuse-

lage, the indicator disc is located on the lower left surface ofthe tailcone.





Limitations

Refer to the applicable airplane manufacturer’s FAA approvedairplane flight manual or approved manual material, markingsand placards, or any combination thereof for all limitations.


Refer to the applicable airplane manufacturer’s FAA approvedairplane flight manual or approved manual material (supple-mentary checklist) as revised, for procedural information.




ICE PROTECTION SYSTEMS 12/99 FOR TRAINING PURPOSES ONLY 10-1

Overview

This chapter describes the ice protection equipment requiredon the Cessna Citation II for flight in icing conditions.

Anti-ice systems are designed to prevent the formation of ice

and should be activated prior to entering icing conditions. Forthis purpose, electrically-powered heating elements are in-stalled in the pitot tubes, static ports, and angle-of-attack sensor(if installed). Protection against windshield icing is accom-

plished primarily using engine bleed air, with alcohol used as abackup anti-ice system for the pilot’s windshield. Engine iceprotection is accomplished by bleed air heating of induction airinlet components and electrical heating of the inboard wingleading edges forward of each engine.

Deice systems are designed to remove ice which has accumu-lated. For this purpose, pneumatically-operated boots are

attached to the leading edges of the stabilizers and the out-board leading edge of each wing.

The Cessna Citation II is approved for flight in icing conditionsas defined by the FAA only when the following ice protectionequipment is installed and checked operational before flight:

Anti-Ice

Heated Pitot Tubes

Heated Static PortsHeated WindshieldBackup Windshield Alcohol SystemEngine Ice Protection

Deice

Wing and Stabilizer Deice Boots

Note: Refer to the FAA-approved Master Minimum EquipmentList (MMEL) for conditions and limitations specific to the iceprotection equipment installed in your airplane.




Anti-Ice Systems

For organizational purposes, this section is divided into pitot-static anti-ice, windshield anti-ice, and engine anti-ice. All anti-ice systems must be activated when operating in visible mois-ture at indicated outside air temperatures (IOAT) between +4°C

(39°F) and -30°C (-22°F).

Pitot-Static Anti-Ice

The pitot tubes and static ports are protected against icing byintegral, electrically-powered heating elements which arecontrolled by the PITOT & STATIC switch on the lower left instru-ment panel. When this switch is in the upper (on) position, 28

VDC power is supplied to the heating elements of the pilot’spitot tube and static ports from the left main bus through the 7.5-amp LH PITOT STATIC circuit breaker, and supplied to the

copilot’s pitot tube and static ports from the right main busthrough the 7.5-amp RH PITOT STATIC circuit breaker.

Independent left (pilot’s) and right (copilot’s) current sensorsmonitor the flow of current to their associated heating elements.On airplanes 550-0550 and after, these current sensors controlthe illumination of an amber, three lens [LH] [RH] [P/S HTR OFF]

annunciator. On airplanes 550-0028 and earlier, the currentsensors control the illumination of a single pitot heat off/fail lighton the left instrument panel; on airplanes 550-0029~0505, thislight is replaced by an amber, single lens [P/S HTR OFF] an-nunciator.

PITOT TUBE STATIC PORTS





When the PITOT & STATICswitch is in the upper “on”position and all heating ele-ments are functioning nor-

mally, the annunciator shouldbe extinguished. When this

switch is in the “on” positionand at least one heating ele-ment is inoperative, the annun-ciator should be illuminated.When the switch is in the“OFF” position or when no

power is being supplied to theheating elements, the annunciator should also be illuminated.On airplanes 550-0550 and after, the corresponding [LH](pilot’s) or [RH] (copilot’s) lens will illuminate in conjunction with

the [P/S HTR OFF] lens to indicate which heating elements aremalfunctioning.

To minimize battery drain and prevent overheating of the ele-

ments during ground operation, the PITOT & STATIC switchshould remain in the “OFF” position except for system testing.To test the system prior to flight, the pitot tube covers should beremoved (if installed), the PITOT & STATIC switch should bepositioned to “on” for a period of 30-seconds and then returnedto “OFF.” If the system is functioning normally, the [P/S HTROFF] annunciator should not have illuminated when the PITOT

& STATIC switch was in the “on” position, the pitot tubes shouldbe hot, and the static ports should be warm.

To reduce the risk of severe burns when checking pitot tubeheat, physical contact with the tube should be minimized andgrasping the tube with any more than a light grip should beavoided. Checking static port heat in high ambient tempera-tures is best accomplished using the back of a finger to com-pare the temperature of each static port to that of the surround-ing fuselage skin. Refer to the appropriate Operating Manual or

Airplane Flight Manual for test procedures specific to the sys-tem installed in your airplane.

Caution: To prevent overheating of the elements, ground opera-tion of the pitot-static ice protection system is limited to 2 min-utes.

PITOT-STATIC HEAT SWITCH




Windshield Anti-Ice

1. Rain Removal Augmenter Door

2. Alcohol Spray Tubes

3. Bleed Air Discharge Nozzle Shroud

Protection against windshield icing is accomplished primarilyusing engine bleed air, with alcohol used as a backup anti-icesystem for the pilot’s windshield. In addition to ice protection,

the bleed air windshield anti-ice system provides rain removaland external defogging capabilities.

Bleed Air Windshield Anti-Ice

The bleed air windshield anti-ice system directs engine bleedair against the windshield to prevent the formation of ice. Major

components of the system include a bleed air control valve,heat exchanger, automatic temperature controls, temperatureand pressure sensors, manual flow controls, and bleed air

discharge nozzles.

Bleed air used by the system is tapped from the supply tubes

between the compressor section of each engine and the envi-ronmental flow control/shutoff valves within the aft fuselage. Onairplanes 550-0484, 0483, 0481 and earlier, the bleed air isrouted through a precooler within each engine nacelle prior toentering the aft fuselage. On airplanes 550-0482, 0485 and

after, availability of bleed air to the system is continuous wheneither or both engines are operating and is not influenced bythe position of the PRESS SOURCE selector. On airplanes 550-

0484, 0483, 0481 and earlier, however, setting the PRESSSOURCE selector to “LH” or “RH” correspondingly results inbleed air being made available to the system from the left

engine or right engine only.

1

2

3





The bleed air tapped from each source is routed through inde-pendent supply tubes to a common cross fitting within the aftfuselage. A check valve in each of these tubes prevents thebackflow of bleed air to either engine when the opposite engine

has failed or is operating at a sufficiently lower RPM. From thecross fitting, bleed air is routed through the bleed air control

valve to the heat exchanger, each also located within the aftfuselage.

1. Heat Exchanger 3. Cross Fitting2. Bleed Air Control Valve 4. Exit Duct

The bleed air control valve is normally-open, electrically-actu-ated closed, and controlled primarily by the W/S BLEED switchon the lower left instrument panel. When this valve is open,

bleed air is routed through the heat exchanger to the rest of thesystem.

12

34




The heat exchanger is an air-to-air type unit which functions toreduce bleed air temperature to that required for system opera-

tion. The heat exchanger is installed within an air duct throughwhich cooling airflow passes and conducts heat from the bleedair. On airplanes 550-0482, 0485 and after, airflow enters thisduct through a screened inlet on the right side of the tailcone,

and exits through a NACA-type exhaust scoop on the leftside of the tailcone below theengine pylon. On airplanes

550-0484, 0483, and 0481 andearlier, airflow enters this ductthrough flush-mounted, NACA-type intake scoops on thedorsal fairing, and exitsthrough a NACA-type exhaust

scoop on the left side of thetailcone below the engine

pylon.

The automatic temperature controller maintains the requiredbleed air temperature by modulating the position of an electri-cally-actuated air control valve, located in the heat exchangerexit duct, which regulates ambient airflow through the heatexchanger. Input signals are provided to the controller by theW/S BLEED switch and by temperature sensors located in thebleed air supply tubing: one downstream of the heat ex-

changer, another upstream of the nozzles.

AIRFLOW OVERBOARD EXHAUST SCOOP





From the heat exchanger,bleed air is routed through apair of mechanically-actuatedcontrol valves which regulatebleed air flow to the dischargenozzles. Each valve is located

within the nose section andmanually positioned by itsassociated LH or RH WIND-SHIELD BLEED AIR control on

the copilot’s lower instrumentpanel. Each discharge nozzle

is enclosed in an aerodynamicshroud and comprised of a manifold which supplies an array ofoutlet tubes that direct bleed air against the windshield. Eachshroud is fitted with a hinged augmenter door for rain removal.

Both augmenter doors are mechanically-operated by a singlePULL RAIN control, located below the copilot’s instrumentpanel. The left (pilot’s) shroud also houses the alcohol dispersanozzle.

Rotating the WINDSHIELDBLEED AIR controls clockwiseprogressively increases bleedair flow to the dischargenozzles; counterclockwiserotation progressively de-

creases bleed air flow to thedischarge nozzles. When

windshield rain removal isrequired, the WINDSHIELDBLEED AIR controls should berotated fully-clockwise to“MAX.” When windshield iceprotection or rain removal isnot required, the WINDSHIELDBLEED AIR controls should be

rotated fully-counterclockwiseto “OFF.”

DISCHARGE NOZZLE VALVE

1. Windshield BleedAir Controls

2. Pull Rain Control Knob

1

2




System Operation

The system is activated by the

three-position (HI/OFF/LOW)W/S BLEED switch on thelower left instrument panel.When this switch is set to “HI”

or “LOW,” the bleed air controlvalve is deenergized openand 28 VDC power is sup-plied from the left main bus

(550-0550 and after) or rightmain bus (550-0505 andearlier) to the automatic tem-perature controller through the 5-amp W/S BLEED TEMP circuitbreaker. When “HI” is selected, the temperature controllermodulates the position of the air control valve to maintain bleed

air temperature at approximately 138°C. When “LOW” is se-lected, the temperature controller modulates the position of the

air control valve to maintain bleed air temperature at approxi-mately 127°C. When set to “OFF,” the temperature controller isdeactivated and the bleed air control valve is energized closed.On airplanes 550-0550 and after, 28 VDC left main bus power issupplied to the bleed air control valve through the 5-amp W/SBLEED circuit breaker. On airplanes 550-0505 and earlier, 28VDC right main bus power is supplied to the bleed air controlvalve through the 5-amp WINDSHIELD BLEED AIR circuitbreaker.

Before activating the system, the WINDSHIELD BLEED AIRcontrols should be rotated clockwise and the PULL RAIN con-trol, located below the copilot’s instrument panel, should bepushed fully-in. When windshield ice protection is required, theW/S BLEED switch should be set to “HI” when IOAT is below

-18°C, or “LOW” when IOAT is above -18°C. Normal systemoperation is indicated by an increase in air noise as bleed air isdischarged from the nozzles.

W/S BLEED SWITCH




Operation of the overheatsensor and [W/S AIR O’HEAT]annunciator may be verified

by rotating the TEST selector,located on the lower left instru-ment panel, to the “W/S TEMP”

position and setting either W/SBLEED switch to “HI” or“LOW,” In this condition, the[W/S AIR O’HEAT] annunciator

should illuminate if the sensoris functional. When illuminatedby the overheat sensor duringsystem operation or testing, the [W/S AIR O’HEAT] annunciatorreceives 28 VDC power from the left main bus through the 5-amp W/S BLEED circuit breaker (550-0550 and after), or from

the right main bus through the 5-amp WINDSHIELD BLEED AIRcircuit breaker (550-0505 and earlier). When illuminated by the

pressure switch, the [W/S AIR O’HEAT] annunciator receives 28VDC power from the left main bus (550-0550 and after) or rightmain bus (550-0505 and earlier) through the 5-amp W/S BLEEDTEMP circuit breaker.

Bleed Air Windshield

Rain Removal

Rain removal is provided bythe windshield anti-ice system

and the augmenter doors oneach discharge nozzleshroud. When bleed air flowfrom discharge nozzles isinsufficient to clear the wind-shield of heavy rain, the aug-menter doors can be openedto provide increased airflow

over the windshield. Both augmenter doors are mechanically-operated by a single PULL RAIN control, located below thecopilot’s instrument panel. When windshield rain removal isrequired, the WINDSHIELD BLEED AIR controls should berotated fully-clockwise to “MAX,” the PULL RAIN control shouldbe pulled fully-out, and the W/S BLEED switch should be set to“LOW.”

Note: Difficulty may be encountered opening the augmenterdoors at airspeeds above 175 KIAS, or when the W/S BLEEDswitch is set to “LOW” prior to operating the PULL RAIN control.






Alcohol Windshield Anti-Ice

The alcohol windshield anti-ice system provides a backup tothe bleed air windshield anti-ice system for the pilot’s wind-shield only. Major components of the system include an alcoholreservoir, an electrically-operated pump, and a six-tube dis-persal nozzle.

1. Fluid Level Sight Gage

2. Alcohol Reservoir

3. Augmenter Door Linkage

The alcohol reservoir is located behind the aft divider within theright nose baggage compartment. A sight gage on the upperreservoir permits fluid level inspection. The sight gage is visiblethrough an inspection window positioned on an access panel.

The access panel is hinged to facilitate servicing and is acces-sible through the right baggage door. If fluid is not visible in thesight gage, the reservoir should be replenished. Reservoircapacity is 0.5 U.S. gallons TT-I-735 isopropyl alcohol only.

1

2

3




The electrically-operated pump, also located behind the aftdivider within the right nose baggage compartment, deliversalcohol under positive pressure to the dispersal nozzle. Thepump incorporates an integral fluid filter through which alcoholpasses before delivery to the dispersal nozzle. To preventaudio interference during system operation, a radio noise filter

is installed in the electrical circuit to the pump. The dispersalnozzle, enclosed within the left (pilot’s) shroud, incorporates six

spray tubes which distribute alcohol over the pilot’s windshield.

The system is activated by thetwo-position (ON/OFF) W/SALCOHOL ANTI-ICE switch onthe lower left instrument panel.When this switch is positionedto “ON,” 28 VDC power is

supplied to the pump from theright main bus (550-0550 and

after) or left main bus (550-0505 and earlier) through the5-amp W/S ALCOHOL circuitbreaker. With the pump ener-gized, alcohol is drawn fromthe reservoir and delivered to the dispersal nozzle. With thealcohol reservoir serviced to capacity, maximum continuousoperation endurance is approximately 10 minutes.

Note: If failure of the bleed air windshield anti-ice system ne-cessitates activation of the alcohol windshield anti-ice system,icing conditions should be exited as soon as practicable.

During preflight inspection, the alcohol spray tubes should bechecked for general condition and cleanliness, and the reser-

voir level should be checked full. Operation of the system canbe tested before flight by positioning the W/S ALCOHOL ANTI-ICE switch to “ON” until alcohol is observed flowing from all sixspray tubes.

W/S ALCOHOL ANTI-ICE SWITCH





1. Outboard LeadingEdge Boot

2. VerticalStabilizer Boot

3. HorizontalStabilizer Boot

Surface Deice System

The surface deice system functions to remove ice accumula-tions from the leading edges of the stabilizers and outboardleading edge of each wing. The electrically-controlled, pneu-matically-operated system consists of inflatable rubber deice

boots, a pneumatic pressure regulator, three control valves, atimer module, two pressure switches, and associated controls,plumbing and circuitry.

1

2

3




The deice boots are essentially fabric-reinforced rubber sheetscontaining built-in, spanwise inflation tubes. Each boot is

bonded by adhesive to the leading edge of the surface beingprotected and features a conductive coating which dischargesstatic electricity. The boots are normally held deflated againsttheir respective leading edge surfaces by suction. When in-

flated by air pressure, the change in boot contour breaks up iceaccumulations to facilitate removal by normal in-flight air forces.

Air pressure for boot inflation and suction for boot deflation isprovided by engine bleed air supplied from the same crossfitting that supplies the windshield anti-ice system. From thiscross fitting, bleed air is routed through the pneumatic pressureregulator which functions to reduce bleed air pressure to ap-proximately 23 PSIG. From the regulator, bleed air is routed

through a cross fitting where its flow is divided into three paths

which independently supply the stabilizer, left wing, and rightwing control valves, also located within the aft fuselage. Bleed

air is continuously supplied to the control valves whenevereither or both engines are operating. Refer to the Bleed AirWindshield Anti-Ice section of this chapter for a description ofthe bleed air source.

The three electrically-actuatedcontrol valves function tocontrol the application ofsuction (when deenergized

closed) or pressure (whenenergized open) to their asso-ciated boots as determinedprimarily by the timer module.When deenergized closed,each control valve functionsas an ejector, producingapproximately 5.5 inHg of

suction by directing bleed airthrough an overboard vent tube. When energized open, theoverboard vent tube is blocked and bleed air inflation pressureis directed to the boots. The timer module, located within theflight compartment left side console, functions to energize anddeenergize the control valves sequentially. The pressureswitches, one located in the stabilizer boot supply line, theother located in the right wing boot supply line, function to

illuminate the white [SURFACE DEICE] (550-0550 and after) or[SURF DEICE] (550-0505 and earlier) annunciator when infla-tion pressure is at least 20 PSI.

DEICE BOOT OVERBOARD VENT TUBES





System Operation

The system is activated by thethree-position SURFACE DE-ICE switch on the lower left

instrument panel. When thisswitch is momentarily actuated

to the upper (on) position, 28VDC power is supplied fromthe right main bus (550-0550and after) or left main bus(550-0505 and earlier) to thetimer module through the 5-

amp SURFACE DE-ICE circuitbreaker, thereby initiating a two-sequence deice cycle.

During the first sequence, the timer energizes the stabilizer

control valve open directing inflation pressure to the stabilizerboots for approximately six seconds. Full inflation and annun-ciator illumination normally occur within approximately two

seconds. During the second sequence, the timer deenergizes

the stabilizer control valve closed and energizes the left andright wing control valves open directing inflation pressure to thewing boots for approximately six seconds. The annunciator willextinguish momentarily between sequences and illuminatewhen the wing boots are fully-inflated. Full deflation of thestabilizer boots normally occurs within approximately twelveseconds following completion of the first sequence. Full defla-

tion of the wing boots normally occurs within approximatelytwelve seconds following completion of the second sequence.Upon completion of the cycle, the timer module and controlvalves are deenergized, the annunciator is extinguished andsuction is applied to all of the boots.

Each momentary actuation of the SURFACE DE-ICE switch to

“on” results in one complete cycle. Though the inflation se-quences last approximately twelve seconds combined, theadditional time required for the deflation of all boots results in

one complete cycle actually lasting approximately twenty-fourseconds. System activation may be repeated as necessaryallowing twenty-four seconds between cycles.

Note: If the boots fail to deflate or if cycle termination is desired,momentary actuation of the SURFACE DE-ICE switch to the

lower “RESET” position overrides the timer module and immedi-ately deenergizes all three control valves closed.

SURFACE DE-ICE SWITCH




The system should be activated when ice accumulations of atleast 1/4” to 1/2” are observed on the leading edge of eitheroutboard wing. Activation of the system with accumulations of

less than 1/4” may result in ice bridging on the wing. Accumula-

tions of greater than 1/2” may exceed the system’s ice removalcapabilities. Operation and/or testing of the system at indicated

outside air temperatures (IOAT) below -40°C (-40°F) may resultin boot cracking or failure of the boots to fully-deflate.

During preflight inspection, the deice boots should be checkedfor general condition and cleanliness. Operation of the systemcan be tested before flight by momentarily actuating the SUR-FACE DE-ICE switch to the upper “on” position and visually

confirming normal inflation and deflation of the wing boots aswell as illumination of the annunciator.

Surface Deice System Schematic (boots deflated)

© PCW





Surface Deice System Schematic (first cycle: empennage boot inflation)

Surface Deice System Schematic (second cycle: wing boot inflation)

© PCW

© PCW




Ice Detection

A wing ice inspection light is

installed on the left side of thefuselage, forward of the wing.The light illuminates the uppersurface and leading edge of

the wing so that these surfacescan be checked for ice accu-mulation during night opera-tions. The light is controlled by

the WING INSP switch on thelower left instrument panel.When this switch is positionedto “ON,” 28 VDC power is supplied to the light from the rightmain bus (550-0550 and after) or left main bus (550-0505 andearlier) through the 5-amp WING INSP circuit breaker.

