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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL VOLUME 2

FlightSafety International, Inc. Marine Air Terminal, LaGuardia Airport Flushing, New York 11371 (718) 565-4100 www.flightsafety.com

Courses for the Citation Jet 3 525B are taught at the following FlightSafety learning center:

Wichita (Cessna) Maintenance Learning Center 1962 Midfield Road Wichita, Kansas 67209 (316) 220-3250 (800) 491-9796 FAX (316) 220-3275

Copyright © 2007 by FlightSafety International, Inc. All rights reserved. Printed in the United States of America.

INSERT LATEST REVISED PAGES, DESTROY SUPERSEDED PAGES LIST OF EFFECTIVE PAGES Dates of issue for original and changed pages are: Original ......0 .......... March 2007 NOTE: For printing purposes, revision numbers in footers occur at the bottom of every page that has changed in any way (grammatical or typographical revisions, reflow of pages, and other changes that do not necessarily affect the meaning of the manual). THIS PUBLICATION CONSISTS OF THE FOLLOWING: Page No.

*Revision No.

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*Zero in this column indicates an original page.

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

NOTICE The material contained in this training manual is based on information obtained from the aircraft manufacturer’s pilot manuals and maintenance manuals. It is to be used for familiarization and training purposes only. At the time of printing it contained then-current information. In the event of conflict between data provided herein and that in publications issued by the manufacturer or the FAA, that of the manufacturer or the FAA shall take precedence. We at FlightSafety want you to have the best training possible. We welcome any suggestions you might have for improving this manual or any other aspect of our training program.

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

CONTENTS VOLUME 2 ATA Chapter Title

Number

LANDING GEAR

32

LIGHTS

33

NAVIGATION

34

OXYGEN

35

PNEUMATICS

36

WATER/WASTE

38

CENTRAL MAINTENANCE SYSTEM

45

AUXILIARY POWER UNIT

49

STRUCTURES

51–57

POWERPLANT

71–80

WALKAROUND LIMITATIONS AND SPECIFICATIONS APPENDIX A—TERMS AND ABBREVIATIONS APPENDIX B—SYMBOLOGY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 32 LANDING GEAR CONTENTS Page INTRODUCTION ................................................................................................................. 32-1 GENERAL ............................................................................................................................ 32-3 LANDING GEAR ................................................................................................................. 32-5 Description..................................................................................................................... 32-5 Components ................................................................................................................... 32-5 Controls and Indications.............................................................................................. 32-17 Operation ..................................................................................................................... 32-19 NOSEWHEEL STEERING ................................................................................................ 32-27 Description................................................................................................................... 32-27 Components ................................................................................................................. 32-27 WHEELS............................................................................................................................. 32-29 Description................................................................................................................... 32-29 BRAKES ............................................................................................................................. 32-31 Description................................................................................................................... 32-31 Components ................................................................................................................. 32-35 Controls and Indications.............................................................................................. 32-41 Operation ..................................................................................................................... 32-41

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CITATIONJET 3 525B

ILLUSTRATIONS Figure

Title

Page

32-1

525B Landing Gear Systems.................................................................................. 32-2

32-2

Main Gear Assembly.............................................................................................. 32-4

32-3

Main Gear Actuators .............................................................................................. 32-6

32-4

Nose Gear Assembly .............................................................................................. 32-8

32-5

Shimmy Damper .................................................................................................. 32-10

32-6

Nose Gear Door Operation .................................................................................. 32-12

32-7

Nose Gear Actuator.............................................................................................. 32-14

32-8

Landing Gear Control Handle.............................................................................. 32-16

32-9

Hydraulic System Schematic ............................................................................... 32-18

32-10

Emergency Extension Bottle Installation............................................................. 32-20

32-11

Regenerative Shuttle Valve .................................................................................. 32-22

32-12

Emergency Extension Handle Assembly ............................................................. 32-24

32-13

Nosewheel Steering.............................................................................................. 32-26

32-14

Main Gear Wheel ................................................................................................. 32-28

32-15

Brake System Schematic...................................................................................... 32-30

32-16

Brake System Controls ........................................................................................ 32-32

32-17

Metering Valve ..................................................................................................... 32-34

32-18

Antiskid Controls and Indications ....................................................................... 32-36

32-19

Digital Control Unit ............................................................................................. 32-38

32-20

Antiskid System Self-Monitoring ........................................................................ 32-40

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 32 LANDING GEAR

INTRODUCTION This chapter describes the landing gear and brake systems for the Citation 525B aircraft (Figure 32-1). System descriptions and operation are included, accompanied by diagnostic information. References for this chapter and further specific information can be found in Chapter 5—“Time Limits/Maintenance Checks,” Chapter 12—“Servicing,” and Chapter 32—“Landing Gear” of the Aircraft Maintenance Manual (AMM).

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Figure 32-1. 525B Landing Gear Systems

32-2

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

GENERAL

NOTES

The landing gear is electrically controlled and hydraulically actuated during normal extension and retraction. Auxiliary extension is manually operated by cable/uplock release and gear free fall. Pneumatic (emergency) gear extension is manually controlled and pneumatically actuated. Selected engine thrust, rudder pedal steering, and brakes are used for nosewheel steering. The nosewheel steering is accomplished by cables connected to the rudder pedals. The main landing gear wheels are a tubeless lock-ring-type assembly constructed of forged aluminum. Each wheel assembly consists of a wheel base assembly and a side rim assembly secured together with a lock ring. The wheels and brakes provide a system for rolling and stopping the aircraft. The brakes can be actuated by either the power brake system or the pneumatic emergency brake system. The power brake system is monitored by an antiskid brake system.

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

TRUNNION SUPPORT

BEARING WASHER ACTUATOR TRUNNION PIVOT PIN

TRUNNION SUPPORT ROLL PIN PIVOT PIN

BOLT DOOR LINKAGE WASHER KEY WAY ROLL PIN

BEARING

BEARING

Figure 32-2. Main Gear Assembly

32-4

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

LANDING GEAR

Landing gear position and warning indicators alert the flight crew of the position of the landing gear.

DESCRIPTION The main landing gear provides the major support for the aircraft while on the ground (Figure 32-2). Each main landing gear strut has a mechanically operated door that is linked to the main gear strut. Each main gear consists of a single wheel assembly and an oil-air (nitrogen) strut. The nose gear tire is chined for water and slush deflection. The nose landing gear provides support for the nose of the aircraft while on the ground. The nose gear has three doors. The aft “spade” door is linked to the nose gear strut and extends and retracts with the strut. The two forward side doors for the nose gear are actuated through a system of link rods and torque tubes and are mechanically opened and closed as the nose gear extends and retracts. The steering system controls the direction of movement of the aircraft on the ground. Extension and retraction of the landing gear is accomplished by actuators powered by the aircraft hydraulic system. The actuators for each gear incorporate internal mechanical downlocks to hold the gear in the extended position. The gear is held retracted by mechanical uplocks that are spring-loaded to lock and hydraulically release. The extension and retraction system includes operating controls, valves, cables, and plumbing.

COMPONENTS Main Gear Assembly Each main landing gear consists of a trunnion, trailing link, and shock strut. The gear is hydraulically operated and retracts inboard. The main landing gear has three attach points in the wing, one on each side of the trunnion and one on the actuator. The landing gear pivots on trunnion pins, which are in the bearing supports. The shock strut assembly is an airover-oil strut containing a metering pin and orifice, which varies the resistance to shock according to severity. The gear has a mechanical latching mechanism in the up position. The actuator has a selflocking provision in the gear-down position. This mechanical lock holds the gear in the down position; hydraulic pressure ensures the security of the gear position.

Main Gear Door The main landing gear door is attached to the wing structure and the main landing gear trunnion. When the gear is extended, the door opens. When the gear is retracted, the door closes.

Components of the landing gear include: • Main gear assembly • Main gear door • Main gear actuators • Landing gear uplocks • Nose gear assembly • Nose gear doors • Nose gear actuator

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

UPLOCK HOOKS LOCKED BALL SEATED

SEQUENCE NOT ACTUATED NO FLOW TO GEAR ACTUATOR

PRESSURE

UPLOCK HOOKS UNLOCKING BALL SEATED

GEAR RETRACT LINE CONNECTION FROM GEAR ACTUATOR RETRACT PORT

UPLOCK HOOK CONNECTION NO FLOW TO GEAR ACTUATOR

PRESSURE

UPLOCK HOOKS UNLOCKED BALL UNSEATED

SEQUENCE ACTUATED

FREE FLOW TO GEAR ACTUATOR

Figure 32-3. Main Gear Actuators

32-6

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Main Gear Actuators The main gear actuators (Figure 32-3) are inboard of each main landing gear. The actuators are normally actuated hydraulically. During emergency (auxiliary) extension, the actuators are operated pneumatically. When the actuator extends, the gear extends to the downand-locked position. The actuator has an integral locking mechanism that locks the landing gear in the full extended position. When hydraulic pressure is applied to the actuator retract port, pressure overcomes an internal spring, allowing the lock to move. While hydraulic pressure continues, the actuator piston applies force on the locks, moving them into the unlock position and allowing the actuator to fully retract. An electrical downlock switch is on the bottom of the actuator. The switch is actuated by the keylocks and electrical circuitry. Position lights in the cockpit indicate to the pilot when the landing gear is down and locked.

uplock mechanism. Manual uplock hook release is accomplished by a cable attached to the uplock mechanism. This cable is routed to a T-handle in the cockpit, allowing manual activation of both the main and nose landing gear uplock hooks simultaneously. An unlock and sequence actuator is at each landing gear uplock hook. The mechanical uplock hook must fully release the landing gear hydraulic pressure passes on to the gear actuator. A check valve in the uplock and sequence actuator prevents the passage of hydraulic fluid until the uplock hook is fully released and the sequence actuator rod has retracted far enough to unseat the check valve. Reverse hydraulic flow during gear retraction unseats the check valve and permits passage of returning fluid. Maintenance for the left and right main landing gear uplock assemblies is the same. The hydraulic uplock and sequence actuators are identical for all three landing gear uplock hooks.

Landing Gear Uplocks Three landing gear uplocks are mechanical hook-type latches that hold the gear in the retracted (up) position. The hooks are springloaded to a locked position as the gear is retracted. Once the gear is fully retracted, an uplock switch opens the retract circuit and releases hydraulic pressure. During normal gear extension, the uplock and hydraulic sequence actuator unlock the hooks to release the gear before directing hydraulic pressure to the gear actuator. During emergency (auxiliary) gear extension, the uplock hooks are manually released by a pull cable overcoming the uplock hook spring. The landing gear uplocks are spring-loaded to lock the gear in the up position and manually or hydraulically actuated to unlock the gear. The uplock hooks automatically lock the gear when the gear unlock rollers contact the unlock hooks. Releasing the gear from the uplock hooks is normally accomplished by an uplock and hydraulic sequence actuator attached to the

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

STEERING GEARS SHIMMY DAMPER

FORWARD DOOR LINKAGE TUBE

STEERING UNIVERSAL JOINT EXTEND/RETRACT ACTUATOR

TRUNNION

AFT DOOR LINKAGE

BONDING JUMPER FORK-NOSE WHEEL

TIE STRAP TORQUE LINKS

Figure 32-4. Nose Gear Assembly

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Nose Gear Assembly The nose landing gear consists of a wheel and tire attached to a shock strut mounted on a trunnion (Figure 32-4). Torque links connect the strut cylinder to the wheel fork. The torque links hold the wheel in alignment with the strut cylinder to steer the aircraft. The gear is hydraulically operated and retracts forward. The nose landing gear has three attach points, one on each side of the trunnion and one on the hydraulic cylinder. The trunnion rotates on teflon-lined bearings. The nose landing gear is hydraulically retracted and extended. The nose gear can also be released for free fall or pneumatically extended. The actuator is attached between the trunnion and the airframe structure. The nose landing gear retracts forward with the hydraulic actuator located on the aft side. When the hydraulic actuator retracts, it pulls the gear to the extended position. Extension of the actuator pushes the gear to the retracted position. The gear has a mechanical latching mechanism in the up position. The hydraulic actuator has an integral locking mechanism in the retracted position to hold the gear in the extended position without hydraulic pressure or mechanical devices.

A shimmy damper is attached to the strut to dampen rapid movements of the nose wheel.

Shock Strut The shock strut has a hydraulic chamber to absorb landing shock. A metering pin, extending into the top of the piston, controls the flow of fluid between the piston and cylinder. The metering pin is tapered. When the strut is fully extended, the small diameter of the metering pin provides a large orifice to allow rapid movement of the strut for the initial contact on landing. As the strut is compressed, the diameter of the metering pin increases, restricting the orifice. The lower volume of fluid passing through the orifice slows the movement of the shock strut. When the strut extends after takeoff, the orifice slows the movement to prevent damage when the limit is reached. The shock strut has an air chamber that provides spring energy and is inflated through an air valve. The compressed air supports the aircraft and acts as a spring to support aircraft weight. An isolation piston moves inside the shock strut piston to separate the air and oil chambers.

One door covering the aft part of the wheel well is linked directly to the nose landing gear strut and remains open when the gear is in the down position. Two doors cover the forward portion of the wheel well opening. These two doors are linked mechanically to the nose landing gear trunnion and are closed when the gear is in either the up or down position. The nose landing gear is used for steering and towing the aircraft. The steering gear assembly, bolted to the top of the cylinder, provides a towing stop. Forcing the nose landing gear into a turning radius beyond the stop shears the bolts attaching the steering gear assembly to the cylinder. The damaged bolts must be replaced prior to the next flight.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

BOLT

NAS1149F1232P WASHER RETAINING RING WASHER NUT

NOSE LANDING GEAR SHIMMY DAMPER

SLEEVE

COTTER PIN

NOSE LANDING GEAR TRUNNION

CHECK VALVE

O-RING

CAP

O-RING AND BACKUP RETAINER

RETAINER CAP

MAKE-UP PISTON

REDUCER FITTING HEAD BEARING

SPRING PISTON

O-RING PLUG

ORIFICE

ORIFICE

PISTON ROD

SCREW

SEALING WASHER

BACKUP RETAINER SCRAPER RING AND LOCK RING

HEAD BEARING HOUSING

O-RING RETAINING RING

RETAINING RING

Figure 32-5. Shimmy Damper

32-10

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

NOTES

Shimmy Damper The nosewheel shimmy damper (Figure 32-5) is a hydraulic unit designed to absorb and dampen nosewheel shimmy during aircraft takeoff, landing, and taxi. The damper cylinder and piston are on a rotating steering gear at the top of the landing gear strut assembly. The piston rod is bolted to a fixed arm on the landing gear trunnion. The nose landing gear shimmy damper incorporates a piston with an orifice to restrict fluid movement when the piston moves through a hydraulic-fluid-filled cylinder. The restricting action dampens rapid movements of the nosewheel. The damper also has a compensating chamber consisting of a makeup piston, spring, check valve, and relief valve. The chamber relieves thermal expansion of the hydraulic fluid. Nosewheel shimmy is transmitted through the shock strut to a steering gear on top of the strut assembly. As the steering gear turns, the barrel of the shimmy damper slides left or right along the piston rod. The fluid within the barrel is forced through an orifice in the damper housing. The restricted fluid flow through the orifice absorbs and dampens the movement of the barrel, steering gear, and nose gear wheel. The shimmy damper is on the nose gear steering assembly (Figure 32-5) with the shimmy damper rod end attached to the trunnion. Access to the shimmy damper is through the nose baggage compartment access panel. In the event of a nose landing gear shimmy damper malfunction or excessive leakage, remove the shimmy damper and refer to the Component Maintenance Manual for repair.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

TORQUE TUBE BELLCRANK

FORWARD DOOR LINKAGES

TRUNNION TO BELLCRANK LINKAGE

TRUNNION

AFT DOOR LINKAGES

FORWARD DOORS

AFT DOOR

HINGE ARMS

EXTENDED POSITION

FORWARD DOOR LINKAGES

TRUNNION TO BELLCRANK LINKAGE

AFT DOOR LINKAGE

AFT DOOR

RETRACTED POSITION

Figure 32-6. Nose Gear Door Operation

32-12

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Nose Gear Doors

NOTES

The two main nose landing gear doors (Figure 32-6) enclose the nose landing gear wheel well when the gear is retracted or extended. When the doors are closed, they fit flush with the fuselage skin. The doors are hinged to the aircraft structure and are actuated by linkage attached to the nose landing gear. The nose landing gear doors are constructed of honeycomb cores, bonded doublers, and skin panels. Two hinges are on each of the three doors, attaching them to the fuselage structure. Electrical bonding of the doors is through bond straps attached to the doors and aircraft structure. Three doors enclose the nose wheel well when the nose landing gear is in the retracted position. Two nose landing gear doors cover the forward portion of the wheel well and the spade door covers the aft portion. The two nose landing gear doors are linked to a torque tube bellcrank, which is linked to the nose landing gear trunnion. The spade door is mechanically linked to the nose landing gear trunnion, below the gear actuator attach fitting, and only closes when the gear is retracted. The left and right nose landing gear doors are opened and closed through a linkage attached to a torque tube. The torque tube is connected to the nose gear through a linkage. When the gear is raised and lowered, the torque tube is rotated, causing the doors to open and close. The spade door is connected directly to the nose landing gear. The left and right nose landing gear doors are normally closed, except during nose landing gear extension or retraction.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

SUPPORT BOX FITTING (UNDER COCKPIT FLOORBOARD PANEL)

NOSE LANDING GEAR ACTUATOR BOND STRAP

ROD END EXTEND LINE DOWNLOCK SWITCH CLAMP BOOT

LOCKNUT

NOSE LANDING GEAR STRUT FITTING

Figure 32-7. Nose Gear Actuator

32-14

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Nose Landing Gear Actuator

NOTES

The landing nose gear actuator (Figure 32-7) is aft of the nose gear assembly. The actuator is enclosed in a cone-shaped housing or boot extending aft from the forward pressure bulkhead. A bolt securing the aft end of the actuator is at the small end of the housing/boot. Access to the bolt is through cockpit floorboard panels. The actuator is normally activated hydraulically. During emergency extension of the gear, the actuator is activated pneumatically. The use of a single shuttle valve in the system at the manifold admits pressure from either source. When the actuator piston retracts, the nose gear extends to the down-and-locked position. The fittings to the actuator are standard AN815 unions. An insert restrictor is in the retract line just outside the nose wheel well. The nose gear actuator has an internal locking mechanism to lock the gear in the full extended (down) position. The nose gear actuator has only the electrical downlock switch to indicate that the actuator internal locking mechanism has the gear down and locked. The downlock switch is not field-adjustable.

Landing Gear Control Valve The landing gear control valve determines which direction hydraulic fluid flows to the actuators. A spool valve within the control valve is positioned by two solenoids and controls the pressure and return flow direction in the system. Deenergized (electrical power off), the spool valve maintains a neutral position blocking the hydraulic inlet port. When the retract or extend solenoids are energized, hydraulic pressure is directed to the gear actuators to extend or retract the gear. The landing gear control valve functions with an 18- to 30-VDC power source. The internal selector spool in the control valve maintains a neutral position until hydraulic inlet pressure repositions the spool. The position of the spool depends on which solenoid (retract/extend) is energized.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

GEAR SELECTOR SWITCH (SI003)

A

SOLENOID PLUNGER

LANDING GEAR CONTROL HANDLE SOLENOID (WI001)

A

LANDING GEAR CONTROL RETRACT SWITCHES (SI053 AND SI054)

LANDING GEAR CONTROL HANDLE PIVOT POINT

CAM

LANDING GEAR CONTROL EXTEND SWITCHES (SI051 AND SI052)

PIN

VIEW A-A

Figure 32-8. Landing Gear Control Handle

32-16

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CONTROLS AND INDICATIONS

Gear Down Indication

Landing Gear Control Handle

Electrical power is present at each safe light when the GEAR CONTROL circuit breaker on left CB panel is engaged. A switch provides an individual ground circuit, which causes the indicator light to illuminate when the landing gear actuator locks are in the down position.

The landing gear normal retraction/extension system is controlled electrically by the action of the landing gear control handle (Figure 328). The landing gear control handle assembly includes retract and extend switches, a gear selector switch, and a control handle locking solenoid. The landing gear control handle locking solenoid prevents the handle from moving to a gear-up position while the aircraft is on the ground. The landing gear control handle actuates electric switches that open and close circuits controlling the retract and extend solenoid on the landing gear control valve. The control handle is spring-loaded in either the retract or extend position. To move the control handle from one position to another, it must be pulled outward to allow a pin on the handle to move over a cam. The pin actuates either the retract or extend switch, depending on which detent the control handle is in. A gear select solenoid on the landing gear control handle assembly has a spring-loaded plunger that prevents movement of the control handle while the aircraft is on the ground. The solenoid is activated by a landing gear squat switch and receives electrical power from the gear position and warning electrical system.

Gear Unsafe Indication The GEAR UNLOCKED indicator has two bulbs in parallel for continued operation should either light burn out. The indicator illuminates when any of the following conditions exists: • Gear selector switch is in the down position and one or more gear are not in the down-and-locked position. • Gear selector switch is in the up position, one or more gear are not up and locked and one or more gear are not down and locked. At least one gear has unlocked from the down position but has not moved to the up-and-locked position. • Aircraft is on the ground, the gear is in the down-and-locked position and the selector switch is in the up position. • Rotary TEST knob is in the LDG GEAR test position. This position provides a separate power source to the GEAR UNLOCKED indicator, and the gear test relay provides a ground regardless of the position of other switches in the system.

A gear selector switch is also on the landing gear control handle assembly and controls the landing gear up or down indicator lights. The gear selector switch is a primary component of the gear position and warning electrical system.

Landing Gear Position Indication The landing gear position system provides a visual indication of status of the landing gear. Three green (gear safe) lights (L, R, and NOSE) are on the LDG GEAR control panel at the bottom of the instrument panel.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

ACTUATOR ASSY, NLG UPLOCK

PNEUMATIC STORAGE AND CONTROL SYSTEM

CHECK/BREATHER VALVE NOSE LANDING GEAR ACTUATOR

LOADING VALVE PRESSURE SWITCH

ACTUATOR ASSY. MLG UPLOCK

ACTUATOR ASSY, MLG UPLOCK

REGEN SHUTTLE VALVE

REL VALVE, SYS CONTROL V, LG REGEN SHUTTLE VALVE

SHUTTLE V DUMP V. PNEUMATIC CONTROL V SPEED BK

LEFT MAIN LANDING GEAR ACTUATOR 49°

68°

RIGHT MAIN LANDING GEAR ACTUATOR 49°

68° LEFT SPEEDBRAKE

0

RIGHT SPEEDBRAKE

SOV, GND

15

0

LEFT FLAP

35 55

FLOW CONTROL ASSY

CHECK VALVE

35 55

MANF ASSY, FLAP RET

REL V, THERMAL DUMP V, SPD BK

CHECK VALVE

FILTER, PRESS

GND SVC PORT PRESS

CONTROL V, FLAP MANIFOLD ASSY, HYD

RESERVOIR LEVEL IND

FILTER, RTN

RESERVOIR VENT V RESERVOIR REL V

CV, RTN GND SVC PORT, RTN

GND SVC PORT O’BD DUMP FLOW SWITCH/CV

FILTER, GEAR RTN

SECONDARY FILTER CV, SUCTION

CV, RTN M

M

PUMP, ENG DR, HYD

FIREWALL SHUTOFF V

Figure 32-9. Hydraulic System Schematic

32-18

15

RIGHT FLAP

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Gear Warning Aural Indication The landing gear warning horn sounds when any of the following conditions exists: • The aircraft airspeed is below 130 knots (241 km/hr), either or both throttles are retarded below 85% N 2, and one or more landing gear are not down and locked. The warning horn may be silenced for this condition by depressing the HORN SILENCE button on the gear control panel. The horn resets if the throttle(s) is advanced to (or above) 85% N 2 . Closing (depressing) the HORN SILENCE button energizes two horn silence relays that stop the warning horn. Two relays (one for each throttle) remain energized as long as the throttles are below 85% N 2 . Advancing the throttles deenergizes their respective relays, allowing the warning horn to sound again if the throttle(s) is moved below 85% N 2 . • The flaps are extended beyond the approach position (15%) and the gear is not down and locked. In this condition, the warning horn cannot be silenced by pressing the HORN SILENCE button. • The aircraft is on the ground, the gears are in the down-and-locked position, and the selector switch is in the up position.

OPERATION Hydraulic Gear Extension When placing the landing gear control handle down, an electrical circuit is completed through the parallel gear downlock switches, the extend solenoid of the landing gear control valve, and the hydraulic system loading valve (Figure 32-9). The loading valve solenoid, when energized, closes the valve and routes full-flow hydraulic fluid to the landing gear control valve. The extend solenoid of the landing gear control valve, when energized, positions the flow ports in the control valve and directs hydraulic pressure to the individual uplock and sequence actuators.

The uplock and sequence actuators unlock the uplocks and direct hydraulic pressure to the extend side of the landing gear actuator pistons until each gear is fully extended and the gear downlock switches are actuated. When all downlock switches are actuated, the electrical circuit opens, removing power from the extend solenoid of the landing gear control valve and to the hydraulic system loading valve solenoid. This causes the hydraulic system loading valve to open. While the landing gear is down and locked, only the internal locks in the gear actuator hold the gear extended.

Hydraulic Gear Retraction When placing the landing gear handle up, an electrical circuit is completed through the parallel gear uplock switches, the retract solenoid of the landing gear control valve, and the hydraulic system loading valve. The loading valve solenoid, when energized, closes the valve and routes full-flow hydraulic fluid to the gear control valve. The retract solenoid winding of the landing gear control valve, when energized, positions the flow ports in the control valve and directs hydraulic pressure to the retract side of the gear actuator pistons, also releasing the actuator internal downlock. Hydraulic pressure continues until each landing gear is retracted and the uplock switches are actuated. When all gear uplock switches are actuated, the circuit is opened. This removes power from the retract solenoid winding of the landing gear control valve and to the hydraulic system loading valve solenoid and causes the hydraulic system loading valve to open. While the landing gear is up and locked, only the uplock hooks hold the gear retracted.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CONTROL VALVE LEVER VENT LINE (OVERBOARD)

SCREW

PNEUMATIC BOTTLE BRACKET LEVER RELEASE BRACKET AUXILIARY GEAR EXTENSION CONTROL CABLE

CLAMP BOLT

RELIEF VALVE WASHER NUT BACKUP RING PACKING O-RING

AUXILIARY GEAR EXTEND LINE

PRESSURE GAGE VENT-FILLER VALVE

BRAKE AIR PRESSURE LINE AIR BOTTLE CONTROL VALVE FILLER VALVE BOTTLE CLAMP EMERGENCY (AUXILIARY) AIR STORAGE BOTTLE MOISTURE BLEED VALVE

Figure 32-10. Emergency Extension Bottle Installation

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FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Emergency Gear Extension The emergency gear extension system consists of two subsystems. Utilizing the cableoperated system (actuated by the T-handle), the uplocks release and the gear free falls to a down-and-locked position. Pneumatic pressure supplied from a storage bottle released into the gear actuators by the T-handle ensures the down-and-locked condition. Emergency operation of the landing gear is restricted to extending the gear. The operation of two manual controls is required: • Pulling the emergency (auxiliary) Thandle • Rotating the T-handle clockwise/counterclockwise to lock the T-handle out Pulling the T-handle overrides (compresses) the springs on the gear uplock hooks, releasing the hooks and allowing the landing gear to fall from the wheel wells under gravity. When rotating the T-handle to the locked position, a groove in the T-handle control assembly aligns with the mounting shaft, permitting pulling the emergency (auxiliary) control knob (nitrogen). Pulling the emergency (auxiliary) control knob actuates the emergency (auxiliary) air storage bottle. High-pressure air is released to the extend side of the landing gear actuating pistons, ensuring that the gear is driven to the downand-locked position. At the same time, air pressure positions the dump valve, permitting hydraulic fluid from the actuators to return directly to the hydraulic reservoir.

The emergency extension lines remain pressurized until the high-pressure air is released by depressing a button and positioning the control handle on the air storage bottle to the normal position. The control handle on the bottle latches in the release position, requiring that the control handle on the emergency (auxiliary) air storage bottle be reset to return the handle to the normal position. The high-pressure air storage bottle supplies air to the auxiliary brake system, as well as to the auxiliary landing gear extension system. When high-pressure air is released into the landing gear system, the landing gear must be cycled before the next flight to purge the air and restore the proper hydraulic fluid quantity.

Emergency Air Storage Bottle The emergency (auxiliary) air storage bottle is on the forward side of the forward pressure bulkhead (Figure 32-10). The emergency storage bottle holds 75–100 cubic inches (1,229–1,639 cubic centimeters) of nitrogen at 2,000 psi (13,790 kPa). A relief valve ruptures and releases excessive pressure at 4,000 psi (27,580 kPa). Emergency gear extension lines are connected to a vent line while the air storage bottle control lever is in the normal position. The vent line provides a route for venting the air extension chambers of the main gear actuators during hydraulic operation of the gear. The relief valve is not reusable after rupture, and either the relief valve and/or the complete emergency (auxiliary) air storage bottle must be replaced.

FOR TRAINING PURPOSES ONLY

32-21

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Figure 32-11. Regenerative Shuttle Valve

32-22

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

NOTES

Dump Valve When the landing gear is extended by the emergency (pneumatic) system, return hydraulic fluid from the gear actuators is forced back to the hydraulic reservoir through the dump valve. The dump valve is connected to the gear retract line, hydraulic system return line, and pneumatic extension line. Pneumatic gear extension pressure of 200 psig (or more) at the dump valve inlet port opens the dump valve to pass hydraulic return fluid with 0–1,500 psig at the inlet port.

Regenerative Shuttle Valve The regenerative shuttle valve (Figure 32-11) assists with manual gear release (or free fall). The regenerative shuttle valve eliminates forcing the retract fluid through the system and back into the pressurized reservoir. The shuttle valve has a normally open path from the retract side of the actuator back to the extend side of the actuator. Consequently, during free fall, the retract fluid is forced out of the actuator, through the regenerative shuttle valve, and into the extend side of the actuator. There is no back pressure to overcome and the extend side of the actuator draws the retract fluid out of the actuator. The valve incorporates a shuttle valve to accomplish normal operation. A direct path from retract to extend exists while the valve is static. When a gear retract is commanded, the retract pressure shuttles the valve to block the passageway to the extend side and the pressure is routed directly to the retract side of the actuator.

FOR TRAINING PURPOSES ONLY

32-23

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CONTROL HANDLE

NOSE LANDING GEAR UPLOCK CABLE WASHER

MOUNTING SHAFT

WASHER PIN

NUT SCREW

PIN

PULLEY BRACKET

PULLEY

MAIN LANDING GEAR UPLOCK CABLE

PULLEY CABLE ASSEMBLY PULLEY

PULLEY

TURNBUCKLE

PIN

SCREW

Figure 32-12. Emergency Extension Handle Assembly

32-24

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Emergency Gear Extension Handle

NOTES

The emergency (auxiliary) gear control consists of a T-handle (Figure 32-12) to unlock the gear uplock hooks and a round collar to release high-pressure air from a storage bottle. The gear blowdown collar cannot be pulled until the T-handle has been pulled and turned to the locked position. The T-handle is connected to the uplock hooks by a cable system. The gear blowdown knob is connected to the air bottle valve lever by cable. The emergency (auxiliary) extension of the landing gear is manually and pneumatically actuated. Operation of the emergency gear extension is accomplished in two steps: 1. The pull-type T-handle connects directly by cable to each landing gear uplock hook. Pulling the T-handle unlocks the uplock hooks and allows the landing gear to free fall to the extended (down) position. 2. The pull-type knob behind the auxiliary T-handle connects by cable to a control lever on a emergency (auxiliary) air storage bottle. Pulling the round knob actuates the storage bottle control lever. High-pressure air is routed to a shuttle valve, which introduces air into the hydraulic supply line, extending the actuators. Hydraulic fluid on the return side of the pistons is directed through a dump valve, which allows the fluid to return directly to the hydraulic reservoir.

FOR TRAINING PURPOSES ONLY

32-25

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

BOLT

FITTING

BUNGEE

STEERING ARM

WASHER SELF-LOCKING NUT

COTTER PIN

WASHER NUT COTTER PIN

NUT CABLE CLEVIS UPPER CONTROL CABLE BOLT

WASHER

BELLCRANK SUPPORT SCREW LOWER CONTROL CABLE

NUT TURNBUCKLE COTTER PIN TURNBUCKLE CLEVIS WASHER

BOLT

Figure 32-13. Nosewheel Steering

32-26

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

NOSEWHEEL STEERING DESCRIPTION Selected engine thrust, rudder pedal nosewheel steering, and brakes provide directional control of the aircraft. Nosewheel steering is through cables connected to the rudder pedals at both pilot stations. The nosewheel steering turning limit is limited by rudder stops. The turning limit is approximately 20° either side of center. A bungee spring allows the nose gear to turn past the limits of the rudder control cables (Figure 32-13). The centerline of the steering universal joint aligns with the centerline of the trunnion supporting bolts. When the nosewheel retracts, the lower half of the steering universal joint remains in position while the upper half, pivoting with the strut, is moved to the center position and automatically centers the nosewheel. With the nosewheel fully retracted, the upper half of the steering universal joint and the nosewheel remain stationary. The lower half of the steering universal joint moves freely, permitting normal operation of the rudder pedals.

Steering Universal Joint The steering universal joint transmits steering control to the wheel. Upon retraction, it automatically centers the nose gear. With the nose gear retracted, it allows the rudder pedals to operate without moving the nosewheel.

Steering Gears Steering gears on top of the trunnion transmit steering control to the strut. Bolts in the trunnion stop the gears when the turning limit is reached. The bolts are designed to shear when the turning radius is exceeded while the aircraft is being towed. A shimmy damper connected to the front gear prevents steering oscillations.

NOTES

COMPONENTS Bellcrank The bellcrank is on a mounting bracket on the left side of the nose wheel well. Cables from the rudder pedals connect to the top and bottom ends of the bellcrank, with the bungee connected to the upper end.

Bungee The bungee is a spring-loaded rod transmitting steering control from the bellcrank to the steering arm. The spring allows the nose gear to turn past the limits of the control cables when the aircraft is being towed.

FOR TRAINING PURPOSES ONLY

32-27

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

WHEEL BASE ASSEMBLY THERMAL FUSE PLUG

SIDE RIM ASSEMBLY LOCK RING

INFLATION VALVE

INSERT

HUBCAP (WITH ANTISKID DRIVE CLIP)

HEAT SHIELD OVERINFLATION PLUG

PREFORMED O-RING SEAL

Figure 32-14. Main Gear Wheel

32-28

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

WHEELS

The aircraft uses a 22” x 8.25”–10” tubeless 12-ply rated ribbed tread tire.

DESCRIPTION

Nose Landing Gear Wheel

Main Gear Wheel The main landing gear wheels are a tubeless lock-ring assembly constructed of forged aluminum. Each wheel assembly consists of a wheel base and a side rim assembly secured together with a lock ring (Figure 32-14). An inf l a t i o n va l ve i s i n t h e o u t e r w h e e l b a s e assembly. The main landing gear wheel base assembly and side rim are sealed when assembled together by a preformed packing (O-ring) in the mating surface of the wheel base assembly. The wheel rotates on tapered roller bearings. The bearing cups shrink-fit into the hub of each wheel half. The bearings are protected against dirt, moisture, contamination, and loss of lubricant by a steel-reinforced rubber bearing seal on the inboard side and the hub cap on the outboard side. The inboard bearing and seal are positioned against the brake housing bushing. A preformed seal prevents grease from entering the cavity between the halves.