WING ICE INSPECTION LIGHT




Availability of bleed air to the inlet lip and core inlet stators isthrough independent pressure regulating/shutoff valves whichare electrically-actuated closed, and controlled primarily by theLH and RH ENGINE anti-ice switches on the lower left instru-ment panel. Although these valves are electrically considerednormally-open, they are pneumatically considered normally-

closed and require at least 8 PSI bleed air pressure to open.Additionally, each inlet lip valve requires that its corresponding

THROTTLE lever be positioned above 60% N2 to open. Thestator valve incorporates a position switch and the inlet lipincorporates a temperature switch, each associated primarilywith system malfunction indication.

Inboard Wing

Leading Edge Anti-Ice

The inboard wing leading

edge anti-ice system operatesin conjunction with the engine

bleed air anti-ice system anduses electrically-heated pan-els to prevent the formation ofice on the upper wing surfaceforward of the engines. Eachremovable panel features ahighly-polished exterior thatforms a 61” section of itsassociated inboard wing leading edge. Five independent,

spanwise heating elements, a high temperature switch, lowtemperature switch, and a temperature sensor are bonded tothe interior of each panel. A Kevlar insulation shield provides athermal barrier between the heated panel and the wingstructure.

An independent temperature control circuit is provided for eachpanel to maintain operating temperatures between 54°C and78°C nominal. Each circuit includes a temperature controller,control relay, and power relay which function to regulate theflow of current to the heating elements. With the system acti-vated, current flow to the heating elements will occur only whenthe control relay and power relay are energized closed.

INBOARD WING ANTI-ICE PANEL





The temperature sensor in each panel is positioned where thehighest operating temperatures exist. The temperature control-ler cycles the system on and off according to input signalssupplied by the temperature sensor. To accomplish this, thetemperature controller energizes and deenergizes the controlrelay as a function of sensed temperature. When approximately

54°C is sensed, the control relay is energized closed, therebysupplying current to the heating elements. When approximately78°C is sensed, the control relay is deenergized open, therebyinterrupting current flow to the heating elements.

The low temperature switch in each panel is also positioned

where the highest operating temperatures exist, while the hightemperature switch is positioned where lower operating tem-peratures exist. In the event of temperature controller failure, thehigh temperature switch energizes and deenergizes the power

relay as a function of sensed temperature. When approximately74°C is sensed, the power relay is deenergized open, therebyinterrupting current flow to the heating elements. When approxi-mately 68°C is sensed, the power relay is energized closed,

thereby supplying current to the heatingelements.

Note: Although the control relay and power relay appear tohave overlapping temperature activation ranges, the location ofthe temperature sensor and high temperature switch ensuresthat neither senses the same temperature simultaneously.

28 VDC power is supplied to each panel by its associated leftor right main bus through a 175 amp current limiter. Eachheating element is provided with a circuit breaker and currentsensor. The five current sensors for each panel are wired inseries and independently monitor the flow of current to theirassociated heating elements. The current sensors and the lowtemperature switch are associated primarily with system mal-

function indication.




System Operation

The engine bleed air and

inboard leading edge anti-icesystems are simultaneouslyactivated by the LH and RHENGINE anti-ice switches on

the lower left instrument panel.When these switches are inthe upper (on) position, 28VDC power is supplied to the

inboard wing leading edgeheating elements and theignition system. After a fivesecond time delay, the pressure regulating shutoff valves foreach inlet lip and core inlet stator will be deenergized open ifthe THROTTLE levers are positioned above 60% N2 and at

least 8 PSI bleed air pressure is available. If the system isfunctioning normally, each heated leading edge panel will

draw approximately 150 amps and the consumption of bleedair for inlet lip and core inlet stator heating will increase indi-cated ITT and decrease engine RPM.

Malfunction Indication

Indication of engine ice protection system malfunction is pro-vided by the amber [LH] and [RH] [ENG ANTI-ICE] annuncia-

tors (550-0550 and after) or [L ENG ICE FAIL] and [R ENG ICEFAIL] annunciators (550-0505 and earlier). Illumination of the

corresponding annunciator(s) will occur under the followingconditions when the LH and RH ENGINE anti-ice switches arein the upper (on) position:

1. by current sensor when at least one heating elementis inoperative

2. by low temperature switch when leading edge temperatureis below approximately 16°C

3. by high temperature switch when leading edge temperatureis above approximately 74°C

4. by temperature controller when temperaturesensor malfunctions

5. by position switch when core inlet stator valve fails to open

6. by temperature switch when inlet lip temperature isbelow 104°C

ENGINE ANTI-ICE SWITCHES





Engine Ice Protection Schematic

Note: Illumination of these annunciators during the five secondsfollowing system activation is normal before the pressure regu-lating shutoff valves for each inlet lip and core inlet stator open.

Limitations



Refer to the applicable airplane manufacturer’s FAA approvedflight manual or approved manual material (supplementarychecklist) as revised, for procedural information.

© PCW




Chapter 11Limitations

Table of Contents

Overview .............................................11-1

Airspeed Limitations ............................11-2

Operating Limitations...........................11-3

Weight Limitations..........................11-3

Center of Gravity Limits ..................11-3

Takeoff and Landing Limitations...... 11-4

Flight Load Factor Limitations ......... 11-4

Enroute Limitations ........................11-5

Approved Operations ...................... 11-5Engine Operating Limitations................11-6

Engine Fan ....................................11-6

Battery and StarterCycle Limitations .................................11-7

Battery Limitation ...........................11-7

Prolonged Ground Operations..........11-7

Oil Limitations ....................................11-8

Approved Oils ................................11-8

Fuel Limitations ..................................11-9

Approved Fuels ..............................11-9

Fuel Temperature andDensity Limitations.......................11-10

Maximum Fuel Imbalance ............. 11-10




Hydraulic Fluid Limitations ................. 11-10

Flight Crew Limitations ...................... 11-11

Cabin Limitations ..............................11-11

Pressurization Differential .................. 11-11

Pressurization Source Selector ...........11-11

Icing Limitations ................................11-12

Thrust Reversing Limitations .............. 11-13

Oxygen System .................................11-13

Autopilot ...........................................11-14

HF/ADF System ................................11-14

Baggage Limitations .......................... 11-14

Baggage CompartmentWeight Limitations........................11-14

Baggage CompartmentVolume Limitations.......................11-14

Limitations, continued

Table of Contents



LIMITATIONS 12/99 FOR TRAINING PURPOSES ONLY 11-1


Overview

This chapter provides a comprehensive listing of operationallimitations for the safe operation of the Citation II airplane, itsengines, systems, and equipment.

Note: The limitations given in this section are for training pur-poses only. Consult your Pilot’s Operating Handbook for limita-tions specific to the year, model and serial number of your

airplane.




Airspeed Limitations

Airspeed 550-0627 and After 550-0626 and Earlier 550-0626 and Earlier

(not incorporating SB550-32-14) (incorporating SB550-32-14)

Maximum Operating SpeedMMO (Mach)

Above 28,000 Feet N/A 0.705 Mach 0.705 MachAbove 30,500 Feet 0.705 Mach N/A N/A

Maximum Operating SpeedVMO (Knots)

14,000~28,000 Feet N/A 277 KIAS 277 KIAS

Below 30,500 Feet 262 KIAS N/A N/A

14,000~30,500 Feet N/A 262 KIAS 262 KIAS(11,000 LB ZFW)

Below 14,000 Feet N/A 262 KIAS 262 KIAS

Maneuvering SpeedVA

Maximum FlapExtended SpeedVFE (Knots)

15° Flaps 202 KIAS 202 KIAS 202 KIAS40° Flaps 176 KIAS 176 KIAS 176 KIAS

Maximum Landing GearOperating SpeedVLO (Knots)

Extend 250 KIAS 176 KIAS 250 KIASRetract 200 KIAS 176 KIAS 200 KIAS

Maximum Landing GearExtended SpeedVLE (Knots) 262 KIAS 176 KIAS 277 KIAS

Maximum Speed BrakeOperating SpeedVSB (Knots) No Limit No Limit No Limit

Minimum ControllableAirspeedVMCA (Knots) 77 KIAS 77 KIAS 77 KIAS

Minimum ControllableGround SpeedVMCG (Knots) 62 KIAS 62 KIAS 62 KIAS

Maximum TireGround Speed 165 KIAS 165 KIAS 165 KIAS

Autopilot OperationAbove 14,000 Feet N/A 277 KIAS/0.705 Mach 277 KIAS/0.705 MachBelow 14,000 Feet N/A 262 KIAS 262 KIAS

Above 30,500 Feet 262 KIAS/0.705 Mach N/A N/ABelow 30,500 Feet 262 KIAS/0.705 Mach N/A N/A

Per Sec II of FAA Approved Airplane Flight Manual

R-6/9/00





Operating Limitations

Weight Limitations

550-0550 ~ 550-0626:

Maximum Ramp Weight ................... 13,500 LBS

Maximum Takeoff Weight ................. 13,300 LBSMaximum Landing Weight ................ 12,700 LBS

Maximum Zero Fuel Weight ..............11,000 LBS

(550-0505 and earlier: 9500 LBS –standard,

11,000 LBS optional)

550-0627 and after:

Maximum Ramp Weight ................... 14,300 LBS

Maximum Takeoff Weight ................. 14,100 LBS

Maximum Landing Weight ................ 13,500 LBSMaximum Zero Fuel Weight ..............11,000 LBS

Note: Maximum takeoff and landing weights may be addition-ally restricted due to altitude, temperature and field length.

Center of Gravity Limits

550-0626 and earlier:Forward Limit:8540 LBS or less............................... 276.10 inches

aft of reference datum.13,300 LBS or less ............................279.80 inches

aft of reference datum.

12,500 LBS or less ............................279.20 inchesaft of reference datum.

Aft Limit: ............................................285.8 inchesaft of reference datum.

550-0627 and after:Forward Limit:8540 LBS or less............................... 276.10 inches


14,100 LBS or less ............................280.40 inchesaft of reference datum.

Aft Limit:14,100 LBS or less ............................285.80 inches


Note: It is the responsibility of the pilot to ensure that the air-plane is loaded properly. Refer to Weight and Balance DataSheet for proper loading instructions.

R-6/9/00




Takeoff and Landing Limitations

Maximum Altitude...................................14,000 Feet

Max Tailwind Component (<0626) .........10 KnotsCrosswind Component (>0627) .............23 Knots

Maximum RunwayWater/Slush Accumulation .....................0.4 Inches

Maximum AmbientTemperature ........................................... ISA + 39°C (130°F)

Minimum AmbientTemperature ........................................... -54°C (-65°F)

Note: Autopilot and yaw damper must be OFF for takeoff andlanding. Vertical navigation system must be OFF below 500 feetAGL.

Flight Load Factor Limitations

550-0626 and earlier at 13,300 LBS maximum takeoff weightFlaps Up ...........................................+ 3.8G, -1.52GFlaps Down .......................................+ 2.0G, 0.0GLanding.............................................+ 3.5G

550-0627 and after at 14,100 LBS maximum takeoff weight

Flaps Up ...........................................+ 3.8G, -1.52G

Flaps Down .......................................+ 2.0G, 0.0GLanding.............................................+ 3.38G at 13,500 LB

landing weight

Note: These accelerations limit the angle-of-bank in turns andseverity of pullup maneuvers.

Note: This airplane is certificated in the normal category. Thenormal category is applicable to aircraft intended for non-aero-batic operations. Aerobatic maneuvers and spins are prohib-ited. No intentional stalls are permitted above 25,000 feet or atany altitude with engine speeds between 61.0% and 65% N1.

R-6/9/00





Enroute Limitations

550-0626 and earliernot incorporating SB55-54-4:

Maximum Operating Altitude ............43,000 FeetTemperature Limits ...........................ISA +39°C*

Generator LoadUp to 25,000 Feet ..........................400 AmpsAbove 25,000 Feet ........................250 Amps

incorporating SB55-54-4:Maximum Operating Altitude ............43,000 FeetTemperature Limits ...........................ISA +39°C*Generator Load

Up to 35,000 Feet ..........................400 AmpsAbove 35,000 Feet ........................325 Amps

550-0627 and after:

Maximum Operating Altitude ............43,000 FeetTemperature Limits ...........................ISA +39°C*Generator Load

Up to 35,000 Feet ..........................400 Amps

Above 35,000 Feet ........................325 Amps

*Note: Maximum enroute operating temperature limit is ISA+39°C ambient temperature adjusted for ram rise or indicated

outside air temperature (IOAT), whichever is less.

Approved Operations

The Citation II is approved for the following types of operationwhen the required equipment is installed and operational asdefined within the Federal Aviation Regulations:

1. VFR day

2. VFR night

3. IFR day including Category I and Category II approaches

4. IFR night including Category I and Category IIapproaches

5. Flight into known icing conditions

R-6/9/00




100%=32,760 100%=15,904

N2 N1

Thrust

Setting

Time Limit

Minutes

Maximum

ITT C RPM % RPM %

Oil Pressure

PSIG (2)

Oil Temp.

C

Takeoff 5 700 (4) 31,450 96 16,540 104 (6) 70 - 85 (5) 10 - 121

Maximum

Continuous Continuous 680 31,450 96 16,540 104 (6) 70 - 85 0 - 121

Maximum

CruiseContinuous 670 31,450 96 16,540 104 (6) 70 - 85 0 - 121

Idle Continuous 58016,000

(min)

49.0

(3)--- ---

35

(min)-40 - 121

Starting (6) --- (1) --- --- --- --- ----40

(min)

Transient

(<0626)

Acceleration

(>0627)

------

700 (4)700

31,45031,450

9696

16,54016,540

104104

(5)---

0 - 1210 - 121

Engine Operating Limitations

Number of Engines ...........................2Engine Manufacturer ........................Pratt & Whitney Canada, Inc.Engine Model ....................................JT15D-4Engine Type ......................................Medium-bypass, axial-flow turbofan

Engine Bypass Ratio.........................2.7 to 1Engine Thrust Rating ........................2500 LBS each

1. Maximum ITT limited to 2-seconds during engine start.

2. Normal oil pressure is 70 to 85 PSIG at engine speeds above 60% N2. Oil pressures under 70PSIG are undesirable, and are allowed only under emergency conditions in order to complete a

flight. Oil pressures below 35 PSIG are unsafe and require engine shut down, or landing as soon aspossible using minimum power required to sustain flight.

3. Idle turbine RPM is 49, ±0.5% with ignition on. A minimum decrease of 0.5% will be noted with

ignition off.4. ITT indications in excess of 700°C during takeoff or in excess of 680°C for more than 5 minutes

require reference to the Engine Maintenance Manual.5. The maximum transient oil pressure can be 95 PSIG for 90-seconds.

6. Refer to the appropriate thrust setting charts for percent fan RPM (N1) setting.

Engine Fan

To ensure accurate fan speed thrust indication, the fan must beinspected for damage prior to each flight.

R-6/9/00





Battery and Starter Cycle Limitations

Whether powered by battery, external power unit, or cross startwith generator assist, starter operation is limited to three enginestart attempts per 30-minute period, with a minimum 30-secondrest period between cycles.

Battery cycling is limited to three engine start attempts perhour.

Battery Limitation

1. If battery limitation is exceeded, a deep cycle, including acapacity check, must be accomplished to detect possiblecell damage. Refer to Chapter 24 of the MaintenanceManual for procedure.

2. Three generator assisted cross starts are equal to one bat-tery start.

3. If an external power unit is used for start, no battery cycle iscounted.

4. Use of an external power source with voltage in excess of28 VDC or current in excess of 1000 amps, may damagethe starter.

Note: Starting ITT exceeding 500°C should be investigated inaccordance with Maintenance Manual.

Note: If the BATT O’HEAT (BATT O’TEMP, 550-0627 and after)annunciator illuminates during ground operation, do not take off

until after the proper maintenance procedures have been ac-complished.

Prolonged Ground Operations

Continuous engine ground static operation up to and includingfive minutes at takeoff thrust is limited to ambient temperatures

not to exceed ISA + 39°C. Continuous ground operation of thestarter-generator above 325 amps is prohibited. Limit ground

operation of pitot/static heat to two minutes to preclude dam-age to the pitot/static heater. Operation in the GND bleed modeat power settings greater than 70% N2 for the right engine isprohibited.




Oil Limitations

Approved Oils

Mobile Jet Oil II or 254, Exxon Turbo Oil 2380, Aeroshell Tur-bine 560 or 500, Castrol 5000, and Royco Turbine Oil 560 or500. In addition, The engine should be serviced with approvedsynthetic oils listed in the most current revision of P&WC SB7001.

Caution: When changing from an existing lubricant formulationto a “third generation” lubricant formulation (Aero Shell/Royco

Turbine Oil 560 or Mobile Jet 254) the engine manufacturerstrongly recommends that such a change should only be madewhen an engine is new or freshly overhauled. For additionalinformation on use of third generation oils, refer to the engine

manufacturers pertinent oil service bulletins.

Note: Do not mix types or brands of oil.

Should it be necessary to replenish oil consumption loss when

oil of the same brand (as contents in tank) is unavailable, thenthe following requirements apply:

1. The total quantity of added oil does not exceed two USquarts in any 400-hour period.

2. If it is required to add more than two US quarts of dissimilaroil brands, drain and flush complete oil system and refill withan approved oil in accordance with Engine MaintenanceManual instructions.

Should oils of non-approved brands or of different viscosities

become intermixed, drain and flush complete oil system andrefill with an approved oil in accordance with Engine Mainte-nance Manual instructions.

Note: Minimum starting oil temperature is -40°C.





Fuel Limitations

Approved Fuels

The following approved fuels comply with the latest revision ofPratt & Whitney Canada Specification 204 and Pratt & Whitney

Canada Service Bulletin 7144R14.

Approved fuels are JET A, JET A-1, JET B, JP-4, JP-5, or JP-8,

all with 0.15% PFA55MB anti-icing additive in solution. Whenpreblended fuel is not available, anti-icing additives conformingto MIL-I-27686E (Ethylene Glycol Monomethyl Ether (EGME)) orMIL-I-85470 (Diethylene Glycol Monomethyl Ether (DIEGME))specifications such as “Prist” may be introduced directly intothe nozzle fuel stream during servicing. Concentrations of lessthan 0.06% (20 fluid ounces of additive per 260 gallons of fuel

or more) may be insufficient to prevent fuel system icing ormicrobiological contamination. Conversely, concentrations of

more than 0.15% (20 fluid ounces of additive per 104 gallons offuel or less) could cause damage to internal components of thefuel system or erroneous fuel quantity indications.

Caution: EGME and DIEGME are aggressive chemicals andshould not exceed 0.15% of fuel volume. Improperly handled,these materials will damage the epoxy primer and sealantsused in the fuel tanks, O-ring seals, and any part of the

airplane’s exterior finish with which it comes in contact.

Warning: Anti-icing additives containing EGME or DIEGME areharmful if inhaled, swallowed, or absorbed through the skin,and will cause eye irritation. Refer to all instructions and warn-ings regarding toxicity and flammability before using thesematerials.

All grades of aviation gasoline (AVGAS) conforming to MIL-G-

5572 specifications are approved for use under emergencycircumstances only. If used during flight, boost pumps shouldbe activated and airplane altitude should not exceed 18,000feet. Use of AVGAS is limited to no more than 3500 US gallonsor 50 hours of engine operation during any period betweenengine overhaul. For record keeping purposes, 1 hour of en-gine operation may be considered equivalent to 70 US gallons.




Fuel Temperature and Density Limitations

Approved Fuel Types

Jet A, A-1,

A-2, Jet B, Aviation

JP-5, -8 JP-4 Gasoline

Minimum Fuel

Temperature(Takeoff) -40°C -54°C -54°C

(Starting) -40°C -54°C -54°C

Maximum

Fuel Temperature +50°C +50°C +32°C

Maximum Altitude 43,000’ 43,000’ 18,000’

Fuel Control Density

(Adjustment forOptimum Engine

Acceleration) 0.81 0.79 0.73

Maximum Fuel Imbalance

Maintaining fuel load symmetry during servicing is unneces-sary; however, the maximum permissible asymmetry is 200LBS during normal flight operations and 600 LBS in anemergency.

Hydraulic Fluid Limitations

The only approved hydraulic fluids are Skydrol 500A, B, B-4, C,or LD-4 or Hyjet W, Hyjet III, IV, or IVA.





Flight Crew Limitations

Minimum flight crew required for Category I operations is onepilot who holds a C-500 type rating and who satisfies require-ments of FAR 61.58 for two-pilot operation, and one copilot whoholds a multi-engine rating and satisfies requirements of FAR

61.55. Category II operation requires a pilot and copilot whoboth satisfy requirements of FAR 61.3.

Cabin Limitations

For takeoff and landing, all seats must be upright and outboard.The seat adjacent to the emergency exit must be fully trackedtoward the rear of the airplane to ensure unobstructed accessto the emergency exit.

To meet smoke detection criteria, the cabin (OVHD) fan mustbe operating any time the aft cabin privacy curtain is closed. If

the fan is inoperable, the curtain must remain open unless thetoilet is in use.

Pressurization Differential

Normal (both valves) ..............................0.0 to 8.8 PSI ±0.1 PSI

Pressurization Source Selector

On airplanes 550-0481 and earlier, 0483 and 0484, operation inBOTH HI mode is not approved for takeoff, landing or at highpower settings.




Icing Limitations

All anti-ice systems must be activated when operating in visiblemoisture at indicated outside air temperatures (IOAT) between+4°C (39°F) and -30°C (-22°F). The surface deice systemshould be activated when ice accumulations of at 1/4” to 1/2”

are observed on the leading edge of either outboard wing. Acti-vation of the system with accumulations of less than 1/4” mayresult in ice bridging on the wing. Accumulations greater than

1/4" may exceed the system's ice removal capabilities. Opera-tion and/or testing of the system at IOAT below -40°C (-40°F)may result in boot cracking or failure of the boots to fully de-flate.

The aircraft must be clear of all deposits of snow, ice, and frostadhering to the lifting and control surfaces immediately prior to

takeoff.

Prolonged flight in severe icing conditions should be avoidedas this may exceed the capabilities of the aircraft ice protectionsystems.

Note: Isopropyl alcohol conforming to TT-I-735 should be usedfor windshield ice protection.





Thrust Reversing Limitations

During landing roll, reverse thrust power must be reduced toidle (thrust reverser levers at the idle reverse detent position)when airplane speed reaches 60 KIAS.

Maximum reverse thrust is limited to 94% N1 at ambient tem-peratures above -18°C or 92% N1 at ambient temperaturesbelow -18°C.

Maximum allowable thrust reverser deployed time is 15 minutesin any 1-hour period.

Deployment of thrust reversers is prohibited when the aircraft isoperating on sod, dirt, or gravel runways.

The drag chute may not be released while thrust reversers aredeployed.

Oxygen System

The standard diluter demand oxygen mask qualifies as aquick-donning mask only if it is positioned around the neck.

The optional crew oxygen mask is a sweep-on diluter demandmask with selectable pressure breathing. The sweep-on mask

qualifies as a quick-donning mask only if it is properly stowed.

Note: Headsets, eyeglasses or hats worn by the crew may in-terfere with the quick-donning capabilities of the optional oxy-gen masks.




Autopilot

During autopilot operation, either the pilot or copilot must beseated in the flight compartment with seat belt fastened.

The autopilot torque monitor must be functionally tested; if

torque monitor functional test is not successful and/or if the[AP TORQUE] annunciator does not illuminate, autopilot opera-tion is prohibited above 14,500 feet.

Continued autopilot operation is prohibited following abnormaloperation or malfunctioning prior to corrective maintenance.

HF/ADF System

The ADF bearing information may be erratic when keying the

HF transmitter. Should this occur, disregard the ADF bearingduring periods of transmission.

Baggage Limitations

Baggage Compartment Weight Limitations

Maximum nose baggagecompartment load .................................. 350 LBS

Maximum cabin baggagecompartment load .................................. 400 LBS

Maximum tailcone baggagecompartment load .................................. 200 LBS

Baggage Compartment Volume Limitations

Maximum nose baggagecompartment volume .............................17 cubic feet

Maximum cabin baggagecompartment volume .............................34 cubic feet

Maximum tailcone baggagecompartment volume .............................13 cubic feet




Chapter 12Emergency Procedures

Table of Contents

Overview .............................................12-1

Engine Emergency Procedures..............12-2

Engine Fire .........................................12-9

Inadvertent Thrust ReverserDeployment ......................................12-10

Electrical System AbnormalProcedures (550-0626 and earlier) ..... 12-13

Loss of Both Generators ................... 12-16

Battery Overheat ...............................12-18

AC Power Failure...............................12-20

Autopilot Hardover.............................12-21

Environmental SystemAbnormal Procedures.........................12-22

Emergency Descent ........................... 12-23

Spins ...............................................12-26

Ditching ............................................12-27

Forced Landing .................................12-28

Electrical System AbnormalProcedures (550-0627 and after) .......12-29

Battery Overheat ...............................12-32

Loss of Both Generators ................... 12-36

AC Power Failure...............................12-38

Environmental SystemAbnormal Procedures.........................12-41

Emergency Descent ........................... 12-43

Autopilot Hardover.............................12-45




EMERGENCY PROCEDURES 12/99 FOR TRAINING PURPOSES ONLY 12-1

Overview

This chapter provides procedures recommended by the manu-facturer for various emergency situations that may be encoun-tered during operation of the Cessna Citation II. Theprocedures are per the Operating Manual(s) for the corre-

sponding serial number range(s) of airplanes, however, theiruse in this manual is for training purposes only.

The appropriate Operating Manual, FAA approved AirplaneFlight Manual (AFM), Pilot’s Check List, and/or related publica-tions should be refered to for normal, abnormal, and emer-gency procedures specific to your airplane.

A thorough understanding of the airplane’s systems and theirinterrelationships is essential to successfully respond to emer-

gency situations. It is suggested that the correspondingchapter(s) in this manual be referenced for specific systemsdescriptions.

Immediate action or “commit to memory” items for each proce-dure are identified by bold type within a shaded box. Notes,Cautions, and Warnings are provided to amplify theprocedures.




Engine Emergency Procedures

Engine Failure or Fire During Takeoff

(without thrust reversers)

SPEED BELOW V1 – TAKEOFF SHOULD

NORMALLY BE ABORTED:

1. Brakes .......................................... AS REQUIRED

2. Throttles ....................................... IDLE

3. Speed Brakes................................ EXTEND

IF ENGINE FIRE:

4. Accomplish Engine Fire Procedures.

IF ENGINE FAILURE:4. Accomplish Engine Failure/Precautionary

Shutdown Procedure.

Note: To obtain maximum braking performance from the antiskid system, it is required that the pilot apply continuous maxi-mum effort (no modulation) to the brake pedals.

Note: The takeoff field lengths assume that the pilot has maxi-mum effort applied to the brakes at the scheduled V1 speed

during the aborted takeoff.

SPEED ABOVE V1 – TAKEOFF SHOULD NORMALLY

BE CONTINUED:

1. After establishing a positive rate of climb, retract landing

gear (and climb at V2, airplanes 550-0626 and earlier).

2. At 400 feet, retract the flaps at V2 + 10 and

accelerate to VENR.

IF ENGINE FIRE:3. Accomplish Engine Fire Procedures.

IF ENGINE FAILURE:

3. Accomplish Engine Failure/PrecautionaryShutdown Procedure.





Engine Failure or Fire During Takeoff (with thrust reversers)

SPEED BELOW V1 – TAKEOFF SHOULD

NORMALLY BE ABORTED:


2. Throttles ....................................... IDLE


4. Thrust Reverser............................. DEPLOY on

unaffected engine.

5. Reverser Indicator Lights.................CHECK illumination ofARM, UNLOCK, and DEPLOY lights.

6. Thrust Reverser................................REVERSE power onthe unaffected engine.

IF ENGINE FIRE:


IF ENGINE FAILURE:


Note: To obtain maximum braking performance from the anti-skid system, it is required that the pilot apply continuous maxi-mum effort (no modulation) to the brake pedals.

Note: The takeoff field lengths assume that the pilot has maxi-mum effort applied to the brakes at the scheduled V1 speedduring the aborted takeoff.

SPEED ABOVE V1 – TAKEOFF SHOULD

NORMALLY BE CONTINUED: (550-0627 and after)

1. After establishing a positive rate of climb, retract

landing gear.

2. At 400 feet, retract the flaps at V2 + 10 and

accelerate to VENR.

IF ENGINE FIRE:


IF ENGINE FAILURE:





Engine Failure/Precautionary Shutdown

1. Throttle (affected engine).................OFF

Any one or more of the following indications might sug-

gest a precautionary shutdown: abnormal or rising ITT,engine vibration, fluctuating or abnormally high or low N1or N2, abnormal oil pressure or temperature, or erraticfuel flow. Circumstances will normally dictate whether tocontinue to operate the engine with possible furtherdamage or shut it down.

2. Ignition (affected engine) .................OFF

3. Engine Synchronizer ........................OFF

4. Generator (affected engine) ............OFF

5. Electrical Load .................................REDUCE as required.

Airplanes 550-0114 and earlier, 0127, 0296 and after,

and 0115~0126 and 0128~0295 incorporating SB550-54-4: 400 amps maximum up to 35,000 feet, 325 ampsmaximum above 35,000 feet.

Airplanes 550-0115~0126 and 0128~0295 not incorpo-rating SB550-54-4: 400 amps maximum up to 25,000feet, 250 amps maximum above 25,000 feet.

6. Fuel Crossfeed.................................AS REQUIREDDo not exceed asymmetric fuel load of 600 LBS(550-0626 and earlier) or 200 LBS (550-0627 and after).

IF NO FIRE:

7. Firewall Shutoff .................................OPEN

8. Fuel Boost Pump..............................ON

Note: If no fire hazard exists, leave firewall shutoff open andturn boost pump ON to prevent damage to engine fuel pump. Ifengine windmills with firewall shutoff CLOSED or with no indica-tion of oil pressure, refer to engine maintenance manual.

9. Refer to Emergency Restart, One Engine or Single EngineApproach and Landing Procedures.





Engine Failure During Coupled Approach

1. Power (operating engine) .............. INCREASE as required.

Only a small power increase will be required to maintain

approach speed and correct rate of descent.

2. Autopilot and Yaw Damper ............ OFF

3. Airspeed ....................................... VREF + 10 KIAS

Accelerate to VREF + 10 before raising flaps.

4. Rudder Trim .................................. TRIM toward operating

engine. The yaw change will be relatively small since

the operating engine is at an approach power setting.

5. Flaps ........................................... T.O. & APPR

6. Throttle (affected engine).................OFF

IF ENGINE FIRE:

7. Accomplish Engine Fire Procedure.

8. Passenger Advisory Lights ..............PASS SAFETY

9. Passenger Seats ..............................CHECK full upright,outboard and positioned aft or forward to clear exitdoors.

10. Seats, Seat Belts andShoulder Harnesses ........................SECURE

Check Seats locked in the desired position. Check seatbelts snug and shoulder harnesses latched to thebuckle.

11. Fuel Crossfeed.................................CHECK

12. Ignition (operating engine)...............ON

13. Landing Gear ...................................DOWN and LOCKED

14. Anti-Skid...........................................CHECK ON

15. Annunciator Panel............................CHECK

With one engine shut down by the throttle, the appropri-ate [OIL PRESS WARN] and [HYD FLOW LOW] annun-ciators (550-0550 and after), or [OIL PRESS LO] and

[HYD PRESS LOW] annunciators (550-0505 and earlier),and [GEN OFF] annunciator will be illuminated. If low fuepressure causes automatic boost pump activation priorto shut down by the throttle, the appropriate [FUELBOOST ON] annunciator will also be illuminated. If auto-matic boost pump activation does not occur prior to shutdown by the throttle, the appropriate [FUEL LOWPRESS] (550-0550 and after) or [FUEL PRESS LO] (550-0505 and earlier) annunciator will be illuminated instead.




With one engine shut down by the firewall shutoff switch,the appropriate [OIL PRESS WARN], [HYD FLOW LOW],and [FUEL LOW PRESS] annunciators (550-0550 andafter), or [OIL PRESS LO], [HYD PRESS LOW], and[FUEL PRESS LO] annunciators (550-0505 and earlier),and the [F/W SHUTOFF] and [GEN OFF] annunciators

will be illuminated. If the boost pump is automaticallyactivated by low fuel pressure, the appropriate [FUEL

BOOST ON] annunciator will also be illuminated. Shouldthis occur, the corresponding FUEL BOOST switchshould be positioned to OFF. If the [MASTER WARNING]light is flashing, it should be extinguished to reduce dis-traction.

16. Flaps ...............................................LAND(when landing assured)

At the pilot’s discretion, flaps may be left at T.O. & APPRor lowered to LAND. If T.O. & APPR flaps are used,maintain VREF + 10 KIAS or “on speed” angle-of-attack (ifoptional AOA indicator installed). LAND flaps are usedunder most conditions since little pitch change is en-countered when they are selected and touchdownspeed can be reduced.

17. Airspeed ..........................................VREF

18. Pressurization ..................................CHECK ZERODIFFERENTIAL

Passing approximately 500 feet above ground level(AGL), check the cabin differential pressure near zero. Ifit is in excess of about one half PSI, select a higher cabinaltitude and adjust RATE to ascend the cabin. Differen-tial pressure should be at zero for landing. Any pressureexisting at touchdown will be dumped by the outflow

valves (actuated by the left main gear squat switch) andmay cause discomfort.

If landing above 12,000 feet pressure altitude, turn the

OXYGEN CONTROL VALVE to CREW ONLY and turnpressurization bleed air OFF to preclude passenger

mask deployment.

19. Speed brakes ..................................RETRACTED prior to 50 feet.

Note: Do not allow Turbine RPM (N2) to be less than 49%.





Emergency Restart – One Engine

FOLLOWING SHUTDOWN, WITH STARTER ASSIST:

1. Throttle .............................................OFF

2. Generator .........................................GEN

3. Firewall Shutoff .................................CHECK OPEN4. Ignition .............................................ON

5. Start Button ......................................PRESS momentarily.

Generator cross start is disabled with weight off the leftmain gear squat switch to preclude generator damagefrom excessive N2 on the operating engine.

6. Throttle .............................................IDLE at 8-10% N2

7. Engine Instruments ..........................MONITORMaximum start ITT 700°C for two seconds.

8. Ignition .............................................NORM

9. If Start Does Not Occur....................PRESS starterdisengage switch.

FOLLOWING SHUTDOWN – WINDMILLING WITH AIRSPEED

ABOVE 200 KIAS:

1. Throttle .............................................OFF

2. Firewall Shutoff .................................CHECK OPEN

3. Ignition .............................................ON

4. Boost Pump .....................................ON

Associated engine ignition and boost pump switchesmust be selected ON since automatic sequencing andselection of these functions does not occur when thestart button is not utilized.

5. Throttle .............................................IDLE

With airspeed maintained above 200 KIAS, throttleshould be brought to IDLE. An N2 of 8-10% is not

required.

6. Engine Instruments ..........................MONITOR

After engine stabilizes:

7. Boost Pump and Ignition..................NORM

It may be necessary to select the associated generatorRESET position momentarily to reinstate the generatorfollowing a windmilling airstart. Maximum start ITT

700°C for two seconds.

8. Generator .........................................GEN




Emergency Restart – Two Engines

1. Ignition ......................................... BOTH ON

2. Boost Pumps ................................ BOTH ON

Engine ignition and boost pump switches must be se-

lected ON since automatic sequencing and selection of

these functions does not occur when the start button is

not utilized.

3. Throttles ....................................... IDLE

Throttles remain at idle for attempted immediate

light-off.

4. If Altitude Allows........................... INCREASE AIRSPEED

to 200 KIAS. Possibilities of immediate start are in-

creased if airspeed is over 200 KIAS.

5. Firewall Shutoff .................................CHECK OPEN

6. All Anti-ice Switches ........................OFFThey are turned OFF to minimize engine bleed air loss.

IF NO START IN TEN SECONDS:

7. Either Start Button ............................PRESS momentarily.Attempt a starter assist restart if altitude and time permit.

Maximum Glide – Emergency Landing (550-0626 and earlier)

1. Airspeed ..........................................BEST GLIDE AT

9500 LBS - 120 KIAS. Increase speed 3 KIAS for each500 LB increase in weight.

2. Flaps ...............................................UP

3. Speed Brakes ..................................RETRACT

4. Landing Gear ...................................UP

5. Transponder ....................................EMERGENCY

6. ATC ...............................................ADVISE

7. Passenger Advisory Switch .............PASS SAFETY

8. Shoulder Harness ............................SECURE

9. Landing Gear, Speed Brakes,and Flaps .........................................AS REQUIRED

for landing anticipated.





Engine Fire (engine fire annunciator/switch illuminated)

1. Throttle (affected engine) .............. IDLE

IF ANNUNCIATOR REMAINS ILLUMINATED:

2. Engine Fire

Annunciator/Switch ...................... LIFT COVER and PUSH.

Cuts off fuel to engine, hydraulic fluid supply to engine-

driven pump, trips the generator field, positions a valve

to allow both bottles to be fired into the affected engine

and illuminates the bottle armed lights.

3. Either Illuminated Bottle

Armed Annunciator/Switch ........... PUSH

4. Ignition .............................................NORMIf ignition is ON, return the switch to NORM.

5. Throttle (affected engine).................OFF6. Reduce Electrical Load....................AS REQUIRED

Airplanes 550-0114 and earlier, 0127, 0296 and after,and 0115~0126 and 0128~0295 incorporating SB550-54-4: 400 amps maximum up to 35,000 feet, 325 ampsmaximum above 35,000 feet.


7. Boost Pump .....................................OFFIf pump is ON, return the switch to OFF.

IF FIRE WARNING ANNUNCIATOR ILLUMINATED

AFTER 30-SECONDS:

8. Remaining Illuminated BottleArmed Annunciator/Switch ..............PUSH

9. Land as soon as possible.

IF ANNUNCIATOR EXTINGUISHED AND SECONDARY

INDICATIONS ARE NOT PRESENT:2. Land as soon as possible.




Inadvertent Thrust Reverser Deployment

Inadvertent Deployment of Thrust Reversers During Takeoff

SPEED BELOW V1 – TAKEOFF SHOULD BE ABORTED:


2. Throttles ....................................... IDLE


4. Thrust Reversers ........................... BOTH DEPLOY

5. Reverser Indicator Annunciators .....CHECK illumination ofARM, UNLOCK and DEPLOY annunciators.

6. Thrust Reversers ..............................REVERSE power onboth engines.

SPEED ABOVE V1 – TAKEOFF SHOULD NORMALLY

BE CONTINUED:

1. Emergency Stow Switch ................ ACTUATE on

affected engine.

2. After establishing a positive rate-of-climb, retract landing

gear. Do not exceed 125 KIAS until thrust reverser stows.

3. At 400 feet, retract flaps at V2 + 10 and accelerate. Do notexceed 200 KIAS after thrust reverser stows.

4. Land as soon as practical.

IF THRUST REVERSER WILL NOT STOW:

5. Thrust Reverser Circuit Breaker .......CHECK in

6. Throttle (affected engine).................CUTOFF

7. Airspeed ..........................................MAINTAIN 150 KIASor below.

8. Refer to Abnormal Procedures for landing.





Inadvertent Deployment of Thrust Reversers In Flight

1. Reverser Indicator Annunciators .... CHECK illumination of

ARM, UNLOCK and DEPLOY annunciators.

2. Affected Throttle ........................... CHECK idle


affected engine.

4. Airspeed ....................................... REDUCE to 125 KIAS

(115 KIAS with flaps extended) or below. After thrust

reverser stows, do not exceed 200 KIAS.

5. Reverser Indicator Annunciators:UNLOCK and DEPLOYAnnunciator......................................EXTINGUISHED

ARM and HYD PRESS ONAnnunciator......................................ILLUMINATED

Note: If thrust reverser is stowed, engine may be operated nor-mally. Thrust reverser cannot be used during landing if emer-gency stowed.