The nose landing gear wheel consists of two castaluminum halves. The two wheel halves are secured together with bolts, washers, countersunk washers, and self-locking nuts. An O-ring is on one wheel half, providing an air seal at the junction of the two wheel halves. One wheel half has an inflation valve. Each wheel half is individually balanced at the time of manufacture, permitting the wheel halves to be assembled in any position relative to one another, and allows interchanging of wheel halves without the need for rebalancing. The wheel rotates on two tapered roller bearings. The roller bearing cup shrink-fits into the hub of each half. The bearings are protected from dirt by bearing seals built into the bearing. The aircraft utilizes an 18” x 4.4” DD tubeless dual deflector rib chine tread tire with 10-ply rating.

Inserts over bosses in the inner wheel half engage drive slots in the brake disk, rotating them as the wheel turns. The inserts are attached to the bosses with screws, washers, and nuts. An overinflation plug is 180° from the inflation valve on the wheel base assembly. Thermal fuse plugs are in the drive lugs of the wheel base assembly. The thermal fuse plugs melt at a predetermined temperature to prevent tire overinflation caused by a brake overtemperature condition. The main gear wheel assembly has inspection requirements at specific intervals. For the inspection requirements and more detailed information concerning the main landing gear wheel, refer to the BFGoodrich wheel and brake publication.

FOR TRAINING PURPOSES ONLY

32-29

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

RUDDER PEDAL ASSY

MIXER ASSY

LEGEND

BRAKE CABLES

METERED BRAKE PRESSURE ANTISKID PRESSURE PNEUMATIC PRESSURE CABLE LINES

PRESSURE INPUT

BRAKE METERING RETURN VALVE

PNEUMATIC BOTTLE

OVERBOARD VENT

PARKING BRAKE KNOB

TO EMERGENCY GEAR EXTEND SYSTEM

PRESSURE INPUT

ANTISKID CONTROL VALVE

RETURN

RH WHEEL AND BRAKE ASSY PARKING BRAKE VALVE LH WHEEL AND BRAKE ASSY

Figure 32-15. Brake System Schematic

32-30

FOR TRAINING PURPOSES ONLY

EMERGENCY BRAKE VALVE

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

BRAKES

NOTES

DESCRIPTION The Citation 525B aircraft uses hydraulically powered main landing gear brakes. Crew inputs to the brake metering valve are mechanically transmitted via a series of cables from the toe brakes on the rudder pedals (Figure 32-15). The brake metering valve regulates hydraulic pressure to the brakes based on pilot or copilot input. An electronic antiskid system monitors the main gear wheel speeds and reduces brake pressure as necessary to optimize stopping distance and prevent wheel lockup. A parking brake valve traps pressurized fluid in the brake lines and is controlled by a knob in the cockpit. The pneumatic brake system is a backup system that supplies pressure to the brake assemblies during hydraulic brake system failure. It supplies pressurized nitrogen from a bottle in the nose directly to shuttle valves on the brakes and is controlled by a lever in the cockpit.

FOR TRAINING PURPOSES ONLY

32-31

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

BATT O'TEMP

LH ENG FIRE

CABIN ALT

OIL PRESS WARN L R

> 160° ENG CTRL SYS FAULT L R

BOTTLE 1

RUDDER BIAS

ARMED PUSH

F/W SHUTOFF L

CABIN DOOR DOOR SEAL

GEN OFF

R

L

FLAPS >35°

BAGGAGE SMOKE NO TAKEOFF

HYD FLOW HYD PRESS LOW ON SPD BRK L R EXTEND

R

BAGGAGE DOOR FWD AFT

EMER EXIT STBY P/S HTR OFF

EMER PRESS ON FRESH AIR

P/S HTR OFF R

W/S AIR O'HEAT AOA HTR FAIL

FUEL GAUGE L R

AFT J-BOX LMT CB

GROUND IDLE

L

PWR BRK FUEL LOW FUEL FUEL FUEL LOW FUEL FLTR LOW PRESS PRESS BOOST ON TRANSFER LEVEL BYPASS ANTISKID L R L R L R L R INOP

BLD AIR O'HEAT L R

ENG ANTI-ICE L R

WING O'HEAT

WING ANTI-ICE

L

L

R

ENG T2 HTR FAIL L R

RH ENG FIRE

R

BOTTLE 2

VIDEO FAIL AUDIO FAIL

TAIL DEICE TAIL DEICE FAIL PRESS AIR DUCT L R O'HEAT

ARMED PUSH

Collins

B/C

FLC

APPR

FD

VNAV

Collins

N1

ALT

%RPM

0.0 VS

80

AP PTCH

80 60

10

ADF2

SPKR

CAB

HDPH

PA

70

USABLE FUEL 4710 LBS

1000

25

500

23 0

0

1200

TA/RA REL

N763CJ NEXT PAGE

TILT

N

RESET

RANGE

HDG 137

RNG

DTK DTK

PWR

Collins

°M

R

ON

START DISG NORM

WINDSHIELD BLEED AIR

OFF

ENG ON

NOT ARMED DIM

O F F ENG ON

OFF

MAX

OFF

BEACON

ANTI-COLL

NAV

R SLEW

OFF

OFF

OFF

ADF1

ADF2

MUTE

SPKR

CAB

HDPH

PA

PRESET FMS2

RANGE

ADF

USH AUTO TILT

TERR LX/RDR TERRAIN

TAS 0

RAT 22 °C

SAT 24 °C

ACTIVATED WHEN LIT

10 OXY

ISA+11 °C

5 BRT DIM

ENT

0.0kt

0 NT

ELT

TFC TA ONLY

ADF GS

DEFAULT NAV

15

GPS

0

CRS

PSI x100

20

ON

ARM TEST/RESET SELECT ON WAIT 1 SECOND SELECT ARM

OBS

MSG

FPL

VNAV

PROC

PUSH CRSR

DIR T EC

Collins

AHRS REV

DADC REV

NORM

NORM

GPWS FLAP OVRD

GPWS G/S

ACTIVE

CANCELED

NORM REV TO PFD

TERR NORM TERR INHIB

GPWS TEST

REV TO MFD

O N

LDG GEAR

OFF

MAX

FLOOD LTS WING INSP

INSTR

NIGHT DIM ON

STBY

OFF DIM

DIM

L ENG

R ENG

FADEC CH B

FADEC CH B

FADEC RESET

FADEC RESET

LIGHTS

EL

DIM

DIM

SEAT BELT

L

R O F F

OFF

NORM SILENCE

LANDING

TAIL FLOOD

PASS SAFETY O F F

OFF

NOSE

UP

DIM

PANEL LIGHT CONTROL L SLEW

AUTO

MKR

VOX

P

MANUAL

EXTERIOR LIGHTS

L AHRS SLAVE MANUAL

DME2

ST

SH PU

DISPLAY PFD/MFD

ARMED

TILT

Collins

NAV NRST

BRT DIM

EMERGENCY LIGHTING

RIGHT

TAIL AUTO

WING/ENGINE L WING/ENG R WING/ENG O F F

OFF

LOW

HF

DME1

FORMAT

BRIEFING SELECTION

PLAY

NORM LEFT

WINDSHIELD ALCOHOL ON WING XFLOW BLEED HI O F F

OFF

COM3

NAV2

INPH V BOTH ID

S

GS

ENR

BRIEFING

NORM ANTI-ICE / DE-ICE

PITOT & STATIC

COM2

NAV1

15

CLR

TFC TA ONLY

ADF

COM1

29.92

138 12

MENU

n m

20 nm –1

ADF INTERCOM CALL

PBS - 250

ON

D

005°M DIS

AUTO OFF

OFF

L

TERR LX/RDR

°M

TRK

KCVG CVG

n m

2.5

PULSE LT TA/RA

IGNITION R

RE SS TO TEST

NORM

DISENGAGE

HDG 137

VOR2 CRS 245

GCS

WPT

S

DIS

ISA+12 °C

4

GPS 500

GARMIN

137 15

12

WPT

TTG – – : – – – – – – NM

USH AUTO TILT

TFC TA ONLY

SAT 24 °C

BRT DIM

TEST

ENGINE START L

R

O F F

ELECT

NAV/BRG

P

33

RAT 22 °C

TAS 0

2

1330.0

0

RADAR

VOR1 CRS 166

GCS

TERRAIN

30

ON

0

USH

1

EMERGENCY USE ONLY

EMER

FUEL BOOST L

ADF

2 00

20

W

RESET

GS 0

O F F

OFF

OFF

R ENG

25

°C FUEL

0 PPH

E

6 3

TERR LX/RDR

ADF

AVIONICS POWER ON

FUEL TRANSFER

R TANK

0 23

DATA

24

OFF

L ENG

0

RADAR

5 STBY FLT DISPLAY ON

R GEN

BATT

RUDDER BIAS

L TANK

0.0

T O TA L H O U R S

1

60

1340 20 3 00

10 MENU ADV

1/2

TCAS

ATC 1

2

21

ADF

DC AMPS

DC POWER L GEN

10

REFS

DME-H

108.10

114.00

– – – – NM

W

PRESET FMS1

DC AMPS

700 600 400 200

30

24

R GEN

DC VOLTS FIRE WARN LDG GEAR BATT TEMP AOA

60

––

25 SEL

NAV2

2000

1 0 0 0 0 4

5 00

IDENT

OFF

124.650

126.900

FUEL QTY LBS

0.0 OIL °C

1500

50

21

VOLTAGE SEL BATT

L GEN

OFF

0.0 N2 % OIL PSI

800

0.0

NAV/BRG

FORMAT

TEST

ANNU ANTI SKID OVER SPEED W/S TEMP

P

P

USH

15 S

– – – – NM

20

ENG

ITT °C 1000 900

N1 % –––.– 100 T T O 90 O

DATA

ELECT

29.92

137 12

––––– 80

BRT

Collins

MENU ADV

4

MASTER WARNING RESET

PTCH

USH

STD

COM2

2

S

HDG 170

VOR1 CRS 166

ROLL

BARO

ENG

REFS

1

2 00

20

MASTER CAUTION RESET

N584CJ3

1330.0

6

ADF1

MUTE

VS

3

MKR

VOX

ADF

HDG

P

DME2

ST

ALT

VNAV

Collins

NEXT PAGE

CRS

FLC

FD

TA/RA REL

N763CJ

M

P

HF

DME1

1200

B/C

APPR

S

COM3

NAV2

33 30

1/2 BANK

NAV

1/2

TCAS

80

0

COM2

N

HDG

DME-H

108.10

Collins

1

INPH V BOTH ID

3

2

1360 40 3 00

10

NAV1

24 W

124.650

NAV1

ATC 1

4

5 00

––

COM1

USH

STD

20

.4

.2

114.00

BARO

––––– .6

ET

126.900

IDENT

OFF

COM1

HDG 360 S 21

P

ROLL

1000

12 15

E 6

N A V

10

10

138

25 SEL

A D F 1

1500

13 80 60

E

FT

10

10

1 40 9

N

LT

CRS 360

STD

R

60

1.0 ANGLE OF .8 ATTACK

17:20 GMT

BRT

Collins

0.0

L

Collins

EMERGENCY USE ONLY

1/2 BANK

NAV

E

HDG

33

MASTER CAUTION RESET

30

MASTER WARNING RESET

N584CJ3

PULL RAIN

CONTROL LOCK PULL

RECOG/ TAXI

ANTISKID ON

LH

AIR SOURCE SELECT BOTH

AIR CONDITIONING

RH

FL

GEAR UNLOCKED

N O R M

SET ALT EXER

M A X

L

AUTO O F F

PUSH DOWN

OFF

SET ALT

RATE

EMER BRAKE PULL 15 10

AUX GEAR CONTROL

5

1. PULL & TURN TEE HANDLE 45° CW TO UNLOCK 2. PULL ROUND KNOB TO BLOW GEAR DOWN

20

4 5 6

25 7 8

0 DIFF 9 PRESS

45

0

30

L

TRIM NOSE DOWN

T O

NOSE UP

LOW TEMPERATURE SELECT M A N U A L

NORMAL

DADC REV

NORM

NORM

PANEL LIGHTING PFD

LOW

DIM

LOW

A U T O

MCT CRU

COLD MANUAL

HOT

F L A P S

T H R O T T L E

IDLE

IDLE

OFF

OFF

0°

UP

TO

CRU

SPEED BRAKE RETRACT

AHRS REV

R SLEW

O F F

R

MCT

T H R O T T L E

L SLEW

AUTO

MAINTENANCE DOWNLOAD

DEFOG HI A U T O

HOT

COLD

MANUAL

EMER OFF

FAN FWD HI

AFT FLOOD H I

35 40

CABIN ALT X 1000FT

TO

A

PARKING BRAKE HANDLE

3 2 1

COMPRESSOR ON

PRESS SYSTEM SELECT MANUAL UP M A N U A L AUTO DOWN GND IDLE HIGH

R AHRS SLAVE

R

FRESH AIR

FAN RECOG/ TAXI

DEPRESSURIZE CABIN BEFORE LANDING

PARK BRAKE – PULL

COCKPIT AIR DIST

CABIN DUMP

TAKEOFF AND APPROACH 200 KIAS

15°

LAND 161 KIAS

35°

GROUND FLAPS

55°

GROUND USE ONLY ENGINE SYNC OFF

ON

ON MUST BE OFF FOR TAKEOFF AND LANDING

EXTEND

ANTI-SKID SWITCH DISPLAY PFD/MFD

AHRS REV

DADC REV

NORM

NORM

NORM REV TO PFD

REV TO MFD LDG GEAR

NOSE

UP

DIM

L ENG

R ENG

LIGHTS TAIL FLOOD

PASS SAFETY

R

L O F F

O F F SEAT BELT

OFF

NORM SILENCE

LANDING

RECOG/ TAXI

RECOG/ TAXI

FADEC CH B

FADEC CH B

FADEC RESET

FADEC RESET

ANTISKID ON

LH

RH

GEAR UNLOCKED PUSH DOWN

OFF

EMER BRAKE PULL

AUX GEAR CONTROL 1. PULL & TURN TEE HANDLE 45° CW TO UNLOCK 2. PULL ROUND KNOB TO BLOW GEAR DOWN

EMERGENCY BRAKE LEVER

DETAIL A

Figure 32-16. Brake System Controls

32-32

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Parking Brake The parking brake valve is downstream of the antiskid control valve and is engaged by pulling the parking brake knob just under the pilot instrument panel (Figure 32-16). The parking brake valve has a check valve in each of the right and left brake lines. Pulling the parking brake knob engages the check valves and traps existing or subsequent pressure applied to the brakes. The parking brake valve also has thermal relief valves to accommodate pressure rise due to fluid expansion when the parking brake is engaged shortly after heavy braking. The thermal relief valves open and relieve pressure in excess of 1,200 + 50 psig, and then reseats to maintain a trapped pressure of at least 600 psig.

tem, pull and hold the lever in a position that provides the desired deceleration. The nitrogen storage bottle has a volume of 95 + 5 cubic inches.

NOTES

Emergency Brake An emergency (auxiliary) brake control handle is immediately below the instrument panel on the left side (Figure 32-16). Pneumatic nitrogen is from the emergency (auxiliary) air storage bottle, which also provides pneumatic power for emergency (auxiliary) landing gear extension. The main landing gear wheel brake assembly uses the pneumatic pressure for emergency braking. The pneumatic system is totally independent of the hydraulic brake system. Dedicated pneumatic lines are routed from the emergency brake valve to the brake shuttle valves. The emergency brake system operates by a lever below the pilot instrument panel. The system uses compressed nitrogen to supply equal pressure to both brake assemblies according to pilot demand, up to a maximum of 838 + 38 psi. The pneumatic system operates without antiskid and without differential braking. Pulling back on the emergency brake lever increases brake pressure in proportion to lever position. Brake pressure is reduced and residual nitrogen is vented overboard when the s p r i n g - l o a d e d l ev e r m o v e s f o r w a r d . Repeatedly increasing and decreasing brake pressure depletes the nitrogen supply. For the most efficient use of the emergency brake sys-

FOR TRAINING PURPOSES ONLY

32-33

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Figure 32-17. Metering Valve

32-34

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

COMPONENTS

Brake Fluid Reservoir

Brake Cable Control System

A pressurized fluid reservoir is on the forward side of the forward pressure bulkhead. The brake system hydraulic fluid reservoir is on the right side of the forward pressure bulkhead. The reservoir consists of:

The brake mechanical control components originate at the rudder pedal rudder/brake bellcranks and terminate at the cable interface with the brake metering valve (Figure 32-17). Forward brake control cables originate at the rudder/brake bellcranks and terminate at the interface with the link assemblies, springs, and cable assemblies. Aft brake control cables consist of cables originating at the cable assemblies and terminating at the cable clevises on the brake metering valve control arms.

• Reservoir tank • Two sight gauges • Filler plug • Connections for a vent line • Bleed/return line • Brake pump supply line

The main landing gear brakes operate by depressing the rudder pedals. Rudder/brake pedal input activates the brake metering valve via a cable interconnect assembly. Pressing the top of the individual rudder/brake pedal actuates the associated left/right input arm on the brake metering valve through cable movement. Braking is controlled independently from either cockpit crew position. The cable system that transmits pedal deflection to the brake metering valve allows pedal inputs by the pilot but does not cause the copilot pedals to move, and vice versa. If the pilot and copilot pedals are depressed simultaneously, the brake system accepts the highest input. Since the brake system is a cable-controlled power brake system, the braking feel force at the pedals is created by springs in the mixer, the springs within the brake metering valve, and a proportional hydraulic feedback force generated by the brake metering valve.

Brake Metering Valve The brake metering valve (Figure 32-17) is in the left wing root fairing and is operated by control cables. The brake metering valve initiates all braking action except emergency braking. During normal operation, the accumulator provides pressurized fluid to the brake metering valve, regulating pressure (0–1,000 +50/–20 psi) to the brake assemblies in proportion to the brake pedal deflection.

FOR TRAINING PURPOSES ONLY

32-35

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A BATT O'TEMP

LH ENG FIRE

DC POWER (BATT) SWITCH

CABIN ALT

OIL PRESS WARN L

> 160° ENG CTRL SYS FAULT L

BOTTLE 1

R

RUDDER BIAS

ARMED PUSH

F/W SHUTOFF L

CABIN DOOR DOOR SEAL

R GEN OFF

R

L

FLAPS >35°

BAGGAGE SMOKE NO TAKEOFF

HYD FLOW HYD PRESS LOW ON SPD BRK L R EXTEND

R

BAGGAGE DOOR FWD AFT

EMER EXIT STBY P/S HTR OFF

EMER PRESS ON FRESH AIR

P/S HTR OFF R

W/S AIR O'HEAT AOA HTR FAIL

FUEL GAUGE L R

AFT J-BOX LMT CB

GROUND IDLE

L

PWR BRK FUEL LOW FUEL FUEL FUEL LOW FUEL FLTR LOW PRESS PRESS BOOST ON TRANSFER LEVEL BYPASS ANTISKID L R L R L R L R INOP

BLD AIR O'HEAT L R

ENG ANTI-ICE L R

WING O'HEAT

WING ANTI-ICE

L

L

R

ENG T2 HTR FAIL L R

RH ENG FIRE

R

BOTTLE 2

VIDEO FAIL AUDIO FAIL

TAIL DEICE TAIL DEICE FAIL PRESS AIR DUCT L R O'HEAT

ARMED PUSH

Collins

B/C

FLC

APPR

FD

VNAV

Collins

N1

ALT

%RPM

0.0 VS

80

AP PTCH

80

0

DME1

DME2

MKR

ADF1

ADF2

ST

VOX

MUTE

SPKR

HDPH

P

2

MENU ADV

4

DATA P

USH

USABLE FUEL 4710 LBS

1000

25

500

23 0

0

1200

TA/RA REL

N763CJ NEXT PAGE

6 3

N

GS 0

RAT 22 °C

TAS 0

SAT 24 °C

°M

O F F

START DISG NORM

ANTI-ICE / DE-ICE WINDSHIELD WING/ENGINE BLEED HI ALCOHOL ON WING XFLOW L WING/ENG R WING/ENG

LOW

OFF

OFF

ENG ON

LEFT

NOT ARMED

EXTERIOR LIGHTS

L SLEW

BEACON

ANTI-COLL

NAV

AUTO

R SLEW

OFF

OFF

OFF

DIM

OFF

MAX

FLOOD LTS

OFF

AHRS REV

DADC REV

NORM

NORM

OBS

MSG

GPWS FLAP OVRD

GPWS G/S

ACTIVE

CANCELED

NORM REV TO PFD

ADF1

ADF2

MUTE

SPKR

CAB

HDPH

PA

RAT 22 °C

TERR LX/RDR TERRAIN

VNAV

ACTIVATED WHEN LIT

10 OXY

ISA+11 °C

5 BRT DIM

DEFAULT NAV

CRS

GPS

FPL

ELT

TFC TA ONLY

SAT 24 °C

15

PROC

PUSH CRSR

0

PSI x100

20

ON

ARM TEST/RESET SELECT ON WAIT 1 SECOND SELECT ARM

USH

DIR T EC

TERR NORM TERR INHIB

GPWS TEST

LDG GEAR

L ENG

R ENG

FADEC CH B

FADEC CH B

FADEC RESET

FADEC RESET

LIGHTS

EL

STBY

TAIL FLOOD

PASS SAFETY

OFF DIM

NOSE

UP

DIM

DIM

R O F F

OFF

SEAT BELT

DIM

NORM SILENCE

LANDING L

O F F OFF DIM

CONTROL LOCK PULL

RECOG/ TAXI

ANTISKID ON

LH

AIR SOURCE SELECT BOTH

AIR CONDITIONING

RH

FL

GEAR UNLOCKED

N O R M

SET ALT EXER

M A X

L

AUTO O F F

PUSH DOWN

OFF

SET ALT

RATE

EMER BRAKE PULL 15 10

5

1. PULL & TURN TEE HANDLE 45° CW TO UNLOCK 2. PULL ROUND KNOB TO BLOW GEAR DOWN

3 2 1

20

4 5 6

25 7 8

0 DIFF 9 PRESS

45

0

30

L

B

TRIM NOSE DOWN

T O

LOW TEMPERATURE SELECT M A N U A L

NORMAL

CRU T H R O T T L E

MCT CRU

TAKEOFF AND APPROACH 200 KIAS

PANEL LIGHTING PFD

LOW

DIM

LOW

A U T O COLD MANUAL

HOT

15°

OFF

OFF

35°

GROUND FLAPS

ANTISKID INOP AND PWR BRK LOW PRESS

NORM

F L A P S

T H R O T T L E

IDLE

EXTEND

DADC REV

NORM

0°

UP

TO

IDLE

SPEED BRAKE RETRACT

AHRS REV

R SLEW

O F F

R

LAND 161 KIAS NOSE UP

L SLEW

AUTO

MAINTENANCE DOWNLOAD

DEFOG HI A U T O

HOT

COLD

MANUAL

EMER OFF

FAN FWD HI

AFT FLOOD H I

35 40

CABIN ALT X 1000FT

TO MCT

ROTARY TEST SWITCH

COMPRESSOR ON

PRESS SYSTEM SELECT MANUAL UP M A N U A L AUTO DOWN GND IDLE HIGH

R AHRS SLAVE

R

FRESH AIR

FAN RECOG/ TAXI

DEPRESSURIZE CABIN BEFORE LANDING PULL RAIN

COCKPIT AIR DIST

CABIN DUMP

AUX GEAR CONTROL

55°

GROUND USE ONLY ENGINE SYNC OFF

ON

ON MUST BE OFF FOR TAKEOFF AND LANDING

DETAIL A ANTI-SKID SWITCH DISPLAY

PFD/MFD

AHRS REV

DADC REV

NORM

NORM

NORM REV TO PFD

REV TO MFD LDG GEAR

NOSE

UP

DIM

L ENG

R ENG

LIGHTS PASS SAFETY

TAIL FLOOD

SEAT BELT

R O F F

OFF

NORM SILENCE

LANDING L

O F F RECOG/ TAXI

RECOG/ TAXI

FADEC CH B

FADEC CH B

FADEC RESET

FADEC RESET

ANTISKID ON

LH

RH

GEAR UNLOCKED PUSH DOWN

OFF

EMER BRAKE PULL

AUX GEAR CONTROL 1. PULL & TURN TEE HANDLE 45° CW TO UNLOCK 2. PULL ROUND KNOB TO BLOW GEAR DOWN

EMERGENCY BRAKE LEVER

DETAIL B

Figure 32-18. Antiskid Controls and Indications

32-36

MKR

VOX

REV TO MFD

O N OFF

MAX INSTR

NIGHT DIM ON

WING INSP

PARK BRAKE – PULL

DME2

ST

P

MANUAL

PANEL LIGHT CONTROL L AHRS SLAVE MANUAL

DME1

Collins

DISPLAY PFD/MFD

ARMED

TAS 0

GS

ENT

0.0kt

0 NT

P

NRST

BRT DIM

EMERGENCY LIGHTING

RIGHT

O F F ENG ON

NAV2

BRIEFING SELECTION

NORM

TAIL AUTO

O F F

O F F OFF

NAV1

FORMAT

ADF

Collins

TFC TA ONLY

ADF

HF

INPH V BOTH ID

GS

ENR

BRIEFING

NORM

WINDSHIELD BLEED AIR PITOT & STATIC

PLAY

RE SS TO TEST

NORM

ON

COM3

CLR

NAV

ADF

R

COM2

15 S

MENU

n m

20 nm –1

INTERCOM CALL

PBS - 250

ON

COM1

29.92

138 12

ADF

DIS

AUTO OFF

OFF

L

PRESET FMS2

RANGE SH AUTO TILT

D

005°M

n m

TERR LX/RDR

TILT

°M

TRK

KCVG CVG

2.5

PULSE LT TA/RA

IGNITION R

GCS RNG

DTK DTK

PWR

Collins

BRT DIM

TEST

DISENGAGE

HDG 137

VOR2 CRS 245

RADAR

WPT

S

DIS

ISA+12 °C

4

GPS 500

GARMIN

137 15

12

PU

33

RESET ENGINE START L

R

ELECT

NAV/BRG

WPT

TTG – – : – – – – – – NM

RANGE

2

2 00

20

1330.0

0

SH AUTO TILT

TFC TA ONLY

30

ON

0

USH

1

EMERGENCY USE ONLY

EMER

FUEL BOOST L

ADF

DATA

E

TILT

HDG 137

VOR1 CRS 166

GCS

TERRAIN

O F F

OFF

RESET

AVIONICS POWER ON

STBY FLT DISPLAY ON

R GEN

BATT

OFF

R ENG

25

°C FUEL

0 PPH

RADAR

TERR LX/RDR

ADF

FUEL TRANSFER

R TANK

0 23

MENU ADV

1/2

TCAS

W

DC POWER L GEN OFF

L ENG

0

ATC 1

5

RUDDER BIAS

L TANK

0.0

10

108.10

114.00

1340 20 3 00

24

W/S TEMP

700

0.0

1

– – – – NM

FORMAT

ADF

DC AMPS

2000 1500

600 400 200

30

W

PRESET FMS1

DC AMPS

OIL °C

OIL PSI

800

2

60

–– REFS

DME-H

T O TA L H O U R S

4

5 00

10

IDENT

25 SEL

NAV2

1 0 0 0 0

20

60

BRT

OFF

124.650

126.900

FUEL QTY LBS

0.0

21

R GEN

DC VOLTS FIRE WARN LDG GEAR BATT TEMP AOA

0.0 N2 %

50

24

OFF

80

STD

ITT °C 1000 900

70

ELECT

21

VOLTAGE SEL BATT

L GEN TEST

ANNU ANTI SKID OVER SPEED

–––––

ENG

NAV/BRG

S

– – – – NM

PTCH

SH PU

Collins

15

12

ROLL

BARO

M

N1 % –––.– 100 T T O 90 O

MASTER WARNING RESET

N584CJ3

1330.0

Collins

29.92

137

VOR1 CRS 166

ADF

HDG

COM2

2 00

20

MASTER CAUTION RESET

Collins

ENG

REFS

1

S

HDG 170

INPH V BOTH ID

VS

6

NAV2

PA

ALT

VNAV

TA/RA REL

N763CJ NEXT PAGE

CRS

FLC

FD

3

NAV1

CAB

1200

B/C

APPR

S

HF

33 30

1/2 BANK

NAV

1/2

TCAS

80

1360 40 3 00

10 COM3

N

2 1

COM2

3

HDG

DME-H

108.10

4

5 00

10

––

COM1

USH

STD

20

60

24 W

124.650

NAV1

ATC 1

P

.6 .4

.2

114.00

BARO

–––––

ET

126.900

IDENT

OFF

COM1

HDG 360 S 21

PU

FT

1000

12 15

E 6

N A V

10

10

138

25 SEL

A D F 1

1500

13 80 60

E

LT

10

10

1 40 9

N

GMT

ROLL

CRS 360

STD

R

60

1.0 ANGLE OF .8 ATTACK

17:20

BRT

Collins

0.0

L

Collins

EMERGENCY USE ONLY

1/2 BANK

NAV

E

HDG

33

MASTER CAUTION RESET

30

MASTER WARNING RESET

N584CJ3

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Motor/Pump Assembly The brakes operate by an independent, closed center hydraulic system with a separate reservoir, pump/electric motor, and accumulator. The hydraulic pack assembly has a direct current (DC) electric motor driving a hydraulic pump that provides power to the hydraulic brake system. The hydraulic pack assembly includes: • Relief valve for backup system protection • Filter assembly to prevent contaminating brake hydraulic fluid in the event of a component failure • Pressure switch to indicate a pressure fault condition • Check valve to prevent reverse flow of the hydraulic fluid

valve is the same type used in servicing the landing gear strut. The accumulator pressure gauge monitors pressure in the accumulator.

Speed Transducer The antiskid speed transducer, also referred to as the wheel speed generator, consists of: • Rotor shaft • Coil assembly • Bobbin assembly • Bearings contained in a housing assembly A drive coupling is attached to the rotor shaft, which is driven directly by hubcap rotation.

The motor is controlled by pressure and gear selector switches. A pressure switch near the fluid end of the accumulator senses brake system pressure and commands the pump on and off accordingly. There is no cockpit switch for the brake pump. The pump is powered on anytime the gear handle is in the down position and the accumulator pressure is below 1,175 ± 75 psig. When the accumulator pressure reaches 1,500 ± 50 psig, power is removed from the pump. A separate low-pressure switch built into the pump monitors the system for low pressure. If the system pressure drops below (1,100 ± 50 psig) and the gear handle is down, the low-pressure switch causes the PWR BRK LOW PRESS annunciator (Figure 32-18) to illuminate. The landing gear extend switch disables the control relay, preventing the motor from operating unless the landing gear control handle is in the extend position.

NOTES

Accumulator The accumulator provides a 25-cubic-inch fluid reserve (under pressure) for the power brake function. An accumulator charging valve and pressure gauge are components of the accumulator system. The pressure gauge and charging valve are adjacent to and on opposite sides of the brake reservoir. The charging

FOR TRAINING PURPOSES ONLY

32-37

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Figure 32-19. Digital Control Unit

32-38

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Digital Control Unit

Antiskid Control Valve

The digital antiskid control unit consists of a circuit board in a cast-aluminum box (Figure 32-19). All electrical connections are made on the antiskid control unit.

Hydraulic system pressure from the brake metering valve is supplied to the antiskid control valve according to pilot input. The antiskid valve is a three-way pressure-control servo valve. Its function is to relieve brake pressure if an impending skid is detected. Each antiskid control valve contains two electrohydraulic servo valves, for the left and right brake independently.

The digital antiskid control unit receives the output signals of the left and right antiskid speed transducers and converts these signals to DC voltage directly proportional to wheel speed. The control unit monitors each wheel independently and compares the difference between the changing reference velocity and the instantaneous wheel velocity signals. When either tire skids in excess of the optimum required for effective braking, a brake pressure reduction is generated for that wheel. The wheel velocity signal, in conjunction with the reference deceleration control, continuously updates the reference velocity circuit.

NOTES

A full-time system integrity monitor circuit is built into the system and operates on a high-low voltage principal. In the event of a short (open fault) occurring in the valve circuit (wheel speed generator circuit or an antiskid control circuit power fault is sensed), the ANTISKID INOP annunciator illuminates. A test circuit, provided as part of the digital antiskid control unit, is activated anytime the switch is positioned to ON for ground testing of the system. The test circuit is also activated automatically when the LDG GEAR lever is in the DOWN position during flight and the ANTI-SKID switch is positioned to ON. Finally, a self-test can be initiated by selecting ANTISKID on the rotary TEST knob. The test circuit monitors the electrical function of the antiskid system and illuminates the ANTISKID INOP annunciator, alerting the crew of an antiskid system fault. If an antiskid system fault is indicated, the antiskid system can be selected OFF and normal braking occurs without antiskid protection.

FOR TRAINING PURPOSES ONLY

32-39

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Figure 32-20. Antiskid System Self-Monitoring

32-40

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CONTROLS AND INDICATIONS ANTI-SKID Switch The cockpit controls for the antiskid system consist of a single ANTI-SKID toggle switch, which has two positions: ON and OFF. The toggle switch is just right of the gear handle (see Figure 32-18). When the switch is selected ON, it delivers 28 VDC to the digital antiskid control unit. The antiskid system provides: • Maximum braking efficiency under all runway conditions • Touchdown protection, preventing braking until adequate wheel spin-up occurs, and locked wheel crossover protection that prevents adverse differential braking

Fault Display Unit The fault display unit consists of five rotary flags to aid in the troubleshooting of ANTISKID INOP annunciations. Flagged conditions include: • Left antiskid speed transducer fault • Right antiskid speed transducer fault • Servo valve fault • Digital antiskid control unit fault and squat switch disagree

brake independently. Therefore, a single wheel skid results in the reduction of brake pressure at the skidding wheel only. Antiskid protection is available unless the touchdown protection mode is active.

Antiskid System Self-Monitoring The antiskid system performs continuous integrity checks on the wheel speed transducer circuits, the antiskid servo valve circuit, and the regulated power to the control box. If a fault is detected during continuous monitoring, the ANTI SKID INOP annunciator illuminates and a signal is sent to the antiskid fault display unit. The fault display unit is in the right nose compartment on the forward side of the forward pressure bulkhead (Figure 3220). The fault display unit consists of five rotary flags to aid in the troubleshooting of an ANTI SKID INOP annunciator illumination. There is one rotary flag for each of the following conditions: • Left transducer fault • Right transducer fault • Valve (servo) fault • Control unit fault • Squat disagree

Antiskid Protection

Both main gear squat switch signals are monitored and compared. If the signals disagree for more than 13 seconds, the squat switch disagree flag is tripped. However, this fault does not cause the ANTI SKID INOP annunciator to illuminate.

Antiskid protection is provided to allow maximum braking efficiency, which in turn minimizes landing distances. If the pilot applies enough brake pedal force to cause slippage between the tires and the runway, the wheel speed transducer data received by the control box indicates a sudden deceleration for the slipping wheel. The control box determines the severity of the impending skid and sends the appropriate current signal to the antiskid servo valve in order to reduce brake pressure accordingly. Dual servo valves reduce pressure for either

A low supply voltage causes the ANTI SKID INOP annunciator to illuminate but does not result in a tripped flag on the fault display unit. The continuous monitor function evaluates the voltage supplied to the control box. Anytime the input voltage is less than 7.0 + 1 volt, the ANTI SKID INOP annunciator illuminates, but the control box flag is not tripped. This feature alerts the crew that the antiskid system is either switched off or is unavailable due to insufficient power. In addition to con-

OPERATION

FOR TRAINING PURPOSES ONLY

32-41

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

tinuous monitoring, the antiskid control box performs a dynamic self-test, which is initiated by any of the following events: • Initial power-up of the antiskid system • Transition of the gear handle to the down position • Selection of ANTI SKID on the rotary TEST knob During a dynamic self-test, a signal is sent to illuminate the ANTI SKID INOP annunciator. Upon successful completion of the test, the ANTI SKID INOP annunciator extinguishes. If a fault is detected during the self-test, the annunciator remains illuminated. A dynamic selftest performed in the air requires approximately 3 seconds, and a dynamic self-test performed on the ground requires approximately 6 seconds. The dynamic self-test is inhibited if wheel speed is greater than 15 ± 5 knots.