IF THRUST REVERSER WILL NOT STOW:


8. Throttle (affected engine).................CUTOFF


10. Refer to Single Engine Approach andLanding Procedures.




Thrust Reverser UNLOCK Annunciator Illuminated In Flight


affected engine.

2. Thrust Reverser Levers .................. CHECK thrust

reverser levers at stowed (full forward) position.

IF ANNUNCIATORS WILL NOT EXTINGUISH:


4. Maintain 200 KIAS or below.


Thrust Reverser ARM Annunciator Illuminated In Flight

(550-0626 and earlier)

1. Thrust Reverser Levers .................... CHECK stowed(full forward)

2. Emergency Stow Switch ..................VERIFY OFF

IF ARM ANNUNCIATOR STILL ILLUMINATED:


4. HYD PRESS ON Annunciator ...........CHECK

IF HYD PRESS ON ANNUNCIATOR ILLUMINATED

(T/R ISOLATION VALVE IS OPEN):5. Affected Thrust Reverser

Circuit Breaker .................................PULL

6. Land as soon as possible (affected T/R should beinoperative).

IF HYD PRESS ON ANNUNCIATOR NOT ILLUMINATED

(PROBABLE PRESSURE SWITCH PROBLEMS):


Note: With a thrust reverser circuit breaker pulled, the emer-gency stow system of the opposite reverser is deactivated.





Electrical System Abnormal Procedures

(550-0626 and earlier)

Electrical Fire or Smoke

1. Oxygen Masks and

Oxygen MIC Switches ................... AS REQUIRED.

Ensure selector is on 100% oxygen when masks are used.

Ensure oxygen MIC switch is in MIC OXY MASK position.

2. Pressurization Source Selector: .......NORM(airplanes 550-0481, and 550-0485~0626)

Pressurization Source Selector: .......BOTH HI(airplanes 550-0481 and earlier, 550-0483, and 550-0484)

KNOWN SOURCE OF FIRE:

3. Isolate faulty circuit. Pull circuit breaker to remove powerfrom faulty equipment.

UNKNOWN SOURCE OF FIRE:

3. Flood Lights .....................................FULL BRIGHT

4. Battery Switch ..................................EMER

Will have COMM 1, NAV 2, copilot’s HSI and cockpitfloodlights after generators turned off.

5. Generators .......................................OFFWith the battery switch in the emergency position and

the generators off, power is supplied for approximately30-minutes to COMM 1, NAV 2, overhead floodlights andcopilot’s HSI.




Caution: When landing with emergency power (Battery Switch-EMER and both generators off), the following are not available:

a. The landing gear normal operation is not available. Thelanding gear must be lowered using the blowdown systemand the landing gear warning lights will not illuminate.

b. The flaps will not operate. If not previously in landing posi-tion, a flaps inoperative landing must be made.

c. The anti-skid/power brake system is inoperative. Only theemergency brake system is available.

d. The engine anti-ice valves will be open. Refer to anti-ice onthrust charts.

e. The outside air temperature gage is not reliable, so usecaution when applying power (except for go-around whereground temperatures can be used).

f. All engine instruments except the vertical tape N1 will beinoperative. The vertical tape N1 will indicate erraticallybelow approximately 50% N1, but will give reliable indica-tions above 50% N1.

6. MIC Selector ....................................EMER/COMM 1

Must be in the EMER/COMM 1 position to transmit whenoperating on emergency battery, and pilot must wear

headset to receive.

7. Receiver Select ................................COMM 1 to HDPH

(550-0356~0626)(required only if AUTO SELECT is OFF)

8. All Electrical Switches ......................OFF

9. Windshield Bleed AirManual Valves..................................OFF

With electrical power lost, the windshield bleed airshutoff valve will fail open. The bleed air manual valvesare closed to prevent an excessive volume of hightemperature air from reaching the windshield.

10. DC Power RH Bus............................PULL No. 1,2,3 CB’s

11. RH CB Panel Circuit Breaker(LH Panel) ........................................PULL

12. AC Inverter No. 1 Circuit Breaker(LH Panel) ........................................PULL

13. Land as soon as practical (within 30-minutes).





IF SEVERITY OF SMOKE WARRANTS:

13. Initiate Smoke Removal and/or Emergency Descentprocedures. Land as soon as possible.

WHEN LANDING ASSURED:

14. LH Generator ...................................ON

15. Landing Gear ...................................DOWN

16. Flaps ...............................................LAND

17. Airspeed ..........................................VREF

Note: Right thrust reverser (if applicable) will be inoperative.

IF FIRE OR SMOKE STARTS AGAIN:

18. LH Generator ...................................OFF

Note: Anti-skid system will be inoperative. Power brakes will beavailable until accumulator discharges. Multiply landing dis-tance by 1.6. Be prepared to use the emergency brakesystem.




Loss of Both Generators

1. Generators .......................................RESET then GENAttempt to reset both generators.

IF ONLY ONE GENERATOR COMES ON:

2. Electrical Load .................................REDUCE as required.Airplanes 550-0114 and earlier, 0127, 0296 and after,

and 0115~0126 and 0128~0295 incorporating SB550-54-4: 400 amps maximum up to 35,000 feet, 325 ampsmaximum above 35,000 feet.


IF NEITHER GENERATOR COMES ON:


3. Battery Switch ..................................EMERWith the battery switch in emergency position and thegenerators off, power is supplied for approximately 30-minutes to COMM 1, NAV 2, overhead floodlights, volt-meter and copilot’s HSI.

On airplanes with single EFIS, with the battery switch inthe emergency position and the generators off, power issupplied for approximately 30-minutes to COMM 1, NAV2, Copilot’s HSI and DG 2. and copilot’s audio panels,

copilot’s attitude indicator, voltmeter, cockpit floodlightsand standby gyro. Airplanes 550-0682 and after alsohave the RH pitot static heater on the emergency bus.

On airplanes with dual EFIS, with the battery switch inemergency position and the generators off, power issupplied for approximately 30-minutes to COMM1, NAV2, copilot’s RMI, NAV 2 repeater indicator, pilot’s andcopilot’s audio panels, voltmeter, cockpit floodlights andstandby gyro. Airplanes 550-0682 and after also havethe RH pitot static heater on the emergency bus.







b. The flaps will not operate. If not previously in the landingposition, a flap inoperative landing must be made.

c. The Anti-skid/power brake system is inoperative.Only the emergency brake system is available.




4. MIC Selector ....................................EMER/COMM 1

Must be in EMER/COMM 1 position to transmit whileoperating on emergency battery and headset worn to

receive. Volume is controlled at the radio since normalamplification is bypassed.

5. Receiver Select ................................COMM 1 to HDPH(550-0356 and after)Required only if Auto-Select is OFF


With electrical power lost, the windshield bleed airshutoff valve will fail open. The bleed air manual valves

are closed to prevent an excessive volume of high tem-perature air from reaching the windshield.



7. Battery Switch ..................................BATT


9. Flaps ...............................................LAND

10. Airspeed ..........................................VREF




Battery Overheat

BATT O’TEMP Annunciator Illuminated

(temp between 145° and 160°F)

1. Battery Switch .............................. EMER

In EMER position the battery will be disconnected from

the generators and will no longer be charged by them.

All electrical equipment will continue to receive power

since the generators are still on line. The Emergency DC

bus is powered by the battery. Battery voltage may now

be read with the voltage selector in BATT and generator

bus voltage with the voltage selector in LH GEN or RH

GEN. Individual generator voltages can be read by se-

lecting one (LH or RH) GEN and turning the other gen-

erator off.

2. Amperage ..................................... NOTE decrease

3. If battery voltage is 1 volt less than generator voltage in 30-seconds to 2-minutes, monitor battery overheat annunciatorfor possible change. In thirty seconds to two minutes afterdisconnect, battery voltage should read at least one voltless than the generators. Rotate the voltage selector to LHGEN and RH GEN position to read generator voltage. Bat-tery voltage will be indicated when the voltage selector is inthe BATT position.

IF BATT O’TEMP (BATT O’HEAT, 550-0505 and earlier)ANNUNCIATOR EXTINGUISHED:


IF NO AMP DECREASE OR BATT O’TEMP (BATT O’HEAT,

550-0505 and earlier) ANNUNCIATOR FLASHES:


5. Generators .......................................OFFSince the battery has continued to overheat, it may bebecause the battery is still being charged through afailed battery relay. Turning both generators OFF tripsthe generators and opens the power relays, isolatingeach generator from its bus. The BATT O’TEMP annun-ciator will extinguish immediately when the generatorsare turned off if the battery relay is not stuck.

With the battery in the emergency position and the gen-erators off, power is supplied for approximately 30-min-utes to COMM 1, NAV 2, overhead floodlights andcopilot’s HSI.





Caution: When landing with emergency power (Battery Switch-EMER and both generators off), the following are not avail-

able:






f. All engine instruments except the vertical tape N1 will beinoperative. The vertical tape N1 will indicate erratically

below approximately 50% N1, but will give reliable indica-tions above 50% N1.

IF NORMAL DC POWER LOST(BATTERY RELAY NOT STUCK):

6. Generators .......................................GEN

(BATT O’TEMP annunciator will illuminate untilbattery cools).

7. Battery Switch ..................................OFF

Caution: With the battery switch off, all power from the emer-gency bus will be removed. After landing, refer to maintenancemanual for proper maintenance procedures as damage to thebattery may have occurred.

8. Land as soon as practical

IF NO DC POWER LOST (BATTERY RELAY STUCK):

6. Mic Selector .....................................EMER/COMM 1(headphones required to receive audio)

7. Receiver Select ................................COMM 1 to HDPH(required only if Auto Select is OFF)(550-0356 and after),


9. DC Power LH and RH Bus ...............PULL No. 1,2,3 CB’s10. Land as soon as practical.





11. DC Power LH and RHCircuit Breakers ...............................RESET


13. Flaps ...............................................LAND

14. Airspeed ..........................................VREF

Caution: After landing, refer to maintenance manual for propermaintenance procedures as damage to the battery may haveoccurred.

AC Power Failure

Both INVERTER 1 FAIL and INVERTER 2 FAIL –

Annunciators Illuminated

(Airplanes with dual AC Busses – 550-0550~0626):

1. Inverter 1 and Inverter 2Circuit Breakers ...............................RESET

2. Battery Switch ..................................EMERIf the inverters will not come back on line after the circuitbreakers have been reset, complete the flight by usingthe copilot’s air-driven attitude indicator or the standbygyro horizon (if installed). Placing the battery switch toEMER will provide AC power from the copilot’s C-14Dstatic inverter to power the copilot’s compass system

and NAV 2.

AC Power and/or Distribution Failure (AC FAIL Annunciator

Illuminated After MASTER WARNING Has Been Reset,

INVERTER FAIL 1 or 2 Annunciators Extinguished –

Airplanes 550-0550~0626):

1. Check the right sub-circuit breaker panel for disengagedAC BUS circuit breaker(s)





Caution: If circuit breaker(s) is/are disengaged, operate withloss of bus as re-engagement may result in further damage tothe electrical system.

Note: Depending on which buses have been lost, the followingequipment will be lost:

a. 26 VAC Bus 1: pilot’s ADI, pilot’s RMI, pilot’s HSI (HDG andNAV flags in view), pilot’s rate-of-turn indicator, copilot’sRMI compass card and NAV 1 bearing pointer, and ADF 1.

b. 115 VAC Bus 1: pilot’s flight director, autopilot, yawdamper, radar, pilot’s attitude gyro, and VNAV computer/

controller.

c. 26 VAC Bus 2: NAV 2, copilot’s ADI (optional dual flightdirector installation), copilot’s HSI, copilots RMI, pilot’s RMIcompass card and NAV 2 bearing pointer, and ADF 2 (op-

tional). Operation of the following equipment can be rein-stated by placing the BATT switch to EMER: NAV 2,bearing pointer and DG of copilot’s HSI.

d. 115 VAC Bus 2: Copilot’s flight director (optional), air datacomputer and pilot’s altimeter. The auto air data computerand pilot’s altimeter. The auto pilot will only operate in basicautopilot modes due to loss of valid signal from the air datacomputer.

Autopilot Hardover

1. Autopilot/Trim

Disengage Switch ......................... PRESS

Press switch on either yoke. Flight director modes re-

main selected.

2. Maximum Altitude Losses during Autopilot Malfunction:

550-0161 and earlier 550-0162 and after

a. Cruise 450 feet at 43,000 feet 550 feet at 43,000 feet

b. Climb 50 feet at 10,000 feet 300 feet at 17,000 feetc. Maneuvering 110 feet at 43,000 feet

d. ILS 37 feet (autopilot must 34 feet (autopilot must

be off at 100 feet) be off at 90 feet)




Environmental Systems Abnormal Procedures

Rapid Decompression

1. Oxygen Masks .............................. DON

Check oxygen selector on 100%.

2. Emergency Descent ...................... AS REQUIRED

3. Pass Oxygen ................................. ENSURE passengers

are receiving oxygen. Visually check mask drop

when cabin reaches 13,500 ± 600 feet If masks are

not down, drop them by the PASS OXY MASK

switch on the left console (OXYGEN CONTROL

VALVE, 550-0550~0626).

4. Oxygen MIC Switches ................... MIC OXY MASK

Switch to MIC OXY MASK in order to use microphone inoxygen mask.

5. Transponder ....................................EMER 7700

6. Refer to use of Supplemental Oxygen Procedures in theAbnormal Procedures.





Emergency Descent

1. Throttles ....................................... IDLE

2. Speed brakes ................................ EXTEND

3. Initiate moderate bank.

4. Airplane Pitch Attitude .................. 15° NOSE DOWN


6. Maximum Airspeed..........................VMO /MMO

MMO (above 28,000 feet) ..................0.705 MACH

VMO (14,000-28,000 feet) .................277 KIAS

VMO (below 14,000 feet) ...................262 KIAS

Use reduced speed if structural damage has occurred.

7. Transponder ....................................EMERGENCY 7700

Environmental System Smoke or Odor

1. Oxygen Mask and OxygenMIC Switches ...................................AS REQUIRED.

Oxygen selector on 100% and MIC oxygen switch inMIC OXY MASK position in order to use MIC inoxygen mask.

2. Cabin (ovhd) Fan .............................OFF

3. Defog Fan ........................................OFF

Both cabin and cockpit defog fans off to prevent fur-

ther circulation of smoke through the aircraft and pos-sibly identify them as the source.

4. Pressurization SourceSelector ............................................ISOLATE source by

selecting LH.

Note: Pressurization source selector must remain in each posi-tion long enough to allow adequate system purging to deter-mine the source of smoke. If smoke has not begun to clear in a

minute, switch to another source.

IF SMOKE CONTINUES:

5. Pressurization Source Selector ........RH(allow time for smoke to dissipate)

IF SMOKE STILL CONTINUES (AIR CYCLE MACHINE SEAL MAY

BE LEAKING):

6. Pressurization Source Selector ........EMER(control cabin pressure with LH throttle)




Smoke Removal

Note: No action is normally required; however, if smoke isintense:

1. Oxygen Mask...................................DON. Check oxygen

selector is on 100%.

2. PASS OXY MASKS...........................MANUAL DROP(OXYGEN CONTROL VALVE, 550-0550~0626)

3. CREW OXY PRIORITY Valve............CHECK NORMAL(550-0505 and earlier)

4. Ensure passengers are receiving oxygen. Visually checkmask drop when cabin reaches 13,500 ± 600 feet. If masksare not down, drop them by PASS OXY MASKS or OXYGENCONTROL VALVE (550-0550~0626) switch on the left con-sole.

5. Oxygen Mic Switches ......................MIC OXY MASK

Switch must be in this position to use microphone in theoxygen mask.

6. Passenger Advisory Annunciator.....PASS SAFETY

7. Cabin Altitude Selector ....................SET to higher

cabin altitude.

8. Pressurization Source Selector ........BOTH HI(550-0481 and earlier, 0483, and 0484)

9. Emergency Dump Switch ................DUMP

This switch manually opens the normal dump valve torapidly depressurize the airplane, allowing the smoke toclear. All smoking material should be extinguished.

10. Refer to use of Supplemental Oxygen Procedures in theAbnormal Procedures.

IF SMOKE PERSISTS OR IT CANNOT BE VERIFIED THAT THERE

IS NO FIRE:

11. Land as soon as possible.





Overpressurization

1. Cabin Altitude Selector ....................SET to higher cabin al-titude. Cabin altitude will be descending and differen-tial pressure will be increasing. Attempt to decreasethe differential pressure by calling for a higher cabin

altitude forcing the outflow valve to open.

2. Rate Control .....................................INC

Go to full increase to program the controller to work asrapidly as possible.

3. Pressurization Source Selector ........LH or RHAttempt to control cabin pressure with the appropriate

throttle by reducing power, thereby letting a smalleramount of air into the aircraft to pressurize the cabin.

IF UNABLE TO CONTROL:

4. Oxygen Masks .................................DON


5. PASS OXY MASK .............................MANUAL DROP(OXYGEN CONTROL VALVE, 550-0550~0626)

6. CREW OXY PRIORITY Valve............CHECK NORMAL(550-0505 and earlier)

7. Assure passengers are receiving oxygen. Visually checkmask drop when cabin reaches 13,500 ± 600 feet. If masksare not down, drop them by the PASS OXY MASKS or OXY-

GEN CONTROL VALVE (550-0550~0626) manual switch onthe left console.

8. Oxygen MIC Switches .....................MIC OXY MASK

Switch to MIC OXY MASK in order to use microphonein oxygen mask.

9. Passenger Advisory Light ................PASS SAFETY

10. Pressurization Source Selector ........OFF

11. Descend

IF STILL OVERPRESSURIZED:12. Emergency Dump Switch ................DUMP

This switch manually opens the normal dump valve to

rapidly depressurize the airplane. All smoking materialshould be extinguished.

13. Refer to the Use of Supplemental Oxygen Procedures in theAbnormal Procedures.




Spins

Intentional spins are prohibited and were not conducted duringflight tests of the aircraft. Should a spin occur, the followingprocedures are recommended:

1. Power to idle on both engines.

2. Neutralize yoke and apply full rudder opposite the directionof rotation.

3. Approximately 1/2 turn of spin after applying rudder, pushyoke forward.

4. Remove rudder input as rotation slows so that rudder iscentered when rotation stops.

5. Pull out of the dive with smooth steady control pressure.

6. Indicated airspeed, or angle of attack if installed, should be

closely monitored during the pullout to avoid a secondarystall.





Ditching

Good crew coordination is essential to the success of anyditching. Radio contact should be attempted giving aircraftidentification, position, heading, altitude, and the transponderset on 7700 and the locator beacon set on EMER. Passengers

should be briefed and don life jackets keeping them uninflateduntil outside the airplane.

Plan on an approach to parallel any uniform swell pattern and

attempt to touch down along a wave crest or just behind it. Ifthe surface wind is very strong or the water surface rough andirregular, ditch into the wind on the back side of a wave.

Gear should be left up with flaps in the LAND position. The LDGGEAR circuit breaker can be pulled to silence the gear warning

horn. Speed should be maintained at VREF with the rate of de-scent at 200-300 feet per minute. Ditch while power is availableif possible, so that the most desirable approach can be made.

Touchdown should be slightly nose high and throttles cut offjust before water contact. Passengers and crew exit throughthe emergency escape hatch inflating life jacket when clear.

1. Radio ...............................................MAYDAY

Identify airplane, position, heading altitude and IAS.

2.Transponder ....................................7700

3. Locator Beacon ...............................EMER


Prevents water from entering through bleed valves.


Check aft facing seats full aft and all seats upright andoutboard.

6. Passenger Life Jackets ....................ON

Life jackets should not be inflated until outside airplane.

7. Gear ...............................................UP8. Flaps ...............................................40°

9. Speed ..............................................VREF

10. Rate of Descent ...............................200-300 feet per min.




11. Plan an approach to parallel any uniform swell pattern andattempt to touchdown along a wave crest or just behind it. Ifthe surface wind is very strong or the water surface roughand irregular, ditch into the wind on the backside of a wave.Airplane should touch down nose high with a minimum rateof descent.

12. Throttles ...........................................CUTOFFjust prior to contact.

13. Emergency Exit ................................OPEN

Forced Landing

All considerations for a successful forced landing are similar to

those for ditching. Attempt to establish radio contact, squawkthe emergency code, and brief passengers.

The approach should be made with gear DOWN, flaps in LANDposition, speed VREF, and a 200-300 feet per minute rate ofdescent.

If possible, establish an abeam position with gear extendedand altitude sufficient to enable a safe landing to be made in

the event of power loss.

Just before touchdown, place throttles in cut-off and turn off thebattery.

Touchdown should be made in a normal landing attitude andemergency braking employed if necessary.

1. Radio ...............................................MAYDAY

Identify airplane, position, heading altitude and IAS.

2. Transponder ....................................7700

3. Locator Beacon ...............................EMER


Brief passengers as thoroughly as possible.

5. Gear ...............................................DOWN

6. FIaps ...............................................40°

7. Speed ..............................................VREF

8. Rate of Descent ...............................AS REQUIRED to effecttouchdown in selected landing area

9. Throttles ...........................................CUTOFF just prior tocontact





Electrical System Abnormal Procedures

(550-0627 and after)

Electrical Fire or Smoke

1. Oxygen Masks .............................. DON and 100%

oxygen

2. Oxygen MIC Switches ................... AS REQUIRED

Ensure selector is on 100% oxygen when masks are

used. Ensure oxygen MIC switch is in MIC OXY MASK

position.

3. Pressurization SourceSelector ............................................NORM

KNOWN SOURCE OF FIRE:

4. Isolate faulty circuit. Pull circuit breaker to remove powerfrom faulty equipment.

UNKNOWN SOURCE OF FIRE:



6. Generators .......................................OFFWith the battery switch in the emergency position andthe generators off, power is supplied for approximately

30-minutes to COMM 1, NAV 2, overhead floodlights,

copilot’s HSI and DG 2, copilot’s attitude indicator, thevoltmeter and both audio panels. Additionally, thestandby gyro battery provides at least 30-minutes ofpower to operate the standby gyro indicator. On air-planes 550-0682 and after, the RH pitot-static heater isalso on the emergency bus. In dual EFIS installations,the CP HSI and CP ADI are not on the emergency bus,

and the CP RMI is on the emergency bus. In dual EFISinstallations when emergency battery power only is avail-able, attitude information is provided only by the standbygyro horizon.






b. The flaps will not operate. If not previously in landing posi-tion, a flaps inoperative landing must be made.

c. The anti-skid/power brake system is inoperative. Only theemergency brake system is available.


e. The outside air temperature gage is not reliable, so usecaution when applying power (except for go-around where

ground temperatures can be used).f. All engine instruments except the vertical tape N1 will be

inoperative. The vertical tape N1 will indicate erraticallybelow approximately 50% N1, but will give reliable indica-tions above 50% N1.

7. Windshield Bleed AirManual Valves..................................OFF or MINIMUM for

clear vision through the windshield.With electrical power lost, the windshield bleed air

shutoff valve will fail open. The bleed air manual valvesare closed to prevent an excessive volume of high tem-

perature air from reaching the windshield.

8. DC Power RH Bus............................PULL No. 1,2,3 CB’s

9. RH CB Panel Circuit Breaker(LH Panel) ........................................PULL

10. AC Inverter No. 1 Circuit Breaker(LH Panel) ........................................PULL

11. Land as soon as practical (within 30-minutes)

IF SEVERITY OF SMOKE WARRANTS:

11. Initiate Smoke Removal and/or Emergency Descentprocedures. Land as soon as possible.






12. LH Generator ...................................ON


14. Flaps ...............................................LAND

15. Airspeed ..........................................VREF

Note: Right thrust reverser (if applicable) will be inoperative.

IF SMOKE OR FIRE STARTS AGAIN:

16. LH Generator ...................................OFF

Note: Anti-skid system will be inoperative. Power brakes will beavailable until accumulator discharges. Multiply landing dis-tance by 1.6. Be prepared to use the emergency brake

system.




Battery Overheat

BATT O’TEMP Annunciator Illuminated

(temp between 145° and 160°F)

1. Battery Switch .............................. EMER

In EMER position the battery will be disconnected from

the generators and will no longer be charged by them.

All electrical equipment will continue to receive power

since the generators are still on line. The Emergency DC

bus is powered by the battery. Battery voltage may now

be read with the voltage selector in BATT, and genera-

tor bus voltage with the voltage selector in LH GEN or

RH GEN. Individual generator voltages can be read by

selecting one (LH or RH) GEN and turning the other gen-

erator off.

2. Amperage ..................................... NOTE decrease

3. If battery voltage is 1 volt less than generator voltage in 30-seconds to 2-minutes, monitor battery overheat annunciator

for possible change. In thirty seconds to two minutes afterdisconnect, battery voltage should read at least one voltless than the generators. Rotate the voltage selector to LHGEN and RH GEN position to read generator voltage. Bat-tery voltage will be indicated when the voltage selector is inthe BATT position.

IF BATT O’TEMP ANNUNCIATOR EXTINGUISHED:4. Battery Switch ..................................BATT





IF NO AMP DECREASE OR BATT O’TEMP

ANNUNCIATOR FLASHES:



6. Generators .......................................OFF

Since the battery has continued to overheat, it may bebecause the battery is still being charged through afailed battery relay. Turning both generators OFF tripsthe generators and opens the power relays, isolatingeach generator from its bus. In standard single EFIS

installations, the emergency DC bus is now powered bythe battery, with COMM 1, NAV 2, voltmeter, pilot’s andcopilot’s audio panels, copilot’s HSI and DG 2, copilot’sattitude indicator and cockpit floodlights available to thecrew. In optional dual EFIS installations, the emergencyDC bus is powered by the battery with COMM 1, NAV 2,

copilot’s RMI, the NAV 2 repeater indicator, pilot’s andcopilot’s audio panels, voltmeter, and cockpit floodlights

available to the crew. On airplanes 550-0682 and afterthe RH pitot-static heater is also on the emergency bus(single and dual EFIS). Emergency battery power will beavailable for approximately 30-minutes under normalconditions. Additionally, the standby gyro battery pro-vides at least 30-minutes of power to operate thestandby gyro indicator. The BATT O’TEMP light willextinghish immediately when the generators are turned

off if the battery relay is not stuck.






b. The flaps will not operate. If not previously in the landingposition, a flaps inoperative landing must be made.

c. The Anti-skid/Power Brake system is inoperative.Only the emergency brake system is available.




IF NORMAL DC POWER LOST

(BATTERY RELAY NOT STUCK):

7. Generators .......................................GEN (BATT O’TEMPlight will come back on until battery cools).

8. Battery Switch ..................................OFF

9.Land as soon as practical.


IF NO DC POWER LOST (BATTERY RELAY STUCK):


clear vision through windshield.

8. DC Power LH and RH Bus ...............PULL No. 1,2,3 CB’s







10. DC Power LH and RHCircuit Breakers ...............................RESET


12. Flaps ...............................................LAND

13. Airspeed ..........................................VREF











e. The outside air temperature gage is not reliable, so usecaution when applying power (except for go-around where

ground temperatures can be used).f. All engine instruments except the vertical tape N1 will be

inoperative. The vertical tape N1 will indicate erraticallybelow approximately 50% N1, but will give reliable indica-tions above 50% N1.


clear vision through windshield. With electrical powerlost, the windshield bleed air shutoff valve will fail open.

The bleed air manual valves are closed to prevent anexcessive volume of high temperature air from reachingthe windshield.





8. Flaps ...............................................LAND

9. Airspeed ..........................................VREF




AC Power Failure

Both INVERTER 1 FAIL, INVERTER 2 FAIL and

AC FAIL Annunciators Illuminated

SINGLE EFIS INSTALLATION:

1. Inverter 1 and Inverter 2 CB’s ..........RESET


If the inverters will not come back on the line after thecircuit breakers have been reset, complete the flight byusing the copilot’s attitude indicator or the standby gyrohorizon. Placing the battery switch to EMER will provideAC power from the copilot’s C-14D static inverter topower the copilot’s compass system and NAV 2. TheEFIS system will be inoperative with electrical systemfailure. With BATT in EMER, NAV 2 and compass infor-

mation will be displayed on the copilot’s HSI, and atti-tude information on the copilot’s attitude indicator andthe standby gyro.

OPTIONAL DUAL EFIS INSTALLATION:

1. Inverter 1 and Inverter 2 CB’s ..........RESET


If the inverters will not come back on the line after thecircuit breakers have been reset, complete the flight byusing the standby gyro horizon. Placing the battery

switch to EMER will provide AC power from the copilot’sC-14D static inverter to power the DG 2 and the copilot’s

RMI, NAV 2, and the NAV 2 repeater indicator.

AC Power and/or Distribution Failure (AC FAIL Annunciators

Illuminated After MASTER WARNING Has Been Reset,

INVERTER 1 and 2 Annunciators Extinguished)

1. Check the right sub-circuit breaker panel for disengagedAC BUS circuit breaker(s).

Caution: If circuit breaker(s) is/are disengaged, operate withloss of bus as re-engagement may result in further damage tothe electrical system.





Note: Depending on which buses have been lost, the followingequipment will be lost:

a. 26 VAC Bus 1: pilot’s ADI, pilot’s RMI, pilot’s HSI (HDG andNAV flags in view), pilot’s rate-of-turn indicator, copilot’sRMI compass card and NAV 1 bearing pointer, and ADF 1.

b. 115 VAC Bus 1: pilot’s flight director, autopilot, yawdamper, radar, pilot’s attitude gyro, and VNAV computer/ controller.

c. 26 VAC Bus 2: NAV 2, copilot’s ADI (optional dual flightdirector installation), copilot’s HSI, copilot’s RMI, pilot’s RMIcompass card and NAV 2 bearing pointer, and ADF 2 (op-tional). Operation of the following equipment can be rein-stated by placing the BATT switch to EMER: NAV 2,bearing pointer and DG of copilot’s HSI.

d. 115 VAC Bus 2: Copilot’s flight director (optional), air datacomputer and pilot’s altimeter. The auto air data computer

and pilot’s altimeter. The autopilot will only operate in basicautopilot modes due to loss of valid signal from the air datacomputer.




EFIS Symbol Generator Failure

(Standard Two-Tube Installation)

RED X AND SG FAIL ON EADI OR BOTH DISPLAYS BLANK:

1. The EFIS system is inoperative. Refer to the copilot’s atti-tude indicator, the standby gyro horizon and the copilot’s

HSI for attitude and compass information.

Symbol Generator Overtemperature (SG HOT Light On)

(Standard EFIS Installation)

1. EFIS 1, EADI 1 and EHSI 1DC Circuit Breakers .........................PULL

The SG HOT annunciator has probably illuminated dueto an overtemp condition caused by failure of the SymbolGenerator internal fan. Continued use of the SymbolGenerator without the fan may lead to its failure. Remov-ing power from the symbol generator will allow it to cool,but restoring power will likely result in another overtempindication. Consideration should be given to leaving the

circuit breaker disengaged, using the copilot’s flightinstruments to complete the flight and, if necessary, re-storing power to the symbol generator for the approachand landing. Pulling the EADI and EHSI circuit breakerswill enable the display tubes to cool. Refer to thecopilot’s attitude indicator, the standby gyro horizon and

the copilot’s HSI for attitude and heading information.





Environmental Systems Abnormal Procedures

Overpressurization

1. Cabin Altitude Selector ....................SET to higher cabin al-titude. Cabin altitude will be descending and differen-

tial pressure will be increasing. Attempt to decreasethe differential pressure by calling for a higher cabinaltitude forcing the outflow valve to open.

2. Rate Control .....................................INC

Go to full increase to program the controller to work asrapidly as possible.

IF STILL OVERPRESSURIZED:

3. Pressurization Source Selector ........LH or RH

Attempt to control cabin pressure with the appropriatethrottle by reducing power, thereby letting a smaller

amount of air into the aircraft to pressurize the cabin.

IF UNABLE TO CONTROL:

4. Oxygen Masks .................................DON. Check oxygenselector on 100%.

5. PASS OXY MASK .............................MANUAL DROP

6. Ensure passengers are receiving oxygen. Visually checkmask drop has occurred.

7. Oxygen MIC Switches .....................MIC OXY MASKSwitch to MIC OXY MASK in order to use microphone inoxygen mask.



10. Descend

IF STILL OVERPRESSURIZED:


This switch manually opens the normal dump valve torapidly depressurize the airplane. All smoking material

should be extinguished.




Rapid Decompression

CAB ALT 10,000 FEET LIGHT WILL ILLUMINATE:

1. Oxygen Masks .............................. DON


2. Emergency Descent ...................... AS REQUIRED

3. Pass Oxygen ................................. ENSURE passengers

are receiving oxygen. Visually check mask drop when

cabin reaches 13,500 ± 600 feet If masks are not

down, drop them by placing the OXYGEN CONTROL

VALVE on the left console to MANUAL DROP.

4. Oxygen MIC Switches ................... MIC OXY MASK

Switch to MIC OXY MASK in order to use microphone in

oxygen mask.

5. Transponder ....................................EMER 7700





Emergency Descent

1. Throttles ....................................... IDLE

2. Speed brakes ................................ EXTEND

3. Initiate moderate bank.

4. Airplane Pitch Attitude .................. 15° NOSE DOWN


6. Maximum Airspeed..........................VMO /MMO

MMO (above 30,500 feet) ..................0.705 MACH

VMO (below 30,500 feet) ...................262 KIAS

Use reduced speed if structural damage has occurred.

7. Transponder ....................................EMERGENCY 7700

IF DESCENT INTO ICING CONDITIONS IS REQUIRED:

8. Throttles ...........................................AS REQUIREDMaintain sufficient power for anti-icing (engine anti-ice

lights remain OFF).

Environmental System Smoke or Odor

1. Oxygen Mask ...................................DON and 100%2. Oxygen MIC Switches .....................AS REQUIRED

Oxygen selector on 100% and MIC oxygen switch in MICOXY MASK position in order to use MIC in oxygen mask.

3. Cabin (OVHD) Fan ...........................OFF

4. Defog Fan ........................................OFF

Both cabin and cockpit defog fans off to prevent furthercirculation of smoke through the aircraft and possibly

identify them as the source.

5. Pressurization SourceSelector ............................................ISOLATE source by

selecting LH.

IF SMOKE CONTINUES:6. Pressurization Source Selector ........RH(allow time for smoke to dissipate)

IF SMOKE STILL CONTINUES (AIR CYCLE MACHINE SEAL

MAY BE LEAKING):

7. Pressurization Source Selector ........EMER(control cabin pressure with LH throttle)




Smoke Removal

Note: No action is normally required; however, if smoke isintense:

1. Oxygen Mask ...................................DON

Check oxygen selector is on 100%.

2. OXYGEN CONTROL VALVE ............MANUAL DROP

3. Ensure passengers are receiving oxygen. Visually checkmask drop has occurred.

4. Oxygen Mic Switches ......................MIC OXY MASK

Switch must be in this position to use microphone in theoxygen mask.


6. Cabin Altitude Selector ....................SET to highercabin altitude. Selecting a higher cabin altitude willcause the outflow valves to open and increase the rateof airflow to clear the smoke.


This switch manually opens the normal dump valve to

rapidly depressurize the airplane, allowing the smoke toclear. All smoking material should be extinguished.

IF SMOKE PERSISTS OR IT CANNOT BE VERIFIED THAT THERE

IS NO FIRE:8. Land as soon as possible.





Autopilot Hardover

1. Autopilot/Trim

Disengage Switch ......................... PRESS

Press switch on either yoke. Flight director modes re-

main selected.

2. Maximum Altitude Losses during Autopilot Malfunction:a. Cruise ..............................................550 feet at 43,000 feetb. Climb ...............................................300 feet at 17,000 feetc. ILS ...............................................34 feet

(autopilot must be off at 90 feet)

Maximum Glide – Emergency Landing

1. Airspeed ..........................................BEST GLIDE AT

9500 LBS-120 KIAS. Increase speed 3 KIAS for each500 pounds increase in weight.

2. Flaps ...............................................UP

3. Speed brakes ..................................RETRACT

4. Landing Gear ...................................UP

5. Transponder ....................................EMERGENCY, 7700

6. ATC ...............................................ADVISE


8. Shoulder Harness ............................SECURE

9. Landing Gear, Speed brakes,and Flaps .........................................AS REQUIRED

for landing anticipated




Appendix 1Standard Operating Procedures

Table of Contents

Standard Operating Procedures

Overview ......................................... A1-1

Engine Start ................................... A1-2

Taxiing ........................................... A1-3

Takeoff ........................................... A1-4

After Takeoff Climb ......................... A1-5

Cruise ............................................ A1-6

Descent.......................................... A1-7

Before Landing ............................... A1-8

Landing (without thrust reversers) .... A1-9

Landing (with thrust reversers) ....... A1-10

After Landing ................................ A1-11

Shutdown .....................................A1-11

Standard Operating Procedures (crew briefings)

Before Takeoff .............................. A1-12

Takeoff ......................................... A1-12

Climb ........................................... A1-13

Cruise .......................................... A1-14

Descent........................................ A1-15

Precision Approach ........................A1-16

Precision Missed Approach ............ A1-18

Non-Precision Approach ................. A1-19

Non-Precision Missed Approach...... A1-21

Visual Approach ............................ A1-22

Landing ........................................ A1-23




STANDARD OPERATING PROCEDURES 12/99 FOR TRAINING PURPOSES ONLY A1-1

Overview

This chapter provides standard operating procedures as wellas the abbreviated cockpit checklist in expanded form. Thestandard operating procedures are per the Operating Manual,however, their use in this manual is for training purposes only.

The appropriate Operating Manual, FAA approved AirplaneFlight Manual (AFM), Pilot’s Check List, and/or related publica-tions should be referred to for procedures specific to the year,

model and serial number of your aircraft. For each expandedphase-of-flight procedure, the call, action and response fromthe Pilot in Command and the Second in Command, (hereafterreferred to as PIC and SIC) is delineated.

It is the responsibility of the pilot in command to ensure thatthe aircraft is safely loaded and properly configured for flight.

It is also necessary to understand and utilize the variousgraphs and tables which outline the performance characteris-

tics of your airplane, its weight and balance data, and theforms which are used to determine load placement.

All Notes, Cautions, and Warnings are provided to amplifyeach procedure.



FOR TRAINING PURPOSES ONLY CITATION II APPENDIX 1 12/99A1-2

Engine Start

Either engine may be started first. If the door is secured prior tobattery start initiation, it is recommended that the left engine bestarted first. Spool up will be slightly faster due to less line lossbecause the battery is mounted on the left side of the tailcone

compartment. Due to foreign object ingestion hazard, the leftengine should not be running during boarding or deplaning. Iflast minute boarding and use of BLEED AIR GND is antici-

pated, the right engine should be started first.

With external power in use, the GEN switches can be off untilstarting is complete. It may not be possible to bring the genera-tors on the line until the external power unit is removed. In anycase, electrical equipment should not be turned on until bothGEN OFF lights are extinguished.

An overcurrent and overvoltage protection system is provided

during use of an external power unit (EPU). The control unitmonitors the external power unit voltage and will deenergize theexternal power relay if the voltage is above 32.5 volts. Duringan engine start using the external power unit, a signal is ap-plied by the current to the control unit. If the signal indicatesmore than 1200 amps after two seconds the control unit willdeenergize the external power relay and terminate the start.External power cannot be reapplied to the airplane until the

current has been interrupted after the start termination for the

current protection or until the voltage is reduced below 32.5volts for the voltage protection.

Should automatic start sequencing not terminate, the boostpump, ignition and associated lights will remain on. The starter,however, will discontinue cranking due to speed sensing whichgoverns at approximately 40 to 43 percent N2. Depressing theSTARTER DISENGAGE button will terminate the automatic start

sequence. This button is illuminated any time the PANEL LIGHTCONTROL master switch is ON.