Locked Wheel Crossover Protection Locked wheel crossover protection prevents inadvertent turning of the aircraft due to differential braking. The velocities of the two wheels are compared in order to determine if one wheel is locked. If the velocity of one wheel falls to less than 30% of the other wheel, the control box sends a full dump command to the antiskid servo valve controlling the slower wheel. The full dump remains in effect until the velocity of the slower wheel increases above the 30% threshold. The locked wheel crossover feature is inactive at wheel speeds below 25 knots to allow for taxiing.

Touchdown Protection Touchdown protection prevents the application of brake pressure prior to wheel spin-up. During landing, the wheels must be allowed to spin up to provide the antiskid system a reference velocity to which individual wheel speeds can be compared. Touchdown protection is active only when an AIR signal is sensed by both main gear squat switches. In touchdown protection mode, the control box commands the antiskid servo valves to dump all brake pressure. The full dump command remains active for 3 seconds after weight on wheels (WOW) or until wheel spinup occurs. Under normal conditions, the wheels spin up almost immediately after touchdown. Therefore, the system incorporates a spin-up override feature. When the velocity of a wheel exceeds 59 ± 2 knots, touchdown protection is overridden and brake pressure application is allowed to that wheel. Each wheel is independent in regard to spin-up override. The touchdown protection mode is overridden for each wheel (independently) only when the speed of a given wheel is in excess of 59 ± 2 knots. The wheel spin-up override remains active until the wheel velocity falls below 15 ± 2 knots.

32-42

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 33 LIGHTS CONTENTS Page INTRODUCTION ................................................................................................................. 33-1 GENERAL ............................................................................................................................ 33-3 FLIGHT COMPARTMENT LIGHTING.............................................................................. 33-5 Primary Lighting............................................................................................................ 33-5 Secondary Lighting........................................................................................................ 33-7 PASSENGER COMPARTMENT LIGHTING ..................................................................... 33-9 Description..................................................................................................................... 33-9 Components ................................................................................................................... 33-9 Controls and Indications.............................................................................................. 33-11 CARGO AND SERVICE COMPARTMENT LIGHTING................................................. 33-13 Description................................................................................................................... 33-13 Components ................................................................................................................. 33-13 EXTERIOR LIGHTING ..................................................................................................... 33-15 Description and Operation........................................................................................... 33-15 EMERGENCY LIGHTING................................................................................................ 33-17 Description................................................................................................................... 33-17 Components ................................................................................................................. 33-17 Controls and Indications.............................................................................................. 33-17

FOR TRAINING PURPOSES ONLY

33-i

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

ILLUSTRATIONS Figure

Title

Page

33-1

Cockpit Lighting Systems...................................................................................... 33-2

33-2

Instrument Light Inverters...................................................................................... 33-4

33-3

Instrument Floodlight Under the Fire Tray............................................................ 33-6

33-4

Cabin Lights........................................................................................................... 33-8

33-5

Entry Light Switches ........................................................................................... 33-10

33-6

Nose Baggage Compartment Light Switch ......................................................... 33-12

33-7

Aft Baggage Compartment Light Switches......................................................... 33-12

33-8

Left Switch Panel................................................................................................. 33-14

33-9

Lights Subpanel ................................................................................................... 33-14

33-10

Emergency Lighting Switches ............................................................................. 33-16

FOR TRAINING PURPOSES ONLY

33-iii

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 33 LIGHTS

INTRODUCTION This chapter describes the lights on the CitationJet 3 525B aircraft. Information is given on the flight, passenger, and cargo and service compartments, as well as exterior and emergency lighting. Further information can be found in Chapter 33—“Lights” in the Aircraft Maintenance Manual AMM).

FOR TRAINING PURPOSES ONLY

33-1

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

WINDSHIELD BLEED AIR RIGHT LEFT

EMERGENCY LIGHTING NOT ARMED S

T

AHRS REV

DADC REV

NORM

NORM

NORM REV TO PFD

REV TO MFD

S TO TE

O N

R

ES

DIM

DISPLAY PFD/MFD

ARMED

P

OFF

OFF

MAX

OFF

MAX

DIM

PANEL LIGHT CONTROL FLOOD LTS

EL

INSTR

NIGHT DIM ON

STBY

OFF DIM PULL RAIN

DIM

SEAT BELT

DIM CONTROL LOCK PULL

R

L O F F

OFF

RECOG/ TAXI

EMER BRAKE PULL

DETAIL A

Figure 33-1. Cockpit Lighting Systems

33-2

R ENG

FADEC CH B

FADEC CH B

FADEC RESET

FADEC RESET

LANDING

TAIL FLOOD

PASS SAFETY O F F

DIM

L ENG LIGHTS

FOR TRAINING PURPOSES ONLY

RECOG/ TAXI

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

GENERAL

NOTES

Controls for most of the lights in the model 525B aircraft are in the cockpit Figure 33-1). The flight compartment section provides system description, operation, and maintenance practices for flight compartment primary instrument lights, cockpit floodlights, and map lights. The passenger compartment section provides information for the indirect fluorescent cabin lights, passenger reading and entrance lights, and passenger advisory signs. The cargo and service compartments section provides a description of illumination for the forward and aft baggage compartments. The exterior section describes the systems used to illuminate the outside of the aircraft. The standard exterior lighting system package consists of: • Position and anticollision lights • Landing/taxi/recognition lights • Wing inspection lights • Ground recognition beacon The emergency lighting section describes systems used to provide illumination during a primary electrical power failure, abnormal conditions, and aircraft evacuation.

FOR TRAINING PURPOSES ONLY

33-3

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Figure 33-2. Instrument Light Inverters

33-4

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

FLIGHT COMPARTMENT LIGHTING PRIMARY LIGHTING Description Instruments with internal lighting are included in the primary flight compartment lighting system. Instruments that have internal lights use 5-VDC power and others require 28 VDC. Switch and control panels are illuminated by electroluminescent panels and include the instrument panel, side consoles, and control pedestal. Instruments and switches that do not contain internal lights or an electroluminescent panel can be seen with the use of the secondary flight compartment lighting system.

Components

Controls and Indications NIGHT DIM Switch The NIGHT DIM switch is on the lower pilot instrument panel see Figure 33-1). The switch has two positions: ON and OFF. The NIGHT DIM switch supplies 28 VDC to each of the flight compartment lighting inverters Figure 33-2). Each of the inverters is controlled by a voltage-control rheostat on the instrument panel. The rheostats increase or decrease the intensity of the light from the components that each of the inverters controls.

INSTR/EL/STBY Rheostats The INSTR, EL, and STBY rheostats are on the lower pilot instrument panel see Figure 331). The INSTR rheostat controls the system instruments. The EL rheostat controls the voltage to all the electroluminescent panels in the flight compartment. The STBY rheostat controls the standby instruments.

Internally lit instruments are divided into two groups: • Various system indication instruments on the instrument panel, side consoles, and control pedestal

NOTES

• Standby instruments grouped at the top center of the instrument panel The instruments with internal lights receive 5VDC electrical power from one of two inverters in the forward baggage compartment. One inverter provides power to the system instruments and the other provides power to the standby instruments. Electroluminescent panels show the functions and positions of switches and controls in the cockpit. The electroluminescent panels receive 40–60 VAC, 400 Hertz electrical power from an inverter inside the radome.

FOR TRAINING PURPOSES ONLY

33-5

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Figure 33-3. Instrument Floodlight Under the Fire Tray

33-6

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

SECONDARY LIGHTING Description The flight compartment secondary lighting system includes floodlights, overhead map lights, and windshield ice-detection lights. All of these lighting systems use 28-VDC electrical power.

Components

Ice-Detection Lights There are two windshield ice detection light assemblies on the glareshield. They point toward the windshield and indicate to the pilot or copilot if ice is accumulating on the windshield. The windshield ice detection lights are on a ny t i m e t h e N I G H T D I M s w i t c h o n t h e PANEL LIGHT CONTROL subpanel is in the ON position.

Floodlights

NOTES

There are two floodlights in the flight compartment overhead panel and one instrument floodlight under the fire tray Figure 33-3). The floodlights use 28 VDC and are controlled by the FLOOD LTS rheostat on the instrument panel. The floodlights illuminate full bright during engine start procedures. They receive electrical power from the emergency lighting battery pack.

Map Lights There are two map lights in the flight compartment overhead panel. The map lights illuminate the approach chart that can be clipped onto the pilot and copilot control wheel. Each map light is protected by a 5-amp PANEL LTS 2 circuit breaker on the left CB panel. Light intensity is controlled by each map light ON–OFF rheostat on either the left or right side consoles.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Figure 33-4. Cabin Lights

33-8

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

PASSENGER COMPARTMENT LIGHTING

Forward/Aft Ordinance Sign Assemblies

DESCRIPTION The passenger compartment light system includes: • Cabin indirect lighting • Passenger reading and entry lighting • Ordinance signs

The forward and aft ordinance sign assemblies are in the cabin overhead console and each displays two messages: NO SMOKING and FASTEN SEAT BELTS. The NO SMOKING message is illuminated continuously. The FASTEN SEAT BELTS message is illuminated when the PASS SAFETY switch is in the SEAT BELT (down) position. A second aft ordinance sign instructs an occupant in the lavatory to RETURN TO SEAT. It is controlled by the PASS SAFETY switch.

• Exit signs The passenger safety chime operates with the ordinance signs to alert passengers when they are not allowed to smoke or when they must fasten their seat belts or return to their seats.

• Dropped-aisle lighting

COMPONENTS Passenger Reading Lights Passenger reading lights are in the cabin overhead consoles. Each of the reading lights can be directionally adjusted and is controlled with a switch in the bezel adjacent to the light assembly.

EXIT Signs The EXIT signs receive electrical power from the emergency lights circuit when the PASS S A F E T Y t o g g l e s w i t c h i s i n t h e PA S S SAFETY (up) position or when the door courtesy switch is in the ON position. The EXIT signs also illuminate when the emergency lights are actuated.

Entry Light

Indirect Lighting Indirect lighting strips are above the headliner assembly. The indirect cabin lighting system uses light-emitting diode (LED) strips. They are controlled by switches near the cabin entry door on the refreshment center. The refreshment center switches provide ON–OFF control of the lighting system and can also control the intensity of the lights.

Dropped-Aisle Lighting The dropped-aisle lighting strips are on either side of the dropped aisle between the cabin seats. The system can be activated by either the entry light switch, the cabin light switch on the refreshment center, or the emergency lighting system.

An entry light is adjacent to the EXIT sign above the cabin door. The light is illuminated anytime the EXIT signs are illuminated.

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Figure 33-5. Entry Light Switches

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CONTROLS AND INDICATIONS

NOTES

Entry Light Switch Pressing the entry light switch (Figure 33-5) at the main entry doorpost illuminates the following: • Main door entry light and exit sign light • Emergency exit light and exit sign light • Two right wing walk lights • Light in the vanity • Aisle lights The entry light switch on the refreshment center switch panel controls the same lights. Power to the system is from the hot battery bus through the EMER LTS circuit breaker on the aft J-box.

Refreshment Center Switch Panel The switch panel on the refreshment center includes three pushbutton-type switches that control the indirect LED lighting: CABIN LIGHT, DIM, and BRIGHT. When the aircraft is powered on, all three switches default to off and the pushbuttons are amber in color. The CABIN LIGHT pushbutton fully illuminates and extinguishes the LED indirect lighting. When the pushbutton turns green, the LED lights are illuminated. The buttons are pressed once to illuminate and again to extinguish. If there is a fault in any circuit, the color of the words for the affected switch changes to red. The DIM and BRIGHT pushbuttons change the intensity of the indirect lights. The pushbutton is held until the desired level of intensity is reached (infinite dimming, not stepped).

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Figure 33-6. Nose Baggage Compartment Light Switch

Figure 33-7. Aft Baggage Compartment Light Switches

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CARGO AND SERVICE COMPARTMENT LIGHTING DESCRIPTION Cargo and service compartment lighting consists of nose and aft baggage compartment lights.

COMPONENTS Nose Baggage Compartment Light

Aft Baggage Compartment Light The aft baggage compartment lighting system consists of a single 28-VDC toggle-type switch (Figure 33-7), microswitch, and aft baggage compartment light assembly. The single toggle-type switch controls 28 VDC to the aft baggage compartment light assembly. When the aft baggage compartment door is closed, another microswitch interrupts power to the aft baggage compartment light. This cutout switch prevents power drain if the toggle switch is inadvertently left in the ON position.

The nose baggage compartment lighting system consists of one 28-VDC light, an illuminated manual switch (Figure 33-6), and two actuating switches. Two normally open actuating switches are at each aft door hinge assembly. These three switches control the nose baggage compartment light. Each is in the power input side of the circuit and controls the 28-VDC power supplied to the light. The switches extinguish the light when both doors are closed and the main toggle switch is in the ON position.

NOTES

The illuminated manual switch connects 28 VDC to the light. The manual switch is positioned to OFF during daylight hours or when the light is not desired. In this position, power is supplied to an internal lightbulb illuminating the lens of the manual switch. The illuminated switch provides easy location in the event the manual switch is in the OFF position. The two actuating switches are in the left and right nose baggage compartment door hinge assemblies. Either switch interrupts the circuit by connecting power through the switch when the nose baggage door is open. Power to the light is interrupted when either nose baggage compartment door is closed and the manual illuminated switch is in the ON position.

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

VOLTAGE SEL BATT

L GEN

R GEN

TEST OFF ANNU ANTI SKID OVER SPEED W/S TEMP

DC VOLTS FIRE WARN LDG GEAR BATT TEMP AOA

DC AMPS

DC AMPS

DC POWER L GEN OFF

RUDDER BIAS

STBY FLT DISPLAY ON

R GEN

BATT

O F F

OFF EMER

RESET

AVIONICS POWER ON

RESET

TEST

OFF

FUEL TRANSFER ENGINE START

FUEL BOOST OFF L L TANK

R TANK

L ENG

R ENG

ON

L

R

O F F NORM

PITOT & STATIC

DISENGAGE

IGNITION R

L

NORM

NORM

ANTI-ICE / DE-ICE WINDSHIELD WING/ENGINE BLEED HI ALCOHOL ON WING XFLOW L WING/ENG R WING/ENG

R

LOW

NORM

TAIL AUTO O F F

O F F

O F F OFF

ON

START DISG

OFF

OFF

ENG ON

L AHRS SLAVE

ENG ON

MANUAL

EXTERIOR LIGHTS

MANUAL

L SLEW

BEACON

ANTI-COLL

NAV

WING INSP

AUTO

R SLEW

OFF

OFF

OFF

OFF

PARK BRAKE – PULL

Figure 33-8. Left Switch Panel

Figure 33-9. Lights Subpanel

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EXTERIOR LIGHTING DESCRIPTION AND OPERATION Position and Anticollision Lights The red and green forward-position lights for the Citation 525B aircraft are behind clear tempered glass covers in assemblies on each wingtip of the aircraft. The aviation red forward-position light is on the left wingtip and the aviation green forward-position light is on the right wingtip of the aircraft. The rearposition light is an aviation white light assembly on the farthest aft surface of the vertical tail. All three position light assemblies are technical standard order (TSO) units that meet the requirements of FAA-TSO-C30 for forward- and rear-position lights. A single NAV control switch (SI055) (Figure 33-8) on the left switch panel supplies 28VDC power to the three lamps (FL001, FR001, and FV001). The position light circuits are protected by a NAV circuit breaker (HZ058) on the right J-box in the tail cone.

Landing/Taxi/Recognition Lights The Citation 525B aircraft is equipped with two lamps illuminated for landing, taxi, and recognition purposes. The landing lights have two 450-watt sealed beam lamps under the belly fairing. These lamps are protected behind tempered glass covers and are located such that the pilot compartment is shielded from glare and halation from the lamps. The landing lights are controlled by separate toggle switches (SI065 and SI066) on the LIGHTS subpanel (Figure 33-9). When the toggle switches are set to the L and/or R positions, separate relays are engaged, supplying power to each lamp (FL003 and FR003). The relays are KZ002 and KR001 for the right side and KZ001 and KL001 for the left side.

When L and/or R toggle switches (SI065 and SI066) are selected to the RECOG/TAXI (down) position, separate relays are engaged, supplying power to each lamp (FL003 and FR003). The taxi/recognition lights are dimmed by having their power supplied through individual resistors: RR001 for the right side and RL001 for the left side. Each RECOG/TAXI light assembly circuit (Figure 33-9) is protected by circuit breakers. The left and right belly fairing assemblies are p r o t e c t e d b y i n d iv i d u a l 2 0 - a m p L & R LAND/REC circuit breakers (HZ056 and HZ059) on the J-box in the tail cone.

Wing Inspection Lights A wing inspection light is on the left side of the fuselage, just above and slightly forward of the wing leading edge. The wing inspection light assembly is equipped with a lamp situated in the assembly so that the light is directed to the outboard leading edge of the left wing. The wing inspection light is used by the crew for detection of wing ice accumulation during nighttime flight in icing conditions. A WING INSP control switch (SI025) in the anti-ice switch section of the left instrument panel supplies 28VDC power to the lamp (FC002) (Figure 33-8). The wing inspection light is protected by a 5amp WING INSP circuit breaker (HZ060) on the right J-box in the tail cone.

Ground Recognition Beacon The Citation 525B aircraft have a ground recognition beacon for added safety during taxiing. The ground recognition beacon assembly is on the top of the vertical tail for optimum line-of-sight visibility. The beacon consists of a red light assembly (FV002) with 150-watt quartz halogen lamp, strobing at a rate of 45–55 flashes per minute. The BEACON switch (SI024) on the left switch panel (Figure 33-8) provides the 28-VDC power to the beacon. The ground recognition beacon is protected by a 5-amp BEACON circuit breaker (HZ062) on the right J-box in the tail cone.

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Figure 33-10. Emergency Lighting Switches

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EMERGENCY LIGHTING DESCRIPTION The emergency lighting system receives electrical power from a battery system and is automatically energized by G-switches near the battery packs. The system also uses hot battery bus power and can be turned on by the pilot with the use of the PASS SAFETY switch on the left instrument panel. A toggle switch on the EMERGENCY LIGHTING switch panel is used to reset the emergency lighting system after the 2-G switches have been operated (Figure 33-10). An escape hatch light is adjacent to the EXIT sign above the emergency exit door. It is illuminated anytime the EXIT signs are illuminated.

COMPONENTS Right Wing Walk Lights Emergency lighting includes two walk lights over the right wing. The lights are illuminated anytime the emergency lighting is illuminated. The exterior emergency over-the-wing exit lights are on the right side of the fuselage immediately under the emergency exit and forward of the wing. The lights provide visibility so that passengers can evacuate the aircraft through the emergency exit and off of the wing.

CONTROLS AND INDICATIONS The PASS SAFETY switch on the LIGHTS subpanel is a three-position switch. When the switch is in the PASS SAFETY (up) position, all the emergency lights illuminate. When the switch is selected to the SEAT BELT (down) position, only the seat belt light and RETURN TO SEAT lights illuminate. The three-position toggle switch on the EMERGENCY LIGHTING subpanel (Figure 33-10) also controls the emergency lighting in the cabin. When the switch is in the ON position, all of the emergency lighting illuminates using normal airframe battery power (through the crossfeed/emergency buses) and the NOT ARMED indicator immediately to the left of the switch illuminates. When the toggle switch is in the ARMED position, the NOT ARMED indicator extinguishes and the emergency lighting system is armed to illuminate if one of the following conditions exists: • One of the 2-G lateral force inertia switches next to each emergency battery pack closes. • All airframe electrical potential is less than 19 VDC.

Electrical power for emergency lights comes from two emergency lighting battery packs or from aircraft battery power. The forward emergency lighting battery pack is between the structural formers at FS 172.70 and FS 184.93 on the right side of the cabin. The aft emergency lighting battery pack is under the left cabin floorboard immediately forward of FS 317.55.

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CHAPTER 34 NAVIGATION CONTENTS Page INTRODUCTION ................................................................................................................. 34-1 GENERAL ............................................................................................................................ 34-3 DATA BUSES........................................................................................................................ 34-7 Description..................................................................................................................... 34-7 COLLINS PRO LINE 21 INTEGRATED AVIONICS PROCESSOR SYSTEM............... 34-11 Description................................................................................................................... 34-11 COLLINS PRO LINE 21 ATTITUDE HEADING SYSTEM ............................................ 34-13 Description................................................................................................................... 34-13 Components ................................................................................................................. 34-13 PITOT-STATIC SYSTEM .................................................................................................. 34-15 Description................................................................................................................... 34-15 Diagnostics .................................................................................................................. 34-15 REDUCED VERTICAL SEPARATION MINIMUM AIRSPACE .................................... 34-17 Description................................................................................................................... 34-17 Aircraft Approval......................................................................................................... 34-19 Operator Authorization ................................................................................................ 34-25 COLLINS PRO LINE 21 ELECTRONIC FLIGHT INSTRUMENT SYSTEM ................ 34-31 Description................................................................................................................... 34-31 Controls and Indications.............................................................................................. 34-31 ENGINES INDICATION SYSTEM................................................................................... 34-35 Description................................................................................................................... 34-35 Components ................................................................................................................. 34-35

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FLIGHT MANAGEMENT SYSTEM ................................................................................ 34-37 Description................................................................................................................... 34-37 DEPENDENT POSITION DETERMINING SYSTEMS .................................................. 34-39 Description................................................................................................................... 34-39 COLLINS RADIO SENSOR SYSTEM ............................................................................. 34-41 Description................................................................................................................... 34-41 Components ................................................................................................................. 34-41 TRAFFIC COLLISION AVOIDANCE SYSTEM.............................................................. 34-43 Description................................................................................................................... 34-43 Controls and Indications.............................................................................................. 34-43 ENHANCED GROUND PROXIMITY WARNING SYSTEM.......................................... 34-47 Description................................................................................................................... 34-47 Operation ..................................................................................................................... 34-47 RTA-800 WEATHER RADAR SYSTEM........................................................................... 34-49 Description................................................................................................................... 34-49 Operation ..................................................................................................................... 34-49 EMERGENCY LOCATOR TRANSMITTER SYSTEM .................................................. 34-51 Description................................................................................................................... 34-51 Operation ..................................................................................................................... 34-51

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ILLUSTRATIONS Figure

Title

Page

34-1

Electrostatic Sensitive Device Warning ................................................................. 34-2

34-2

Data Bus Communications..................................................................................... 34-4

34-3

Collins Pro Line 21 Simplified Diagram (525B) ................................................... 34-6

34-4

Sample Citation 525B Avionics Print .................................................................... 34-8

34-5

Integrated Avionics Processor System—Right Nose Area .................................. 34-10

34-6

Attitude Heading Computer................................................................................. 34-12

34-7

Air Data Computer............................................................................................... 34-14

34-8

RVSM Flight Envelope ........................................................................................ 34-16

34-9

Proper Documentation ......................................................................................... 34-18

34-10

Equipment Requirements..................................................................................... 34-20

34-11

Approval Process ................................................................................................. 34-22

34-12

Program Manual Requirements ........................................................................... 34-24

34-13

Report Height Keeping Errors ............................................................................. 34-26

34-14

RVSM Worldwide ................................................................................................ 34-28

34-15

AFD-3010 Adaptive Flight Display ..................................................................... 34-30

34-16

Display Control Panel ......................................................................................... 34-32

34-17

Adaptive Flight Display—Engine Indicating System.......................................... 34-34

34-18

Control Display Unit (CDU)................................................................................ 34-36

34-19

Citation 525B Cockpit ......................................................................................... 34-38

34-20

RTU-4210 ............................................................................................................ 34-40

34-21

TCAS Processor................................................................................................... 34-42

34-22

TCAS Antenna ..................................................................................................... 34-44

34-23

EGPWS Information............................................................................................ 34-46

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

34-24

Radar Antenna...................................................................................................... 34-48

34-25

C406-2 Emergency Locator Transmitter.............................................................. 34-50

TABLE Table 34-1

34-iv

Title

Page

Acronyms ............................................................................................................. 34-52

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 34 NAVIGATION B

2 1

C

9

NAV

6

3

N 3 3

PLAN

A

B

FPL

F

G

AFIS

K

D

E

H

I

J

M

N

O

INTRODUCTION This chapter describes the avionics systems used on the Citation 525B aircraft for navigation systems. The discussion is meant to be for familiarization purposes and does not cover all the different instrument options available for the aircraft. References for this c h a p t e r a n d f u r t h e r s p e c i fi c i n f o r m a t i o n c a n b e f o u n d i n C h a p t e r 5 — “ Ti m e Limits/Maintenance Checks,” Chapter 12—“Servicing,” and Chapter 34—“Navigation” of the Aircraft Maintenance Manual (AMM) .

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CAUTION ELECTROSTATIC SENSITIVE DEVICES DO NOT OPEN OR HANDLE EXCEPT AT A STATIC FREE WORK STATION

Figure 34-1. Electrostatic Sensitive Device Warning

34-2

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

GENERAL

NOTES

Electrostatic Discharge Protection With the increase of integrated avionics systems, protection against static electricity must be provided to prevent damage to the electronic systems (Figure 34-1). This assures that the information displayed to the crew is not corrupted or inaccurate. Electrostatic discharge is the most common cause of degradation (or destruction) of many electronic components, particularly integrated circuits (ICs), transistors, and semiconductors. Handle electrosensitive devices (ESDs) with extreme care. A rate/approved wrist strap attached to the same ground potential as the desired circuit card, logic module, or component places a technician at the same potential, eliminating a discharge of electricity (and damage to equipment). A typical discharge of electrostatic voltage is not seen or heard until it is in excess of 10,000 volts. This means that damage can occur without any indications to the operator until the device or component ceases to function. Most digital electronic components function on 5 VDC. Therefore, 100 volts of induced static electricity is more than enough to damage a component. ESDs are clearly marked and all necessary precautions should be taken. To find the factory suggested handling procedures of ESDsensitive items refer to Chapter 20—“Standard Practices Airframe” in the AMM.

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Figure 34-2. Data Bus Communications

34-4

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Analog Signals and Digital Signals An analog circuit is any circuit in which the output voltage and current values are considered significant over a continuous period of time. Analog = continuous change of state. A digital circuit is any circuit in which the output currents or voltages are interpreted as having two values. Digital = two changes of state. Generally, digital systems offer faster and more precise calculations than analog systems, and require less power to do so.

There are two methods of data transmission down a data bus or communications line: serial and parallel: • Serial data transmission—Information is sent down the bus single file. This is the slowest means of data transmission. • Parallel data transmission—Information is sent side by side. Each bit of information arrives at the same time making this the fastest means. In the Citation 525B aircraft, the most common means of data transmission is inside the avionics boxes using parallel busing. Outside of the avionics boxes, the preferred method is serial communications.

Analog systems still fill a gap where digital technology may fall short and would not be as practical (e.g., high-power applications).

NOTES

A discrete signal is a positive switch, it can either be a change from open to short or no voltage to voltage. This type of a discrete signal is commonly used with the squat switch. A discrete signal is considered to be an analog signal even though it is not used to transmit data.

Data Communication Data communication is a means by which avionics units communicate with each other to carry out programmed functions. Modern avionics units are capable of controlling other avionics units, sending and/or receiving information, and making complex decisions. To allow these digital units to communicate with each other, a communications line has to be connected to each unit. This is known as busing (Figure 34-2).

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DATA BUSES

On most shielded wires, the shield is grounded at both ends of each wire segment. An open shield at one end of a wire segment nullifies the HIRF, EMI, and EMC protection of the wire.

DESCRIPTION The line replaceable units (LRUs) and line replaceable modules (LRMs) in the Collins Pro Line 21 avionics systems communicate with each other using data-bus lines (Figure 343). Physically the data buses consist of two wires that are twisted together and shielded from interference. High-energy radiated fields (HIRF) and electromagnetic interference (EMI) is unwanted energy interfering with aircraft electronics, causing a disruption of normal operation. Coupling may occur through the aircraft wiring or directly into the equipment itself. This unwanted energy may come in the form of a lightning strike or interference from other transmitters. Electromagnetic comparability (EMC) is a condition when a signal transmitted by an onboard transmitter (or other electrical/electronic component) affects other system(s) in the aircraft.

Shield grounding may be at a connector backshell or equipment rack. In some cases, the shields are bonded to a backshell or equipment rack with a band clamp. Some wire bundles are enclosed in a shield overbraid to provide additional HIRF/EMI protection. The shield overbraid is grounded at both ends with a band clamp. Splice shielded wires by using a braided solder sleeve splice. HIRF, EMI, and EMC protection is designed with consideration for the wire bundle in which wires are routed. Relocation of a wire bundle may cause a change in the common mode impedance between wire conductors and the aircraft fuselage.

CAUTION Wire bundles should not be rerouted in a manner that changes the relative distance between the aircraft skin (or structure) and the wire bundle.

EMC caused by onboard transmitters occurs due to improper bonding of an access panel or other element common to the skin of the aircraft. The transmitted signal creates skin currents and is reradiated at the point of improper bonding and may bleed back into other system(s). EMC caused by electronic equipment is a condition when the equipment case or wire shields connected to the equipment is improperly bonded. Without proper bonding the signal is radiated into other equipment or wiring. Protection against HIRF, EMI, and EMC is accomplished through specific wire routing, proper grounding of equipment, and use of shielded wires (with the shield grounded at both ends of each wire segment).

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34-7

34-8

PI421 (REF) PFD2 (PFD-PS) WHT BLU WHT BLU

AFT PRESS BLKHD AVN/ELEC JA309 (REF) DISCONNECT *0 RA-FADEC 429 *8

PI413 (REF) PFD1 (PFD-P2)

FOR TRAINING PURPOSES ONLY

ALT ALERT WARN 51 OH/MOA ALERT 50 OVERSPEED WARN 53

JC004 E AVN ELEC DISCONNECT C 9-203439-9 H CONNECTOR S2039-6 SOCKET (34) WHT BLU GRN PI908 M24318/2-2 CONNECTOR 205960-1 RETAINER (2)

1 2 3 10

5 6 7

REV DSC1/PFD REV 35 REV DSC3-ENG_ENB 37 C XB900 (REF) WHT BLU

CC

8

CDC11 (REF)

15

ADC SOURCE SEL 40 AMC SOURCE SEL 49 MS27484T16F8S CONNECTOR M85049/47N16 BACKSHELL 28 VDC PRI POWER G 28 VDC HEATER PWR B

9 RH CB PANEL

PI414 PFD1 (PFD-P3)

PF900 (REF)

2

1

16 16

*S L

2

1

HEATER PWR GND D POWER GROUND E

16 16

FF HH

CHASSIS GROUND F

16

J

75 LB-FADEC 429 A 78 LC-FADEC 429 B

GRN BLU WHT

51 60 53 66 54 88

ALT ALERT WARN OH/MDA ALERT OVERSPEED WARN LCD BRT CTRL L LCD BRT CTRL IN +25 VCD LCD BRT

49 40 52 35

AHC SOURCE SEL ADC SOURCE SEL POWER DOWN REV DSC1/PFD REV

PILOT LIGHTING PANEL 1 GRN 2 BLU WHT 3

RI400 S3495-1K RHEOSTAT CM3590-1 KNOB PFD/MFD DIMMER

D7

4 CDC10 (REF) J

WHT BLU

JI908 M24308/4-2 CONNECTOR 205817-1 RETAINER (2)

LH INSTR PNL DISCONNECT POWER DOWN 62

SV PNL LTG H 32 SV PNL LTG L 33

90 RA-FADEC 429 A 91 RA-FADEC 429 B

WHT BLU

LB-FADEC 429 A 75 LB-FADEC 429 B 75

+28 VDC LCD BRT 65 LCD BAT CTRL IN 55 LCD BAT CTRL L 66

PI416 (REF) MFD (MFD-P2)

WHT BLU WHT BLU

AVN/ELEC JA306 (REF) DISCONNECT 15 LB-FADEC 429 16

RA-FADEC 429 A RA-FADEC 429 B LB-FADEC 429 A LB-FADEC 429 B ALT ALERT WARN OH/MDA ALERT OVERSPEED WARN REV OSC3/ENG_ENB REV OSC1/PFD REV

SI402 S3308-3 SWITCH CM3590-2 (REF) MFD KNOB D6 D7 MORM PFD/MFD REV SWITCH D/C D8 PFD SI319 S382-3 A1 A2 MORM ADC REV SWITCH REV SI312 A3 MORM S382-3 A1 AHRS REV A2 REV SWITCH A3 B5

D/C

PFD PFD HTR

CDC10 (REF)

CCS10 (REF)

Figure 34-4. Sample Citation 525B Avionics Print

REF COMM

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

RA-FADEC 429 A 90 RA-FADEC 429 B 91

90 91 75 76 51 60 53 37 35

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

ARINC 429 Data Bus

NOTES

An aeronautical radio incorporated (ARINC) 429 bus system is comprised of transmitters and receivers connected by shielded/twisted wire pairs. ARINC 429 is the most common standard data bus used by the Collins systems. The ARINC 429 consists of a 32-bit, binary coded decimal data word. The first 8 bits make up a label that categorizes the data, (e.g., pitch attitude information). Bits 9 and 10 make up the source destination identifier (SDI), which identifies either the left or right system. Bits 11 through 29 contain pertinent information (e.g., actual pitch attitude, in degrees, of the aircraft). Bits 30 and 31 make up the sign status matrix (SSM) and defines the overall system status. The remaining bit (32) is an odd parity bit used by the avionics input/output processors to ensure data integrity. Each ARINC 429 transmitter can communicate with up to 20 receivers. Data flows only one way over an ARINC 429 bus. Bidirectional transmission between two LRUs must be accomplished by using two sets of transmitters, receivers, and twisted pair wires.

RS-422 Data Bus The RS-422 data bus is an electrical specification as defined by the Electronics Industries Association (EIA). It is used where bidirectional communications are needed (e.g., between the displays and display controllers). The data buses consist of a pair of shielded twisted wires.

RS-232 Data Bus The RS-232 data bus is an electrical specification as defined by EIA. The RS-232 bus describes any connection between the avionics system and a personal computer. Refer to Figure 34-4 for a sample of a Citation 525B aircraft avionics wiring diagram.

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Figure 34-5. Integrated Avionics Processor System—Right Nose Area

34-10

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COLLINS PRO LINE 21 INTEGRATED AVIONICS PROCESSOR SYSTEM DESCRIPTION Collins Pro Line 21 integrated avionics processor system (IAPS) provides the integration function required to connect and manage the various avionics systems in the aircraft (Figure 34-5). The IAPS consists of one ICC-3001 card gage. The ICC-3001 card gage provides a mounting surface for: • Two CSU-3000 configuration strapping units—Provide a matrix of configuration shunts that program the IAPS. • IEC 3001 environmental controller— Monitors temperature sensors, controls cooling fans and heaters to maintain a stable operating environment. • Two IOC-3000 IO processor cards— Provide a data management function by acting as a central data collection and distribution point for:

• Tw o P W R - 3 0 0 0 p ow e r s u p p l i e s — Provide independent power to FGC3000 flight guidance computers and IOC-3000 IO processor cards. • Two FMC-3000 flight management cards—Fully functional flight management system utilizing navigation sensors for flight planning and progress information. • MDC-3110 maintenance diagnostics computer—Provide information that can be utilized to troubleshoot aircraft systems for faults. • Tw o F G C - 3 0 0 0 f l i g h t g u i d a n c e computers—Provide the autopilot and flight director functions. The empty ICC-3001 IAPS card cage is an LRU. Each CSU, FGC, IEC, IOC, FMC, MDC, and PWR cards installed in the card cage are individual LRMs. If any of these modules fail, that module, and not the entire IAPS assembly, can be replaced.

bus inputs from the air data com° Data puter (ADC)

° ° ° ° ° °

Attitude heading computers (AHCs) Flight guidance computers (FGCs) Flight management computer COMM/NAV/pulse radios Radio altimeter Engines data concentrator units (DCUs)

IOC-3000 IO processor cards output data to weather radar assembly, primary COMM-NAV radios, ADC, flight control computers, flight management computer, and engine DCUs.