Prior to taxiing, the Second in Command (SIC) will complete theTakeoff Data Card insuring the latest information is used fordata computation. Computations for takeoff thrust setting, V1,VR, V2, and VENR speeds, takeoff field length, and climb thrustare based on the runway length available, runway gradient,

field temperature, field pressure, wind, icing conditions andrunway condition.





Taxiing

Taxiing on one engine may be advisable at light weights toreduce brake wear, particularly in very cold weather whenidle thrust is relatively high. Turning capability into the liveengine is reduced however, and consideration should be

given to the direction of anticipated turns in deciding whichengine to operate. Peak exhaust velocity to generate thenecessary thrust will be higher on one engine. Maneuvering

in close quarters may dictate use of both engines.

Ground operations in visible moisture with an outside ambi-ent air temperature from -30°C to +4°C require that ENGINEANTI-ICE be “ON” and the engines run at or above 65% N2one minute out of every four.

Note: The anti-ice system must be checked prior to takeoff ifflight into icing conditions is expected. Approximately 70%

turbine speed is required to provide adequate engine bleedair to extinguish the ENGINE ANTI-ICE light in two minutesor less.

Note: When operating in visible moisture and ambient airtemperature is between +4°C and -30°C, position groundidle switch to “HIGH”, turn pitot and static heat and engineLH and RH anti-ice systems “ON”. If temperature is above -

18°C, turn W/S BLEED air switch to “LOW”. If temperature is

-18°C or below, turn W/S BLEED air switch to “HI”. CheckW/S bleed air valves “MAX”. For sustained ground operation,the engines should be operated one out of every four min-utes at 65% turbine RPM or above.




Takeoff

Monitoring the engine instruments, apply power slowly whilereferencing the cabin vertical speed indicator. Very rapidthrust application can cause a pressure surge due to in-creased airflow into the cabin. A rolling takeoff may be used

with sufficient runway available, but it should be rememberedthat Flight Manual takeoff field length data and takeoff N1settings assume a static runup.

Directional control is normally maintained with nose gearsteering and rudder; and upwind (wing down) aileron incrosswind conditions. It is suggested that the copilot performthe engine instrument monitoring function and set thethrottles enabling the pilot to direct his full attention to air-plane control. N1 should be closely observed, and throttle

corrections made as necessary to ensure symmetrical thrustapplication. Large differential power changes, particularly at

the higher thrust settings, can induce significant yaw.

It is recommended that the copilot verbally state when take-off thrust is set, a cross-check of airspeed indicators at 70knots is made, and when reaching V1 and VR. Positive backpressure is required to rotate the Citation II and it should beaccomplished precisely at Vr. Early or late rotation will de-grade takeoff performance. It should be done smoothly,

however, so that a decrease in airspeed does not occur.

Should an emergency situation occur at a speed below V1,takeoff should normally be aborted. Proceed with a normaltakeoff should the emergency situation occur at a speedabove V1. Single engine rotation is approximately 7° to 10°pitch attitude. Normal rotation angle is 10 to 12° nose upwith both engines operating. Procedures for abort and singleengine takeoff are outlined in Chapter 12 – Emergency Pro-

cedures.





After Takeoff Climb

Using indicated temperature and the MULTI ENGINE NOR-MAL CLIMB thrust setting graph in your FAA ApprovedAirplane Flight Manual (or in the Abbreviated checklist),determine climb N1.

During climb, observe the differential pressure/cabin altitudeand cabin vertical speed gages for proper programming and

comfortable rate. Periodic checks of time to climb remaining,cabin altitude and rate of cabin ascent will provide the re-quired information to determine any adjustments necessary.As an example, passing 20,000 feet with a cabin altitude of4000 feet and an estimated climb time remaining of 10minutes to 35,000 feet (8000 feet cabin altitude), wouldrequire a cabin climb rate of 400 feet-per-minute to attain

planned cruise and cabin altitudes concurrently. With RATEset too low, maximum differential pressure may be reached

before cruise altitude. This takes control of the system awayfrom the crew because the outflow valve will relieve as nec-essary to maintain maximum differential. A RATE setting toohigh may be uncomfortable and will result in programmedcabin altitude being reached before cruise flight level. Athorough understanding of DIFF PRESS/CABIN ALT gageinterpretation will aid the crew in smooth operation of thepressurization system.




Cruise

Climb thrust is normally maintained upon level off until accel-eration to the desired cruise mode takes place. Thrust isthen adjusted to the appropriate setting. If engine RPM doesnot automatically synchronize at desired cruise setting, turn

the engine synchronizer switch to “OFF”, allowing the syn-chronizer actuator to center; roughly synchronize the engineswith the throttles and turn the synchronizer switch to “FAN”

or “TURB”. For the maximum range case, thrust necessary tomaintain optimum angle of attack diminishes with fuel burnoff because of increased performance and lower airspeedrequirements as weight decreases.

Although the Citation II is not operationally restricted inrough air, flight in severe turbulence should be avoided. If

severe turbulence is encountered, it is recommended thatthe igniters be turned “ON” and airspeed maintained at

approximately 180 KIAS. Maintain a constant attitude, avoidabrupt or large control inputs, and do not chase airspeedand altitude indications. Use of the autopilot in the SOFTRIDE mode is recommended.

A comfortable cabin temperature is normally maintained withthe AUTO TEMP SELECT in the 12 to 2 o’clock position.During daylight, the crew environment may not be an accu-

rate reference to cabin comfort level due to solar heating

taking place through the wide expanse of cockpit windows.An approximate indication of airflow warmth into the cabincan be determined by placing a hand over an open crewfoot warmer outlet. The foot warmers are an extension fromthe same source as the cabin under floor ducting and canbe used as a reference for AUTO TEMP SELECT adjust-ments to maintain a comfortable cabin.

The DEFOG FAN should be turned on and foot warmersclosed approximately 15 minutes before descent to reducecondensation on the windshield and cockpit side windows.





This is particularly important when a rapid let down into an

area of high humidity is anticipated after cold soaking ataltitude.

The engine bleed air anti-ice must be activated when operat-ing in visible moisture at temperatures of from -4°C to -30°C

indicated by OAT and any time icing is occurring. The pitotand static anti-ice is normally operated during flight.

Detailed instructions of the engine anti-ice and surface deicesystems are found in Chapter 10 – Ice Protection and in theFAA Approved Airplane Flight Manual.

Descent

Once destination altimeter setting is known, field pressure

altitude can be determined because each .10 inches ofmercury deviation from 29.92 equates to 100 feet difference

between field elevation and pressure altitude. An altimetersetting above standard gives a pressure altitude below fieldelevation and the inverse is also true. As an example, de-scending to a field elevation of 350 feet with a reportedaltimeter of 29.77 would result in a field pressure altitude of500 feet. The cabin altitude should be set to 700 feet toinsure depressurization prior to touchdown. Rate is normallyadjusted to give a 300-500 FPM cabin rate of descent.

Monitor the differential pressure/cabin altitude and cabinvertical speed gages. A high cabin altitude and low differen-tial pressure indicates an insufficient rate of descent anddepressurization will occur when cabin and airplane altitudeare identical. High cabin descent rates will occur whencabin and airplane altitude are identical. High cabin descentrates may be uncomfortable and may result in programmedcabin altitude being reached well before landing. Optimum

comfort is realized by spreading cabin descent required overthe majority of airplane let down time.




Before Landing

Consistently comfortable and safe landings are bestachieved from a stabilized approach. The point at which theairplane should be stabilized with airspeed at VREF to VREF +KIAS, full flaps, and the desired descent rate is normally

coincident with commencing the final descent to landing.Under instrument conditions, this usually occurs at the finalapproach fix inbound. During visual approaches, this would

be a point approximately equal to a turn on to base leg,adjusted for the altitude difference between the traffic patternand field elevation.

After passing the instrument approach fix outbound or near-ing the airport traffic area, airspeed should be reducedbelow 202 KIAS and the flaps extended to the APPR (15°)

position. Approaching the final instrument fix inbound (onedot from glideslope intercept on an ILS), or a downwind

abeam position, extend the landing gear below 176 KIAS. Atthe point where final descent to landing is begun, extendFULL FLAPS, establish the desired vertical rate, and adjustpower to maintain VREF to VREF + 10 KIAS indicated air-speed.

Power management during the approach/landing phase isrelatively easy in the Citation II because an N1 setting in the

60-65% range will normally result in desired indicated air-

speeds for the various configurations. Depending on airtraffic control requirements, thrust necessary for the entireapproach can often be set during descent keeping in mindthat fan (N1) RPM will decrease slightly for a fixed throttlesetting with a decrease in altitude for indicated airspeed.Using a sea level airport with zero wind at a typical landingweight (10,000 LBS), a throttle setting that results in about60% N1 in close will give approximately level flight indicated

airspeeds of 160 knots clean and 140 with flaps APPR. Gearextended, flaps FULL and commencing an average descent(500 FPM) will result in approximately VREF airspeed. Higherfield elevations, landing gross weights and/or headwindcomponent will require a greater power setting.





For maneuvering prior to final approach, minimum airspeeds

of VREF + 30, VREF + 20 and VREF + 10 should be maintainedclean, flaps APPR and flaps LAND respectively to provide anadequate margin above stall.

Speed control on final should be precise for optimum land-

ing performance and this is best accomplished byestablishing Vref airspeed well before crossing the threshold.In gusty wind conditions, it is recommended that one half

the gust factor in excess of 5 knots be added to Vref.

Approaching within approximately 50 feet of airport elevationpower should be gradually reduced to counter the accelera-tion induced by ground effect. Wind velocity and directionwill dictate the rate at which the throttles are retarded. Invery high surface headwind conditions, as an example, it

may be necessary to maintain at or near approach poweruntil close to touchdown. With a tailwind, a fairly rapid power

reduction may be necessary in the final descent to landingphase for accurate speed control. In ground effect, whereinduced drag is reduced, leaving approach power on willcause the airplane to float to a longer touchdown than de-sired. Retarding the throttles gradually in the final descentwill normally result in idle thrust being reached just beforetouchdown.

Landing (without thrust reversers)

Touchdown, preceded by a slight flare, should occur on themain wheels. Check thrust at idle and extend thespeedbrakes while lowering the nosewheel. With the nose-wheel on the runway, the drag chute (if installed) may bedeployed by depressing the safety button on the drag chutehandle and then pulling up firmly to full travel (approximately27°). This will require approximately 35 LBS force. Braking

should be commenced according to runway length availableto reduce brake wear. Normally with excess runway, brakingis begun after aerodynamic deceleration to below 80 KIAStakes place. Apply smooth, gradually increasing pressureuntil a comfortable turn off speed is reached. For maximumbraking performance, immediately after touchdown andwheel spin-up, apply brakes firmly and hold to full stop (donot modulate brake pedals).




Landing (with thrust reversers)

Touchdown, preceded by a slight flare, should occur on themain wheels. Check thrust at idle and extend the speedbrakes while lowering the nose, apply wheel brakes anddeploy the thrust reversers. The airplane pitches slightly

upward during the deployment of the reversers. Therefore,slight nosedown elevator pressure should be used duringthrust reverser deployment especially at high speeds such

as a refused actuation of the thrust reversers to eliminate thepossibility of FOD and improve directional control. To avoidpossible jamming of the throttle lockout cams, do not exceedapproximately 15 LBS force on the thrust reverser leversduring deployment. Check illumination of the thrust reverserlights.

Caution: On any airplane, do not attempt to restow reversersand take off once reversers have started to deploy. On

airplanes not incorporating SB550-78-03, throttle linkagedamage may occur resulting in loss of power or flameout.

Once the thrust reversers are deployed, move the thrustreverser levers aft to a maximum reverse thrust. For conve-nience, “Stops” have been installed on the thrust reverserlevers and are set to provide 92% fan speed (N1) + 2% or -2% at -18°C at sea level. This will allow the pilot to keep his

attention on the landing rollout instead of diverting his atten-

tion on the landing rollout instead of diverting his attention tothe reverse power settings, except in an abnormal ambienttemperature condition.

At 60 KIAS, return the thrust reverser levers to the idle re-verse detent position leaving the thrust reversers deployedfor aerodynamic drag. Thrust reversing and braking shouldbe commenced according to runway length. Normally, with

excess runway, braking is begun after thrust reverser decel-

eration is below 60 knots. For maximum braking performance,

immediately after touchdown and wheel spin-up, applybrakes firmly and hold to full stop (do not modulate brakepedals ). The thrust reversers should not be used for touchand go landings; a full stop landing should be made oncethe reversers are selected.





After Landing

It is recommended that the checklist be delayed until theairplane is clear of the runway. Turn off anti-collision andrecognition lights. The rotating beacon (if installed) may beleft “ON”.

Do not advance the throttles while the thrust reversers arebeing stowed. On airplanes not incorporating SB550-78-03

the automatic engine power retard system will activate caus-ing misrigging of the throttle linkage system. This wouldresult in only partial takeoff power or possibly a flameout ifthe throttle was placed in idle position. On airplanes incorpo-rating SB550-78-03 there is danger of the throttle beingrapidly returned to idle position, which could cause injury. Toavoid activating the automatic retard system, do not advance

the primary throttle after moving the reverse thrust lever tostow until the UNLOCK light is out.

Shutdown

Always check cabin differential pressure at zero beforeopening the door. Any pressure existing due to malfunctionof the left main gear squat switch or outflow valves couldcause the door to open rapidly presenting a hazard to per-sonnel in the vicinity.

For deplaning at night, the battery switch may be left inBATT to make available all cabin lighting until passengersand cabin baggage are disembarked. Turning the EXTERIORWING INSP LIGHTS switch “ON” provides additional illumina-tion in front of the cabin door. An illuminated courtesy lightswitch located on the forward door post is wired to the hotbattery bus and turns on the emergency exit lights and oneaft baggage compartment light.

When securing the airplane, install the engine and pitot tubecovers. Check the BATT, passenger advisory and courtesylight switches off. Closing the door extinguishes integralcourtesy light switch illumination. All doors and the noseavionics compartment can be key locked. A locking pin canbe installed in the internal emergency exit door handle toprevent access from the outside. This pin must be removed

prior to flight.




Crew Briefing: The pilotshould review crew coordina-tion with respect to flap

setting, ice protection proce-dures, takeoff power setting,

"V" speeds and airspeedlimits, as well as normal and

emergency procedures.

Call: “Set takeoff power”

ACTION: Positive backpressure should beapplied and arotation to anapproximate 9°pitch attitude begun

ACTION: Check engineinstruments

Call: “Takeoff power set”

Call: “Airspeed alive”

ACTION: 70 knots crosscheck(both airspeedindicators)

At: V1

Call: “V1”

At: VR

Call: “Rotate”

Takeoff

Note: Should an emergency situation occur at a speed belowV1, takeoff should normally be aborted. Proceed with a normaltakeoff should the emergency situation occur at a speedabove V1. Single engine rotation is approximately 7° to 10°pitch attitude. Procedures for abort and single engine takeoffare outlined in Chapter 12 – Emergency Procedures.

Call: “BeforeTakeoff

checklist” ACTION: Complete BeforeTakeoff checklist

Call: “Before Takeoff checklist complete”

Before Takeoff

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Call: “Gear up”

ACTION: Retract flapsCall: “Flaps Up”

At: VENR

Call: “Set climb power”

Call: “Climb checklist”

At: 1.5% N1 or 1% N2

ACTION: Select FAN or TURBswitch as required

Call: “Engine Sync set”

At: Positive rateof climb

Call: “Positive rate”

ACTION: Retract gear

Call: “Gear selected up”

ACTION: Check indication

Call: “Gear indicating up”

At: V2 + 10 KIAS

Call: “V2 + 10 400 feet”

Call: “Flaps indicating up”

At: VENR

Call: “Best climb”

ACTION: Check

Call: “Climb power set”

ACTION: Complete climbchecklist

Call: “Climb checklistcomplete”

ACTION: Check engine instru-ments within limits

Call: “Engine Sync set”

Climb

Crew Briefing: Using bothindicated temperature andthrust setting graph (see

Abbreviated Checklist),determine climb N1.

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At: Transition Altitude

Call: “29.92 set.

Transition Altitudechecklist”

Call: “29.92 set”

ACTION: Complete TransitionAltitude checklist

Call: “Transition Altitudechecklist complete”

Climb (continued)

Cruise

Call: “Cruise checklist”

ACTION: Complete Cruisechecklist

Call: “Cruise checklistcomplete”

Crew Briefing: If engine RPMdoes not automatically syn-chronize at desired cruise

setting, turn the enginesynchronizer switch to OFF,

allowing the synchronizeractuator to center. Roughly

synchronize the engines withthe throttles and turn the

synchronizer switch to FAN orTURB as required.

Note: Although the engine is not operationally restricted inrough air, flight in severe turbulence should be avoided. Ifsevere turbulence is encountered, it is recommended that theigniters be turned ON and airspeed maintained at approxi-

mately 180 KIAS. Maintain a constant airspeed, avoid abruptor extended control inputs, and do not chase airspeed andaltitude indications. Use of the autopilot in the SOFT RIDEmode is recommended.

Caution: Do not operate deice boots when indicated OAT isbelow -40°C (-40°F).

Note: The engine bleed air anti-ice must be activated whenoperating in visible moisture at temperatures from +4°C to-30°C indicated OAT and any time icing is occurring.

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Call: “Descent checklist”

Call: “Set altimeter.Transition Levelchecklist”

ACTION: Complete Descent

checklistCall: “Descent checklist

complete”

At: Transition altitude

Call: “18,000 ft”

ACTION: Set altimeter

Call: “Altimeter set”

ACTION: Complete TransitionLevel checklist

Call: “Transition Levelchecklist complete”

Descent

Crew Briefing: When practi-cable, review approach and

missed approach procedures.Determine N1 and V2 for use

in the event of a missedapproach. Set proper NAVfrequencies and required

heading and courseinformation. Check runway

requirements based on grossweight and destination field

information.

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Call: “Approach checklist”

At: Airspeed below202 KIAS

Call: “Flaps 15°”

Call: “Localizer captured”

Call: “Gear down”

Call: “Before Landingchecklist”

At: Glideslope intercept

Call: “Glideslopecaptured”


ACTION: Complete Approach

checklistCall: “Approach checklist

complete”

ACTION: Select 15° flaps

Call: “Flaps selected 15°

Call: “Flaps indicate 15°”

Call: “Localizer captured”

At: One dot fromglideslope intercept

Call: “One dot to go”

ACTION: Extend gear

Call: “Gear indicates

down”ACTION: Complete Before

Landing checklist

Call: “Before Landingchecklist complete”

Call: “Glideslope

captured”


Call: “Flaps selected 40°”


Precision Approach

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Call: “Check”

Call: “Check”

Call: “Check”

Call: “Looking”

At: 1,000 ft above DH

Call: “1,000 ft to

minimums”

At: 500 ft above DH

Call: “500 ft tominimums”

At: 200 ft above DH


At: Point of visualcontact

Call: “Runway at ____o’clock”

At: Decision height

Call: “Minimums”

Precision Approach (continued)

Note: Missed approach procedures must be executed imme-diately if either of these parameters have not been met atDecision Height:

1. a normal approach to the intended runway cannot beestablished or,

2. adequate visual reference cannot be maintained.

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Call: “Missed approach”

ACTION: Set go-around powerand rotate to 10°


Call: “Gear up”

At: DH

Call: “Minimums. Missed

approach”








Call: “Gear indicates up”

Precision Missed Approach

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At: Localizer/coursecapture

Call: “Localizer/coursecapture”


Call: “Landing checklist”



complete”




Call: “Localizer/coursecapture”

At: Approach to FAF

Call: “_____ minutes/_____ miles to FAF”

ACTION: Extend gear

Call: “Gear selecteddown”

Call: “Gear indicatesdown”

ACTION: Complete Landingchecklist

Call: “Landing checklistcomplete”

Non-Precision Approach

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At: FAF

Call: “Final fix”

ACTION: Begin adequatedescent rate(approx. 1,000 fpm)


Call: “Check”

Call: “Check”

Call: “Check”

Call: “Looking”

Call: “Final fix”

ACTION: Start timing, setminimum descentaltitude, and checkaltimeters




Call: “Altimeters check”

At: 1,000 ft above MDA

Call: “1,000 ft tominimums”

At: 500 ft above MDA


At: 200 ft above MDACall: “200 ft to

minimums”


Call: “Runway at _____o’clock”

At: MDA

Call: “Minimums”

Non-Precision Approach (continued)

Note: Missed approach procedures must be executed imme-diately if either of these parameters have not been met atMinimum Descent Altitude:

1. a normal approach to the intended runway cannot beestablished or,

2. adequate visual reference cannot be maintained.

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Call: “Missed approach”

ACTION: Disconnect autopilot(if engaged), set go-around power androtate to 10°


Call: “Gear up”

At: MDA

Call: “Minimums. Missed

approach”








Call: “Gear indicates up”

Non-Precision Missed Approach

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At: Abeam touchdownpoint


At: Base turn


ACTION: Set fuel flow to400lb/ENG.,begin descent

Call: “Before Landingchecklist”



complete”




ACTION: Extend gear

Call: “Gear selecteddown”

Call: “Gear indicatesdown”




ACTION: Complete BeforeLanding checklist

Call: “Before Landing

checklist complete”

Visual Approach

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ACTION: Disengage autopilotand yaw damper

At: Touchdown

Call: “Extendspeedbrakes”

Call: “Deploy thrustreversers”

ACTION: Monitor VREF andannunciator panel,verify landing gearand flap indication

At: 100 ft abovetouchdown

Call: “100 feet”

ACTION: Retract speedbrakes(as required)

At: 50 ft abovetouchdown

Call: “50 feet”

ACTION: Extend speedbrakes

Call: “Speedbrakesextended”

ACTION: Deploy thrustreversers

Call: “Two deployed”

At: 60 KIAS

Call: “60 knots”

ACTION: Reverser levers toidle reverse

Landing

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Appendix 2Flight Profiles

Table of Contents

Flight Profiles

Normal Takeoff ............................... A2-1

ILS Flight Director/Autopilot Approach .......................... A2-2

Non-Precision Approach ................... A2-3

Radar Approach .............................. A2-4

Circling Approach ............................ A2-5

Steep Turns....................................A2-6

Acceleration and Deceleration .......... A2-7

High Rate of Descent Recovery........ A2-8Emergency Descent .........................A2-9

Visual Approach and Landing ......... A2-10

Flaps Up Landing .......................... A2-11

Single Engine ILS ApproachLanding, and Go-Around................. A2-12

Takeoff Engine FailureAfter V1 ....................................... A2-13



F L I G H T P R O F I L E S

1 2 /

9 9

F O R

T R A I N I N G P U R P O S E S

O N L Y

A 2 - 1

Takeoff

Takeoff power .......... SET

(prior to 60 KIAS) Airspeed .................. 70 KIAS

Rolling Takeoff

Takeoff power ....... SET

Rotate smoothly to 10°

Rate of climb ........ POSITIVE

Gear..................... UP

Close-In T

Flaps .

Airspe

Bank a

S

Roll Out

Airspeed .................V2 + 20 KIAS MIN.

Climb power ...........SET

Accelerate

Flaps ......................RETRACT

After-takeoff/ climb checklist ........ COMPLETE

1

2

3

4

5



F O R

T R A I N I N G P U R P

O S E S

O N L Y

C I T A T I O

N I I A P P E N D I X

2

1 2 / 9 9

A 2 - 2

Approach Preparations

Approach procedure ............... REVIEW

Go-around procedure .............. REVIEW

Airspeed bug ......................... SET VREF

Avionics ................................. CHECK

Before landing checklist ......... COMPLETE

Radar Vectors

Airspeed .......... VREF

+ 20 KIAS MIN.

Flaps ............... T/O & APPR.

Hdg ................. ON

Nav ................. ARM

Apr .................. ARM

Alt ................... ON

1

Localizer ......... TRACK

Gear................ DOWN

Flaps ............... LAND

Hdg ................. LIGHT EXTINGUISHED

Glideslope ....... CAPTURE

Alt ................... ON

3

Localizer ..... CAPTURE2

Autopilot/FD (CoupledApproach) ............BEGINS DESCENT

LOC and GSGlideslope ...........COUPLED

Alt .......................OFF

4

Airspeed ....... VREF

MIN.5

Missed Approach

Go-around button...... PUSH

Autopilot .................. OFF Go-around light ......... ON

Go-around power ..... SET

Command bars......... 7

Rotate ..................... 10°

Flaps ....................... 15°

6

7




1 2 /

9 9

F O R


O N L Y

A 2 - 3

5

Procedure Turn

Flaps ...................T/O & APPR.

Airspeed ..............VREF

+ 20 MIN.

Approach Prep

Approach procedur

Go-around procedu

Airspeed bug ........

Avionics ...............

Before landing chec

1

Descent ........ 1,000 FPM RECOMMENDED2

Note: Maximum use of flight director/autopilot if desired:

Press APR button for VOR approach

Press NAV button for LOC only approach

Press BC button for back course localizer approach

Descent ........ LEVEL


+ 20 MIN.

Gear ............. DOWN

Flaps ............ LAND

3

Runway in Sight

Circling approach ....... INITIATE


Maintain normaldescent to landing

6

4

Mi

Minimumdescent altitude



F O R


O S E S

O N L Y

C I T A T I O


2

1 2 / 9 9

A 2 - 4

Approach Prepa

Approach procedure ..

Go-around procedure .

Airspeed bug ............

Avionics .................... Before landing checkli

Flaps .............. T/O & APPR.

Airspeed ......... VREF

+ 20 KIAS MIN.

Intercept final approach at 30° to 45°

1

At Glideslope Intercept


Gear ................ DOWN

Flaps ............... LAND

2

Maintain normaldescent to landing

3




1 2 /

9 9

F O R


O N L Y

A 2 - 5

Approach Pr

Approach proce

Go-around proc

Airspeed bug .

Avionics .........

Before landing

Caution: The visual cues used when flying normal traffic patterns can bemisleading due to the lower altitudes associated with circling approaches.Common mistakes are flying the downwind leg too close to the runway, beginningthe downwind-to-final turn too soon, and descending below the MDA too early.

1

At Radio Fix

Gear................ DO Flaps ............... T/

Airspeed .......... VRE

2

Abeam "Key" Point

Timing ............. BEGIN

MDA ................ MAINTAIN

3Turn to Final

Bank angle....... 30° MAX.

Flaps ............... LAND


+ 10 MIN.

15 SECONDS

4

Normal descent to landing

Airspeed .......... VREF + WIND

5

6 Missed Approach

Go-around power ........ SET

Flaps ......................... T/O & APPR.

Gear .......................... UP

Flaps .........................RETRACT ON FLAP/SPEED SCHEDULE



F O R


O S E S

O N L Y

C I T A T I O


2

1 2 / 9 9

A 2 - 6

Airspeed ....

Bank angle .(increase thr

Altitude ......

1

2

1

2

3




1 2 /

9 9

F O R


O N L Y

A 2 - 7

ACCELERATE

DECELERATE

MANEUVER

Airspeed bug ........ VREF

Airspeed ............... STABILIZE AT 250 KIAS

Elevator trim ......... SET

Climb power ......... SET

Airspeed .......... STABILIZE JUSTBELOW V

MO

2

1

Airspeed

VMO

aural

Power...

Speed br

Timing ..

Altitude .

3

Airspeed

Timing ..

STABILIZ

3



F O R


O S E S

O N L Y

C I T A T I O


2

1 2 / 9 9

A 2 - 8

1 Airspeed bug ........... VREF

Gear........................ DOWN

Flaps ....................... LAND

Power ...................... IDLE

Rate of descent ....... CHECK

2

Recovery ..........

Go-around power

Attitude ............

Engine accelerati

Altitude lost ......

3

4

Warning: Attempting a landing from a high rate of descent is extremely hazardous.Selecting the proper flare height to arrest the descent rate at touchdown whileairspeed is rapidly decreasing would be difficult if not impossible. Idle thrustapproaches and high rates of descent near the ground must be avoided.




1 2 /

9 9

F O R


O N L Y

A 2 - 9

Oxygen masks............. Emergency descent ....

Pass. O2 valve .............

MIC OXY MASK switch..

Rapid decompression c

1

Throttles ......................................... IDLE

Speed brakes .................................DEPLOY

Moderate bank ................................ INITIATE

Pitch ................................................15° NOSE DOW

Emergency descent checklist ..........COMPLETE

2

3

Roll wings level

ATC ......................... ADVISE

Transponder ............ 7700 EMERGENCY

Altimeter setting ...... OBTAIN

Altitude .................... CALL

MEA .......................CONFIRM

4 2000 FT Above Desired Altitude

Rate of descent ....... REDUCE SMOOTHLY

5

1000 FT Above Desired Alti

Speed brakes .......... RET

6

14,000

Leve

Crew

Ignit

Airsp

Dete

7

Note: If structural damage is suspected, limit airspeed.Speedbrakes may be used depending on type of damage.

R - 2 / 8 / 0 1

Rapid Decompression

Emergency Descent

Passengers ................................... CHECK



F O R


O S E S

O N L Y

C I T A T I O


2

1 2 / 9 9

A 2 - 1 0

Approach Prep

Approach procedure

Go-around procedure

Airspeed bug .........

Avionics ................. Before landing chec

Base Turn

Flaps .............. LAND

Fuel flow......... SET (400 LB/ENG)

Descent .......... BEGIN

2

Abeam

Fla

Ai

Abeam

Ge

1

Bank angle .............. 30° MAX.

Final approach ........ CLEAR

Airspeed .................VREF

+ 10 KIAS MIN.

3

Landing Assured

Airspeed .................. VREF

Maintain normal descent to landing

4 To

5




1 2 /

9 9

F O R


O N L Y

A 2 - 1 1

Approach Prep

Approach procedure..

Go-around procedure .

Airspeed bug ........... Avionics...................

Before landing checkl

Gear................ DOWN

Airspeed .......... New VREF

+ 10 MI

1

2 Base Turn

Descent ........... BEGIN (300-500 FPM)

Bank angle....... 30° MAX.

3 Rollout

Airspeed .......... NEW VREF

Altitude ............ CALL

Stabilized in slot

4 Maintain normaldescent to landing

5 Missed Approach

Go-around power.......

Attitude ....................

Rate of climb ............

Gear .........................

R - 6 / 9 / 0 0



F O R


O S E S

O N L Y

C I T A T I O


2

1 2 / 9 9

A 2 - 1 2 Approach Preparations

Engine failure checklist ..................... COMPLETE

Approach procedure ......................... REVIEW

Go-around procedure ........................ REVIEW

Airspeed bug ...................................SET VREF

Avionics ........................................... CHECK

Single engine landing checklist ......... COMPLETE Flaps .....

Airspeed

1

Airspeed ............ VREF

+ 10

Gear .................. DOWN

2

When Runway Assured

Flaps ...........LAND

3

Go-Around

Go-around pow GA button ....

Flaps ...........

Rate of climb

Gear.............

Flaps ...........FLAP/SPEED

4




1 2 /

9 9

F O R


O N L Y

A 2 - 1 3

Roll Out

Airspeed ...............................VEN

Flaps .................................... UP

Climb power ......................... SE

Engine failure checklist .......... CO

After takeoff checklist ........... CO

Takeoff

Takeoff power .......... SET (prior to 60 KIAS)

Airspeed .................. 70 KIAS

1

Airspeed .......... V1 /V

R

Attitude ........... 10° NOSE UP

2

Engine Failure

Rudder ...................AS REQUIRED

Wings ..................... LEVEL

Rate of climb .......... POSITIVE

Gear .......................UP

3

Straight C

Airspe

Flaps Obst

Airspe

Climb

Engine

After t

4

Close-In

Airspe

Flaps

Bank

5

6

R - 6 / 9 / 0 0



F O R


O S E S

O N L Y

C I T A T I O


2

1 2 / 9 9

A 2 - 1 4

Airspeed ........... DECELERATE to 160 KIAS

Power ............... 50% N1

1

Airplane Configuration

Flaps ............... UP

Gear ............... UP

Maintain altitude withpitch trim to .6 AOA

2

Maintain altitude with back pressure at anapproximate rate of 1 per second —PITCH UP

3

At first indication of

(buffet or shaker):

Power ................

Release back presto 7° NOSE UP

Airspeed .............

4

5

R - 7 / 3 / 0 2




1 2 /

9 9

F O R


O N L Y

A 2 - 1 5

n Airspeed ........... DECELERATE to 160 KIAS

n Flaps ................. T/O & APPR.

n Power ............... 50% N1

n Bank angle ........ 20° either direction

1


n Flaps ............... T/O & APPR.

n Gear ............... UP

n Maintain altitude withpitch trim to .6 AOA

2

n Maintain altitude with back pressure at anapproximate rate of 1 per second —PITCH UP

3

At first indication of an

(buffet or shaker):

n Power ....................T

n Bank angle .............

n Release back pressu7° NOSE UP

n Airspeed ................A

4

5

R - 2 / 8 / 0 1



F O R


O S E S

O N L Y

C I T A T I O


2

1 2 / 9 9

A 2 - 1 6

Airspeed ........... DECELERATE to 160 KIAS

Flaps ................. T/O & APPR.

Power ............... 65% N1

Gear ................. DOWN

1


Flaps ............... FULL DOWN

Gear ............... DOWN

Flaps ............... DOWN

Power ............. REDUCE to 65% N1

Maintain altitude with pitch trim

2

Maintain altitude with pitchtrim to .6 AOA

Maintain altitude with backpressure at an approximaterate of 1° per second

3

At first indicati

(buffet or shak

Power .......

Flaps .........

Release bacto 6° NOSE

4

5

R - 7 / 3 / 0 2




1 2 /

9 9

F O R


O N L Y

A 2 - 1 7

Accelerate aircraft with takeoff power

Prior to V1, ABORT the takeoff

1


Flaps ............... T/O & APPR.

Power ....................REDUCE to IDLE

Brakes ................... APPLY

Speed brakes ........DEPLOY

Thrust reversersor drag shoot .........DEPLOY

2

Maintain centerlineorientation

Advise tower ofthe abort

3

Exit runway as inst5

R - 2 / 8 / 0 1




Appendix 3Glossary

Table of Contents

Glossary of Abbreviations

and Terminology .................................A3-1

Weight and Balance Terminology ........ A3-1

Performance and

Flight Planning Terminology .................. A3-5

Airspeed Terminology........................... A3-7

Meteorological Terminology ................ A3-11

Powerplant Terminology ..................... A3-13



APPENDIX 3 12/99 FOR TRAINING PURPOSES ONLY

Cessna Citation ll Technical Manua

A3-1

Glossary of Abbreviations and Terminology

The following glossary is a reference of abbreviations andterminology applicable to airplane operation, non-specific to asingle aircraft manufacturer or type.

Weight and Balance Terminology

Standard Empty Weight

The empty weight of an airplane in standard configura-tion as documented by the manufacturer, including theweight of unusable fuel, full engine oil, and full operatingfluids.

Basic Empty Weight

The standard empty weight of an airplane plus all op-

tional equipment installed, used as a basis for loadingdata determination.

Payload

The total weight of the crew, passengers, baggage,cabinet contents, and cargo.

Maximum Payload

The maximum permissible weight of the crew, passen-gers, baggage, cabinet contents, and cargo as deter-

mined by structural limitations.

Usable Fuel

The weight of that portion of the total fuel load which isavailable for consumption as determined in accordancewith applicable regulatory standards.

Unusable Fuel

The weight of that portion of the total fuel load which isnot available for consumption as determined in accor-dance with applicable regulatory standards, including

the weight of residual fuel.

Residual Fuel

The weight of all undrainable fuel remaining onboardafter the airplane has been defueled for weighing pur-

poses in accordance with specified procedures.




Maximum Useful Load

The maximum permissible combined weight of payloadand usable fuel as determined by structural limitationsassociated with ground handling. Equivalent to maxi-mum ramp weight minus basic empty weight.

Ramp Weight

The basic empty weight of an airplane plus payload andusable fuel prior to engine start. Ramp weight must not

exceed the maximum ramp weight.

Maximum Ramp Weight

The maximum permissible ramp weight of an airplane asdetermined by structural limitations associated withground handling. Equivalent to maximum takeoff weightplus the weight of fuel consumed during engine start,

taxi, run-up, and takeoff roll to VR.

Zero Fuel Weight

The ramp weight of an airplane excluding the weight ofusable fuel. Zero fuel weight must not exceed the maxi-mum zero fuel weight.

Maximum Zero Fuel Weight

The maximum permissible ramp weight of an airplaneexcluding the weight of usable fuel as determined by

associated structural limitations. Any weight in excess ofthis value must be fuel.

Takeoff Weight

The weight of an airplane upon lift-off from the runway.Takeoff weight must not exceed the maximum takeoffweight.

Maximum Takeoff Weight

The maximum permissible takeoff weight of an airplaneas determined by associated structural limitations.

Landing Weight

The weight of an airplane upon runway touchdown.Landing weight must not exceed the maximum landingweight.





A3-3

Maximum Landing Weight

The maximum permissible landing weight of an airplaneas determined by associated structural limitations.

Maximum Weight

The maximum permissible weight of an airplane asdetermined by associated design, structural, and perfor-mance limitations.

Center of Gravity (CG)

An imaginary point at which the weight of an airplane isconsidered to be concentrated, used to determineproper weight and balance for flight.

CG Limits

The extreme fore and aft limits of CG range within which

an airplane must be operated at a specific weight.

Reference Datum

An imaginary vertical plane perpendicular to an arbitrarypoint along the airplane’s longitudinal axis from which allhorizontal distances are measured to determine properweight and balance for flight.

Arm The horizontal distance in inches from the referencedatum to the center of gravity of an object or compo-

nent. Arm measurements are expressed in positive (+)inches aft of the datum and negative (-) inches forwardof the datum.

Moment

A measurement of the rotational force about theairplane’s CG, obtained by multiplying the weight of anobject or component by its respective arm (weight x arm

= moment).

CG Arm

The horizontal distance in inches from the referencedatum to the airplane’s center of gravity, obtained bydividing the total sum of all moments by the total sum oftheir respective weights (total moment ÷ total weight =CG Arm).




Approved Loading Envelope

Those combinations of airplane weight and center ofgravity which define the limits beyond which loading isnot approved.

StationThe horizontal distance in inches from the referencedatum to any position along the airplane’s longitudinalaxis.

Jack Points

Specific points on the airplane identified by the manu-facturer as suitable for supporting the airplane on jacks.

Leveling Points

Specific points on the airplane identified by the manu-

facturer as suitable for leveling the airplane.

Tare Any weight indicated by a scale before a load is applied.





A3-5

Performance and Flight Planning Terminology

Accelerate-Go Distance

The distance required to accelerate an airplane to take-off decision speed (V

1) and upon experiencing an en-

gine failure, continue accelerating to takeoff rotationspeed (V

R), then rotate, climb, and accelerate to engine

failure speed (VEF) before reaching 50 feet AGL for thepurposes of obstacle clearance.

Takeoff Flight Path

The minimum takeoff climb gradient required to clearobstacles beyond the end of the runway, measuredhorizontally from the distance at which the airplane hasreached 50 feet AGL, and vertically from the surface ofthe runway.

Accelerate-Stop Distance

The distance required to accelerate an airplane to take-off decision speed (V

1) and upon experiencing an en-

gine failure, bring the aircraft to a complete stop on theremaining runway using maximum effective braking.

Maximum Effective Braking

The maximum amount of braking pressure which can beapplied to the toe brakes without locking the wheels.

Clearway

An area beyond the end of a runway not less than 300feet on either side of the extended center line of therunway, at an elevation no greater than the end of therunway, clear of all fixed obstacles, and under the con-trol of the airport authorities.

Balked Landing

An aborted landing.

Balked Landing Transition SpeedThe minimum speed at which transition to a balked

landing climb should be attempted from 50 feet AGL.




Demonstrated Crosswind

The maximum 90° crosswind component under whichadequate control of the airplane during takeoff andlanding was actually demonstrated during certificationtesting.