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Figure 34-6. Attitude Heading Computer

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COLLINS PRO LINE 21 ATTITUDE HEADING SYSTEM DESCRIPTION The Collins Pro Line 21 attitude heading system (AHS) senses the aircraft attitude, heading, three-axis rate and acceleration movements. The system consists of: • Two AHC-3000 attitude heading computers (Figure 34-6) • Two FDU-3000 flux detector units • Two ECU-3000 external compensation units

The right AHC-3000 attitude heading computer provides pitch, roll, and stabilized magnetic heading data information for the right FGC. The AHC receives data from the ADS, the right flux detector unit, and the right ECU. The AHC outputs data to the right IO processor card and right FGC.

Flux Detector Units The FDU-3000 flux detector units are in the vertical stabilizer of the aircraft. The units sense the horizontal component of the earth’s magnetic field. The flux detector units provide magnetic flux measurements and outputs to the AHC-3000 attitude heading computers, which compute the aircraft magnetic heading angle.

External Compensation Units

COMPONENTS Attitude Heading Computers AHC-3000 attitude heading computers are installed in precision-leveled mounts. The AHC3000 attitude heading computer replaces conventional vertical gyros, directional gyros, three rate gyros, and three linear accelerometers. The system provides a visual indication of the aircraft attitude and direction to the pilot multifunction display (MFD), primary flight display (PFD), and the copilot PFD (if installed).

The ECU-3000 external compensation units provide electrical compensation for any flux detector errors caused by the aircraft altering the earth’s magnetic field. The ECUs are remotely mounted and contain electrically erasable programmable read only memory (EEPROM), which stores compensation data for hard iron errors and flux detector misalignment. The compensation data is obtained through an automatic compass swing procedure.

The left AHC-3000 attitude heading computer provides pitch, roll, and stabilized magnetic heading data information for the left FGC and optional traffic alert and collision avoidance system. The AHC receives data from the air data system (ADS), the left flux detector unit, and the left external compensation unit (ECU). The AHC outputs data to the left IO processor card, left FGC, and optional traffic alert and collision avoidance system.

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TRI-COLOR LED • RED = ADC FAIL • AMBER = DATA LOAD MODE (SERVICE CENTER ONLY) • OFF = NORMAL OPERATION

Figure 34-7. Air Data Computer

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PITOT-STATIC SYSTEM DESCRIPTION The standard pitot-static system incorporates two ADC-3000 air data computers (1ADC, left; 2ADC, right) (Figure 34-7). The pitot-static system is two independent pitot and two independent static systems. Pitot and static systems connect pitot hoses, tubes, and static ports to the flight data instruments. Two pitot tubes are on the lower surface of the aircraft nose section. The pitot systems provide pitot pressure to their respective airspeed indicators. The pitot system also provides pitot pressure for the ADCs.

A red IAS failure flag in the airspeed tape, ALT in the altitude, and/or VS in the vertical speed indication alerts the crew a failure of the ADC has occurred.

DIAGNOSTICS Aircraft static systems are required by FAR 91.411 to have altimeter and static system tests. Persons and facilities authorized to perform altimeter and static systems tests are also identified in FAR 91.411. Pressure actuated (barometric pressure) encoding altimeter and static system tests are described in Appendix E of FAR Part 43.

NOTES

The ADC is equipped with a tri-colored light emitting diode (LED) built-in test equipment (BITE). If the BITE indicator turns red, the ADC has failed. If the indicator turns amber, the ADC is in data load mode (this should be seen in the field). If the LED is off, the ADC is operational. The ADS relays information to the IOC in the IAPS to be shared with the flight guidance system (FGS). The ADS also provides information to the AHS for enhanced accuracy and the PFDs and MFD for display to the crew. The altitude information given to the crew is corrected for the changes in barometric pressure. The correction is set with the display control panel. Rotating the BARO knob adjusts the pressure while pressing the center of the knob standardizes the barometric correction to either 29.92 Hg (standard day) or 1013 HPA. The ADC also gives Mach speed information to the crew through a digital window directly below the airspeed indication. This is generated by using the temperature probe on the right side of the aircraft nose to calculate true airspeed (TAS) and then comparing that to the speed of sound at sea level on an ambient day.

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Figure 34-8. RVSM Flight Envelope

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REDUCED VERTICAL SEPARATION MINIMUM AIRSPACE DESCRIPTION Within reduced vertical separation minimum (RVSM) airspace, air traffic control (ATC) separates aircraft by a minimum of 1,000 feet vertically between flight level (FL) 290 and FL 410 inclusive (Figure 34-8). RVSM airspace is special qualification airspace in which the administrator must approve the operator and the aircraft used by the operator. ATC notifies operators of RVSM by providing route-planning information.

RVSM Flight Envelope A full RVSM flight envelope includes: • Range of Mach number • Weight divided by atmospheric pressure ratio • Altitudes over which an aircraft is approved to be operated in cruising flight within RVSM airspace The altitude flight envelope extends from FL 290 upward to the lowest altitude of the following FL 410 (RVSM altitude limit):

RVSM Grouped Aircraft RVSM grouped aircraft is defined as a group o f a i r c r a f t ( a p p r ove d a s a g r o u p b y t h e Administrator) in which each of the aircraft satisfy the following: • The aircraft have been manufactured to the same design and have been approved under the same type certificate, amended type certificate, or supplemental type certificate. • The static systems of each aircraft are installed in a manner and position that is the same as those of the other aircraft in the group. The same static source error correction is incorporated in each aircraft of the group. • The avionics units installed in each aircraft meet the minimum RVSM equipment requirements and are manufactured to the same manufacturer specification and have the same part number (or of a different manufacturer or part number if the applicant demonstrates that the equipment provides equivalent system performance).

RVSM Nongrouped Aircraft RVSM nongrouped aircraft is defined as an aircraft that is approved for RVSM operations as an individual aircraft and is not part of a specific RVSM group.

• Maximum certificated altitude for the aircraft • Altitude limited by cruise thrust, buffet, or other flight limitations

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Figure 34-9. Proper Documentation

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AIRCRAFT APPROVAL

NOTES

An operator may be authorized to conduct RVSM operations if the Administrator finds that its aircraft comply with RVSM aircraft requirements. The applicant for authorization shall submit the appropriate data package for aircraft approval (Figures 34-9 through 34-11). The package must consist of: • An identification of the RVSM aircraft group or the nongroup aircraft • Definition of the RVSM flight envelopes applicable to the subject aircraft • Documentation that establishes compliance with the applicable RVSM aircraft requirements • Conformity tests used to ensure that the aircraft approved with the data package meet the RVSM aircraft requirements

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Figure 34-10. Equipment Requirements

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Altitude-Keeping Equipment (All Aircraft)

Altimetry System Error Containment (All Aircraft)

To approve an aircraft group or a nongroup aircraft, the Administrator must find that the aircraft meets the all requirements (Figure 34-10):

To approve group aircraft for which type certification application was made on or before April 9, 1997, the Administrator must find that the altimetry system error (ASE) is contained:

• The aircraft must be equipped with two operational independent altitude measurement systems. • The aircraft must be equipped with at least one automatic altitude control system that: the aircraft altitude within a ° Controls tolerance band of ±65 feet about an acquired altitude when the aircraft is operated in straight and level flight under nonturbulent, nongust conditions. the aircraft altitude within a ° Controls tolerance band of ±130 feet under nonturbulent, nongust conditions for aircraft for which type certification application occurred on or before April 9, 1997 that are equipped with an automatic altitude control system with flight management/performance system inputs.

• At the point in the basic RVSM flight envelope where mean ASE reaches its largest absolute value, the absolute value may not exceed 80 feet. • At the point in the basic RVSM flight envelope where mean ASE plus three stand a r d d ev i a t i o n s r e a c h e s i t s l a rg e s t absolute value, the absolute value may not exceed 200 feet. • At the point in the full RVSM flight envelope where mean ASE reaches its largest absolute value, the absolute value may not exceed 120 feet. • At the point in the full RVSM flight envelope where mean ASE plus three stand a r d d ev i a t i o n s r e a c h e s i t s l a rg e s t absolute value, the absolute value may not exceed 245 feet.

• The aircraft must be equipped with an altitude alert system that signals an alert when the altitude displayed to the flight crew deviates from the selected altitude: more than ±300 feet for aircraft for ° By which type certification application was made on or before April 9, 1997 more than ±200 feet for aircraft for ° By which application for type certification is made after April 9, 1997.

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Figure 34-11. Approval Process

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Altimetry System Error Containment Grouped Aircraft To approve group aircraft for which type certification application is made after April 9, 1997, the Administrator must find that the ASE is contained: • At the point in the full RVSM flight envelope where mean ASE reaches its largest absolute value, the absolute value may not exceed 80 feet. • At the point in the full RVSM flight envelope where mean ASE plus three stand a r d d ev i a t i o n s r e a c h e s i t s l a rg e s t absolute value, the absolute value may not exceed 200 feet.

Altimetry System Error Containment (Nongrouped Aircraft) To approve a nongroup aircraft, the Administrator must find that the ASE is contained:

Traffic Alert and Collision Avoidance System Compatibility Traffic alert and collision avoidance system compatibility includes all aircraft with RVSM operations approval after March 31, 2002 (unless otherwise authorized by the Administrator) if the aircraft is equipped with TCAS II in RVSM airspace. It must be a TCAS II that meets TSO C-119b (Version 7.0), or a later version. If the Administrator finds that the applicant’s aircraft comply with this section, the Administrator notifies the applicant in writing.

Operating Restrictions If the applicant demonstrates that the aircraft otherwise comply with the ASE containment requirements, the Administrator may establish operating restrictions on that applicant’s aircraft. This may restrict the aircraft from operating in areas of the basic RVSM flight envelope.

• For each condition in the basic RVSM flight envelope, the largest combined absolute value for residual static source error plus the avionics error may not exceed 160 feet.

NOTES

• For each condition in the full RVSM flight envelope, the largest combined absolute value for residual static source error plus the avionics error may not exceed 200 feet.

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Figure 34-12. Program Manual Requirements

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OPERATOR AUTHORIZATION Authority for an operator to conduct flight in airspace where RVSM is applied is issued in the Operations Specifications, a Letter of Authorization, or Management Specifications issued under Subpart K of Part 91, as appropriate (Figure 34-12). To issue an RVSM authorization, the Administrator must find that the operator’s aircraft is approved in accordance with Section 2 of Appendix G of Part 91.

• Procedures for returning noncompliant aircraft to service. An applicant who operates under Part 121 or 135 or under Subpart K of Part 91 must submit RVSM policies and procedures that enable them to conduct RVSM operations safely.

An applicant for authorization to operate within RVSM airspace shall apply in a form and manner prescribed by the Administrator.

• In a manner prescribed by the Administrator, the operator must provide evidence that it is capable to operate and maintain each aircraft or aircraft group for which it applies for approval to operate in RVSM airspace, and each crewmember has an adequate knowledge of RVSM requirements, policies, and procedures.

The application must include an approved RVSM maintenance program outlining procedures to maintain RVSM aircraft in accordance with the requirements of Appendix G.

• Adequate procedures for the notification to the flight crew when their aircraft are determined to be non-RVSM capable for dispatch.

Program Manual Requirements Each program manual must contain criteria for: • Periodic inspections, functional flight tests, and maintenance and inspection procedures, with acceptable maintenance practices, for ensuring continued compliance with the RVSM aircraft requirements.

• Adequate provisions for auditing all outsourcing of maintenance on a regular basis to ensure conformance to RVSM maintenance program requirements and the operator’s Continuing Analysis and Surveillance Program, or an equivalent program for Part 91 operators conducting flights in RVSM airspace, under its approved maintenance program.

• A quality assurance program for ensuring continuing accuracy and reliability of test equipment used for testing aircraft to determine compliance with the RVSM aircraft requirements. • Initial and recurrent training requirements in the use of special test apparatus for performing geometrical inspections (e.g., skin waviness measurements), and other special requirements relating to the maintenance of height-keeping equipment and certification requirements. • Established provisions to ensure that all RVSM maintenance personnel are properly trained, qualified, and authorized to maintain the aircraft for RVSM operations.

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Figure 34-13. Report Height Keeping Errors

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Reporting Altitude-Keeping Errors

NOTES

Each operator shall report to the Administrator each event in which the operator’s aircraft has exhibited the following altitude-keeping performance (Figure 34-13): • Total vertical error of 300 feet or more • ASE of 245 feet or more • Assigned altitude deviation of 300 feet or more

Removal or Amendment of Authority The Administrator may amend operation specifications or management specifications issued under Subpart K of this part to revoke or restrict an RVSM authorization, or may revoke or restrict an RVSM letter of authorization, if the Administrator determines that the operator is not complying, or is unable to comply. Examples of reasons for amendment, revocation, or restriction include, but are not limited to: • An operator’s committing one or more altitude-keeping errors in RVSM airspace • Failing to make an effective and timely response to identify and correct an altitude-keeping error • Failing to report an altitude-keeping error

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Figure 34-14. RVSM Worldwide

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Airspace Designation

RVSM in the Gulf of Mexico

Refer to Figure 34-14.

RVSM may be applied in the Gulf of Mexico in the following areas:

Northern Atlantic Track

• Gulf of Mexico High Offshore Airspace

RVSM may be effective in the Minimum Navigation Performance Specification (MNPS) airspace within the Northern Atlantic Track (NAT). The MNPS airspace within the NAT is defined by the volume of airspace bet w e e n F L 2 8 5 a n d F L 4 2 0 ( i n c l u s iv e ) . Extending between latitude 27° north and the North Pole. Bounded in the east by the eastern boundaries of control areas Santa Maria Oceanic, Shanwick Oceanic, and Reykjavik Oceanic. Bounded in the west by the western boundaries of control areas Reykjavik Oceanic, Gander Oceanic, and New York Oceanic, excluding the areas west of 60° west and south of 38° 30 minutes north.

• Houston Oceanic ICAO FIR • Miami Oceanic ICAO FIR • Atlantic High Offshore Airspace • San Juan ICAO FIR

RVSM in Europe RVSM may be enforced in all European Civil Aviation Control (ECAC) controlled areas.

NOTES

RVSM in the Pacific RVSM may be applied in the Pacific in the following ICAO Flight Information Regions (FIRs) Anchorage Arctic, Anchorage Continental, Anchorage Oceanic, Auckland Oceanic, Brisbane, Edmonton, Honiara, Los Angeles, Melbourne, Nadi, Naha, Nauru, New Zealand, Oakland, Oakland Oceanic, Port M o r e s b y, S e a t t l e , Ta h i t i , To ky o , U j u n g Pandang and Vancouver.

RVSM in the West Atlantic Route System RVSM may be applied in the New York FIR portion of the West Atlantic Route System (WATRS).

RVSM in the United States RVSM may be applied in the airspace of the 48 contiguous states, District of Columbia, and Alaska, including that airspace overlying the waters within 12 NM of the coast.

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Figure 34-15. AFD-3010 Adaptive Flight Display

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COLLINS PRO LINE 21 ELECTRONIC FLIGHT INSTRUMENT SYSTEM

CONTROLS AND INDICATIONS Primary Flight Display Each PFD depicts: • Aircraft attitude

DESCRIPTION

• Flight director commands

Collins Pro Line 21 electronic flight instrument system (EFIS) interfaces with and displays data information from the:

• Flight control system annunciations • Heading

• Air data system

• Course

• Attitude heading system

• Bearing

• Engines indicating system

• Vertical speed

• Flight control system

• Airspeed

• Te r r a i n awa r e n e s s wa r n i n g s y s t e m (TAWS)

• Baro-corrected altitude

• Traffic alert and collision avoidance system (optional)

• Preselect altitude

• Radio altitude • Minimum descent-reporting altitude

Data information is displayed on liquid crystal displays referred to as adaptive flight displays (AFDs) (Figure 34-15). Collins Pro Line 21 EFIS standard installation configuration consists of: • Three AFD-3010 adaptive flight displays • Two DCP-3000 display control panels • CKP-3000 course knob panel • CHP-3000 course heading panel Two AFD-3010 adaptive flight displays are utilized as PFDs. One AFD-3010 adaptive flight display is utilized as a MFD. The pilot PFD is on the left instrument panel and the copilot PFD is on the right instrument panel. The MFD is on the center radio panel.

• Decision height • Temperature • Optional TCAS II advisory information

Multifunction Display The MFD depicts radar return and comprehensive navigation information in the horizontal situation indicator or the present position map format. The MFD also allows engines indication system information, flight management, checklist, diagnostic, TAWS and optional TCAS II pictorial information to be viewed by the pilot.

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BARO PUSH

BAROMETRIC SET KNOB (PRESS TO STANDARDIZE)

STD

ENG

COMPRESS ENGINE INFO ALLOWING MORE ROOM FOR PLAN OR MAP INFO

REFS

BRING REFERENCE PAGES INTO VIEW

MENU ADV

DATA H

PUS

USED WITH LINE SELECT BUTTONS ON AFD TO SET VARIOUS ITEMS

T

S

EL C E

NAV/BRG

BRING NAV BEARING COMMANDS ONTO AFD FOR LINE SELECTION

RADAR

BRING RADAR COMMANDS ONTO AFD FOR LINE SELECTION

GCS

TURN GROUND CLUTTER SUPPRESSION ON MOMENTARILY

TILT

RANGE H

PUS AUTO TIL T

RADAR TILT AND RANGE CONTROL

Collins

Figure 34-16. Display Control Panel

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DCP-3000 Display Control Panel

NOTES

The pilot and copilot use the display control panels (Figure 34-16) to: • Set barometric pressure correction • Secondary engines display • V-speed reference setting • Navigation source selection • Bearing source selection • Weather radar control • Display range selection The display control panels provide a display control output to the MFD and left PFD. The left display control panel receives discrete input from the course heading panel and external switches. The right display control panel receives discrete input from the course heading panel, course knob panel, and external switches. A course-heading panel provides heading and preselect altitude data to the left and optional right display control panel. The pilot uses the course heading panel to: • Select course • Select heading • Preselect altitude references for the FGS • Provide left side course data to the left display control panel • Cancel an altitude alert annunciation The course-heading panel provides discrete output to both display control panel. A course knob panel provides right side course data to the right display control panel.

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Collins

ITT °C 1000 900

N1 % –––.– 100 T T O 90 O

700

50 600 400 200

30

0.0

HDG 137

VOR1 CRS 166

OIL °C

12

FUEL QTY LBS 2000 1500

0

0 °C FUEL 0 PPH

23

25

25

1000 500

23 0

0

0

137 15 S

E

TTG – – : – – – – – – NM

OIL PSI

0.0

800

70

0.0

0.0 N2 %

5

2.5 TERR LX/RDR ADF

TFC TA ONLY

ADF

BRT DIM

MFD - MULTIFUNCTION DISPLAY AFD - ADAPTIVE FLIGHT DISPLAY ACT AS AFD OR PFD 8 X 10 IN LDC - LIQUID CRYSTAL DISPLAY EIS - ENGINE INDICATING SYSTEM

Figure 34-17. Adaptive Flight Display—Engine Indicating System

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ENGINES INDICATION SYSTEM DESCRIPTION

•

ITT—Gauge indicates the temperature between the first and second compressor stages in degrees centigrade. The display of ITT consists of an analog scale and pointer for each engine.

• N2—Standardized display of engine turbine rpm measured against a fixed 100% value. The N 2 displays consist of digital readouts for each engine.

The engine indication system (EIS) consists of two DCU-3010 data concentrator units, two full authority digital engine controls (FADECs), and the electronic flight information system (EFIS) (Figure 34-17).

• Oil pressure—Analog and part time digital display for each engine.

COMPONENTS

• Oil temperature—Analog and part time digital display for each engine.

Data Concentrator Units The left and right DCUs collect and format aircraft data for display on the EFIS system. The DCU receives oil pressure, oil temperature, fuel flow, fuel quantity, and fuel temperature inputs from the aircraft engine and other systems. The data inputs are concentrated and processed for transmission on ARINC 429 buses. The DCU outputs engine data to the EFIS displays and maintenance, diagnostic, and aircraft data to the IAPS data concentrators.

• Fuel flow—Displays in pounds per hour (PPH). Optional kilograms per hour (KPH) may be displayed. Consists of digital readouts for each engine. A fuel flow PPH legend is displayed between the left and right digital readouts. • Fuel quantity—Consists of an analog and digital display for the fuel in each wing tank. Fuel quantity is in pounds (LBS). Optional kilograms (KGS) may be displayed. The fuel quantity digital readout is displayed below the associated analog scale. The fuel quantity digital readout has the same source of data as the fuel quantity analog pointer.

FADECS There are two dual channel FADECs that receive N 1 , N 2 , and interturbine temperature (ITT) information from the aircraft engines. The information is converted to ARINC 429 and sent to the EFIS displays and maintenance, diagnostic, and IAPS data concentrators.

Engine Indication Display The engine indication display shows the following indications to the crew: • N 1 — Turbine speed (N 1 ) gauge indicates engine fan rpm. N 1 is measured against a fixed 100% value and is expressed in percentages. The N 1 displays consist of an analog and digital display for each engine.

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Collins

ACT FPLN

1/1

ORIGIN

DIST

KCEA

DEST A LT N

RO U T E

O RIG RWY VIA

TO

COMPLETED

FLT N O

< C O P Y AC TI V E < S E C FP L N

PERF

[

INIT> [

G P S N OT AVA I L A B L E

MSG

DEP ARR

LEGS

PERF

MFD MENU

EXEC

MFD ADV

MFD DATA

DIR

FPLN

IDX

1

2

3

A

B

C

D

E

TUN

4

5

6

H

I

J

K

L

PREV

NEXT

F

G

CLR DEL

M

N

BRT DIM

7

O

P

Q

R

S

T

U

0 +/– V

W

X

Y

Z

SP

/

8

9

Figure 34-18. Control Display Unit (CDU)

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FLIGHT MANAGEMENT SYSTEM

NOTES

DESCRIPTION The Citation 525B aircraft is standard equipped with a Collins FMC-3000 flight management system (FMS). The FMC-3000 is a circuit card assembly in the IAPS. Only one FMS is standard and if selected, there is only one FMC-3000 in the left side of the IAPS . If the optional second Collins FMS is installed, there is a second FMC-3000 in the right side of the IAPS. The FMC-3000 receives inputs from the navigation equipment in the aircraft to plot present position, follow a flight plan, determine performance and progress of the aircraft while flying a flight plan, and provide information as guidance to the flight director and autopilot. The FMC-3000 allows for storage of the navigation database, custom database, and the performance database. The setup and control of the FMS is accomplished through the control display unit (CDU). The FMS is automatically powered whenever the avionics switch is turned on.

CDU-3000 The CDU-3000 is in the pedestal area in the cockpit of the aircraft and provides control functions for the FMS and the radio sensor system (RSS) (Figure 34-18). The CDU has an liquid crystal display (LCD) screen, line select keys on either side of the screen, program keys for mode selection, and an alpha/numeric keypad for data entry.

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Figure 34-19. Citation 525B Cockpit

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DEPENDENT POSITION DETERMINING SYSTEMS DESCRIPTION This section describes subsystems that provide position information and are dependent on ground installations or orbital satellites. These navigation aids include: • Low frequency navigation systems (below 30 MHz) • Very high frequency (VHF) navigation systems (30 MHz–300 MHz) • Ultra high frequency (UHF) navigation systems (300 MHz–3, 000 MHz)

Low Frequency Navigation Systems The low frequency navigational systems include automatic direction finder (ADF) navigation systems and a Collins ADF 4000 digital ADF radio. These low frequency systems operate at frequencies between 30 kHz and 30 MHz. The ADF is a radio receiver that: • Automatically and continuously determines bearing to a radio station being received • Monitors aural reception of audio transmissions from the radio station Operations and tuning of the ADF are discussed in Chapter 23—“Communications.”

VHF Navigation Systems The VHF navigation systems include navigation receivers. VHF systems operate at frequencies between 30 MHz and 300 MHz. The navigation receivers, designated as NAV 1 and NAV 2, contain the: • Collins NAV 4000 VOR/ILS—Features 200 channels, with 50 kHz receiver selectivity, and operates in the ranges of 108.00 MHz to 117.95 MHz. • Marker beacon • Glide slope—Features 40 channels, with 150 kHz receiver selectivity, and operates in the ranges of 329.15 MHz to 335.00 MHz.

UHF Navigation Systems UHF navigation systems include a Collins DME 4000 distance measuring equipment (DME) and a Collins TDR 94D transponder. UHF systems operate at frequencies between 300 MHz and 3,000 MHz. The DME system is a pulse system that transmits and receives pulses of UHF electrical energy. This information is used to measure distance and is displayed on the DME indicator, as well as analog and EFIS systems. The transponder system provides and receives electronic pulses from the ground station to the aircraft. Operations and tuning is discussed in Chapter 23—“Communications.”

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Figure 34-20. RTU-4210

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COLLINS RADIO SENSOR SYSTEM DESCRIPTION The radio sensor system (RSS) in the Citation 525B aircraft is comprised of all Collins equipment. Control of the RSS is accomplished by a radio tuning unit (RTU) (Figure 34-20) which is interfaced with all the radio/audio components to provide a complete Collins system. Operations of the RTU are discussed in Chapter 23—“Communications.” The RSS uses a radio interface unit (RIU) to complete the system interfacing.

Transponder Two TDR-94D diversity mode S transponders (XNPDR) are in the Citation 525B aircraft. The No. 1 system is in the nose baggage area on the left side of the aircraft. The No. 2 system is in the tail cone above the baggage area. The TDR-94D is capable of modes A, C, and S functions allowing for bearing distance and identification to be received and transmitted. The information can be shared with a TCAS II system.

COMPONENTS

A mode S strapping module is in the tail cone aft of the baggage compartment. This module allows for specific indent information to be delivered to the XNPDRs and the TCAS II system.

Navigation Receivers

The RTU controls the transponders.

The navigation receivers are Collins NAV4000 receivers that contain a VOR/LOC/ glideslope receiver, marker beacon receiver, and an ADF receiver. The No. 1 system is in the nose baggage area on the left side of the aircraft. The No. 2 system is in the tail cone above the baggage area.

NOTES

If only one ADF is installed, the No. 2 system contains a NAV-4500 (no ADF) instead of a NAV-4000 unit.

Distance Measuring Equipment Two DME-4000 units are in the Citation 525B aircraft. The No. 1 system is in the nose baggage area under the floor. The No. 2 system is in the tail cone above the baggage area. The DME provides distance, time-to-station, ground speed, and station identification. Tuning is accomplished automatically with either the FMS or the NAV receiver.

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Figure 34-21. TCAS Processor

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TRAFFIC COLLISION AVOIDANCE SYSTEM DESCRIPTION The Citation 525B aircraft is equipped with the Collins TTR-4000 TCAS II system as an option (Figure 34-21). The TTR-4000 is Change 7 compliant and meets all worldwide requirements. The TTR-4000 is under the floor on the left side of the nose area and provides: • Surveillance • Threat detection • Collision avoidance tracking • Interface with the audio and electronic display system (EDS) Resolutions are presented in verbal and visual form. Visual resolutions are presented on the PFD screen as a vertical fly to arc in the vertical speed scale. Audio resolutions are announced over the speakers and headphones by the aircraft audio system. An optional speaker override switch prevents the resolution audio from being heard over the speakers, which may scare the passengers. When activated the resolution can only be heard through the headsets; therefore, the headsets must be worn all the time.

If the test PASSES, the green PASS annunciator momentarily illuminates and the audio system in the aircraft announces “TCAS system test OK.” If the self-test FAILS, the the red FAIL annunciator momentarily illuminates and the audio system announces “TCAS system test fail.”

Annunciators The TTR-4000 is equipped with fault analysis to recognize problems with antennas, XNPDRs, radio altimeter, heading, and improper RA or TA sensing. An annunciator lamp on the front of the TTR-4000 is assigned to each of these areas: • XNPDR annunciator—Indicates the link between the TCAS and the XNPDR has failed. • LOWER and UPPER ANT annunciator—Indicate an ANT failure. • RAD ALT annunciator—Indicates a failure with the communications between the TCAS and the RAD ALT has failed. • HDNG annunciator—Indicates the TCAS is not receiving heading information from the AHRS. • RA and TA annunciator—Indicate an erroneous resolution or traffic advisory.

The traffic symbols may be activated and displayed on the MFD screen through the line select switches on the bezel of the screen.

CONTROLS AND INDICATIONS Self Test The TTR-4000 is equipped with a built-in selftest. To activate the test, press the button on the front of the TTR-4000 itself. Annunciators on the front of the TTR-4000 momentarily illuminate. Once the self-test is completed, all annunciators extinguish except either the PASS or FAIL annunciators.

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Figure 34-22. TCAS Antenna

34-44

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TCAS Antennas

NOTES

The TTR-4000 utilizes two antennas: one on the top of the aircraft and one on the bottom of the aircraft. Both are directional antennas that use a set of four coaxes to connect to the TTR (Figure 34-22). Each coax is then connected to the antenna coils, creating an impedance matching between coax and antenna. Coaxes are color coded for installation and must be matched to the color code on both the antenna and the TTR. Disconnecting the coaxes at the TTR and measuring impedance between the center conductor and the shielding of the coax connector can perform a check of the coaxes. Refer to the Collins Manual for specifications. Only the original equipment manufacturer (OEM) can replace the coaxes.

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Figure 34-23. EGPWS Information

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ENHANCED GROUND PROXIMITY WARNING SYSTEM

NOTES

DESCRIPTION The enhanced ground proximity warning system (EGPWS) is a terrain awareness and warning system (Figure 34-23). The system incorporates terrain alerting and display functions that provide the flight crew with aural and visual warnings if the projected flight path could result in impact with the terrain.

OPERATION The EGPWS is interfaced with: • Radio altimeter • ADCs • Angle-of-attack • Attitude heading reference system • Landing gear position sensors • Flap position sensors • Selected decision height • Glide-slope navigation receivers The EGPWS system utilizes information from these systems to determine flight path and possibility of impending danger. Data from the EGPWS system goes to the IAPS. Visual warnings are indicated on instrument panel annunciators and on the AFDs. Voice warning messages are announced on the aural warning system. Operation of the EGPWS is automatic when the avionics power switches are in the ON position and all related systems are valid. The system is operational from altitudes of 50–2,450 feet (4.57–746.76 meters) above ground level as determined by the radio altimeter. The EGPWS has six operational modes, all of which are valid for ground proximity warning.

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Figure 34-24. Radar Antenna

34-48

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RTA-800 WEATHER RADAR SYSTEM

NOTES

DESCRIPTION The RTA-800 weather radar system is a fully integrated radar system that consists of a single unit on a precisely aligned surface in the radome (Figure 34-24). The RTA-800 receiver/transmitter/antenna is a solid-state unit with an attached 14-inch flat plate antenna. The system is an X-band radar that detects precipitation and ground feature returns in front of the aircraft. The radar video can be displayed on the MFD. The radar outputs weather reflectivity data on a ARINC 453 very-highspeed bus to the AFDs.

OPERATION The RTA-800 weather radar system detects wet precipitation along the flight path and ahead of the aircraft. The display control panels provide radar mode and range control. The display range is selectable up to 300 NM. The system operates on a nominal output of 25 W. Scan and tilt control circuits cause the motor to sweep the antenna horizontally and vertically. The transmitter generates X-band pulse signals output by the antenna. When the transmitted signals encounter wet precipitant, part of the signal reflects back to the antenna. The receiver captures this signal return, supplying raw data containing range, azimuth, and intensity to the IAPS. The IAPS converts the information into video data to be presented on the AFDs. The RTA-800 weather radar system may be operated in a split mode, where the radar functions as two independent radars, each updating on alternate sweeps of the antenna. In this mode, each AFD controls one of the two radar channels.

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Figure 34-25. C406-2 Emergency Locator Transmitter

34-50

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EMERGENCY LOCATOR TRANSMITTER SYSTEM DESCRIPTION An Artex ELT C406-2 emergency locator transmitter (ELT) operates in a wide range of environmental conditions, enabling rescue teams to locate the aircraft in the event of a crash (Figure 34-25). The Artex C406-2 ELT system consists of: • Transmitter with integral battery pack and G-switch

A low-pass filter is on the ELT C406-2 support assembly and is installed between two coax connectors. The low-pass filter eliminates COMM 1 and COMM 2 radio transmission radiation from the ELT. A interconnect between the ELT and the GPS allow for transmission of position. An interconnect between the ELT and the RSS allow for transmission of the mode S identifier.

OPERATION The ELT C406-2 ELT system activates: • Automatically upon impact

• Low pass filter

• Manually with the G-switch

• Antenna

• Manually by either one of the two remotely mounted control switches

• Cockpit ELT control panel • NAV interface

The G-switch operates and activates the transmitter as a result of crash accelerations parallel to, or coincident with, the longitudinal axis of aircraft, moving generally in a forward direction.

• Cable assembly • Antenna coax cable The transmitter, along with its integral battery pack and G-switches, is in the tail cone. The system activates automatically in the event of aircraft impact or through the cockpit panel switch. When the aircraft electrical system is on, a microprocessor in the transmitter utilizes power from the aircraft electrical system. Power from the transmitter integral alkaline battery pack is utilized for the system test sequence and activating/sustaining the system in the event of an emergency. The ELT C406-2 system utilizes an antenna on the upper fuselage. The antenna is coaxial and is connected to the transmitter. Controlling devices for the system include a G-switch in the transmitter and a two-position cockpit panel switch on the right instrument panel. This allows the flightcrew to activate and reset or test the system. An ON–OFF toggle switch on the transmitter is positioned to ON for normal system operation and to OFF during maintenance or servicing.

The pilots can manually activate the transmitter by placing the remote switch on the right instrument panel to the ON position. When activated, the ELT transmits simultaneously on emergency frequencies 121.50, 243.00, and 406.025 MHz, utilizing a swept tone at three sweeps-per-second. The ELT C406-2 system also incorporates a complete self-analysis program with test routines transmitted at reduced power. The test sequence checks the system microprocessor, antennas, and transmitter. The test routine can be activated by the remote switch by placing the switch to the ON position for 1 second then returning the switch to the ARM position.

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Table 34-1. ACRONYMS ACRONYM

DESCRIPTION

ACRONYM

DESCRIPTION

Air Data Computer

FGS

Flight Guidance System

ADF

Automatic Direction Finder

FMS

Flight Management System

ADI

Attitude Direction Indicator

GPB

General Purpose Bus

ADS

Air Data System

AFD

Adaptive Flight Display

ADC

AFMS

GS

Aircraft Flight Manual Supplement

Ground Speed or GlideSlope

HIRF

High Intensity Radiated Fields

HSI

Horizontal Situation Indicator

IAPS

Integrated Avionics Processing System

AGL

Above Ground Level

AHC

Attitude Heading Computer

IAS

Indicated Airspeed

AHS

Attitude Heading System

ICC

Integrated Card Cage

Attitude Heading Reference System

IEC

Internal Environmental Control

ALI

Altitude Indicator

IOC

Input-Output Concentrator

ALT

Altimeter

LCD

Liquid Crystal Display

AOA

Angle Of Attack

LRM

Line Replaceable Module

AP

AutoPilot

LRU

Line Replaceable Unit

AP

DISC Autopilot Disconnect

MFD

Multifunctional Display

AutoPilot Panel

MSI

Mach Speed Indicator

Aeronautical Radio Incorporated

MSP

Mode Select Panel

ASI

Airspeed Indicator

NAV

Navigation

ATC

Air Traffic Control

NM

Nautical Mile

ATT

Attitude

PFD

Primary Function Display

Back Course

PWR

Power Supply

Control Display Unit

RAC

Radio Altimeter Converter

Commercial Standard Data Bus

RAD

ALT Radio Altimeter

CHP

Course Heading Panel

RIU

Radio Interface Unit

CKP

Course Knob Panel

RSS

Radio Sensor System

COM

Communications

RTA

Receiver-Transmitter-Antenna

CPA

Closest Point of Approach

RTU

Radio Tuning Unit

CSU

Configuration Strapping Unit

SDI

Source Destination Identifier

DCP

Display Control Panel

SSM

Sign Status Matrix Servo-Motor True Airspeed

AHRS

APP ARINC

B/C or BC CDU CSDB

DCU

Data Concentrator Unit

SVO

DME

Distance Measuring Equipment

TAS

ECU

External Compensation Unit

EFIS

Electronic Flight Information System

TCAS TDR

Traffic Collision Avoidance System Transponder

EDC

Engine Concentrator Unit

VS

EIS

Engine Indicating System

VSI

Vertical Speed Indicator

FD

Flight Director

VHF

Very High Frequency

FDU FGC

34-52

WXR

Flux Detector Unit Flight Guidance Computer

XNPDR

Vertical Speed

Weather Radar System Transponder

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 35 OXYGEN CONTENTS Page INTRODUCTION ................................................................................................................. 35-1 GENERAL ............................................................................................................................ 35-3 STORAGE SYSTEM ............................................................................................................ 35-5 Components ................................................................................................................... 35-5 Controls and Indications.............................................................................................. 35-11 Operation ..................................................................................................................... 35-13 Limitations................................................................................................................... 35-15 Diagnostics .................................................................................................................. 35-17

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ILLUSTRATIONS Figure

Title

Page

35-1

Oxygen System Schematic..................................................................................... 35-2

35-2

Oxygen Bottle Installation ..................................................................................... 35-4

35-3

Oxygen Regulator Details ...................................................................................... 35-6

35-4

Cabin Oxygen System and Oxygen Controls......................................................... 35-8

35-5

Oxygen Pressure Gauge....................................................................................... 35-10

35-6

Oxygen Control Valve Schematic ........................................................................ 35-12

35-7

Oxygen Filler Valve and Placard.......................................................................... 35-16

TABLE Table 35-1

Title

Page

Oxygen Filling Pressures ..................................................................................... 35-14

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CHAPTER 35 OXYGEN

INTRODUCTION This chapter presents information on the oxygen system for the Citation 525B aircraft, with the discussion limited to the storage and delivery of breathable oxygen into the cabin area. General maintenance considerations are included, with an introduction to functional and operational checks. References for this chapter and further information can be found in Chapter 5—“Time Limits/Maintenance Checks,” Chapter 12—“Servicing,” and Chapter 35—“Oxygen” of the Aircraft Maintenance Manual (AMM).

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

SOLENOID TO PASSENGER DISTRIBUTION SYSTEM 29 VDC ALTITUDE PRESSURE SWITCH TO COPILOT FACE MASK

MANUAL CONTROL VALVE (NORMAL POSITION)

PILOT FACE MASK

NORMAL CREW ONLY MANUAL DROP

SHUTOFF KNOB PRESSURE REGULATOR

OXYGEN PRESSURE OVERBOARD

CYLINDER PRESSURE GAGE OVERBOARD DISCHARGE INDICATOR

LEGEND HIGH-PRESSURE OXYGEN LOW-PRESSURE OXYGEN OXYGEN CYLINDER

MECHANICAL LINKAGE

FILLER VALVE AND PROTECTIVE CAP

Figure 35-1. Oxygen System Schematic

35-2

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GENERAL

NOTES

This chapter provides maintenance information on systems that store, regulate, and deliver oxygen for the occupants of the aircraft. The oxygen system for the Citation 525B aircraft consists of components that supply oxygen to the flight and passenger compartments: • Oxygen cylinder • Pressure regulator with shutoff valve • High-pressure relief valve • Oxygen filler valve • Masks • Passenger service units • Oxygen control valve • Altitude pressure sensor • Pressure gauge The oxygen system conforms to Military Specifications MIL-0-2710, Type I, Aviator’s Breathing Gaseous Oxygen. An additional publication addressing the oxygen system is Department of Transportation (DOT) 49 CFR171-190, Hazardous Materials Regulations o f t h e D e p a r t m e n t o f Tra n s p o r t a t i o n . Application for copies should be addressed to: Superintendent of Documents Government Printing Office Washington, D.C. 20402 A schematic of the oxygen system is illustrated in Figure 35-1.

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Figure 35-2. Oxygen Bottle Installation

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STORAGE SYSTEM COMPONENTS Oxygen Cylinder

be removed and overhauled by the manufacturer or an FAA-approved overhaul station. A ruptured green disc must be replaced when replacing the cylinder/regulator.

Oxygen Filler Valve

A 50-cubic-foot (1,407-liter) oxygen cylinder (Figure 35-2) is in the right side of the nose baggage compartment. The standard oxygen cylinder is a Type 3 FC 1850 cylinder manufactured under DOT-E 8162. Cylinder construction consists of continuous filaments of Kevlar ® impregnated with epoxy resin, wound longitudinally and circumferentially over a seamless aluminum liner.

The cylinder fill valve is remotely located from the cylinder/regulator on the bulkhead aft of and immediately inside the right nose baggage door. The fill valve has a check valve and filter. A pressure sealing cap prevents contaminants from entering the oxygen system.

NOTES

Pressure Regulator and Shutoff Valve T h e p r e s s u r e r eg u l a t o r i s a s i n g l e - s t a g e ON–OFF unit that delivers 70 ± 10 psi (483 ± 69 kPa) controlled outlet pressure with cylinder inlet pressure between 200–1,850 psig (1,379–12,755 kPa). The pressure regulator is threaded to the cylinder and confines the pressure to the cylinder itself. An integral low-pressure relief valve provides protection against overpressurization of the low-pressure system. When the regulator is in the OFF position, the low-pressure lines are vented to atmosphere. The regulator contains ports for fill, pressure gauge, rupture disc, and low-pressure lines. Safety wire holes lock the regulator in either the ON or OFF position.

High-Pressure Relief Valve and Overboard Discharge Indicator The pressure regulator incorporates overpressure protection by means of a rupture disc that opens when the cylinder pressure reaches 3,775 ± 150 psig at 70°F (26,028 ± 1,034 kPa at 21.1°C). The rupture disc fitting is lineconnected to an overboard port beneath the nose baggage compartment. The port is covered by a green disc that is secured by a snap ring. A missing green disc is an indication that the oxygen cylinder has discharged due to overtemperature or overpressure and must

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

MICROPHONE CORD

OXYGEN SUPPLY HOSE HARNESS INFLATION LEVER

N

100% PUSH

EMERGENCY POSITION

PRESS TO TEST

OXYGEN MASK (UNDERSIDE VIEW)

EMERGENCY ROTARY KNOB

Figure 35-3. Oxygen Regulator Details

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Crew Oxygen Outlet Connectors

NOTES

Oxygen outlet connectors for the pilots are identical bayonet-type connectors. One connector is on the pilot console and one on the copilot console. The crew oxygen mask bayonet hose fitting plugs directly into the connector and initiates oxygen flow through the connector. When the mask bayonet hose fitting is removed from the connector, the connector closes, preventing oxygen from escaping. A dust cover protects the connector when the mask is not plugged in.

Crew Masks A diluter-demand oxygen mask has an integral microphone and oxygen regulator. The oxygen regulator has a lever allowing manual selection of diluter demand, normal, or 100% oxygen. A flow indicator assures crew members that oxygen is being received. The oxygen mask has a purge button on the front o f t h e m a s k . W h e n t h e p u r g e bu t t o n i s pressed, the mask is cleared of contaminants if the mask is donned in an emergency smokefilled environment. The standard EROS ™ mask operates similar to the diluter-demand mask, except that mask emergency pressure is tested with a PRESSTO-TEST button. It can be changed to emergency by rotating the same button to the EMER setting.

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

NORMAL CREW ONLY

MANUAL DROP

FROM OXYGEN CYLINDER REGULATOR PASSENGER MASK WITH PINTLE PIN INSTALLED

MANIFOLD VALVE

PASSENGER DROP BOX

PASSENGER MASK WITH PINTLE PIN REMOVED TO ADDITIONAL DROP BOXES

LEGEND STATIC PRESSURE LOW-PRESSURE OXYGEN

Figure 35-4. Cabin Oxygen System and Oxygen Controls

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Passenger Oxygen Masks

NOTES

The passenger oxygen masks are a constantflow type enabling passengers to receive the same quantity of oxygen regardless of aircraft or cabin altitude. The flow rate is 4.5 liters per minute. The masks consists of: • Face plates • Economizer bags • Length of plastic tubings • Lanyard cords with pintle pins attached • Head straps The passenger oxygen mask stowage/dropout boxes are in the overhead consoles. A passenger oxygen mask is provided for each passenger seat. When the door actuator is energized, the stowage/dropout box opens, releasing the mask. Oxygen does not flow until the oxygen service valve an each mask is released.

Oxygen Service Valve An oxygen service valve is inside each oxygen mask container. The valve supplies oxygen flow to the respective mask when the lanyard cord is pulled, removing the pintle pin from the oxygen service valve. The oxygen service valve has a lever to shut off the flow without replacing the pintle pin.

Door Actuator A door actuator connects to the supply system line inside each mask stowage compartment. The door actuators unlock and open the mask container door, allowing the mask to drop. The actuator exerts a force of approximately 20 pounds at 70-psi (483 kPa) operating pressure.

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OXYGEN CONTROL VALVE

DISCONNECT B A T T E R Y NORMAL

CREW ONLY

NORMAL

MANUAL DROP

PILOT OUTLET VALVE OFF

MIC OXY MASK

EMER OXY

OXYGEN CONTROL VALVE

MIC OXY MASK

NORMAL INTERIOR MASTER

PHONE

PILOT SIDE CONSOLE

MIC HEAD SET

MIC OXY MASK

EMER OXY MIC

PHONE

MIC OXY MASK

HEADSET

COPILOT SIDE CONSOLE

10 OXY 1

15

0

PSI 1000

20

Figure 35-5. Oxygen Pressure Gauge

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MIC HEADSET

MIC HEAD SET

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CONTROLS AND INDICATIONS

Altitude Pressure Sensor

Pressure Gauge

An altitude pressure sensor on the pressurization printed circuit board (PCB) in the cabin senses cabin pressure. It opens the electrical solenoid at a pressure altitude of 14,500 ± 500 feet (4,419 ± 152 m). The sensor closes the electric solenoid at a minimum pressure altitude of 12,000 feet (3658 meters).

An oxygen pressure gauge is on the right side of the instrument panel in the flight compartment. The gauge is electrically illuminated and the range markings are: • Yellow arc ................................ 0–400 psi (0–2,758 kPa) • Green arc ..................... 1,600–1800 psi (11,032–12,411 kPa)

NOTES

• Red line .................................. 2,000 psi (13,790 kPa)

Oxygen Control Valve The oxygen control valve is on the pilot side console and has three positions: CREW ONLY, NORMAL, and MANUAL DROP. The oxygen control valve has an electrical solenoid for automatic operation and controls only the flow of oxygen to the passenger masks. The crew masks receive oxygen directly from the cylinder/ regulator. The three positions of the valve are as follows: • CREW ONLY position—Prevents oxygen from flowing into the cabin plumbing. • NORMAL position—Oxygen pressure is available to the solenoid-operated oxygen control valve. This control valve is normally closed and is actuated by the altitude pressure sensor. When the oxygen control valve is actuated to the open position, oxygen flows to the passenger oxygen system. • MANUAL DROP position—The solenoid-operated oxygen control valve is manually bypassed and oxygen pressure is directed to the passenger oxygen system. This valve is normally in the closed position.

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

SOLENOID TO PASSENGER DISTRIBUTION SYSTEM

NORMAL CREW ONLY

Figure 35-6. Oxygen Control Valve Schematic

35-12

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OPERATION The oxygen system has a high-pressure side and a low-pressure side. Oxygen is stored in the high-pressure side in a cylinder in the nose compartment (to the right of the wheel well). The low-pressure side is controlled by a pressure regulator that attaches directly to the oxygen cylinder. The regulator may be turned ON or OFF with the control handle. With the handle in the OFF position, the regulator functions as a shutoff valve and vents the low-pressure side of the system (supply to the cabin) internally through the regulator. With the handle in the ON position, the regulator provides a constant 70 ± 10 psi pressure to the low-pressure side of the system. The low-pressure side supplies oxygen to the crew outlets and to the oxygen dropout boxes in the cabin.

In the event of a failure, or at the discretion of the crew, the passenger system can be manually actuated by selecting the passenger oxygen select valve to MANUAL DROP. The MANUAL DROP position opens the passenger oxygen select valve and drops the passenger masks. After deployment of the passenger masks, oxygen flow is initiated by pulling the lanyards attached to the pintle pin. The control of flow into the passenger masks is achieved by a precision orifice between the supply line and the mask. When the passenger oxygen select valve is positioned to CREW ONLY, no oxygen flows to the cabin, but the crew system remains operational.

NOTES

The low-pressure side of the system can be split into two subsystems: one for the cabin and one for the cockpit. The oxygen control valve separates the two halves. Oxygen is always available to the crew through the crew distribution lines, but is blocked from entering the passenger system by the control valve. The OXYGEN CONTROL VALVE is a combination solenoid-operated or manual valve. Typically, the valve is left in the NORMAL position. While the valve is in NORMAL, if the cabin altitude increases to a pressure altitude of 14,500 ± 400 feet, a pressure altitude transducer applies electrical power to the solenoid of the passenger oxygen select valve. When the passenger oxygen select valve opens, it allows 70-psi pressure to flow into the cabin oxygen system. This pressure is sufficient to open the doors on the oxygen dropout boxes and deploy the masks.

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Table 35-1. OXYGEN FILLING PRESSURES STABILIZED TEMPERATURE (°F)

FILLING PRESSURE TO ACHIEVE 1,800 PSIG @ 70°F (PSIG)

–10

1,430

0

1,480

20

1,570

40

1,660

60

1,755

80

1,845

100

1,940

120

2,030

140

2,120

NOTE: TO EXTRAPOLATE ADD 5.0 PSIG FOR EACH 1°F.

35-14

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WARNING The oxygen cylinder is shipped with a partial-to-full charge. Care should be taken not to drop the cylinder or otherwise damage the cylinder or the pressure regulator.

WARNING

With the oxygen pressure gauge line connected to the regulator, the bayonet probe (welded on the end of the oxygen pressure gauge line) holds the needle valve in the open position. Therefore, the gauge indicates pressure whether the regulator is turned on or off. When the nut from the oxygen pressure gauge line is removed, the pressure closes (or seats) the needle valve, terminating oxygen flow.

The entire oxygen system must be kept free of moisture as the cooling produced by expansion of the compressed oxygen or low operating temperatures causes water to freeze in the small orifices of the system.

When installing the oxygen line into the regulator, insert the bayonet probe on the end of the oxygen pressure gauge line into disconnect. Tighten the nut from the oxygen pressure gauge line onto disconnect. As the nut is tightened, the bayonet probe opens (or unseats) the needle valve housed in the disconnect. Oxygen will flow to the oxygen pressure gauge.

Refer to Table 35-1 for oxygen filling pressures.

WARNING

LIMITATIONS Except for the purpose of the pressure regulator replacement, the oxygen cylinder must never be completely discharged. A completely discharged oxygen cylinder must be treated as if the regulator has been damaged and returned to the manufacturer or an FAA-approved overhaul station for reconditioning. This is due to the possibility of contaminants being drawn into the regulator by negative pressures within the cylinder caused by temperature variations.

Oxygen supports combustion. Materials that do not normally burn in the atmosphere, readily ignite or explode in the presence of concentrated oxygen.

Never remove fittings or check valves from any port of the pressure regulator when the oxygen cylinder is pressurized. The ports are pressurized at cylinder pressure (except the system distribution port, which is regulated pressure). Positioning the pressure regulator shutoff valve knob to OFF closes only the system distribution port. A missing green disc is an indication that the oxygen cylinder has discharged due to overtemperature or overpressure. The oxygen cylinder must be removed and overhauled by the manufacturer or an FAA-approved overhaul station. A ruptured green disc must be replaced when replacing the cylinder/regulator.

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OXYGEN FILLER AVIATORS BREATHING OXYGEN PER MIL-0-27210 SEE SERVICE MANUAL FOR SERVICING INSTRUCTIONS

CAP

FILLER VALVE

Figure 35-7. Oxygen Filler Valve and Placard

35-16

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DIAGNOSTICS

CAUTION

The oxygen filler valve is inside the right nose baggage door (Figure 35-7). Only breathing oxygen conforming to MIL-0-27210, Type 1, must be used for charging cylinders. If the oxygen bottle has been thermally discharged, the disc at the regulator and indicator must first be replaced before servicing can be accomplished.

Ambient temperature has a directeffect on indicated pressure; therefore, always refer to the oxygen fill chart in Chapter 12 of the AMM .

4. Shut off the oxygen supply and disconnect the charging cylinder. 5. Install the filler cap on the filler valve.

Servicing Precautions Ensure the following safety precautions are adhered to at all times:

NOTES

• Do not service the oxygen bottle while the aircraft is being fueled. • Ensure no unconfined flammable material is near when servicing the oxygen bottle. • Do not direct highly compressed oxygen towards personnel. • Gaseous oxygen containers under pressure must be given extra attention for cleanliness.

Servicing Procedures Before servicing the oxygen system to examine the cylinder for condition and hydrostatic test date. Both the filler valve and the servicing hose connection must be checked for contamination before servicing begins. Service the bottle as follows: 1. Remove the oxygen filler valve cap. 2. Connect the charging cylinder to the filler valve. 3. Slowly charge the oxygen cylinder to 1,800–1,850 psi.

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CHAPTER 36 PNEUMATIC CONTENTS Page INTRODUCTION ................................................................................................................. 36-1 GENERAL ............................................................................................................................ 36-3 DISTRIBUTION ................................................................................................................... 36-5 Components ................................................................................................................... 36-5 Operation ..................................................................................................................... 36-11 PNEUMATIC CABIN ENTRY DOOR SEAL ................................................................... 36-13 Description and Operation........................................................................................... 36-13

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ILLUSTRATIONS Figure

Title

Page

36-1

Pneumatic System Schematic ................................................................................ 36-2

36-2

Precooler Assembly................................................................................................ 36-4

36-3

Pneumatic Temperature Probe ............................................................................... 36-4

36-4

Bleed-Air Plumbing ............................................................................................... 36-6

36-5

23-psi Service Air Regulator .................................................................................. 36-8

36-6

Temperature Control Schematic .......................................................................... 36-10

36-7

Pneumatic Cabin Door Valve ............................................................................... 36-12

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 36 PNEUMATICS BLEED AIR CO

L

R AIR

5

15 20

LV VA E

INTRODUCTION This chapter presents information on the pneumatic system (Figure 36-1) for the Citation 525B aircraft. Discussion in this chapter is limited to the delivery and control of bleed air into the tail cone area. Each user system is covered in detail within the appropriate chapter of this training manual. General maintenance considerations are included, with an introduction to functional and operational checks. For more information on the pneumatic system, refer to Chapter 36—“Pneumatics” in the Aircraft Maintenance Manual (AMM).

FOR TRAINING PURPOSES ONLY

36-1

36-2 BILEVEL FLOW CONTROL VALVE

FOR TRAINING PURPOSES ONLY

TO CABIN DISTRIBUTION SYSTEM

CABIN HEAT EXCHANGER

COMPRESSION AIR PRSOV

MUFFLER DUCT OVERHEAT SENSOR

TO WINDSHIELD ANTI-ICE

DUCT TEMPERATURE SENSOR FRESH-AIR FAN

SERVICE AIR REGULATOR

PNEUMATIC TEMPERATURE PROBE

TO WING ANTI-ICE MANIFOLD

OVERTEMPERATURE SWITCH

PRECOOLER

RAM-AIR VALVE

RAM COOLING AIR

FAN BYPASS COOLING AIR ENGINE BLEED AIR

Figure 36-1. Pneumatic System Schematic

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHECK VALVE

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

GENERAL

NOTES

This chapter provides maintenance information on assemblies that extract, regulate, and distribute bleed air. High-temperature, primary bleed air is extracted from the high-pressure compressor. The bleed air is routed from the nacelle through the pylon, where the temperature is modified by the pylon heat exchanger. Plumbing in the tail cone routes the bleed air into the bleed-air distribution systems. Systems included in this chapter that utilize pneumatic pressure (service air) are precooler controls and the cabin entrance door seal.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

36-4

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

DISTRIBUTION

Maximum Cool Bypass Valve

COMPONENTS Precoolers Precoolers are steel, cross-flow heat exchangers in the pylons, which convert hot engine bleed air (up to 900°F [482°C]) to a temperature suitable for use in the aircraft anti-ice, environmental, and pneumatic systems (Figure 36-2). Precoolers are the primary means of regulating the temperature of the bleed air going to the wing anti-ice system. Precoolers have two fundamental paths: the bleed-air flow path (hot) and the engine fanair flow path (cool). Cooling air is provided by extracting fan bypass air from the engine casing. An intake vent for cooling airflow is on the inboard side of the engine, forward of the nozzle. The air exits the aircraft through the gap between the engine cowling and engine exhaust nozzle.

The maximum cool bypass valve is in the plumbing between the 23-psi regulator and the precooler actuator. The valve is a normally closed, solenoid-operated device and is in parallel with the pneumatic temperature probe. It allows 23-psi air pressure to bypass the pneumatic temperature probe and open the fan air valve when: • Wing anti-ice is selected off • Air-conditioning compressor is on • Aircraft altitude is lower than 30,000 feet The maximum cool bypass valve is an electrically actuated valve. In the absence of electrical power, the maximum cool valve remains closed.

Barometric Sensor

The precooler temperature control system cools hot engine bleed air to a temperature of 500°F.

The barometric sensor is in the empennage above the aft baggage compartment. When it senses aircraft altitude higher than 30,000 feet, it deenergizes the maximum cool bypass valve closed.

Pneumatic Temperature Probe

Precooler Actuator

The pneumatic temperature probe is a pressureregulating valve (Figure 36-3). The set point varies such that when the temperature increases, the pressure-regulating set point decreases, allowing more air to flow through the pneumatic temperature probe. The pneumatic temperature probe has dissimilar materials to ensure that, as the pneumatic temperature probe approaches a set point temperature, pressure regulation relaxes sufficiently to allow greater quantities of air to pass.

A bleed-air temperature control line connects the outlet of the pneumatic temperature probe to the precooler actuator. The precooler actuator is a pneumatic-cylinder configuration spring-loaded in the retracted position. The spring-loading on the precooler actuator holds the door in the closed position and is opened by air supplied by the pneumatic temperature probe. By varying the pressure between the pneumatic temperature probe and the precooler actuator, the precooler door assembly may be regulated to the open and closed positions.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

RH MANIFOLD ASSY

CROSS FITTING CHECK VALVES

TEMPERATURE PROBES

LH MANIFOLD ASSY

Figure 36-4. Bleed-Air Plumbing

36-6

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Overtemperature Switch

Cross Fitting

The overtemperature switch is an electrical switch with contacts that close at 560°F bleedair temperature. The switch is in the bleed-air plumbing between the pneumatic temperature probe and the manifold (Figure 36-4).

The cross fitting accepts bleed-air pressure from either or both operating engines (Figure 36-4). On the aft side of the cross fitting, bleed air connects to a capped pressurization service air port that is used for testing the pneumatic system(s). Additionally, the aft side of the cross fitting supplies bleed air to the service air pressure regulator. The forward side of the cross fitting supplies bleed air to the windshield anti-ice system.

Manifold Assemblies Ram air passes through a precooler, over a bleed-air overtemperature switch, then forward to the aft side of the manifold assembly. Left and right manifold assemblies are in the tail cone. Engine bleed air is routed from the nacelles to the respective manifold assembly. Each manifold assembly provides three individual outlets:

NOTES

• Outboard outlet—Provides bleed air routed forward to the wing pressureregulating and shutoff valves (PRSOVs). The air then goes through a tube assembly, to a cross-flow shutoff valve over the undertemperature switch. The air is then distributed along the wing leading-edge assembly. • Inboard outlet—Provides bleed air to the PRSOVs, into the flow control shutoff valves. Air from the flow control shutoff valves and a fresh-air blower provide air to the cabin distribution system. • Manifold outlet—Plumbing for connection through a check valve to a cross fitting providing bleed-air distribution for the windshield anti-ice system and the 23-psi regulator.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

D

FW

CROSS FITTING

REGULATOR

Figure 36-5. 23-psi Service Air Regulator

36-8

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

23-psi Service Air Regulator

NOTES

The aft side of the cross fitting supplies bleed air to the service air pressure regulator (Figure 36-5), regulating bleed air to 23 ± 5 psig (159 ± 34 kPa). Operational check of the service air pressure regulating shutoff and relief valves is done by connecting a calibrated air pressure gauge to the output side and applying air pressure to the input side. Shop air may be connected to the ground service port to operationally troubleshoot the entrance door seal and deice system. Engine-operated bleed air may be used to operationally troubleshoot all bleed-air systems. However, engine-operated bleed-air pressure is required to operationally troubleshoot the windshield bleed-air system, airfoil anti-ice systems, and/or the engine inlet anti-ice system.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

OUTER TUBE MATERIAL EXPANDS FREELY WITH HEAT PRECOOLER TEMPERATURE CONTROL BYPASS SOLENOID

INNER ROD MATERIAL EXPANDS SLIGHTLY

ORIFICE PNEUMATIC PROBE PNEUMATIC ACTUATOR SPRING-LOADED CLOSED

LEGEND 23-PSI REGULATED BLEED AIR CONTROL PRESSURE

Figure 36-6. Temperature Control Schematic

36-10

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

OPERATION

NOTES

Air from the compressor section of the engine is directed to the precooler. This air is unregulated and very hot, ranging 3 0 0 ° F – 9 0 0 ° F. B l e e d - a i r p r e s s u r e va r i e s 16–300 psi. The bleed-air temperature at the exit of the precoolers is controlled to between 420–500°F by modulating the amount of cooling air that passes across the precooler. This modulation is accomplished with butterfly valves regulating the flow of fan air through the cooling air portion of the precoolers. The actuator on the fan air modulating valve is a pneumatic piston/cylinder arrangement. The valve assembly is spring-loaded in the closed position. Control of the system is by the service air circuit in the form of 23-psi air. The 23-psi air is routed in two paths to the fan air modulating valve. The service air is routed to the pneumatic probe and to the maximum cool bypass valve. No electrical signal is required for the pneumatic probe. The maximum cool bypass valve is actuated when the wing and engine anti-ice systems are off, the air conditioning is running, and the tail cone environment is below 30,000 feet. With the maximum cool bypass valve actuated, the control pressure is routed to the fan air modulating valve, rotating the modulating valve to its fully open position. This provides maximum cooling of the bleed air. The barometric switch prevents the bypass valve from activating (opening) above 30,000 feet altitude.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

PRESSURE SWITCH

DOOR SEAL CONTROL VALVE

23-PSI SERVICE AIR

CHECK VALVE

Figure 36-7. Pneumatic Cabin Door Valve

36-12

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

PNEUMATIC CABIN ENTRY DOOR SEAL DESCRIPTION AND OPERATION

A cabin entry door seal pressure switch monitors the cabin door seal pressure. When pressure exceeds 8.5 psi, the primary door seal pressure switch extinguishes the DOOR SEAL annunciator. As pressure decreases to 5.5 psi, the switch opens and the annunciator illuminates.

The cabin entry door seal is inflated with engine bleed air and consists of the following components:

NOTES

• Check valve • Pressure regulated and relief valve • Spring-loaded door seal control valve • Inflatable door seal • Pneumatic distribution lines and fittings A pneumatic line connects at the pressure regulator and routes through the upper aft pressure bulkhead, along the fuselage floor, to an orifice tee, across to a check valve, and finally to the cabin entry door seal valve (Figure 36-7). The cabin entry door seal control valve is at the forward door frame. The lower forward door lock pin actuates the cabin entry door seal control valve in the door-locked position, allowing bleed air to inflate the seal. A check valve is in the input line to the door seal control valve, preventing the loss of inflation pressure in the event of pneumatic air source loss. When the door is unlocked, the spring-loaded valve deactivates, closing off bleed-air pressure and allowing trapped air in the seal to deflate through a vent in the valve body. When the cabin entry door is opened, the cabin door position switch indicates the condition by illuminating the DOOR NOT LOCKED annunciator.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 38 WATER/WASTE CONTENTS Page INTRODUCTION ................................................................................................................. 38-1 GENERAL ............................................................................................................................ 38-1 POTABLE WATER ............................................................................................................... 38-3 Description..................................................................................................................... 38-3 Components ................................................................................................................... 38-3 Operation ....................................................................................................................... 38-3 WATER/WASTE DISPOSAL ............................................................................................... 38-3 Components ................................................................................................................... 38-3 Operation ....................................................................................................................... 38-5 OPTIONAL FLUSH TOILET............................................................................................... 38-6 Description..................................................................................................................... 38-6 Operation ....................................................................................................................... 38-6

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

ILLUSTRATIONS Figure

Title

Page

38-1

Toilet....................................................................................................................... 38-2

38-2

Toilet Component Locations .................................................................................. 38-4

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 38 WATER/WASTE

INTRODUCTION This chapter presents information on water and waste systems for the cabin area of the Citation 525B aircraft. Information is provided on fixed units and components that store and deliver fresh water for use, as well as fixed components that store and remove water and waste. References for this chapter and further specific information can be found in Chapter 5—“Time Limits/Maintenance Checks,” Chapter 12—“Servicing,” and Chapter 38—“Water/Waste” of the Aircraft Maintenance Manual (AMM).

GENERAL Maintenance information for potable water and waste disposal for the Citation 525B aircraft may vary depending on interior arrangement and options in equipment description and installation. The potable water section describes provisions for drinking water that is available through a refreshment center in the forward cabin.

The waste disposal section describes a nonflush-type toilet and optional flush-type toilet in the aft cabin on the right side. The flush toilet is self-contained and requires a 29-VDC e l e c t r i c a l p ow e r s o u r c e f o r o p e r a t i o n . Information is also described in this section on the relief tube.

FOR TRAINING PURPOSES ONLY

38-1

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

RIGHT AFT DIVIDER

TOILET TISSUE DISPENSER

RELIEF HORN ACCESS DOOR

DRAIN TUBE

DOOR

HEATED DRAIN

TOILET TANK ASSEMBLY

TOILET STRUCTURE

Figure 38-1. Toilet

38-2

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

POTABLE WATER

Relief Tubes

DESCRIPTION

The cockpit relief tube is under the pilot seat on the forward side.

Potable water containers are in the right forward refreshment center.

COMPONENTS An ice container is included in the standard and optional refreshment centers.

The aft relief tube is a standard installation on the aft toilet. The assembly is on the aft side of the toilet structure. On the optional flush toilet, access to the relief tube is through an access door on top of the toilet structure. The relief tubes vent overboard through heated vents.

The right forward refreshment center has an electrically heated stainless steel water container for hot water.

NOTES

OPERATION The right forward refreshment center water container is equipped with a pushbutton drain valve. The valve assembly is spring-loaded in the closed position. A cap on top is removed for filling. The hot water container has a receptacle connecting to a mating plug at the back of the shelf. When the container is in place, the circuit is completed for heat to be activated by a switch on the front of the cabinet.

WATER/WASTE DISPOSAL COMPONENTS Refreshment Center Drain The right forward refreshment center has a drain system for the drip pan below the water dispenser and a drain from the ice container to the outside of the aircraft through a heated outlet.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

LEFT AFT DIVIDER

D

LIFE VEST/ RELIEF HORN STORAGE TOILET TISSUE DRAWER

A

LAP BELT ASSEMBLY

B STORAGE ASSEMBLY

DRAIN HOSE

TOILET TANK ASSEMBLY

HOSE DOOR HEATED DRAIN ASSEMBLY

C DETAIL A

DETAIL B

BOLT

WASHER

SHOULDER RESTRAINT ASSEMBLY

FRAME ASSEMBLY

DETAIL C

DETAIL D

Figure 38-2. Toilet Component Locations

38-4

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Toilets

NOTES

The standard toilet assembly is on the right side of the aft cabin directly under the escape hatch. Waste from the toilet is stored in an integral holding tank between servicing. The right flush toilet is in the aft cabin, aft of the divider on the right side of aircraft. The relief horn and associated plumbing are on the left side of the toilet structure. The belted flush toilet is in the aft cabin, aft of the divider on the left side of aircraft. The relief horn and associated plumbing are on the right side of the toilet structure. Shoulder restraint assemblies are on the forward and aft intercostals with webbing protruding through the reveal and shroud, aft and above the belted toilet. A lap belt assembly is on the toilet structure assembly. A placard with servicing instructions is on the side of the toilet assembly. Waste from the relief tube is drained overboard through the bottom of the aircraft.

OPERATION Water that accumulates in the ice chest (right forward refreshment center) is disposed of by turning a ball valve beneath the ice container to the OPEN position and allowing the water to drain overboard through a heated outlet. Operation of the standard nonflush toilet is by raising and lowering the toilet lid and ring. Servicing is accomplished by removing the holding bucket and replacing the liner bag.

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

OPTIONAL FLUSH TOILET DESCRIPTION A recirculating flushing toilet is optional equipment on the Citation 525B aircraft. The toilet is a completely self-contained unit, requiring only the external connection of 29VDC electrical power. The toilet assembly is designed for permanent installation in the aircraft, requiring only the removal of the holding tank when servicing is desired. The toilet assembly consists of a seat and shroud assembly resting on the toilet mounting plate. The polished stainless steel bowl assembly, the motor and pump assembly, and the PRESS TO FLUSH switch are attached to the mounting plate. The slide assembly is on the bottom flange of the bowl in which the removable holding tank assembly is installed.

A shoulder harness restraint above the toilet structure assembly on the left reveal and a lap belt assembly on the toilet structure assembly are available with the left belted toilet.

OPERATION The flush cycle is initiated by depressing the PRESS TO FLUSH switch on the seat and shroud assembly. The switch applies 29-VDC power to the motor and pump assembly. Flushing continues until the switch is released. During the flush cycle, flushing fluid is pumped from the holding tank to the bowl by the self-priming pump. The flush fluid enters the bowl through a nozzle in the upper rim and washes the inner surface of the bowl in a swirling pattern. Waste is carried to the holding tank through the knife valve below the bowl. When desired, the holding tank may be removed from the toilet for servicing after closing the knife valve.

The removable holding tank assembly consists of a storage tank with the following components: • Knife valve • Flush line quick-disconnect • Carrying handle Extending through the cover of the knife valve is a manually operated actuator to open or close the knife valve, sealing the tank contents prior to removal. The position of the knife valve may be observed through the opening at the bottom of the bowl. The holding tank assembly detaches from the toilet at the front of the unit. Two press-lock fasteners (one on each side of the knife valve) secure the installed tank in the sealed position against the bottom of the bowl. By detaching and draining the flush line at the quick-disconnect, depressing the two Press-Lock fasteners, and pulling the carrying handle, the tank is easily removed for servicing.

38-6

FOR TRAINING PURPOSES ONLY

The information normally contained in this chapter is not applicable to this particular aircraft.

The information normally contained in this chapter is not applicable to this particular aircraft.