Maneuvering Fuel

The usable fuel available for all airplane configurations,provided the maximum sideslip duration is not ex-

ceeded.

Aerobatic Maneuver

An intentional maneuver involving an abrupt changing ofthe airplane attitude, abnormal attitude, or abnormalacceleration, beyond the requirements for normal flight.

Fix Any geographic location which can be identified byvisual reference or radio navigational aids.

Route Segment

A specific portion of a route identified by geographic orradio navigational fixes.

Minimum Obstacle Clearance Altitude (MOCA)

The minimum airplane altitude permissible between fixeswhich meets necessary obstacle clearance require-

ments for a specific route segment.

Minimum En Route Altitude (MEA)

The minimum airplane altitude permissible betweenradio navigational fixes which assures adequate signalreception while meeting necessary obstacle clearancerequirements for a specific route segment.





A3-7

Airspeed Terminology

IAS Indicated Airspeed: Airplane speed as displayed on anairspeed indicator.

CAS Calibrated Airspeed: Airplane speed as displayed on anairspeed indicator corrected for instrument error. CAS isequal to true airspeed in standard atmosphere at sealevel.

TAS True Airspeed: Airplane speed relative to undisturbed airas displayed on an airspeed indicator corrected foraltitude, temperature, and compressibility.

KIAS Indicated Airspeed expressed in knots.

KCAS Calibrated Airspeed expressed in knots.

KTAS True Airspeed expressed in knots.

GS Ground Speed: Airplane speed relative to the ground.

M Mach: Airplane speed expressed in numerical propor-tion to the speed of sound under standard atmosphericconditions (e.g., Mach 1 is equivalent to the speed ofsound (approximately 642 KTS at sea level)).

G Acceleration Force: A measurement of force expressedin numerical proportion to the force of gravity (e.g., 1 Gis equivalent to the force gravity acting on an object atrest (approximately 32ft./sec2), which in turn determinesthe weight of the object). An acceleration force of 2 Gsacting on an airplane is therefore equivalent to doublingits actual weight.

V Velocity: A measurement of airplane speed relative tospecific operating limitations.

V1

Takeoff Decision Speed: The maximum speed to whichan airplane may accelerate before deciding to continueor abort takeoff according to accelerate-stop and accel-erate-go performance limitations.




VR

Takeoff Rotation Speed: The speed at which an airplanemay rotate according to take-off weight and takeoffperformance limitations.

V2

Takeoff Safety Speed: The speed to which an airplane

must accelerate after rotation before reaching 50 feetAGL for the purposes of obstacle clearance.

VEF

Engine Failure Speed: The speed to which an airplanemust accelerate after rotation

before reaching 50 feet

AGL for the purposes of obstacle clearance with oneengine inoperative.

VX

Best Angle-of-Climb Speed: The speed at which anairplane will deliver the greatest gain in altitude in theshortest possible distance over ground.

VXSE

Best Single-Engine Angle-of-Climb Speed: The speed atwhich an airplane will deliver the greatest gain in altitudein the shortest possible distance over ground with oneengine inoperative.

VY

Best Rate-of-Climb Speed: The speed at which an air-plane will deliver the greatest gain in altitude in theshortest possible period of time.

VYSE Best Single-Engine Rate-of-Climb Speed: The speed atwhich an airplane will deliver the greatest gain in altitude

in the shortest possible period of time with one engineinoperative.





A3-9

VMCA

Air Minimum Control Speed: The minimum speed atwhich the airplane is directionally controllable in flightwith one engine inoperative as established by FAAcertification procedures.

Certification is based on the following conditions: Oneengine inoperative and secured, operative engine set fortakeoff power, 5° bank towards the operative engine,gear up, flaps in takeoff/approach position, and rear-

ward most C.G.

At certain variations of aircraft weight and operatingaltitude, stall conditions can be encountered at speedsabove V

MCA as established by the certification procedure

described above. Under these conditions, stall speed

(VS) must be regarded as the air minimum control speed.

VS

Stalling Speed: The minimum steady speed at which theairplane is controllable in flight.

VSO

Landing Configuration Stalling Speed: The minimumsteady speed at which the airplane is controllable inflight when configured for landing.

VSSE

Intentional One-Engine-Inoperative Speed: A speedlimitation above both V

MCA and V

S established to provide

a margin of lateral and directional control when oneengine is intentionally failed for the purposes of pilottraining. Intentional failing of one engine below thisspeed is not recommended.

VMO

Maximum Operating Limit Speed: A speed limitationwhich may not be deliberately exceeded in normal flightoperations. Also expressed as M

MO in reference to the

equivalent Mach limitation.

VNO

Maximum Structural Cruising Speed: A speed limitation

which may not be exceeded except with caution insmooth air conditions.

VNE

Never Exceed Speed: A speed limitation which may notbe exceeded under any condition.




VA

Maneuvering Speed: The maximum speed at whichapplication of full available aerodynamic control will notstructurally over stress the airplane.

VF

Design Flap Speed: The maximum airplane speed at

which the wing flaps may be safely extended to a spe-cific position.

VFE

Maximum Flap Extended Speed: The maximum speed atwhich the airplane may be safely operated with the wingflaps extended to a specific position.

VLE

Maximum Landing Gear Extended Speed: The maximumspeed at which the airplane may be safely operated withthe landing gear extended.

VLO Maximum Landing Gear Operating Speed: The maxi-mum airplane speed at which the landing gear may besafely extended or retracted.

VMCG

Ground Minimum Control Speed: The minimum speed atwhich the airplane is directionally controllable on theground with one engine inoperative as established byFAA certification procedures.





A3-11

Meteorological Terminology

°C Degrees Celsius: A measurement of temperature relativeto a scale which indicates the freezing point of water as0° and the boiling point of water as 100° under standard

atmospheric conditions. Equivalent to (°F - 32) x 0.566.

°F Degrees Fahrenheit: A measurement of temperaturerelative to a scale which indicates the freezing point of

water as 32° and the boiling point of water as 212° understandard atmospheric conditions. Equivalent to (°C X1.8) + 32.

Dry Adiabatic Lapse Rate

The rate at which air temperature decreases with in-creasing altitude under standard atmospheric conditions

without losing or gaining heat energy (approximately3°C/5.5°F per 1000 feet).

Standard Lapse Rate

The average rate at which air temperature decreaseswith increasing altitude without losing or gaining heatenergy (approximately 2°C/3.6°F per 1000 feet), used todetermine freezing levels in the atmosphere relative tosurface air temperatures.

IOATIndicated Outside Air Temperature: A measurement ofstatic air temperature as displayed on an OAT indicatornot corrected for instrument error.

OAT Outside Air Temperature: A measurement of true outsideair temperature obtained from ground meteorologicalsources, or by correcting IOAT for the compressibilityeffects of airspeed and altitude, and used as a basis for

airplane performance determination.

Temperature Compressibility Effects

IOAT error corresponding to airspeed and altitude.




ISA International Standard Atmosphere: A standardizedreference measurement of atmospheric conditions usedto determine airplane performance in nonstandardconditions, assuming the following factors:

1. Air is a dry, perfect gas.

2. Atmospheric pressure at sea level is 29.92 inHg/ 1013.2 mb.

3. Air temperature at sea level is 15°C/59°F.4. Air temperature decreases with increasing altitude

by approximately 3°C/5.5°F per 1000 feet to 35,750feet and zero above that altitude.

Pressure Altitude

A measurement of altitude above the Standard DatumPlane, a theoretical level where atmospheric pressure isequal to standard sea level pressure. Equivalent to

indicated altitude when the altimeter is set to 29.92 inHg/ 1013.2 mb.

Density Altitude

A measurement of pressure altitude used to determineairplane performance in nonstandard atmospheric con-ditions. Equivalent to true altitude under standard atmo-spheric conditions.

Indicated Altitude

An indicated measurement of altitude above the atmo-spheric pressure to which the altimeter is set. Equivalentto true altitude when the altimeter is correctly set to thelocal reported pressure.

True Altitude

An indicated measurement of altitude above Mean SeaLevel (MSL), a level where atmospheric pressure is

equal to sea level pressure as computed from stationpressure corrected for nonstandard conditions, assum-ing the indicating altimeter is correctly set and zero

instrument error. Equivalent to pressure altitude anddensity altitude under standard atmospheric conditions.





A3-13

Station Pressure

A measurement of actual atmospheric pressure (baro-metric pressure) at field elevation.

Local Reported Pressure

A measurement of atmospheric pressure correspondingto sea level pressure as computed from station pressurecorrected for nonstandard conditions. Used to determinealtimeter correction settings required to indicate true

altitude.

Powerplant Terminology

Igniter

A device used to start the burning of the fuel/air mix-ture in a combustion chamber.

Impeller

The main rotor of a radial compressor, which increasesthe velocity of the air which it pumps.

Plenum

A duct, housing or enclosure used to contain air underpressure.

Stator

A row of stationary airfoils that direct the airflow be-tween the rows of rotor blades.

Turbine

A rotating device actuated either by reaction or impulse(or a combination of both), and used to transform someof the kinetic energy of the exhaust gases into shafthorsepower to drive the compressor(s) and accesso-

ries.

Turbofan

A type of gas turbine that converts heat energy intocore and bypass thrust.




Appendix 4Annunciators

Table of Contents

Annunciators

Airplanes (550-0550 and after) ....... A4-1

Annunciators

Airplanes (550-0550 and earlier) .... A4-5

Airplanes (550-0482,550-0485~550-0505...................... A4-5





A4-1

Annunciators

Citation II advisory lights (annunciators) are designed to pro-vide an easily interpreted representation of both normal andabnormal system conditions. The majority of annunciators are

collocated on an annunciator panel located on the centerinstrument panel.

Annunciator Panel

Airplanes 550-0550 and after

AC

FAIL

BATT

O’TEMP

CABIN ALT

10000 FT

OIL PRESS

WARN

LH RH

FUEL LOW

LEVEL

LH RH

FUEL LOW

PRESS

LH RH

HYD FLOW

LOW

LH RH

ENGINE

ANTI-ICE

LH RH

GEN

OFF

LH RH

INVERTER

FAIL

1 2

EMERG

PRESS ON

BLD AIR

GND

POWER

BRAKELOW PRESS

FUEL FLTR

BYPASS

LH RH

FUEL

BOOST ON

LH RH

HYD LOW

LEVEL

HYD PRESS

ON

P/S HTR

OFF

LH RH

SPEED

BRAKEEXTEND

AIR DUCTO’HEAT

ACMO’PRESS

ANTI SKID

INOP

DOOR NOTLOCKED

F/W

SHUT OFF

LH RH

W/S AIRO’HEAT

SURFACEDEICE

ACM

O’PRESS

BLD AIR

GNDSPARE SPARE

KEY:

Red

Amber

White

requires immediate attention, hazardouscondition exists.

requires attention, possible dangerouscondition exists.

safe or normal configuration, routine

action.

1

3 42

R-6/9/00




Annunciator Test

If the system is functioning normally, all annunciators and theMASTER WARNING lights should illuminate when the TESTselector knob on the lower left instrument panel is positionedto "ANNUN" and the BATT switch is in the "ON" position.

AC power bus voltage is above 130 VAC orbelow 90 VAC. Illumination of light triggers themaster warning system, which will illuminatethe MASTER WARNING light.

Steady illumination: battery temperature over145°F; flashing: battery temperature over160°F. Illumination of either annunciator trig-gers the master warning system, which willilluminate the MASTER WARNING light.

Cabin pressure altitude is above 10,000 feet.Illumination of either annunciator triggers themaster warning system, which will illuminatethe MASTER WARNING light.

Oil pressure is below safe limits (35 PSI) in leftor right engine. Illumination of light also triggersthe master warning system, which will illumi-nate the master warning light.

Fuel quantity in left and or right tanks hasreached a level of 169 to 219 pounds.

Low fuel supply pressure to engine-drivenpump. Primary pump failure will automaticallyinitiate boost pump operation as long as FUELBOOST switch is in the NORM position. Light

remaining on indicates failure of both pumps.

Left and/or right hydraulic system flow is belowapproximately 0.35 to 0.55 gallons per minute.

AC

FAIL

BATT

O'TEMP

CABIN ALT

10,000 FT

OIL PRESS

WARN

LH RH

FUEL LOW

LEVEL

LH RH

FUEL LOW

PRESS

LH RH

HYD FLOW

LOW

LH RH





A4-3

Left and/or right engine inlet, stator, or inboardwing anti-ice has failed.

Illuminates twice during the 12- second surfacedeice boot cycle to indicate proper boot infla-tion pressure. Tail boot inflation cannot bechecked visually from the cockpit.

Left and/or right generator is not connected tothe airplane bus. Illumination of both annuncia-tors triggers the master warning system, whichwill illuminate the MASTER WARNING light.

Number 1 or 2 inverter output voltage is above130 VAC or below 90 VAC. Inverter failuretriggers AC FAIL annunciator. Resetting MAS-TER WARNING will extinguish the AC FAILannunciation unless both INVERTER FAIL lightsare illuminated.

Emergency pressurization has been manuallyselected or automatically activated by an aircycle machine overheat.

High flow rate of bleed air has been selectedfrom the right engine for ground operation ofthe air conditioner.

Power brake hydraulic pressure is low.

Bypass on the AMBER fuel low pressure lightpressure is low in left and right systems.

ENGINE

ANTI-ICE

LH RH

SURFACE

DEICE

GEN

OFF

LH RH

INVERTER

FAIL

1 2

EMER

PRESS ON

BLD AIR

GND

POWER

BRAKE

LOW PRESS

FUEL FLTR

BY PASS

LH RH

R-6/9/00




Electric power has been applied to the left and/ or right fuel boost pump.

Hydraulic reservoir is low.

Hydraulic system is pressurized.

Left and/or right pilot heat is off.

Left and right speedbrakes are fully extended.

Ventilation duct temperature exceeds safelimits.

Air cycle machine pressure is over 42 psi.

Anti skid system is inoperative.

Cabin door, aft compartment access door lock,or either nose baggage door lock is/are notlocked.

Left and/or right fuel and hydraulic shutoffvalves are closed.

Bleed air to the windshield exceeds safe tem-perature limits, or >5 PSI in the plumbing withthe system off.

FUEL

BOOST ON

LH RH

HYD LOW

LEVEL

HYD PRES

ON

P/S HTR

OFF

LH RH

SPEED

BRAKE

EXTEND

AIR DUCT

O'HEAT

ACM

O'PRESS

ANTI SKID

INOP

DOOR NOT

LOCKED

F/W

SHUTOFF

LH RH

W/S AIR

O'HEAT

R-6/9/00





A4-5

AC FAILR OIL

PRESS LO

BATT

O’HEAT

CAB ALT

10,000 FT

L HYD

PRESS LOR HYD

PRESS LO

L OIL

PRESS LO

ANTI-SKID

INOP

R GEN

OFF

W/S AIR

O’HEAT

BLEED AIR

GND/HI

L ENG

ICE FAILR ENG

ICE FAIL

L GEN

OFF

PWR BRK

PRESS LO

R FUEL

PRESS LOAIR DUCT

O’HEAT

EMER PRESS

ON

L F/W

SHUTOFF

R F/W

SHUTOFF

L FUEL

PRESS LO

HYD PRESS

ON

R FUEL

BOOST ON

DOOR NOT

LOCKED

ACM EJECTOR

ON

L PRECOOL

FAILR PRECOOL

FAIL

L FUEL

BOOST ON

HYD LEVEL

LO

R FUEL

LEVEL LO

P/S HTR

OFF SURF DEICE

SPD BRAKE

EXTENDED

FUEL FILT

BYPASS

L FUEL

LEVEL LO

1

2 3 4

Annunciator Panel

Airplanes 550-0550 and earlier

Airplanes 550-0482, 550-0485~550-0505

Annunciators

Citation II advisory lights (annunciators) are designed to pro-vide an easily interpreted representation of both normal andabnormal system conditions. The majority of annunciators are

collocated on an annunciator panel located on the centerinstrument panel.

ACM

O’PRESS

BLD AIR

GNDSPARE SPARE

KEY:Red

Amber

White

requires immediate attention, hazardouscondition exists.

requires attention, possible dangerouscondition exists.

safe or normal configuration, routine

action.

1

3 42

R-6/9/00




Annunciator Test

If the system is functioning normally, all annunciators and theMASTER WARNING lights should illuminate when the TESTselector knob on the lower left instrument panel is positionedto "ANNUN" and the BATT switch is in the "ON" position.

AC power bus voltage is above 130 VAC orbelow 90 VAC. Illumination of light triggersthe master warning system, which will illumi-nate the MASTER WARNING light.

Steady illumination: battery temperature over145°F; flashing: battery temperature over160°F. Illumination of either annunciatortriggers the master warning system, whichwill illuminate the MASTER WARNING light.

Cabin pressure altitude is above 10,000 feet.Illumination of either annunciator triggers themaster warning system, which will illuminatethe MASTER WARNING light.

Left and/or right hydraulic system flow isbelow approximately 0.35 to 0.55 gallonsper minute.

Oil pressure is below safe limits (35 PSI) inleft or right engine. Illumination of light alsotriggers the master warning system, whichwill illuminate the master warning light.

Anti-skid system is inoperative.

Bleed air to the windshield exceeds safetemperature limits, or >5 PSI in the plumbingwith the system off.

High flow rate of bleed air has been selectedfrom the right engine for ground operation ofthe air conditioner.

AC

FAIL

BATT

O'HEAT

CABIN ALT

10,000 FT

L HYD

PRESS LO

R HYD

PRESS LO

L OIL

PRESS LO

R OIL

PRESS LO

ANTI-SKID

INOP

W/S AIR

O'HEAT

BLEED AIR

GND/HI

R-6/9/00





A4-7

Air cycle machine pressure is over 42 psi.(550-0482, 550-0485~550-0505)

Left and/or right engine inlet, stator, or in-board wing anti-ice has failed.

Left and/or right generator is not connectedto the airplane bus. Illumination of bothannunciators triggers the master warningsystem, which will illuminate the MASTERWARNING light.

Power brake hydraulic pressure is low.

Ventilation duct temperature exceeds safelimits.

Emergency pressurization has beenmanually selected or automatically activatedby an air cycle machine overheat.

Left and/or right fuel and hydraulic shutoffvalves are closed.

Low fuel supply pressure to engine-drivenpump. Primary pump failure will automati-cally initiate boost pump operation as long

as FUEL BOOST switch is in the NORMposition. Light remaining on indicates failureof both pumps.

Hydraulic system is pressurized.

L ENG

ICE FAIL

R ENG

ICE FAIL

L GEN

OFF

R GEN

OFF

PWR BRK

PRESS LOW

AIR DUCT

O'HEAT

EMER

PRESS ON

L F/W

SHUTOFF

R F/W

SHUTOFF

L FUEL

PRESS LO

R FUEL

PRESS LO

HYD PRESS

ON

ACM

O'PRESS

R-6/9/00



Cabin door, aft compartment access doorlock, or either nose baggage door lock is/arenot locked.

Air cycle machine ejector on during groundoperation with both throttles below 85% or

with pressure applied to wheel brakes. (550-0550 and earlier)

High flow rate of bleed air has been selectedfrom the right engine for ground operation ofthe air conditioner. (550-0482, 550-0485~550-0505)

Air from precooler above 282°, +6 or -6°C.(550-0481 and earlier, -0483, -0484)Spare on (550-0482, 550-0485~550-0505)

Electric power has been applied to the leftand/or right fuel boost pump.

Fluid in the hydraulic reservoir is low.

Left and/or right pilot heat is off.

Illuminates twice during the 12- secondsurface deice boot cycle to indicate properboot inflation pressure. Tail boot inflationcannot be checked visually from the cockpit.

Left and right speedbrakes are fully ex-tended.

ACM EJECTOR

ON

L PRECOOL

FAIL

R PRECOOLFAIL

L FUEL

BOOST ON

R FUEL

BOOST ON

P/S HTR

OFF

HYD

LEVEL LO

SURF DEICE

SPD BRAKE

EXTENDED

DOOR NOT

LOCKED

BLD AIR

GND

cessna citation ii tm [pwd] dec 1999

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