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 51-57 STRUCTURE CONTENTS Page INTRODUCTION ................................................................................................................. 51-1 GENERAL ............................................................................................................................ 51-1 STRUCTURES...................................................................................................................... 51-3 Components ................................................................................................................... 51-3 DOORS.................................................................................................................................. 52-1 Description..................................................................................................................... 52-3 Components ................................................................................................................... 52-3 FUSELAGE........................................................................................................................... 53-1 Description..................................................................................................................... 53-3 Components ................................................................................................................... 53-5 PYLONS................................................................................................................................ 54-1 Description..................................................................................................................... 54-3 STABILIZERS ...................................................................................................................... 55-1 Description..................................................................................................................... 55-3 WINDOWS ........................................................................................................................... 56-1 Description..................................................................................................................... 56-3 Components ................................................................................................................... 56-3 Diagnostics..................................................................................................................... 56-5 WINGS .................................................................................................................................. 57-1 Description and Operation............................................................................................. 57-3

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

ILLUSTRATIONS Figure

Title

Page

52-1

Inflatable Door Seal ............................................................................................... 52-2

52-2

Door Warning Switch............................................................................................. 52-4

52-3

Entry Door Step Assembly .................................................................................... 52-6

52-4

Escape Hatch.......................................................................................................... 52-8

52-5

Nose Baggage Door ............................................................................................. 52-10

52-6

Hydraulic Service Door ....................................................................................... 52-12

53-1

Tail Stinger Assembly ............................................................................................ 53-2

53-2

Nose Radome Assembly ........................................................................................ 53-4

54-1

Pylon Assembly...................................................................................................... 54-2

55-1

Horizontal Stabilizer Assembly ............................................................................. 55-2

55-2

Elevator Assembly.................................................................................................. 55-4

55-3

Rudder Assembly ................................................................................................... 55-6

56-1

Side Window Assembly ......................................................................................... 56-2

56-2

Ultrasonic Test Equipment..................................................................................... 56-4

56-3

Cabin Window Assembly....................................................................................... 56-6

57-1

Wing Assembly ...................................................................................................... 57-2

57-2

Wing Tip Assembly................................................................................................ 57-4

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 51-57 STRUCTURE

INTRODUCTION This chapter describes the structure of the Citation 525B aircraft. Further specific information can be found in Chapter 52—“Doors,” Chapter 53—“Fuselage,” Chapter 54— “Nacelles/Pylons,” Chapter 55—“Stabilizers,” Chapter 56—“Windows,” and Chapter 57—“Wings” in the Aircraft Maintenance Manual (AMM).

GENERAL The Citation 525B aircraft is a twin turbofan, pressurized aircraft with a cabin that holds six to seven passengers and space for luggage and optional equipment. The doors section contains maintenance procedures for the cabin entry door, emergency exit door, nose access doors, tail cone baggage door, external electrical and hydraulic power receptacle doors, the refuel/defuel panel access door, and tail cone access door. The fuselage section provides information about the equipment, passengers, crew and baggage compartments.

The nacelles and pylons section contains information on the structure of the pylons. The stabilizer sections describes the horizontal stabilizer, elevator, vertical stabilizer, and the rudder. The window section provides information on windows in the flight and passenger compartments, including the procedure to examine and repair any window damage. The wing section describes the wing assembly and wing tips as well as the components that make up the wings.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Figure 51-1. Citation 525B Airplane

51-2

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

STRUCTURES

A low wing is found aft on the fuselage and is usual on aircraft with rear attached engines.

COMPONENTS

The wing structure is a semimonocoque design. The wing structure has spars, stringers, ribs, and skin.

Airframe Refer to Figure 51-1. The main landing gear retract inboard into the wing and the nose landing gear retracts forward into the fuselage nose section.

The forward spar is at 9% chord, the center spar is at 31% chord, and the rear spar is at 65% chord. The spars are almost the same in structure and include cap extrusions, stiffeners, and webs.

An entrance door is forward of the wing on the left side. An emergency exit is in the aft cabin area, on the right side of the fuselage. The nose baggage compartment doors are attached to each side of the nose.

The wing is three parts made of the left wing, center wing, and right wing. The wing has a contour that goes under the fuselage and attaches to the fuselage with special bolts and fittings. The wings have integral fuel cells.

The nose section is not pressurized. The cabin section of the fuselage is pressurized.

NOTES

The nose section has a typical bulkhead, stringers, and skin construction. It has sufficient space for luggage and avionics. The radome and forward section of the nose is removable to give access to the avionics equipment. The windshields and adjacent structure are bird-proof to meet FAR 25 requirements. The tail cone is a skin and stringers with a nacelle carry-through structure to support the engines. The tail cone baggage compartment door is used to access the tail cone luggage compartment. The tail cone is not pressurized. The empennage structure is a semimonocoque design. The empennage structure has spars, stringers, ribs, and skin.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 52 DOORS

FOR TRAINING PURPOSES ONLY

52-1

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

NOTE 1: BOND THE PRIMARY DOOR SEAL AROUND THE COMPLETE EDGE OF THE PRIMARY DOOR SEAL WITH RTV157. NOTE 2: IF THE DISTANCE IS MORE THAN 0.34 ± 0.03 INCH (8.64 ± 0.76 mm), ADD BONDTITE TO THE FUSELAGE DOOR FRAME AS NECESSARY

A BOND AREA

0.06 INCH (0.152 MM) INTERFERENCE

PRIMARY DOOR SEAL (NOTE 1)

0.06 INCH (0.152 MM) MINIMUM GAP

RETAINING ANGLE

DETAIL B

SKIN DOUBLER

CHANNEL

CLAMP

PRIMARY INFLATABLE SEAL

RAIN SEAL SECONDARY DOOR SEAL

A SKIN

A

SKIN DOUBLER SECONDARY SEAL

B

VIEW A-A

0.34 ± 0.03 INCH (8.64 ± 0.76 mm) (NOTE 2)

DETAIL A

Figure 52-1. Inflatable Door Seal

52-2

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

DESCRIPTION This section contains the maintenance procedures for the: • Cabin entry door and step • Emergency exit door • Cargo doors

The cabin entry door can be opened from the inside or the outside of the aircraft. Both handles are kept in the stowed position by springs. When one of the two handles is turned, lock pins are retracted into the door. This allows the door to move outward to the open position. A latch on the door frame keeps the door in this position.

• Service doors The cabin entry door section provides a general description of the entry door and it's operation, troubleshooting instruction, and maintenance practices. An entry door step is attached to the cabin entry door frame for passenger and crew access to the aircraft.

NOTES

The emergency exit door section gives a general description and operation of the emergency exit door and its latch mechanism. The cargo doors section gives a general description of the nose and aft baggage doors. The service doors section gives a general description of the hydraulic service door and the electrical power receptacle door.

COMPONENTS Cabin Entry Door The cabin entry door is on the left side of the fuselage at the forward side of the passenger compartment. The cabin entry door is flush with the external skin of the aircraft and includes a window. The inner surface is covered with upholstery panels. The cabin entry door is assembled with frames, pin fittings, and stiffeners. The cabin entry door is attached to the fuselage structure by a single hinge. The cabin entry door uses an inflatable seal installed in a retainer on the door (Figure 52-1). Bleed air from the engine gives pressure to inflate the seal. A bleed-air fitting is attached to the fuselage in the hinge area. In the closed position, a fitting in the hinge connects with the fuselage fitting and air is supplied to the pressure seal.

FOR TRAINING PURPOSES ONLY

52-3

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

A

A

NUT

NUT

LOCK WASHER LOCK WASHER BAGGAGE DOOR STRUCTURE

CAM LOCK SLEEVE

STRIKER SWITCH WASHER

WASHER NUT

SCREW SCREW

WASHER NUT

DETAIL A

Figure 52-2. Door Warning Switch

52-4

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

The cabin entry door latch mechanism has an inner and outer handle, shaft, rocker arm, drive assemblies, pushrod assemblies, rod ends, idlers, clevises, links, pins, bell cranks, spacers, and indicators. The latch mechanism is assembled with pins, washers, cotter pins, bolts, and nuts.

bleed-air valve operates and removes air from the inflatable door seal. The cabin entry door can then be pushed outboard to the full-open position. At the full-open position, the door stop lug attaches to the door stop latch assembly. The spring-loaded door stop latch assembly is engaged to hold the door in the open position.

The cabin entry door lock mechanism includes: • Channels • Bell crank attach angles and brackets— Give the structure support for the bell cranks. • Doublers • Bearing plate support angles • Fillers • Spacers • Latch pin fittings—Attach to the channels at the top and bottom of the door. The forward and aft latch pins are supported by the pin support channels. • Bearing plates—Help the lock pins go into the fuselage door frame sockets. Attach to the bearing plate doublers, pin support channels, or latch pin fittings. There are five indicator windows in the cabin entry door. The indicator windows are on the inside of the door to give a visual indication of the lock mechanism position. There is a window by the two top and bottom locking pins and one window is by the inside door handle.

Push the outside handle on the end of the PUSH handle and the opposite end of the handle lifts out of the socket. Turn the handle clockwise to open the cabin entry door from the outside. To close the cabin entry door, put the entry steps inside the aircraft. Push the door stop lever to release the door stop lug. A door assist chain helps close the door from inside the aircraft. With the door closed, the inside handle is turned clockwise to engage the door lock pins into the fuselage door sockets. When the door lock pins are engaged, the bleed-air valve operates to supply bleed air to the inflatable seal. The door warning switch (SC011) is operated and the CABIN DOOR annunciator extinguishes. The handle is returned to the stowed position. There are four dots on a set of green bars that are aligned in the indicator windows when the door is latched.

A door warning switch (SC011) is in the fuselage door frame. The forward lower locking pin operates the switch (Figure 52-2).

Cabin Entry Door Operation The cabin entry door is opened from the inside with a trigger, which releases the handle from the striker plate. When the handle is turned counterclockwise the lock mechanism retracts the locking pins from the fuselage door frame sockets. The warning switch (SC011) operates and completes a ground circuit that makes the CABIN DOOR annunciator illuminate. A

FOR TRAINING PURPOSES ONLY

52-5

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

SCREW

C

HINGE

A HINGE PINS CHAIN

B

EYEBOLT NUT

SCREW

WASHER SPACERS

HINGE SCREW

EYEBOLT NUT

SHIMS BOLT

DETAIL A

DOOR FRAME SCREW

SPRING PLUNGER

DETAIL C

WASHER

SPRING PLUNGER CLIP

SHIMS NUTPLATE

DETAIL B

Figure 52-3. Entry Door Step Assembly

52-6

STEP ASSEMBLY

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Entry Door Step

NOTES

An entry door step assembly is attached to the fuselage door frame at the cabin entry door in four places. The two upper attach points are cable supports and the two lower attach points are hinges. A bumper assembly is on the forward step frame to prevent damage to the step assembly if the door is accidentally closed when the step assembly is in the down position. The cabin entry door is opened and the step assembly is lowered to operate the entry door step. The step assembly then turns on a pivot to the down position with the hinges. A chain is pulled up to put the step assembly in the stowed position. The step assembly turns on a pivot on its hinges to the stowed position. The detent lever on the door stop is part of the latch and the arm mates with a spring clip to hold the step assembly in the stowed position. When the cabin entry door is closed, the step assembly mates with the cutout in the door.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

B

DOOR FRAME

LATCH STRIKE PLATE

C

DOOR OUTER HANDLE

DETAIL A

PIN LATCH PIN

COTTER PIN SUPPORT ASSEMBLY COVER

CAM RETURN SPRING

INNER HANDLE BOLT

LATCH CAM

GUARD ZEE SUPPORT CAP NUT LOCKING

INNER FRAME NUT SEAL

DETAIL B

SCREW RETAINER

SCREW

DETAIL C

Figure 52-4. Escape Hatch

52-8

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Emergency Exit Door The emergency exit door is on the right side of the fuselage at the aft end of the passenger compartment. The door can be fully removed, and is installed from the interior of the aircraft (Figure 52-4). The emergency exit door has frames, doublers, stiffeners, and an outside skin panel. A rubber pressure seal is around the edge of the hatch. The emergency exit door is locked in position by a latch pin at the top and two retainers at the bottom of the hatch. The latch pin is operated by the inner and outer door handles. The retainers at the bottom of the door are in a locked position. The retainers are placed behind two stop blocks on the emergency exit door frame. The inner surfaces of the emergency exit door handle and housing are coated with Teflon to prevent the buildup of ice. The outer handle housing has a moisture drain.

Emergency Escape Hatch Latch Mechanism A latch mechanism is on the upper half of the emergency exit door. The mechanism has an outer door handle assembly with an attached internal spindle, latch cam, latch pin, cam return spring, and inner door handle assembly. The latch cam has a square key slot, through which the outer door handle spindle is installed. The inner door handle assembly attaches to the outer door handle spindle with a fastener and a nut. The latch pin is attached to the latch cam, which moves to the open or closed position when the emergency exit door handles are turned.

When the emergency exit door is installed, the outer door handle is flush with the door skin and the inner door handle folds into a recess in the interior door panel.

Emergency Exit Door Operation The inner or outer door handle assemblies are turned to disengage the latch pin and open the emergency exit door. The locking pin and plastic cover must be removed from the inner door handle assembly to open the emergency exit door from inside the aircraft. The top of the door is moved inside the aircraft until the emergency exit door clears the door frame. The emergency exit door is raised to clear the stop blocks and remove the door. The inner door handle is turned to the open position to install the emergency exit door from inside the aircraft. The bottom of the door is put into the door frame, and the retainers on the bottom of the door behind the stop blocks. Align the retainer slots in the guides at the stop blocks. Push the top of the door into the opening. Align the door in the door frame and the top of the door within 2 inches of the closed position. Shut the door tight and turn the inner door handle into the closed position. Install the plastic cover and locking pin.

In the closed position the latch pin goes into a cutout in the door frame and engages on a latch striker plate. In this position the cam return spring keeps a constant spring-load on the latch pin.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

DETAIL A

Figure 52-5. Nose Baggage Door

52-10

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Cargo Doors

NOTES

Nose Baggage Doors The nose baggage doors are on the left and right sides of the upper nose cone (Figure 525). The doors are attached to the nose cone by two hinges. Two latch assemblies hold the nose baggage doors in the closed position. The doors are held in the opened position by a gas spring on the aft hinge assembly. The doors are locked by a key lock when closed. To open the nose baggage doors, release the two latch assemblies and lift the doors upward. In the open position, the BAGGAGE DOOR– FWD annunciator on the pilot instrument panel illuminates. The nose baggage doors are held open by a gas spring on the aft hinge assembly. The doors are pulled down and the latch assemblies engage to close the nose baggage doors. When closed, the BAGGAGE DOOR– FWD annunciator extinguishes.

Aft Baggage Door The aft baggage door is on the lower left side of the aircraft tail cone. The door is held closed by two latch assemblies on the top and bottom of the door. The door opens on a dual hinge assembly, which is on the inside forward frame of the door panel. The aft baggage door is held in the opened position by a gas spring. A key lock is used to lock the door. To open the aft baggage door, release the lock to disengage the two latch assemblies and pull the door out. In the open position, the BAGGAGE DOOR–AFT annunciator on the pilot instrument panel illuminates. To close the aft baggage door, push in to the closed position to engage the two latches and the lock the door. The BAGGAGE DOOR–AFT annunciator extinguishes.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

FAIRING PANEL STRIKER

SCREW BONDING JUMPER

BOLT CLIP HYDRAULIC SERVICE DOOR

WASHER BUSHING LATCH

WASHER NUT NUT

WASHER

COTTER PIN

NUT HI-LOK

HINGE ASSEMBLY

DETAIL A VIEW LOOKING OUTBOOARD

Figure 52-6. Hydraulic Service Door

52-12

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Service Doors

NOTES

Hydraulic Service Door The hydraulic service door is on the right tail cone fairing (Figure 52-6). The door provides access to the hydraulic reservoir and hydraulic ground service quick-disconnect receptacles. To open the door, release the two pin latches.

Electrical Power Receptacle Door The electrical power receptacle door is on the left lower side of the tail cone. The door provides access to connect an external electrical power source to the aircraft. The electrical power receptacle door is spring-loaded to keep it in the closed position.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 53 FUSELAGE

FOR TRAINING PURPOSES ONLY

53-1

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

STATIC WICK MOUNTING BASE

WEDGE COVER

TAIL CONE STINGER

WASHER SCREW

DETAIL A

Figure 53-1. Tail Stinger Assembly

53-2

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

DESCRIPTION

NOTES

This chapter provides information about the compartments for equipment, passengers, crew and baggage. The fuselage consists of the: • Forward frame—Nose compartment, frames, stringers, radome, nose wheel well, and nose baggage compartment panels • Main frame—Fuselage primary structure, including the forward pressure bulkhead, aft pressure bulkhead, frames, stringers, and drains • Main frame interior—Fuselage secondary structure, including the floor panels, seat rails, control pedestal, and side consoles • Aft frame—Stringers, engine carry through beams, tail stinger, and aft baggage compartment panels (Figure 53-1).

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

BRACKET

NUTPLATE

BOLT

DETAIL B ROTARY LATCH

A

LIGHTNING DIVERTER STRIP

B

GUIDE PIN WASHER

NUT

EYEBOLT WASHER

NUTPLATE

NUT WEATHER EROSION BOOT

RADOME

WASHER SCREW

DETAIL A

Figure 53-2. Nose Radome Assembly

53-4

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

COMPONENTS

NOTES

Nose Radome The nose radome is attached to the nose of the aircraft. It is shaped for the best anti-ice function and incorporates lightning diverter strips (Figure 53-2). The nose radome is a bonded assembly of prepreg epoxy glass and honeycomb, which uses an adhesive film autoclave bonding, corrosion inhibitor adhesive primer, and room temperature bonding. Guide-pin assemblies align the nose radome with the aircraft nose frame for installation. Rotary latches connect to adjustable eyebolts to attach the nose radome to the aircraft. Two seals ensure a weather tight installation. An outer tubular seal is on the nose radome and an inner seal is on the nose frame. Both are attached with RTV. The surface area of the nose radome is finished with an anti-P static conductive layer before the last finish. The lightning diverter strips and a weather erosion boot are installed after the last finish.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 54 PYLONS

FOR TRAINING PURPOSES ONLY

54-1

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

B

PYLON SKIN FIRE EXTINGUISHING

ELECTRICAL

PRECOOLER ELECTRICAL FUEL HYDRAULICS

RAM AIR INLET

DETAIL A (LEFT SIDE SHOWN, RIGHT SIDE OPPOSITE)

Figure 54-1. Pylon Assembly

54-2

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

DESCRIPTION

NOTES

A pylon is on the left and right side of the tail cone (Figure 54-1). A forward and an aft engine beam go through the fuselage and give the structure for the pylons. Engine mount attach fittings are on the forward and aft engine beams. The pylons contain: • Fuel lines • Hydraulic lines • Engine control cables • Ram-air duct • Bleed-air duct • Bleed-air precooler • Fire-extinguishing agent deployment tubes

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 55 STABILIZERS

FOR TRAINING PURPOSES ONLY

55-1

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

ELEVATOR HORN

ELEVATOR PUSHROD

HORIZONTAL STABILIZER DOUBLER HORIZONTAL STABILIZER BOLT DOUBLER

COTTER PIN NUT WASHER

BOLT

ELEVATOR PUSHROD

WASHER ELEVATOR SECTOR

NUT COTTER PIN

DETAIL A

Figure 55-1. Horizontal Stabilizer Assembly

55-2

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

DESCRIPTION

NOTES

Horizontal Stabilizer The horizontal stabilizer is mounted to the vertical stabilizer with bolts, nuts, and cotter pins in a T-tail configuration (Figure 55-1). The horizontal stabilizer is a semimonocoque structure with the leading edges covered with deicing rubber boots. Five vortex generators are on each side of the vertical fin just under the horizontal stabilizer, one above the other, just in front of the rudder. The vortex generators stabilize airflow across the elevators. The horizontal stabilizer has a 0° dihedral, a sweep of 20°, 0 minutes along the 25° chord line.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A B

C DETAIL A

D E

WASHER ELEVATOR HORN WASHER NUT COTTER PIN

SCREW BOLT ELEVATOR STRUCTURE

BONDING JUMPER

COTTER PIN

ELEVATOR ASSEMBLY

SCREW

TORQUE TUBE

WING BOLT WASHER WASHER

NUT

BONDING JUMPER

ATTACH FITTING

DETAIL E

BALANCE WEIGHT ELEVATOR STRUCTURE

WING BOLT

COTTER PIN

DETAIL B ELEVATOR STRUCTURE

WASHER

WING BOLT COTTER PIN

NUT ATTACH FITTING

ATTACH FITTING

DETAIL D WASHER NUT

DETAIL C

Figure 55-2. Elevator Assembly

55-4

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Elevator

NOTES

The elevators are hinged to the trailing edge of the horizontal stabilizer and connected by torque tubes to the control column (Figure 55-2). Removal of the elevator is necessary only for replacement or repair of a damaged elevator, or inspection of elevator and horizontal stabilizer.

Vertical Stabilizer The vertical stabilizer, horizontal stabilizer, and associated structure make up the empennage. The vertical stabilizer is attached to the empennage structure with bolts, nuts, and cotter pins. The vertical fin is a semimonocoque structure and has a sweep back of 49°, 0 minutes along the 25° cord line.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

D

A

C RUDDER HINGE BOLT

B DETAIL A WASHER

COTTER PIN

NUT BOLT SHIM RUDDER BIAS ARM

CLAMP WASHER

HINGE BOLT

NUT

COTTER PIN BONDING JUMPER WASHER COTTER PIN NUT

DETAIL C

DETAIL B

Figure 55-3. Rudder Assembly

55-6

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Rudder

NOTES

The rudder is a movable airfoil hinged to the vertical stabilizer rear spar (Figure 55-3). Removal of the rudder is necessary only for replacement or repair of a damaged rudder, and inspection of the rudder and vertical stabilizer.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 56 WINDOWS

FOR TRAINING PURPOSES ONLY

56-1

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

RETAINER REMOVAL TOOL

A

RETAINER WINDSHIELD

VIEW A-A

RETAINER REMOVAL TOOL

RETAINER REMOVAL TOOL

A

A

Figure 56-1. Side Window Assembly

56-2

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

DESCRIPTION

NOTES

This section gives the procedure to examine and repair any window damage.

COMPONENTS Flight Compartment Windows The flight compartment windows include the windshield and cockpit side windows. The windshield is laminated with an outer pane of stretched acrylic, two vinyl layers, and an inner pane of stretched acrylic. The pilot side windows are aft of the windshield on each side of the fuselage (Figure 561). Each window consists of a 0.675 inch prestressed acrylic outer pane and two 0.025 inch prestressed polyvinyl frost panes.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

FOUR EQUALLY SPACED MS20426 RIVETS

RADIUS END TO REMOVE SHARP CORNERS

3.00 INCHES (76.2 MM)

10.00 TO 14.00 INCHES (254.0 TO 355.6 MM)

HANDLE 0.187 INCH (4.70)

MATERIAL: 0.050 INCH FULL HARD STAINLESS STEEL-PHENOLIC BLOCK HANDLE

Figure 56-2. Ultrasonic Test Equipment

56-4

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

DIAGNOSTICS

c. The thinnest point of the inner ply is usually not at the same location or area as the thinnest point of the outer ply.

Damage Analysis 1. Use ultrasonic test equipment to measure the window ply(s) thickness in a minimum of three areas within 0.5 inch (12.70 mm) of the damage (Figure 56-2). 2. Use an optical depth micrometer to find the maximum depth of the damage. 3. Subtract the measured depth of the damage from the measured ply(s) thickness to find the remaining window ply(s) thickness. 4. Replace the window if the values are less than the minimum window thickness as shown in Table 801 of the AMM. 5. Complete the inspection that follows as written to make sure that the inspection agrees with all thickness measurements if the test equipment or test equipment operator changes. a. Calibrate the test equipment on material with the same sound velocity as the material in the windshields to be examined.

NOTE The minimum thickness standards shown in Table 801 of the AMM are for the view area only.

Windshield Rework Criteria Crack Cracks are not repairable. Replace the windshield(s) if: • The cracks extend beyond the retainer into the vision area. • There are cracks in three or more adjacent attachment holes.

NOTE If windshield replacement is not necessary, but lesser cracks are in the windshield, it is necessary to change the windshield inspection intervals to every 300 hours or 24 months, whichever occurs first.

Pit, Chip, Gouge, Scratch, or Crazing

NOTE The CJMD156-101 calibration standard is made from the same type of acrylic as the Cessna Citation aircraft acrylic windshield and cabin windows. b. Place the CJMD156-101 calibration standard on or adjacent to the windshield or windows to be inspected for 30 minutes before the calibration of the thickness gage. This allows the reference standard to become the same temperature as the aircraft windshield or window.

Replace the windshield(s) if: • A pit, chip, gouge, scratch, or crazing occurs at the attachment holes of the windshield and causes a depth more than 0.080 inches after a repair of the area. • The depth of the repaired damage is 0.080 inch or less, but greater than 0.010 inch.

NOTE A pit, chip, gouge, scratch, or crazing that is .010 inch or less in depth is serviceable for further flight operations with no repairs. It is necessary to measure and examine the damage at the next windshield phase inspection.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

WINDOW RETAINER

SKIN SEAL

A A A

HI-LOK PIN (NOTE)

FROST PANE

B DETAIL A CAUTION: MAKE SURE THE WINDOW IS FLAT ON THE SKIN AND IS IN THE CENTER OF THE WINDOW OPENING.

WASHER NUT

WINDOW ASSEMBLY

FILLET SEAL, TYPE 1, CLASS B SEALANT WASHER

NOTE: INSTALL THE HI-LOK PIN WET WITH TYPE 1, CLASS A SEALANT. HI-LOK PIN (NOTE) SKIN RETAINER NUT TORQUE TO 2O ± 2 IN-LBS (2.26 ±0 .23 N.m)

FROST PANE

SEAL

VIEW A-A TYPICAL CROSS SECTION

Figure 56-3. Cabin Window Assembly

56-6

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Delamination

Discoloration

Replace the windshield(s) if:

Discoloration does not have an effect on the strength of the windshield and cannot be repaired. If vision is decreased, replace the window.

• There is delamination of the windshield in the critical vision area. • There is delamination of the windshield outside the critical vision area more than 0.75 inch in diameter (0.442 square inches) at an individual location.

Flight Compartment Side Windows Rework Criteria Crack

Distortion Minor optical distortion outside the critical viewing area is permitted. It is necessary to repair the distortion in the critical viewing area.

Replace the window if: • The cracks extend beyond the retainer into the vision area. • There are cracks in three or more adjacent attachment holes.

Discoloration Discoloration does not have an effect on the strength of the windshield and is unrepairable. If vision is decreased, replace the windshield.

Cabin Window Rework Criteria Refer to Figure 56-3 for cabin window assembly components.

NOTE If it is not necessary to replace the window, but if lesser cracks are in the window, it is necessary to change the window inspection intervals to ev e r y 3 0 0 h o u r s o r 2 4 m o n t h s , whichever occurs first.

Crack Replace the window if it has a crack.

Pit, Chip, Gouge, Scratch, or Crazing It is necessary to repair or replace any windowpanes with damage that is more than 0.015 inch in depth.

Delamination If more than 10% of the cabin window area has delamination, it is necessary to replace window.

Distortion Minor optical distortion is permitted.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Pit, Chip, Gouge, Scratch, or Crazing It is necessary to repair or replace any damage in the windowpanes greater than 0.010 inch in depth.

dows, remove only the minimum material necessary to repair the damage (Figure 56-3). Before starting the repair, always determine if the repair meets the criteria listed in the AMM.

NOTE

NOTE

If a window replacement is not necessary, but lesser pits, chips, or gouges are repaired, it is necessary to change the window inspection intervals to every 300 hours or 24 months, whichever occurs first.

It is recommended to contract a company that repairs acrylic windows, to complete the repair procedures.

Replace the window if:

If a scratch repair kit with instructions is available, use the repair kit instructions. If a repair kit is not available, use the procedure in the AMM

• A pit, chip, or gouge that occurs at the attachment holes of the windshield is more than 0.050 inches in depth after rework. • There is a pit, chip, or gouge in three or more adjacent attachment holes.

Delamination If there is delamination in more than 10% of the window area, it is necessary to replace cabin windows.

Distortion Minor optical distortion is permitted.

Discoloration Discoloration does not have an effect on the strength of the windshield and cannot be repaired. If the vision is decreased, replace the windshield.

Approved Repair Repair Procedures Approved repairs for acrylic windows are those permitted for the removal of small damage. This is to make the visibility and/or appearance better and to remove stress concentration points. When repairing win-

56-8

FOR TRAINING PURPOSES ONLY

NOTES

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 57 WINGS

FOR TRAINING PURPOSES ONLY

57-1

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

SPEEDBRAKE RETRACT TUBE HYDRAULIC MANIFOLD

FLAP EXTEND TUBE FLAP RETRACT TUBE

FLAP RETRACT TUBE

SPEEDBRAKE EXTEND TUBE FLAP EXTEND TUBE AUXILIARY LANDING GEAR PNEUMATIC TUBE RIGHT BRAKE PNEUMATIC TUBE

SPEEDBRAKE EXTEND TUBE

FLAP EXTEND TUBE

LANDING GEAR RETRACT TUBE LANDING GEAR EXTEND TUBE

LEFT BRAKE PNEUMATIC TUBE

DETAIL A

Figure 57-1. Wing Assembly

57-2

FOR TRAINING PURPOSES ONLY

FLAP EXTEND TUBE

SPEEDBRAKE RETRACT TUBE

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

DESCRIPTION AND OPERATION

NOTES

Wing Assembly The wing assembly is a single unit that attaches under the fuselage with four attach fittings, one locator bolt, and a yaw plate (Figure 57-1). The entire wing removes and installs as one unit. The wing assembly components make up a one-piece wing with carry-through spars that extend from wing tip to wing tip. The wing spars, (forward, main, and aft) consist of upper and lower extruded spar caps joined by sheet spar webs and extruded angles. Longitudinal support of the wing is provided by ribs constructed of extruded tees, formed caps, and sheet webs. Wing attach fittings consist of four links that connect attach fitting points from the wing to the fuselage. The attach points on the wing are connected to the forward and aft spars and protrude upward through the top skin. The attach fittings on the fuselage mount in the cabin between double bulkheads and protrude downward through the cabin skin to mate with the wing. A locator bolt is positioned on the center line of the aircraft and is installed through the bottom of the cabin to a locator fitting on the top of the wing at the forward spar. An aluminum alloy yaw plate attaches from the top surface of the aft wing skin to the fuselage. The yaw plate prevents longitudinal or lateral movement of the wing assembly.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

AILERON

STATIC WICK SCREW

ANTICOLLISION LIGHT NAVIGATION LIGHT

WING

WING TIP

DETAIL A

Figure 57-2. Wing Tip Assembly

57-4

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Wing Tips

NOTES

The wing tips enclose the outboard ends of the wings and consist of ribs covered with aluminum alloy skin (Figure 57-2). Navigation and anticollision/strobe lights are in the wing tips. An access plate on the underside of the wing tip permits access to the strobe light inverters. The wing tips are attached to the structure with screws. The refueling electrical ground receptacle is on the bottom of the wing tips. Wing tips do not contain fuel. A vented barrier between the wing tips and wings, and a sealed rib inboard of the wing tips prevents fuel in the wing tips. The wing tips have a louvered vent on the bottom that allows bleed air from the leading edge anti-ice tube to escape through a lightening hole and vent overboard.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 71–80 POWERPLANT CONTENTS Page INTRODUCTION ................................................................................................................. 71-1 GENERAL ............................................................................................................................ 71-3 ENGINES .............................................................................................................................. 71-5 Description..................................................................................................................... 71-5 Components ................................................................................................................... 71-7 ENGINE FUEL AND CONTROL ........................................................................................ 73-1 Description..................................................................................................................... 73-3 Components ................................................................................................................... 73-5 Controls and Indications.............................................................................................. 73-13 Operation ..................................................................................................................... 73-19 IGNITION ............................................................................................................................. 74-1 Description..................................................................................................................... 74-3 Components ................................................................................................................... 74-3 ENGINE INDICATING ........................................................................................................ 77-1 Description..................................................................................................................... 77-3 Components ................................................................................................................... 77-3 Operation ....................................................................................................................... 77-3 OIL ...................................................................................................................................... 79-1 Description..................................................................................................................... 79-3 Components ................................................................................................................... 79-3 Controls and Indications ................................................................................................ 79-7 Operation ..................................................................................................................... 79-11 Diagnostic.................................................................................................................... 79-11

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

ILLUSTRATIONS Figure

Title

Page

71-1

FJ44-3A Engine Nacelles....................................................................................... 71-2

71-2

FJ44-3A Engine—Left Side View.......................................................................... 71-4

71-3

FJ44-3A Engine—Right Side View ....................................................................... 71-4

71-4

Bypass Duct Group ................................................................................................ 71-6

71-5

Bypass Valve Lever Installation ............................................................................. 71-8

71-6

Accessory Gearbox Installation ........................................................................... 71-10

73-1

Engine Fuel System Schematic.............................................................................. 73-2

73-2

Fuel Filter Installation ............................................................................................ 73-4

73-3

Fuel Delivery Unit Installation............................................................................... 73-6

73-4

Fuel Metering Unit................................................................................................. 73-8

73-5

Fuel Shutoff at FDU............................................................................................. 73-10

73-6

FADEC Assembly ................................................................................................ 73-12

73-7

Engine Electrical Harness .................................................................................... 73-14

73-8

Throttle Lever Angle (TLA) Positions ................................................................ 73-16

73-9

FADEC Electrical Schematic............................................................................... 73-18

73-10

TLD Fault Indications.......................................................................................... 73-20

74-1

Ignition Switch Panel ............................................................................................. 74-2

77-1

Multifunction Display ............................................................................................ 77-2

77-2

Engine Turbine Speed Indicating (N2) System ...................................................... 77-4

77-3

Interturbine Temperature Indicating (ITT) System ................................................ 77-6

79-1

Oil System Schematic ............................................................................................ 79-2

79-2

Oil Pressure Regulator ........................................................................................... 79-4

79-3

Oil Sensor Installation............................................................................................ 79-6

79-4

Oil Temperature Indicator ...................................................................................... 79-8

79-5

Oil Pressure vs. Oil Temperature Graph .............................................................. 79-10

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 71–80 POWERPLANT

INTRODUCTION This chapter describes the Williams/Rolls Royce FJ44-3A powerplants on Citation 525B aircraft. Sections of this chapter include engines, fuel, ignition, engine indicating, and oil. General maintenance considerations are included in each section, accompanied by functional and operational checks. References for this chapter and further information can be found in Chapter 5—“Time Limits/Maintenance Checks,” Chapter 12—“Servicing,” Chapter 54—“Nacelle/Pylons,” Chapter 71—“Powerplant,” Chapter 73—“Engine Fuel and Control,” Chapter 74—“Ignition,” Chapter 76—“Engine Controls,” Chapter 77— “Engine Indicating,” Chapter 78—“Exhaust,” Chapter 79—“Oil,” and Chapter 80— “Starting” in the Aircraft Maintenance Manual (AMM).

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

OIL INSPECTION DOOR

QUARTER-TURN FASTENER

UPPER COWL DOOR

SEAL

AFT COWL

NACELLE INLET ASSEMBLY

LOWER COWL DOOR

SEAL

INSPECTION DOOR

DETAIL A

Figure 71-1. FJ44-3A Engine Nacelles

71-2

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

GENERAL The Williams International FJ44-3A engine (Figure 71-1) is a two-spool, coronating, axial flow turbofan engine with a medium bypass ratio, mixed exhaust, and high compressor pressure ratio. The engine produces a minimum of 2,780 pounds of takeoff thrust at sea level, flat rated to an ambient temperature of 72°F. Engine dry weight is approximately 525 pounds.

An AGB, driven through a towershaft from the high-pressure spool, supplies power to drive aircraft accessories. Pressurization, cabin heating, windshield defogging, and anti-ice are supplied through ports that deliver high-pressure bleed air. The engine is mounted to the aircraft at three locations, two on the front and one on the rear of the engine.

Other important engine features include:

NOTES

• Dual, airframe-mounted, fully redundant full authority digital engine controls (FADECs) • An accessory gearbox (AGB) with mounting pads for engine-driven accessories • Bleed-air system provisions for aircraft services • Three mounting point locations (two front and one rear) for engine support structure Engine control is provided by FADEC units. Thrust is managed through power lever input to the FADECs, giving an input to the fuel delivery unit (FDU). The FDU is mounted on and driven by the engine gearbox. FADECs are fully redundant. The low-pressure spool of the engine has a three-stage, low-aspect-ratio, foreign-objecttolerant fan with integral blades. The fan is followed by a single-stage axial intermediate pressure compressor in the gas generator flow path. The fan and intermediate pressure compressor are directly driven by two axial flow inserted blade turbine rotors. The high-pressure spool consists of a singlestage, high-pressure ratio centrifugal compressor driven by one uncooled axial turbine that has replaceable blades. A folded annular combustor is incorporated, fed by a rotating fuel slinger that atomizes and delivers fuel uniformly to the primary combustion zone.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

ELECTRONIC PT2/TT2 SENSOR (A/F INLET MOUNTED)

TT2/PT2 HEATER CONNECTOR EXHAUST NOZZLE MOUNTING FLANGE

AFT MOUNT RING

OIL FILL PORT (EITHER SIDE)

OIL SIGHT GLASS (EITHER SIDE) EXHAUST MIXER GEARBOX INLET MOUNTING FLANGE

IGNITER PLUGS (2) FUEL FILTER IMPENDING BYPASS SWITCH

FUEL INLET PORT

FUEL DELIVERY UNIT (FDU) FUEL FILTER

FUEL FILTER TEMPERATURE PORT FUEL SUPPLY PRESSURE PORT

Figure 71-2. FJ44-3A Engine—Left Side View AFT ENGINE ELECTRICAL HARNESS CONNECTION (P7) CHANNEL B (EITHER SIDE)

EXCITER POWER INPUT CONNECTORS (2)

FORWARD MOUNTS (2 PER SIDE)

FAN BYPASS BLEED PORT (EITHER SIDE)

ITT THERMOCOUPLES (6) FORWARD ENGINE ELECTRICAL HARNESS CONNECTION (P4) CHANNEL A (EITHER SIDE)

LP TURBINE OVERSPEED TRIP LP TURBINE DRAIN IGNITER PLUG (BOTH SIDES) CUSTOMER BLEED PORT (BOTH SIDES)

FDU OVERBOARD DRAIN

OIL FILTER IMPENDING GEARBOX MAGNETIC BYPASS INDICATOR OR ELECTRONIC (OPTIONAL) CHIP COLLECTOR

Figure 71-3. FJ44-3A Engine—Right Side View

71-4

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

ENGINES

low-pressure turbine (LPT) noz° First zle, including the No. 3 and No. 4 roller bearings and seals.

DESCRIPTION

• The LP turbine group consists of:

The FJ44-3A engine has the following three compressor sections: • Low-pressure (LP) axial compressor • Intermediate-pressure (IP) axial compressor • High-pressure compressor

(HP)

centrifugal

The LP compressor (fan) and IP compressor are driven by two LP turbines. The HP compressor is driven by a single HP turbine. When the mixture of fuel and compressed air is ignited in the combustor section, the expanding gases drive the turbines.

turbine module (first-stage LP tur° LP bine rotor, second-stage LP turbine nozzle assembly, second-stage LP turbine rotor)

° Rear housing ° Heat exchanger ° Rear case with exhaust mixer • Gearbox module and engine-mounted accessories • Airframe-mounted FADEC and TT2/PT2 sensor

The FJ44-3A engine (Figures 71-2 and 71-3) is comprised of six distinct groups:

NOTES

• The LP shaft module consists of the LP shaft, No. 1 and No. 1.5 bearing supports, No. 1 ball bearing, No. 1.5 roller bearing, and No. 1 carbon seal. • The fan group consists of the spinner, fan rotor, fan housing, fan stator, three-stage IP compressor, and IP stator stages. • The core module is made up of: housing with integral oil ° Interstage tank and first reduction bevel gear high-pressure compressor (HPC) ° The and compressor cover

° ° ° ° ° ° °

HP shaft Pinion gear and No. 2 ball bearing Diffuser assembly Combustor cover assembly Fuel manifold and seal assembly Fuel slinger and seal HP turbine nozzle and primary plate assembly

° HP turbine

FOR TRAINING PURPOSES ONLY

71-5

71-6 REAR BYPASS DUCT ASSEMBLY

FOR TRAINING PURPOSES ONLY

FRONT BYPASS DUCT ASSEMBLY

ITT PROBE PORT (6)

LP TRIP SENSOR HOUSING

REAR MOUNT BALL SOCKET

REAR MOUNT RING

Figure 71-4. Bypass Duct Group

FUEL DRAIN PORT

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

LIFTING SLING MOUNTS

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

COMPONENTS

Rear Mount Ring

Bypass Duct Group The bypass principle allows a portion of the air passing through the fan to bypass the IP compressor, HP compressor, combustor section, and HP and LP turbines. This permits the engine to use high-cyclic temperatures and pressures to produce a low jet exhaust velocity.

The rear mount ring is between the front and rear bypass duct assemblies. The rear mount ball socket changes location depending on which side of the aircraft the engine is on.

NOTES

Decreasing the velocity and temperature of the exhaust gases creates high thermal and propulsive efficiency. In addition, the bypass air decreases the noise level and increases the power-weight ratio for a given engine thrust.

Front Assembly The front bypass duct assembly (Figure 71-4) is formed from sheet aluminum and is seamwelded. Six access holes for interstage turbine temperature (ITT) probes are toward the rear flange. Spacers, retainers, plates, and retaining rings seal the thermocouple bosses. The front duct also houses the removable service adapters, which supply access ports for fuel, HP bleed air, and igniter plugs. A boss is on the front duct for mounting of the fuel nozzle. Fuel drain ports are at the bottom of the duct.

Rear Assembly The rear bypass duct assembly is also formed from sheet aluminum and is seam-welded. The assembly has a fuel drain port on the bottom of the duct. The rear duct contains a trip lever housing and cable attachment as part of the LP trip system that shuts down the engine in the event of shaft separation. The rear duct has a bracket for external mounting of the start nozzle control valve. On some engines, a fan bleed port is also provided.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

IP BLEED VALVE LEVER

ROD END BEARING

CABLE ASSEMBLY

Figure 71-5. Bypass Valve Lever Installation

71-8

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Compressor Sections The engine has three compressor stages: • LP • IP

The turbine compressor sections operate as follows: • The LP turbines drive the fan and IP compressors. • T h e H P t u r b i n e d r iv e s t h e H P compressor.

• HP These stages supply compressed air for combustion. Air supplied by the compressor section is also used for: • Sealing air for oil and labyrinth seals • Cooling air for LP and HP turbine discs. The rotation of the compressor and turbine assemblies moves air into the engine. The diameters of the compressors and stators narrow from front to rear, increasing the pressure and velocity of the airflow.

• The LP shaft passes through the HP shaft and connects the LP turbines with the fan and IP compressor. • The HP shaft connects the HP turbine with the HP compressor. Both shafts rotate in the same direction but not at the same speed (HP shaft speed is slightly more than twice that of the LP shaft).

NOTES

The airflow travels from the fan (LP compressor) to the IP compressor. The splitter fairing, which is part of the IP stator, divides the airflow into core air and bypass air. Core air is routed to the HP compressor and then into the diffuser. The diffused air then enters the combustor section, where it mixes with fuel and burns. Bypass air is ducted around the core and mixes with the exhaust.

Inducer Bleed System The engine has an inducer bleed system to improve the transient characteristics of the engine. This system is driven by the FDU in response to actuation commands from the FADEC. A bleed valve actuation cable connects the FDU to a butterfly bleed valve (Figure 71-5) that vents core flow from the inducer of the HP compressor to bypass.

FOR TRAINING PURPOSES ONLY

71-9

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

GEARBOX ASSEMBLY

STARTER/GENERATOR MOUNTING

HYDRAULIC PUMP MOUNTING

IFCU/FDU MOUNTING OIL FILTER LUBE AND SCAVENGE PUMP MOUNTING

Figure 71-6. Accessory Gearbox Installation

71-10

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Accessory Gearbox Assembly

NOTES

The gearbox provides mounting and drive provisions for the accessories. Oil-wetted shafts provide lubrication for accessory shafts. The gearbox provides internal passages for the engine oil system. The gearbox assembly (Figure 71-6) connects to the interstage housing and is driven off the HP shaft by an accessory drive shaft. The gearbox provides rotational power to four mount pads for engine accessories. These accessories are: • Starter-generator • FDU • Lube/scavenge pump • Hydraulic pump The starter-generator pad is on the forward end of the gearbox. The lube and scavenge pump pad is directly aft of the starter-generator pad. Drive for the starter-generator and lube/scavenge pump is provided by the same drive gear. The fuel control unit pad is adjacent to the oil pad. The hydraulic pad is also forward of and adjacent to the starter-generator pad. The shaft drives are oiled and carbon seals are provided at each mount pad (except the lube/scavenge pump mount pad). The lube/ scavenge pump mount pad uses internal passages to transfer oil to and from the gearbox assembly. The gearbox receives its rotational power from the HP shaft via an accessory drive shaft. Two reduction bevel gears (one at the top and one at the bottom of the accessory drive shaft), reduce the HP shaft rotational speed for accessory operation.

FOR TRAINING PURPOSES ONLY

71-11

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CHAPTER 73 ENGINE FUEL AND CONTROL

FOR TRAINING PURPOSES ONLY

73-1

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

TO AIRCRAFT

MOTIVE FLOW TO AIRCRAFT ELECTRICAL CONNECTION TO FADEC CHANNEL B

ELECTRICAL CONNECTION TO FADEC CHANNEL A

P INDICATOR FILTER BYPASS VALVE

FUEL HEATER

FUEL DELIVERY UNIT

FILTER 20 MICRON ABSOLUTE

LUBE SYSTEM

FUEL INLET FROM AIRCRAFT

GEAR PUMP

PMA BOOST STAGE

GEARBOX DRIVE

FUEL METERING AND CONTROL FUNCTIONS

FUEL FLOW METER START NOZZLE

FUEL SLINGER

200 MICRON ABSOLUTE FILTER

FUEL MANIFOLD

LEGEND AIRFRAME FUEL INLET BOOST PUMP PRESSURE TO PMA PUMP DISCHARGE TO ENGINE FUEL SLINGER TO START NOZZLE MOTIVE FLOW TO AIRFRAME

Figure 73-1. Engine Fuel System Schematic

73-2

FOR TRAINING PURPOSES ONLY

N2 SPEED

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

DESCRIPTION

NOTES

The fuel system (Figure 73-1) is made up of the following components: • Fuel filter • Fuel filter bypass valve • FDU • Fuel manifold • Fuel slinger • Fuel/oil heat exchanger (FOHE) • Fuel nozzle Fuel enters the FDU, where it is pressurized and routed to the FOHE. The fuel is heated (and oil cooled) to prevent icing prior to returning to the fuel filter element. After filtering, fuel is moved by the positive displacement pump to the fuel metering section of the FDU. Metered fuel is split as it leaves the FDU as follows: • Approximately 9 pph is routed to an atomizing nozzle positioned in the burner to enhance cold-day and high-altitude start capability. • The remainder of metered flow passes through the fuel flow meter (airframe provided) and then enters the core engine through the fuel manifold tube. A last-chance filter is in the manifold tube. The manifold tube follows the contour of the diffuser to a fuel distribution manifold. From there, the fuel is supplied to the underside of the fuel slinger. The fuel slinger, rotating with the HP rotary group, ejects the fuel radially through a series of delivery holes into the combustion zone and is ignited.

FOR TRAINING PURPOSES ONLY

73-3

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

FUEL FILTER

FILTER BOWL

DETAIL A

Figure 73-2. Fuel Filter Installation

73-4

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

COMPONENTS

NOTES

Fuel Filter The fuel filter assembly (Figure 73-2) is a replaceable filter element in a stainless steel bowl threaded into the pump housing. The filter is between the boost pump and the gear pump to protect the gear pump from any fuel tank contaminants. The fuel filter element is 20 microns.

Delta P Switch The fuel filter electrical indicator (delta P switch) is on the FDU. In the event that the fuel filter becomes blocked, a filter bypass valve (adjacent to the filter) allows fuel to bypass the fuel filter for continued engine operation. Fuel filter bypass does not occur below a filter differential pressure of 10 psid. An electrical fuel filter differential pressure switch senses the pressure drop across the fuel filter and signals impending bypass (4 ± 0.5 psid) to a warning indicator in the cockpit. The indication of impending bypass is removed when the fuel filter differential pressure drops below 2 ± 1 psid.

Fuel/Oil Heat Exchanger The FJ44-3A engine fuel system uses an FOHE to provide heated fuel to the inlet of the fuel filter element. Ports are ahead of the fuel filter element to allow fuel to be routed to and from the FOHE. Providing warm fuel to the filter precludes the use of an anti-icing additives.

FOR TRAINING PURPOSES ONLY

73-5

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

FUEL SUPPLY MOTIVE SUPPLY HYDRAULIC PRESSURE HYDRAULIC RETURN

LP SHAFT SEPARATION DETECTION DEVICE

FILTER ASSEMBLY

FUEL DELIVERY UNIT

DETAIL A

Figure 73-3. Fuel Delivery Unit Installation

73-6

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Fuel Delivery Unit The FDU (Figure 73-3) is engine-driven through the FDU-to-gearbox attachment and includes the following components: • Engine fuel pump • Main engine fuel filter • Metering components • Permanent magnet alternator (PMA) The FDU is mounted on the engine gearbox and provides fuel metering and transient bleed valve actuation. The FDU also includes the following features: • PMA electrical power generation to the FADEC • Fuel conditioning (heating) • Fuel shutoff in the event of an LP shaft separation

The FDU requires pressurized fuel to function. Pressurization at high altitudes and near-idle power settings is provided by a pressurizing valve. The valve also splits the metered flow between the start nozzle and the engine manifold and provides positive fuel shutoff for both flow paths. Fuel flow measurements taken from between the FDU and the engine manifold must always account for start nozzle flow. An integral dual-coil solenoid fuel shutoff valve causes fuel flow to be terminated when power is removed from the solenoid. The shutoff valve also includes a mechanical shutdown function that is actuated by an LP (or fan) shaft separation mechanism attached to the exterior of the FDU. In the event of an LP shaft separation, the aft movement of the LP turbine rotors triggers a cable release to actuate the mechanical shutdown function.

The fuel pump consists of a centrifugal boost element and an HP positive displacement element and operate as follows:

NOTES

• The boost stage provides charging pressure to the gear stage inlet to accommodate low inlet pressure and suction feed conditions. • The HP gear stage delivers fuel at the required pressures for FDU and engine operation. An HP relief valve is in the FDU, protecting the HP pump element (and other fuel system components) from overpressurization by limiting total system pressure. If HP relief is required, fuel flow from the HP element is bypassed directly back to its inlet (boost pressure). Fuel from the HP pump element is sent to the fuel metering valve. The metering valve schedules all fuel flow from minimum to maximum fuel flow levels in response to commands from the FADEC.

FOR TRAINING PURPOSES ONLY

73-7

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

ELECTRICAL CONNECTOR (PD014, LEFT; PE008, RIGHT)

FUEL PRESSURE SWITCH (SD003, LEFT; SE003, RIGHT)

UNION PACKING

UNION

PACKING ORIFICE

DETAIL A LEFT SIDE SHOWN, RIGHT SIDE TYPICAL

Figure 73-4. Fuel Metering Unit

73-8

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

The use of motive flow by the airframer is required for the FJ44-3A engine. Fuel flow from the HP pump element that is not scheduled to the engine is bypassed back to the inlet of the pump element. This bypassed flow is available as motive flow for airframe jet pump operation.

NOTES

The FDU provides a regulated motive pressure to the aircraft system. Actual motive flow (and pressure) is a function of the jet pump size and aircraft fuel system efficiency. Motive flow shutoff is required for engine starts and is the responsibility of the airframer. The FDU (Figure 73-4) contains an integral PMA that supplies redundant three-phase power to the FADEC. The PMA is geared to the pump shaft and rotates at twice the speed of the pump. Fuel enters the FDU, where it is pressurized and routed to the FOHE. The fuel is heated (oil cooled) to prevent icing and returned to the fuel filter element. After filtering, fuel is moved by the positive displacement pump to the fuel metering section of the FDU. Fuel metering is accomplished via: • A coil stepper motor-driven fuel valve • A pressure regulator • A rotary variable differential transformer (RVDT) feedback device The valve terminates fuel flow when power is removed from the solenoid. The FDU also has a pressurizing valve to seal the fuel supply from the engine combustor when the system is shut down. The engine inducer bleed valve is actuated by a coil solenoid-driven, hydraulically actuated servo built into the FDU. Finally, the FDU provides a regulated motive pressure supply to the aircraft motive ejector system.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

SHUTOFF VALVE PLUNGER FDU SUPPORT BOSS

INNER CABLE

WASHER NUT

OUTER CABLE

CLEARANCE

RESET PIN

DETAIL A

Figure 73-5. Fuel Shutoff at FDU

73-10

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Start Nozzle

NOTES

The engine has a fixed-flow, stationary fuel nozzle for improved high-altitude restart reliability. The nozzle provides approximately 9 pph of additional fuel flow continuously during all engine operation. The nozzle receives HP (P2) metered fuel from the fuel control unit via the start nozzle control valve.

Fuel Manifold The fuel manifold supplies fuel to the fuel slinger in the combustion section. The manifold tube goes through the front bypass duct by way of the right service island and enters the diffuser case. The manifold supplies fuel to the underside of the rotating fuel slinger.

Fuel Slinger The fuel slinger is on the HP shaft between the combustor cover and the primary plate. It is part of the seal and fuel slinger assembly. Fuel is supplied to the underside of the slinger by the fuel manifold. The fuel is ejected radially outward into the combustion zone through a series of holes in the slinger.

LP Shaft Separation Detection Device A mechanical LP (N1) shaft separation detection device (Figure 73-5) on the engine detects LP shaft rearward movement of 0.050 inch (1.27 mm) or greater. This safety device prevents an LP turbine rotor overspeed condition (and possible turbine burst) in the event of LP shaft separation. The device automatically shuts off fuel flow to the fuel control unit via mechanical linkage (Figure 73-5).

FOR TRAINING PURPOSES ONLY

73-11

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Figure 73-6. FADEC Assembly

73-12

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CONTROLS AND INDICATIONS FADEC The Williams FJ44-3A engines use an electronic control system (Figure 73-6) based on one dual-channel FADEC that monitors and controls each engine. This control system consists of two identical FADECs on the aft pressure bulkhead in the empennage. Each FADEC interfaces with its respective engine and the input and reporting systems from the cockpit. Within each FADEC assembly are two completely independent channels: FADEC A and FADEC B. All signals between the FADECs and engines, and between the FADECs and the aircraft, are completely redundant. Independent electrical harnesses are provided for each FADEC system for complete electrical isolation between the two channels. As input is supplied to the FADECs from throttle quadrant dual RVDTs, the FADECs control steady-state and transient operation of both engines. The system modulates fuel flow rate to the engine in response to input from aircraft sensors and measurements of engine operating conditions. FADECs automatically preve n t e n g i n e s f r o m ex c e e d i n g s p e e d a n d temperature operating limits.

NOTE Both the A and B FADECs (for both engines) must be operating with no faults for dispatch.

FADECs automatically alternate at each engine start to ensure proper functioning and reliability. The alternate FADEC must not be selected unnecessarily. The FADECs recall which unit (A or B) last started the engines. When power is applied, the FADECs power up in FADEC A. If FADEC A was used last, the system automatically switches to FADEC B. The FADECs are powered by 28-VDC enginedriven PMAs or by 28-VDC current supplied by the aircraft generators or battery. The PMA supplies current for the engine ignition system after engine start and is the primary source of power to the FADECs. With the engine running, the PMA continues to supply FADEC power, allowing for continued engine operation in the event of total loss of normal generator and battery power.

FADECs also control engine synchronization as commanded by a pilot-controlled switch on the instrument panel. This two-position rotary switch may be positioned to either OFF or ON. Each engine has dual (A and B) FADEC channels with only one FADEC channel in command. The other is fully operating at all times and automatically ensures engine control if the FADEC in command experiences a failure. Either FADEC channel can be selected at any time. Pressing the momentary type switch changes channels each time the switch is pushed to operate the opposite FADEC. The FADEC does not allow selection of a failed FADEC. Pushing a momentary type FADEC reset switch may reset some faults in the selected FADEC.

FOR TRAINING PURPOSES ONLY

73-13

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A

ENGINE ANTI-ICE VALVE (PD005, LEFT; PE005, RIGHT)

EXCITER NO. 1 (PD030, LEFT; PE030, RIGHT) EXCITER NO. 2 (PD031, LEFT; PE031, RIGHT)

GROUND (GD001, LEFT; GE001, RIGHT)

ENGINE TT2/PT2 SENSOR (PD019, LEFT; PE019, RIGHT)

GROUND (GD008, LEFT; GE003, RIGHT)

ENGINE FIREWALL (PD001, LEFT; PE003, RIGHT)

FUEL FLOW TRANSMITTER (PD002, LEFT; PE002, RIGHT)

FUEL FILTER (PD017, LEFT; PE018, RIGHT)

FIRE DETECTOR (PD003, LEFT; PE001, RIGHT)

FUEL PRESSURE SWITCH (PD014, LEFT; PE008, RIGHT) OIL TEMPERATURE SWITCH (PD010, LEFT; PE010, RIGHT)

OIL PRESSURE TRANSDUCER (PD023, LEFT; PE023, RIGHT)

ENGINE BLEED AIR TEMPERATURE SWITCH (AD015/AD016, LEFT; AE013/AE014, RIGHT)

OIL PRESSURE SWITCH (AD005/AD006, LEFT; AE003/AE004, RIGHT) STARTER-GENERATOR MAGNETIC PICKUP (PD012, LEFT; PE006, RIGHT)

DETAIL A LEFT SIDE SHOWN, RIGHT SIDE TYPICAL

Figure 73-7. Engine Electrical Harness

73-14

FOR TRAINING PURPOSES ONLY

INB

RD

D

FW

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

FADEC/Engine Interface Integration between the FADEC system and the airframe is predominately electrical (Figure 73-7). The airframe supplies power, throttle position, and numerous discrete inputs to define the state or condition of the aircraft to the FADEC. The FADEC provides engine data RS-422 data bus and ARINC 429 to and from the aircraft. The FADEC provides a series of discrete outputs to the aircraft providing a control system and engine operating status. The FADEC system supplies an RS-422 serial communication bus. This bus is available to the aircraft via an aircraft-mounted maintenance interface for FADEC diagnostics, maintenance, and data retrieval. A standard stream of engine and control system data is sent to this port and is available for recording by the aircraft. RS-422 between engines is used for FADEC fault accommodation and synchronization. Each FADEC interfaces with the engine through a dedicated set of engine control sensors and actuator electrical interfaces. All interfaces to the engine-controlled actuators are electrically redundant and isolated. This allows either FADEC to control the following engine parameters independently:

• N 2 speed—The N 2 speed sensors consist of two coils and are on the side of the AGB. Rotation of a spur gear in the AGB induces a pulsating voltage in each of the coils in the N 2 sensor. The frequency of this pulsating voltage is a function of HP rotor rotational speed and is interpreted by the FADEC. The N 2 sensor coils are wired identically to the N 1 sensor coils. • ITT—The ITT system consists of two thermocouple harnesses, each having three thermocouples connected in parallel and joined at a sealed intermediate junction box (J-Box). The six thermocouples are mounted externally on the engine case and protrude into the gas path behind the HP turbine. With the six thermocouples connected in parallel, an average temperature is supplied to each channel of the FADEC. • FDU metering valve—An RVDT with excitation power from the FADEC feeds the FDU metering valve position to both channels of the respective FADEC.

• Power supply: ° Aircraft power supplies 28 VDC to the FADEC system during power-up and start and continues to provide backup power in the event of PMA failure. ° The PMA provides primary power for both FADEC channels and related relays, as well as solenoids used for engine control and operation. • N1 speed—Dual N1 speed sensors are externally accessible on the engine case. A phonic wheel on the LP shaft induces a pulsating voltage in each of the oils in each N 1 sensor. The frequency of this pulsating voltage is a function of LP rotor rotational speed and is interpreted by the FADEC. Each sensor is dedicated to a specific channel of the FADEC.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

CRU

MCT

TAKEOFF

60.0° IDLE

53.4° 46.2°

CUTOFF 13.7°

13.4°

49.1° 58.1° 67.1°

Figure 73-8. Throttle Lever Angle (TLA) Positions

73-16

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

• T2P2 sensors—The T2P2 sensor system consists of dual-element probes protruding into the engine inlet air stream immediately in front of the fan case. Each probe contains dual T2 remote temperature devices (RTDs) that provide a dedicated inlet air temperature to each channel of the FADEC. Dual P2 transducers provide inlet air pressure input to each channel of the FADEC. Both probes are electrically anti-iced via the airframe DC bus when engine anti-ice is selected. Refer to Chapter 30—“Ice and Rain Protection” for more details on the probe anti-ice. • Throttle lever angle (TLA)—TLA is an input to the FADEC based on the position of the throttle levers in the cockpit (Figure 73-8). A dual element RVDT, attached to each throttle lever by a linkage, provides the input of TLA to each FADEC. The left engine FADEC channel A receives a TLA input from one element of the left throttle lever RVDT, while channel B receives a TLA value from the other element. The same is true regarding the right engine FADEC and the right thrust lever RVDT. The throttle levers have a gated cutoff position, an idle position, and three detented positions for cruise, maximum continuous thrust, and takeoff.

Airframe Inputs to FADEC

• We i g h t - o n - w h e e l s ( W OW ) — S q u a t switch information from the landing gear PCB (pilot J-box) determines WOW status and supplies an input to the FADEC. Aircraft status of on ground is sent to the FADEC if either main gear squat switch indicates on ground. • ENGINE SYNC switch—With the ENGINE SYNC switch selected ON, an input is supplied to the FADEC requesting the FADEC to synchronize engine speeds. • Anti-ice status—An input is sent to the FA D E C f r o m t h e W I N G / E N G I N E ANTI-ICE switches providing the status of engine and wing anti-ice. • FADEC channel select—The FADEC changes channels automatically at each power-up cycle. The channel swap can be accomplished manually by using the respective FADEC SELECT A/B switch. The FADEC does not allow a forced change over to an unhealthy/less healthy channel. • Gear down and locked—The gear position input to the landing gear No. 1 PCB provides information to the FADEC indicating that the landing gear is down and locked. • Ground idle switch—The ground idle switch indicates to the FADEC what idle speed is selected for the engine.

Several discrete inputs are provided from the airframe to both channels of the FADEC as follows: • ENGINE START switchlight—When the start switchlight is pressed, an input is supplied to the FADEC from the respective left or right start printed circuit board (PCB) to initiate the start logic sequence. • Normal shutdown—Taking the thrust lever into the cutoff position actuates the cutoff switch, which supplies a normal shutdown command to the FADEC.

FOR TRAINING PURPOSES ONLY

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

MATE DUAL CHANNEL FADEC FUEL/OIL HEAT EXCHANGER RS-422 LINK

FUEL SUPPLY

MOTIVE FLOW

AIRFRAME

FUEL METERING BYPASS SWITCH

ANALOG AIRFRAME SENSORS ARINC N1, N2, ITT COMMUNICATION

FDU FUEL PUMP FILTER AND BYPASS METERING UNIT FADEC POWER SUPPLY INSTRUMENTATION PORTS FUEL INLET PRESSURE FILTER TEMPERATURE

AIRFRAME AIRFRAME POWER I/O

3 PHASE POWER FUEL FLOW COMMAND FUEL SUPPLY

FUEL FLOW FEEDBACK

DUAL CHANNEL FADEC

NORMAL/OVERSPEED SHUTDOWN COMMAND BLEED VALVE COMMAND

N2 SHAFT SPEED

TRANSIENT BLEED VALVE LP SHAFT SEPARATION ENGINE

ENGINE INSTRUMENTATION (N1, N2, ITT, PSI, T2)

Figure 73-9. FADEC Electrical Schematic

73-18

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

OPERATION The overall engine control system block diagram is shown in Figure 73-9. FADEC provides automatic control of engine power setting, transient control, and fuel delivery during start. The FADEC provides fuel valve position control to modulate fuel flow, control of the solenoid fuel shutoff valve, and control of the inducer bleed valve state, all integral to the FDU. One of the principle functions of the FADEC is to provide operating limits for critical engine parameters: N 1 , N 2 , and ITT. The N 1 is governed at high power and N 2 at low power. The overspeed function limits both N 1 and N 2 . ITT is held to a start limit and to a fixed margin above the red line during operation. As the engine deteriorates with time, the ITT margin at takeoff N 1 decreases and may eventually reach the redline limit, indicating that maintenance action is necessary. The pilot is responsible to limit ITT, as required. Takeoff thrust is commanded when the TLA is in the takeoff position and scheduled based on inputs of N 1 , TT2, PT2, and three discrete signals: anti-ice 1, anti-ice 2, and TT2 heat compensate. Maximum continuous thrust is commanded when TLA is in the maximum continuous thrust (MCT) position. Maximum cruise thrust is commanded when TLA is in the maximum cruise position. The FADEC is aircraft-installed. The FADEC governs N 1 and N 2 speeds to set engine thrust. At power settings from partial power-up to takeoff, the FADEC governs to N 1 as a function of TLA, engine inlet total pressure (PT2), engine inlet total temperature (TT2), and aircraft bleed air extraction level. The FADEC recognizes five set power setting regions (or flats): shutdown, idle, maximum cruise, maximum climb, and takeoff. Takeoff thrust is achieved when TLA is in the takeoff position. MCT is achieved with the TLA in the MCT position. Maximum cruise thrust is achieved when the TLA is in the MC position. Maximum thrust schedules are also depend-

ent upon the combination of anti-ice discrete inputs from the aircraft. Typically, three thrust schedules are available: • Bleed level 0 is for normal pressurization (0 to ≤ 8 ppm) • Bleed level 1 (BL1) is for the addition of engine anti-ice (=8 to ≤ 22 ppm) • Bleed level 2 (BL2) is for the addition of wing anti-ice (=22 to ≤ 46 ppm) When BL1 or BL2 schedules are selected, the anti-icing, electrical power, and T2P2 heater discrete are selected ON. At low part power settings, the FADEC governs to N 2 as a function of TLA, PT2, TT2, aircraft bleed level, and aircraft WOW indication. WOW and aircraft gear down and locked (GDAL) input descretes control T2 sensor “lock” during takeoff and goaround thrust situations. During a normal takeoff, the T2 sensor inputs are locked until one of the following occurs: • Liftoff (WOW to flight) +19 seconds has elapsed • The landing gear is retracted (GDAL is UP) +19 seconds • Throttle is retarded to below the takeoff position • 120 seconds has elapsed When the pilot advances the throttle to the takeoff position while remaining on the ground (WOW input), the T2 sensor inputs are locked for 120 seconds or until the throttle is retarded to below the takeoff position. The T2 sensor input is also locked while on the ground with anti-ice selected on. If the aircraft is still on the ground when anti-ice is turned off, the T2 sensor unlocks 3 minutes after the anti-ice is selected off. Idle (flight or ground) is commanded when the TLA is in the idle position. Flight idle (FI) versus ground idle thrust is scheduled based on the aircraft WOW discrete (FADEC has an 8second built-in delay for ground idle selection and a 2-second delay for flight idle selection from ground idle). Another discrete input, FI select, can command flight idle independent of the WOW discrete.

FOR TRAINING PURPOSES ONLY

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Figure 73-10. TLD Fault Indications

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FADEC Health and Status Annunciations Each FADEC channel sends four discrete outputs to the aircraft (eight total) indicating the health and status of the units. Each channel transmits the following data: • Maintenance required

Each FADEC has an single connection just forward of the pedestal. This connection enables ground support equipment to monitor FADEC parameters for diagnostic purposes. FADEC fault and exceedance records on each channel (through an RS-422 serial link) can be downloaded. The aircraft interface connection can also be used to update FADEC software versions.

• Time limited dispatch (TLD) • No dispatch (ND)

NOTES

• Channel “active” The TLD annunciators are in the empennage inside the battery access door (Figure 73-10). The unit is required to be inspected at predetermined maintenance intervals (the FADEC must be powered for this inspection). ND is an engine control system fault. The ND (on the annunciator panel) is identified as ENG CTRL SYS FAULT and is tied to the aircraft MASTER CAUTION indicators. The channel “active” annunciator is in the cockpit on the vertical instrument panel. During FADEC power-up, (approximately 10 seconds), all annunciators cycle on and off during the initialization process. If the FADEC has no faults (and is capable of controlling the engine), the ND annunciation remains off and either channel A or B “active” is annunciated. If no electrical current faults and/or faults in FADEC memory are detected, the maintenance and TLD annunciators remain off. If the ND lamp is annunciated, flight is not recommended. A complete listing of faults that cause ND annunciation, maintenance requirements, and TLD annunciations are found in the Engine Maintenance Manual.

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CHAPTER 74 IGNITION

FOR TRAINING PURPOSES ONLY

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VOLTAGE SEL BATT

L GEN

R GEN

TEST OFF ANNU ANTI SKID OVER SPEED W/S TEMP

DC VOLTS FIRE WARN LDG GEAR BATT TEMP AOA

DC AMPS

DC AMPS

DC POWER L GEN OFF

RUDDER BIAS

STBY FLT DISPLAY ON

R GEN

BATT

O F F

OFF EMER

RESET

AVIONICS POWER ON

RESET

TEST

OFF

FUEL TRANSFER ENGINE START

FUEL BOOST OFF L L TANK

R TANK

L ENG

R ENG

ON

L

R

O F F NORM

PITOT & STATIC

DISENGAGE

IGNITION R

L

R

START DISG NORM

NORM

ANTI-ICE / DE-ICE WINDSHIELD WING/ENGINE BLEED HI ALCOHOL ON WING XFLOW L WING/ENG R WING/ENG

LOW

NORM

TAIL AUTO O F F

O F F

O F F OFF

ON

OFF

OFF

ENG ON

L AHRS SLAVE

ENG ON

MANUAL

EXTERIOR LIGHTS

MANUAL

L SLEW

BEACON

ANTI-COLL

NAV

WING INSP

AUTO

R SLEW

OFF

OFF

OFF

OFF

PARK BRAKE – PULL

Figure 74-1. Ignition Switch Panel

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DESCRIPTION The ignition switching system provides 29 VDC to the engine ignition exciters as required for proper engine operation. Components in the system include: • Left and right ignition switches • Igniter control PCBs

The ignition switching system provides 29 VDC to the engine ignition exciters when the applicable L IGNITION or R IGNITION circuit breaker on the left CB panel is engaged and any of the conditions that follow are met: • The IGNITION switch is in the ON position. • The start relay is engaged and the throttle lever is out of the cutoff position.

• FADEC

• The aircraft is airborne, the landing gear is down, and engines are running.

• Ignition exciters • Igniter leads

• T h e FA D E C S s e n s e a n i m p e n d i n g flame out.

• Igniter plugs Ignition switches in the cockpit activate the ignition exciters (Figure 74-1).

Ignition indicators on the left switch panel near the ignition switches illuminate when power is supplied to the engine ignition exciters.

COMPONENTS Igniter Control Printed Circuit Boards

NOTES

Electrical power to the exciters is controlled by logic circuitry on the IGNITER CONTROL PCBs.

Ignition Exciters, Leads, and Plugs The two ignition exciters are high-energy, capacitive discharge devices. The parts for each ignition exciter are contained in a sealed, corrosion-resistant case. The ignition exciters draw current from the aircraft electrical system and send current through each igniter lead, independently, at a rate of one to six sparks per second. The igniter plugs ignite the fuel/air mixture in the combustor. Ignition exciter cases are mounted to the engine interstage housing. The spark rate is one to six sparks per channel, per second. The exciters can operate continuously with input voltage ranging 10–30 volts, 3.5 Joules stored energy.

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CHAPTER 77 ENGINE INDICATING

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Collins

ITT °C 1000 900

N1 % –––.– 100 T T O 90 O

OIL PSI

0.0 OIL °C

800

70

700

50 600 400 200

30

0.0

0.0

HDG 137

VOR1 CRS 166 TTG – – : – – – – – – NM

0.0 N2 %

12

FUEL QTY LBS 2000 1500

0

0 °C FUEL 0 PPH

23

25

25

1000 500

23 0

0

0

137 15

E

S

5

2.5 TERR LX/RDR ADF

TFC TA ONLY

ADF

BRT DIM

Figure 77-1. Multifunction Display

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DESCRIPTION

OPERATION

The engine indicating system is made up of the following parts:

The N 1 LP speed pickup has dual output channels. Each dual element sensor provides an independent and isolated speed input to each of the FADEC channels. The speed sensor signals are routed through the engine electrical harness to the engine-to-aircraft interface connectors. From this point, the airframe electrical harness routes an individual N 1 speed signal to each of the two FADEC channels. The primary source of engine N 1 for the cockpit display is via the FADEC-to-aircraft ARINC communication bus. The FADEC changes the input data from the magnetic pickup and sends ARINC data to the multifunction display (MFD) (Figure 771) and analog N 1 data to the standby N 1 indicators. The fan speed is shown on the MFD.

• LP magnetic speed pickup assembly • Gearbox magnetic speed pickup assembly (HP) • ITT probe harness (six probes) • ITT/N 1 trim box assembly • TT2/PT2 dual temperature/pressure sensor • FADEC engine harness assembly • Engine vibration pickup (test equipment) • Fuel filter electrical indicator switch (delta P) • Oil filter differential pressure (delta P) indicator • Magnetic chip collectors

COMPONENTS Engine Fan Speed (N1) The N 1 speed sensor is a dual-coil magnetic pickup that is electrically isolated, providing a complete electrical separation of the FADEC channels. Each magnetic coil transmits a sine wave for each tooth of the sensor ring. The sensor ring has 18 teeth for each revolution (100% rpm is equal to 18,000 rpm [5,400 Hz]). The N 1 red line is 102.8%–103.9% for 20 seconds or more, or ≥103.9% rpm (103.9% rpm equals 18,702 rpm [5,610.6 Hz]). The output from the magnetic pickup is transmitted to the FADEC.

The N 1 digital readouts show the current fan speed for the left and right engines. The MFD uses the ARINC data from the FADEC to give a vertical tape display of the fan rotational speed in percent rpm. The readouts are shown in numerical values as follows: • Green for data ≤102.8% • Yellow for valid data 102.8%–103.9% for less than 20 seconds • Red for valid data as follows: ° Data 102.8%–103.9% for 20 seconds or more ° Data ≥103.9% The current fan speed is also shown as a digital readout on the N 1 standby indicator. The indication flashes when the value is ≥102.9%. Each channel in the standby indicator uses the data from the FADEC to give a digital display of the fan speed in percent rpm. The standby indicator also receives a channel active signal from the FADEC and decides which channel of data from the FADEC to use.

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Figure 77-2. Engine Turbine Speed Indicating (N2) System

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Engine Turbine Speed (N2) The engine turbine speed (N 2 ) indicating system has a dual-channel cockpit display and a dual-channel FADEC (Figure 77-2). The system includes the wiring from the turbine (HP shaft) dual-channel magnetic speed pickup assembly on each engine to the FADEC. It also includes the ARINC data wiring from the FADEC to the MFD.

The digital display of the turbine speed is shown as green, yellow, or red. The display is green if the speed is ≤100%. The display is yellow if the speed is between 100% and 100.7% for less than 20 seconds. The display is red if the speed is between 100% and 100.7% for more than 20 seconds or more than 100.7%.

NOTES

The N 2 indicator speed is shown as a numeric value on the MFD. The MFD uses a digital LCD. Separate digital displays are supplied for each engine. The HP rotor speed is sensed by a dual-coil magnetic pickup assembly and is sent to the FADEC. The N 2 sensor uses the movement of the gear teeth on the gear box fuel pump drive to make a modified sine wave output. The gearbox (N 2 ) magnetic speed pickup is on the gearbox and reads the HP spool speed through the gearbox. An output frequency of 6,741 Hz is produced at maximum HP shaft speed (100% N 2 ). The engine turbine speed (N 2 speed sensor) is a dual-coil magnetic pickup that is electrically isolated to maintain a complete electrical separation of the FADEC channels. Each magnetic coil transmits a sine wave per tooth of sensor ring. The sensor ring has 51 teeth and is on a shaft in the AGB. This shaft is driven at a gear reduction of 0.1906 or 9.7206 teeth per N 2 shaft revelation (100% rpm is 41,200 N 2 shaft rpm). The N 2 red line is 100.73% rpm, which is equal to 41,500 rpm or 6723.4 Hz. Each channel in the FADEC receives the signal from the related channel of the magnetic speed pickup assembly. The signal is changed to ARINC data and sent to the MFD. The MFD uses the data to make a digital display of the turbine speed in percent rpm.

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Figure 77-3. Interturbine Temperature Indicating (ITT) System

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Interturbine Temperature Indicating There are six type K ITT probes in the first LP turbine nozzle (Figure 77-3). The ITT probes are the same length. The thermocouple leads are joined and averaged at the electrical connector. The averaged signal from probes 2, 4, and 6 is sent to FADEC channel A. The averaged signal from probes 1, 3, and 5 is sent to FADEC channel B. The ITT/N1 trim box assembly provides the capability to: • Adjust ITT cockpit readings so they are approximately equal at equal power settings • Reduce N 1 within the allowable range to match thrust between engines The trim box is on the front bypass duct of the engine. It contains two sets of DIP switches (both ITT and N 1 trim for each channel) that permit changes of up to 20 discrete increments of ITT or N 1 trim.

A chromel-alumel thermocouple transmitting system supplies the average input signals to each FADEC channel. For maintenance information on the ITT probes and wiring harness, refer to the Williams International FJ44-3A Engine Maintenance Manual. The engine ITT thermocouples make an electrical signal as a result of the changes in temperature of the intermediate turbine section. The ITT sensor has six individual thermocouples. These six thermocouples are grouped into two sets of three. Each thermocouple in the group of three is electrically connected in parallel. The chromel-alumel wiring harness makes an average of the electrical signals from the three thermocouples. Each average signal is sent to one of the FADEC channels. The FADEC puts together the two channels of ITT temperatures and prepares an average of the two. The average temperature from the FADEC is sent as ARINC data to the MFD. The MFD shows the data as a vertical tape display of ITT in °C.

The ITT is displayed on the MFD in the center instrument panel.

NOTES

The ITT is shown in degrees Celsius (°C). The range of the display is 200°C–1,100°C. The scale of 600°C–1,100°C is expanded for more precise indication. The ITT is shown in a vertical tape format. Separate displays are supplied for each engine. The left and right signal inputs to the MFD are separate. Red, yellow, and green markings on the temperature scale show the warning, caution, and normal operating ranges, in that order.

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CHAPTER 79 OIL

FOR TRAINING PURPOSES ONLY

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LEGEND OIL SEPARATOR DISCHARGE HP SHAFT AND 1ST REDUCTION LUBE JET

VENT

REGULATED SYSTEM PRESSURE PRESSURE REGULATOR

OIL TANK AIR PRESSURE SCAVENGE PUMP RETURN PRESSURE SCAVENGE PUMP SUCTION

OIL FILL SIPHON BREAK BRG NO. 1.5

LUBE PUMP INLET P PRESSURE TAP

OIL TANK

T TEMPERATURE SENSOR TAP

BRG NO. 3

BRG NO. 1

SCREEN FILTER MAGNETIC PARTICLE COLLECTOR

BRG NO. 4

OIL FILTER P INDICATOR

BRG NO. 2 T

P BYPASS VALVE

TOWER SHAFT BEARINGS INJECTOR NOZZLE

TOWER SHAFT SPLINE 1/2 CAVITY DRAIN

BYPASS DUCT VENT DISCHARGE

MAIN SYSTEM FILTER BYPASS VALVE FUEL

FUEL/OIL HEAT EXCHANGER

GEARBOX LUBE JET FOR 2ND REDUCTION GEAR, TOWER SHAFT, AND GENERATOR SPLINE

GEARBOX

RELIEF VALVE

LUBE AND SCAVENGE PUMP

GEARBOX AIR/OIL SEPARATOR SCAVENGE ELEMENTS

Figure 79-1. Oil System Schematic

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PRESSURE ELEMENT OIL DRAIN

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

DESCRIPTION

Oil Filter

Various components, external to the engine, store and deliver lubricating oil to and from the engine.

The oil filter (Figure 79-1) is a disposable cartridge in a stainless steel bowl that is threaded into the gearbox housing. It connects to the pressure pump outlet through passages in the gearbox.

The “Engine Indicating” section includes information on the engine oil pressure indicating system and the engine oil temperature indicating system. All other information is found in the Williams Maintenance Manual. The oil system (Figure 79-1) is made up of the following components: • Lube and scavenge pump • Oil filter

Oil Filter Bypass An oil filter bypass valve (Figure 79-1) is in the gearbox housing. When differential pressure across the oil filter approaches 25–30 psid, the valve opens, allowing oil to flow even if the filter element becomes totally blocked.

Delta P Indicator

• Oil filter bypass valve • Filter differential pressure (delta P) indicator • FOHE • Reservoir (oil tank) • Oil pressure regulator • Magnetic chip collectors

A filter differential pressure (delta P) indicator is in parallel with the oil filter and bypass valve. As the bypass valve approaches its opening pressure, the delta P indicator button extends, indicating excessive filter contamination. The indicator operates at 15 ± 3 psid. A thermal lockout prevents the indicator from operating because of high Delta P due to cold oil.

COMPONENTS Lube and Scavenge Pumps T h e l u b e a n d s c av e n g e p u m p a s s e m b l y (Figure 79-1) includes one pressure element and two scavenge elements contained in a single housing. One pressure lube and two scavenge pumps are positive displacement pumps. The inlet to each element is protected by a wire mesh screen. All inlet and discharge ports are on the pump mounting flange face facilitating pump replacement. A pressure relief valve is in the pump housing. This valve maintains a 130–140 psi pressure rise across the lube element under normal operating temperatures and flow conditions. The regulator valve also protects against overpressurization.

FOR TRAINING PURPOSES ONLY

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OIL FILLER CAP AND SIGHT GLASS

OIL PRESSURE REGULATING VALVE O-RING O-RING

Figure 79-2. Oil Pressure Regulator

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Oil Cooler

NOTES

The oil cooler is an FOHE constructed of aluminum (Figure 79-2). The oil cooler is mounted to the gearbox and transmits oil in and out through the mounting flange faces. Fuel travels in and out through external lines.

Oil Reservoir (Oil Tank) The oil storage tank (Figure 79-2) is internally vented to the forward bearing cavity (through a 5 psi pressurizing valve) vented to the bypass duct via the gearbox air-oil separator. The reservoir vent valve and bypass duct pressure reference ensure sufficient oil tank pressure for efficient pump operation at all altitudes. Oil level is established by the oil sight glass on the side of the engine interstage housing. The sight gauge is marked with two levels: FULL and ADD. The oil volume between these levels is approximately 1.0 U.S. quart. A gravity-type oil filler port, fitted with an overcenter-lock sealed cap, is on the outboard side of the engine. The quantity of usable oil is approximately 2 quarts at 0° installed pitch angle. The oil tank filler port does not allow the oil tank to overfill even if the oil filler port itself is overfilled.

Oil Pressure Regulator The oil pressure regulator (Figure 79-2) is on the upper right side of the interstage housing, near the oil filler port. The regulator is set at the factory to monitor the main oil pressure of the internal passage, which carries oil to the gearbox. The regulator returns oil to the reservoir tank through a separate internal passage in the interstage housing.

FOR TRAINING PURPOSES ONLY

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A A

A

VIEW A-A HOSE ASSEMBLY

LOW-PRESSURE WARNING SWITCH (SD002, LEFT; SE002, RIGHT)

O-RING

PRESSURE TRANSDUCER (UD005, LEFT; UE005, RIGHT)

CONNECTOR (PD023, LEFT; PE023, RIGHT)

ISOLATOR CLAMP

DETAIL A

Figure 79-3. Oil Sensor Installation

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CONTROLS AND INDICATIONS Oil Pressure Indicating The engine oil pressure indicating system (Figure 79-3) uses resistance-type bulbs, a pressure switch (in the engine oil system), and indications on the instrument panel MFD. An oil pressure transmitter is on the forward cowl bulkhead of each engine compartment. The transmitter is connected to the engine oil system by a line assembly. Oil in the oil transmitter plumbing is pressurized by the engine oil pump during engine operation.

The engine oil pressure transmitter provides resistance signals that are given to the data collection unit (DCU) in the tail cone aft of the tail cone baggage compartment. The analog signal sent to the DCU is placed on a digital bus for display on the MFD. The MFD uses a digital LCD.

NOTES

The oil pressure transmitter contains a pressure-sensitive potentiometer. The output and wiper voltages are in proportion to the oil pressure. Input voltage of 29 VDC to the transmitter is supplied by the aircraft electrical system. The pressure indicating and low pressure warning systems are separate. Oil pressure is displayed in psig in a vertical tape format on the MFD. The range of the tape display is 0–100 psi. Separate tape displays are provided for each engine. A low oil pressure warning switch is in the oil pressure transmitter plumbing adjacent to the oil pressure transmitter. The low-pressure warning switch (SD002 left, SE002 right) closes on increasing oil pressure of 37.0 psig (255 kPa) and opens on decreasing oil pressure of 25.0 ± 2.0 psig (172 ± 13.8 kPa). In the open (low-pressure) position, the lowpressure warning switch supplies a ground for the L–R OIL PRESS WARN annunciators. Low oil pressure is indicated by the red L–R OIL PRESS WARN annunciators and by the MASTER WARNING switchlights.

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A A

A

VIEW A-A

ELECTRICAL CONNECTOR (PD010, LEFT; PE010, RIGHT) TEMPERATURE BULB (UD006, LEFT; UE006, RIGHT)

GASKET O-RING

DETAIL A

Figure 79-4. Oil Temperature Indicator

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CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

Oil Temperature

NOTES

The engine oil temperature indicating system (Figure 79-4) has a temperature transmitter (bulb) at the oil filter housing. Engine oil temperature is shown on the MFD. The engine oil temperature transmitters are resistance-type bulbs that have an element sealed in an insulating material within a housing. The housing has an electrical connector and external mounting threads, which attach with a boss connection near the engine oil filter. The oil temperature transmitter bulb resistance changes with oil temperature, causing the current from the transmitter to change in proportion to the oil temperature. The normal operating range for the engine oil is 10°C–135°C. The engine oil temperature transmitter provides resistance signals that are given to the DCU in the tail cone of the aircraft, aft of the tail cone baggage compartment. The analog signal sent to the DCU is placed on a digital bus for display on the MFD. Oil temperature is shown in °C in a vertical tape format. The range of the tape display is 0°–140°C. There are separate tape displays for each engine.

FOR TRAINING PURPOSES ONLY

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Figure 79-5. Oil Pressure vs. Oil Temperature Graph

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OPERATION Lubricating oil from the reservoir tank passes through a perforated sheet-metal screen and magnetic chip collector at the pump inlet, and then through the lube pump to the main oil filter and cooler. The pressurized oil is then taken, via a manifold in the oil tank, to the top of the tank. At this point, an antisiphon orifice breaks the siphon between the oil tank and the bearing cavities (during engine shutdown), preventing flooding. The flow is then split between the gearbox and the front and rear bearings. The pressure-regulating valve maintains a stable lube pressure at the lube jets, independent of upstream pressure drops from the oil filter, cooler, and lines. The oil for the gearbox passes through a lastchance filter and lubricates the second reduction gear mesh, the lower towershaft spline, and the generator spline. The remaining gear meshes, splines, and gearbox bearings are splash-lubricated. The oil from the front bearings passes through lastchance filters to jet-lubricate the Nos. 1 and 1.5 bearings and supplies oil for the No. 1 bearing cage damper. The remainder of the front bearing oil passes through a last-chance filter and lubricates the No. 2 bearing with two jets, an under-race supply, and damper film. The towershaft bearings are also jet-lubricated, along with the upper towershaft spline and the first reduction gear mesh. Another oil jet supplies cooling for the HP tieshaft. The oil for the rear bearings passes through a last-chance filter and jet-lubricates the Nos. 3 and 4 bearings and supplies oil for the respective dampers. The No. 3 bearing is both under-race and external-jet-lubricated.

livered to the oil reservoir is vented to the front bearing cavity through a pressurizing valve. The bearing cavities are vented to the engine bypass duct through a breather air-oil separator in the gearbox.

DIAGNOSTICS Adjusting Oil Pressure • Start the engine and allow to idle for 5 minutes (oil temperature must be at least 115°F when adjusting pressure). Allow additional idle time, if required. • Loosen the regulating valve adjustment locknut. • Using a flat-tip screwdriver, turn the adjustment screw clockwise to increase engine oil pressure.

NOTE It may be necessary to hold the adjustment locknut while turning the adjustment screw. One full turn of the screw yields approximately a 5–8 psig change in pressure. When adjusting the oil regulator, if five threads are exposed, disassemble the regulator, and clean and replace the internal O-ring. If oil pressure cannot be adjusted without five threads exposed, replace the regulator.

• While holding the adjustment screw, torque the adjustment locknut 30–40 inch-pounds (3.3–4.5 N.m).

Two scavenge elements provide independent scavenging of rear bearing cavity and gearbox sumps. The forward bearing cavity drains down the towershaft annulus into the gearbox scavenge. A screen and a self-sealing magnetic chip collector (before each scavenge pump) removes particles from the oil that are large enough to cause damage. Excess air de-

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Oil System Servicing Drain the engine oil (if required) using a TL37255 oil drain fitting and accomplish the following steps: • Vent the oil tank by removing the oil fill cap. • Remove the self-closing plug from the bottom rear of the lube pump (TANK DRAIN port). • Install the TL37255 oil drain fitting to self-closing fitting and drain the oil tank into a suitable container. • Remove the self-closing plug from the bottom of the gearbox. • Install the TL37255 oil drain fitting to self-closing fitting and drain the gearbox oil into a suitable container. Engine oil may also be drained (if required) by removing the gearbox and the lube and scavenge pump (TANK DRAIN) chip collector fittings. Approved engine lubricating oils are listed below (mixing of approved oils is permissible): • Mobil Jet II (preferred oil) • Mobil 254

Water Wash Policy In the interest of ensuring that the safe service lives for engine parts are not degraded by c o r r o s i o n o r s a n d i n g e s t i o n , Wi l l i a m s International recommends that the following policy be followed regarding water washing of the engine.

Marine Atmosphere Operation in a marine atmosphere is defined as operation over salt water or within 5 miles (8 kilometers) of the ocean. If the aircraft is operated in a marine atmosphere below 4,000 feet for more than 30 minutes, perform the desalinization rinse procedure at the end of flight operations for the day. If the aircraft is operated in any of the flight conditions listed below, perform the desalinization rinse procedure at least once a week: • In a marine atmosphere at altitudes below 4,000 feet for less than 30 minutes • In a smog-laden or sandy environment

Compressor Cleaning Perform the compressor cleaning procedure (using an approved engine cleaning fluid) to improve compressor efficiency by removing normal accumulations of dirt and grime. Perform the compressor cleaning procedure whenever the fan or compressor components appear dirty. Adjust the cleaning interval to suit local operating conditions. If engine performance has gradually degraded (increase in ITT), perform the compressor cleaning procedure.

NOTE

Desalinization Rinse Perform this fresh-water rinse procedure to remove salt or other corrosive substances from the interior of the engine. Sea salt contamination from flying in a marine atmosphere or landing at airports near the ocean may initiate corrosion on metallic engine hardware. Water rinsing the engine, as soon after contamination as possible, minimizes corrosion by dissolving and flushing out contaminants.

79-12

Running the engine immediately after rinsing drives out the water and the majority of the contaminants. This also eliminates moisture that could set up galvanic cells and initiate surface pitting and crevice corrosion.

There are two separate cleaning methods for the powerplant. The desalinization rinse is a water-only wash for removing salt or other corrosive substances from the interior of the engine. The compressor cleaning procedure uses approved cleaning fluids to remove dirt and grime from the gas path and improve compressor efficiency.

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

1. If ambient temperature is below freezing and the engine has been shut down for more than 30 minutes, the engine must be warmed up. Start the engine and run at idle for a minimum of 5 minutes. 2. Set the engine bleed air to OFF/CLOSED. 3. Cap the HP bleed ports. 4. Set the ignition to OFF.

NOTE To ensure that fluid reaches the engine core, initial spray cleaning is performed with the engine static (not running). Rotation of the fan (LP spool) forces the flow outward into the bypass area.

CAUTION To prevent damage to the fan, keep the end of the nozzle outside of the inlet lip.

5. While the engine is static (not running), use pressure-washer spray equipment capable of 50–100 psi. Spray 2.5–3.5 gallons of clean water (drinking purity) into the engine core. Direct the spray toward the base of the spinner and the fan blade root area. Spray 360° around.

6. Have a second person starter-crank the engine in accordance with the duty cycle limits for the starter-generator in the AMM. As soon as the fan starts to rotate, spray 2.5–3.5 gallons of clean water into the engine core flowpath. Direct the spray at a 30° angle relative to the axial engine center line, in the same direction as the rotor rotation. Hold the spray nozzle at the appropriate radius to enter the core. To confirm that water has entered the core, observe water dripping from the engine diffuser drains (aircraft drain mast). If water is not observed, change the direction of spray and the repeat starter-crank. 7. Repeat the starter-crank rinse (one time). 8. After completing the desalinization rinse procedure, start the engine and run at idle for a minimum of 5 minutes to ensure that all water has evaporated or blown from the engine. Shut down the engine.

Compressor Cleaning Accomplish the following steps to clean the engine compressor: 1. With the engine shut down, wipe readily accessible inlet areas, the spinner, and the fan blades free of all dirt and grime using a cloth wetted with cleaning fluid MAT021, MAT-029, or MAT-031.

CAUTION CAUTION Make sure that the spray nozzle is clear of the fan rotor blades during engine starter cranking.

To ensure that cleaning fluids do not enter the FCU/FDU or aircraft bleed system, remove the CDP line from the FCU elbow (if the engine configuration has a CDP line). Set the aircraft bleed to OFF/CLOSED.

2. Set the engine bleed air to OFF/CLOSED. 3. Set the ignition to OFF. 4. Open the lower cowling to access the engine. Refer to the AMM.

FOR TRAINING PURPOSES ONLY

79-13

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

5. Disconnect (separate) the aircraft bleed piping from the engine bleed fittings. 6. Cover the aircraft side of the bleed piping with a plastic bag (or equivalent) to prevent entry of water. Leave the piping disconnected so that water can drain from the engine bleed fittings. 7. If the engine configuration has a gearbox and breather flex hose, remove the heatshield (if required) for access to the CDP elbow. 8. Disconnect the gearbox breather flex hose from the elbow by loosening the hose clamp and pulling off the aft end of the flex tube. Plug the end of the flex tube and elbow with a suitable plastic, water-tight plug or cap. 9. If the engine configuration has a CDP line, disconnect the CDP air-to-elbow tube assembly from the FCU elbow.

NOTE To ensure that fluid reaches the engine core, initial spray cleaning should be performed with the engine static. Rotation of the fan (LP spool) forces the flow outward into the bypass area.

10. While the engine is static, use pressuretype spray equipment (50–100 psi) and spray approximately 1–2 gallons of cleaner MAT-021, MAT-029, or MAT031 into the engine. Direct the spray near the spinner and between the fan blades, concentrating on the core of the engine. Spray 360° around the spinner. Let the engine stand 5 minutes.

CAUTION Make sure that the spray nozzle is clear of the fan rotor blades during engine starter-cranking.

79-14

NOTE A heavy spray volume has a better chance of reaching the core while starter-cranking the engine.

11. Have a second person starter-crank the engine in accordance with the duty cycle limits for the starter-generator in the AMM. As soon as the fan starts to rotate, spray fluid into the engine core. Direct the spray near the fan/spinner to concentrate the cleaner into the core of the engine, as shown in the AMM. Move the spray 360° around the spinner to ensure that the cleaner can reach all areas of the core. 12. Spray some cleaner around the remaining inlet area to clean the bypass duct flow path. Spray 3–4 gallons of cleaner during this procedure. Repeat startercranking as necessary to use up the cleaner (no more than two times total). Allow the engine to stand 10 minutes. 13. Repeat steps 10 and 11 using 4–5 gallons of clean water. 14. If the engine is equipped with a CDP line: Blow out any residual water from the CDP line. Reattach the CDP air-to-elbow tube assembly to the FCU. 15. Remove the plastic bag from the aircraft bleed piping and reconnect the piping to the engine bleed fittings. Refer to the AMM. 16. If the engine configuration has a gearbox breather flex hose, remove the plastic plug or cap from the gearbox breather tube flex hose. Reconnect the flex hose to the elbow and tighten the hose clamp. 17. Ensure that the aircraft bleed system is still set to OFF/CLOSED. Run the engine at idle for 5 minutes to ensure that excess cleaning fluid and water have evaporated or blown from the engine and that there are no bleed leaks. Shut down the engine.

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

APPENDIX A TERMS AND ABBREVIATIONS ADC

Air data computer

ADF

Automatic direction finder

ADS

Air data system

AFD

EICAS

Engine [and airframe systems] indicating and crew alerting system

EIS

Engine [and airframe systems] indicating system

Adaptive flight displays

ELT

Emergency locator transmitter

AFM

FAA-approved Aircraft Flight Manual

EMC

Electromagnetic comparability

AGB

Accessory gearbox

EMI

Electromagnetic interference

AHC

Attitude heading computers

ENT

Enter key

AHS

Attitude heading system

ESD

Electrosensitive devices

ALT

Altitude; altimeter

FAA

Federal Aviation Administration (USA)

AMM

Aircraft Maintenance Manual

FADEC

ARINC

Aeronautical radio incorporated

Full authority digital engine controls

ASE

Altimetry system error

FDU

Fuel delivery unit

ATC

Air traffic control

FGC

Flight guidance computers

BITE

Built-in test equipment

FGS

Flight guidance system

CAS

Crew alerting system

FI

Flight idle

CD

Course-deviation indicator (normally on PFDs)

FIR

Flight Information Regions

FL

Flight level

CDU

Control display unit

FMS

Flight management system

DC

Direct current

FOHE

Fuel/oil heat exchanger

DCU

Data concentrator units

GDAL

Gear down and locked

DME

Distance-measuring equipment

HIRF

High-energy radiated fields

DOT

Department of Transportation

HP

High-pressure

ECAC

European Civil Aviation Control

hPa

Hectopascals, barometric pressure

ECU

External compensation unit

HPC

High-pressure compressor

EDS

Electronic display system

IAPS

EEPROM Electrically erasable programmable read only memory

Integrated avionics processor system

IC

Integrated circuit

EFIS

Electronic flight instrument system

IP

Intermediate-pressure

EGPWS

Enhanced ground proximity warning system

ITT

Interturbine temperature

J-box

Junction box

KIAS

Knots indicated airspeed

EIA

Electronics Industries Association

FOR TRAINING PURPOSES ONLY

APP-1

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

KPH

Kilograms per hour

UHF

Ultra high frequency

LCD

Liquid crystal display

V1

Takeoff safety speed

LED

Light-emitting diode

V2

Takeoff rotation speed

LP

Low-pressure

VHF

Very high frequency

LPT

Low-pressure turbine

WATRS

West Atlantic Route System

LRM

Line replaceable modules

WOW

Weight on wheels

MCT

Maximum continuous thrust

YD

Yaw damper

MFD

Multifunction display

MNPS

M i n i m u m N av i g a t i o n P e r f o rmance Specification

MSL

Mean sea level; altitude above MSL

NAT

Northern Atlantic Track

OEM

Original equipment manufacturer

PCB

Printed circuit board

PFD

Primary flight display

PMA

Permanent magnet alternator

PPH

Pounds per hour

PRSOV

Pressure-regulating and shutoff valves

RIU

Radio interface unit

RSS

Radio sensor system

RTD

Remote temperature devices

RTU

Radio tuning unit

RVDT

Rotary variable differential transformer

RVSM

Revised vertical standards minimums (ATC standard)

SDI

Stination identifier

SSM

Sign status matrix

TAS

Traffic Avoidance System

TAWS

Te r r a i n - Aw a r e n e s s Wa r n i n g System

TLA

Throttle Lever Angle

TLD

Time limited dispatch

TSO

Technical standard order

APP-2

FOR TRAINING PURPOSES ONLY

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

APPENDIX B SYMBOLOGY GROUND

BATTERY +

– HEADSET

BUS

HEATER CAP AND STOW HORN/SPEAKER CAPACITOR

INTEGRATED CIRCUIT Integrated circuits do not necessarily work on the principle of ON–OFF as a switch; instead, some work on high and low voltage. Example: High might be 5.0 volts and low might be 0.5 volts.

CIRCUIT BREAKER

CONNECTOR

CURRENT SENSOR

IN

OUT

Current flowing in coil opens switch to indicate circuit is functioning correctly. IN

CURRENT TRANSFORMER

OUT

Current flowing in wire produces a voltage in coil.

AND GATE — Output is low until both inputs are high; then the output is high.

NAND GATE — Output is high until both inputs are high; then the output is low.

OR GATE — Output is low until either or both inputs are high, then output is high.

IN

OUT

IN

OUT

NOR GATE — Output is high until either or both inputs are high; then output is low.

IN

OUT

INVERTER — Output is low when input is high; output is high when input is low.

DIODE REGULAR — Low resistance forward, high resistance reverse. ZENER — Low resistance forward, high resistance reverse until a specific voltage is applied, then conducts freely. TRANZORB — A tranzorb is similar to a zener, but with higher peak current limit. VARISTOR — High resistance either way until a specific voltage is applied, then conducts freely. Example: V47ZA1 conducts freely above 47 volts.

+ IN

OUT

VARISTOR — Encapsulated for moisture protection. FILTER Passes direct current but opposes pulsating current used to reduce noise in sensitive avionics equipment.

FOR TRAINING PURPOSES ONLY

OPERATIONAL AMPLIFIER (OP AMP) — Amplifies the difference in voltage between the two inputs. The minus input is the inverting input, and the plus input is the noninverting input. If an input is applied to the minus input, with the plus input grounded, the polarity of the output will be opposite to the input. If an input is applied to the plus input, with the minus input grounded, the polarity of the output will be the same as that of the input. TIMER — Changes the output from high to low in a regular pattern.

APP-1

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

A diode is connected across the input wires of relays and solenoid-operated devices such as valves to protect voltagesensitive navigation and electronics equipment. The diode is reverse-biased for normal power and no current flows through the diode. Current flowing through the coil of wire produces a magnetic field to operate the relay or valve. The instant power is removed from the coil, the collapsing magnetic field produces a momentary spike of high voltage, which can be several hundred volts depending on the current and the number of turns of wire in the coil. The diode is forwardbiased for the power generated in the coil and the high-voltage spike is dissipated through the diode. A varistor is used in place of the diode on some relays.

LAMP

MOTOR M Basic symbol for motor.

MOT P M

M

RED BLK

REVERSIBLE MOTOR — Direction of rotation is controlled by reversing power and ground on input wires.

ANNUNCIATOR LOAD

28 DVC CW

CCW

An annunciator relay has a connection on the material contact to indicate by a light or annunciator panel when the relay is energized.

MOT

The contacts of a time delay relay do not move to the energized position usually when power is applied.

REVERSIBLE MOTOR — Direction of rotation is controlled by applying power to either field winding input wire.

MOT

MOT

WHT WHT

HI LOW

NONREVERSIBLE MOTOR — Direction of rotation is controlled by design; input wires may be connected either way.

TIME For some time delay relays, the delay time DELAY is part of the relay design. For some time delay relays, the delay time is controlled by the size of an external resistor.

NONREVERSIBLE MOTOR — Two-speedcontrolled by applying power to either input wire.

Jumper wire gives 0.1 seconds of delay. 3

9

6 1

7 FUSE/LIMITER

4 B

PHONE JACK

The symbol for the solenoid may be a box or a coil; the operation is identical.

160.000 OHM resistor gives a 10-second delay. Other resistors give a delay time between 0.1 second and 1.0 second.

TIME A DELAY 2 5

EXTERNAL RESISTOR

RESISTOR

RELAY

TEMPERATURE CONTROLLED — Resistance changes with the temperature. CURRENT FLOW WITH POWER APPLIED

APP-2

REGULAR — Resistance does not change.

+ – CURRENT FLOW THE INSTANT POWER IS REMOVED

FOR TRAINING PURPOSES ONLY

VARIABLE OR ADJUSTABLE — Resistance changes with mechanical input.

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

RESISTOR (Cont.)

D E C K A C C 1

RHEOSTAT — Type of variable resistor with two wires.

POTENTIOMETER — Type of variable resistor with three wires.

D E C K B C 1

D E C K C 1

1 10 9 8 2 GRAYHTII 3 7 4 5 6

2 2 2 3 3 3

SOLENOID/SOLENOID VALVE MOTOR OPERATED — Limit switches stop power when limit of travel is reached.

E F CLOSED B CLOSE

B CLOSE D CLOSED

M

E MOTOR

M

F ANNUNCIATOR (MOTOR ON) A OPEN C OPENED

SHOWN OPEN

SIDE VIEW REAR VIEW WITH TYPICAL TERMINAL NUMBERING C = Common terminal for each deck 1 – 10 = Switch Terminal Position EXAMPLE: AC = Common Terminal of deck A B1 = Switch position 1 on deck B

G MOTOR A OPEN C D OPENED

SHOWN OPEN MOTOR OPERATED — Limit switches stop power when limit of travel is reached.

1NO TWO-STAGE — Two-pole, four-pole, or sixpole switch designed so that all movable 1NC 2NC switch contacts do not move simultaneously.

1C 2C

SHUNT

2NO

OPPOSITE

2-POLE

CENTER* 1NO 1C

1NC 2NO

2C

2NC 3NC

3C

SPLICE ENVIRONMENTAL

Handle position is reference to flat side of the mounting threads.

3NO 4NC *NOTE: Some switches of this type do not have a center position. 4NO

4C

SWITCH

TO FLAT

4-POLE SINGLE-POLE/SINGLE-THROW (SPST) SINGLE-POLE/DOUBLE-THROW (SPDT) — May have OFF position in the center.

1NO 1C

1NC 2NO

2C DOUBLE-POLE/DOUBLE-THROW (DPDT) — May have OFF position in the center. Dashed line indicates all parts move simultaneously.

2NC 3NO

3C

3NC 4NC

4C ROTARY OR MULTIPOSITION TWO-POLE ROTARY — On rotary or multiple pole switches controlled by a knob, the poles (or decks) are identified on wiring diagrams as A, B, C, with A being the part on the knob or shaft end.

4NO 5NC

5C

5NO 6NC

6C

TO FLAT Contact 1 of 2-Pole NO Contacts 1 and 2 of C 4-Pole Contacts 1, 2, and 3 of 6-Pole NC NC Contact 2 of 2-Pole Contacts 3 and 4 of C 4-Pole Contacts 4, 5, and 6 of 6-Pole NO

CENTER* OPPOSITE NO

NO

C

C NC

NC

NC

NC

C

C NO

NO

C IS THE COMMON TERMINAL NC IS THE NORMALLY CLOSED TERMINAL NO IS THE NORMALLY OPEN TERMINAL

6NO 6-POLE

FOR TRAINING PURPOSES ONLY

APP-3

CITATIONJET 3 525B MAINTENANCE TRAINING MANUAL

SWITCH (Cont.)

WIRE PRESSURE-OPERATED

TWISTED

TEMPERATURE-OPERATED

SHIELDED

TERMINAL STRIP

Q12A20 TRANSFORMER

TRANSISTOR

B

B

Transistor contacts are identified as base, collector, and emitter. Flow of current C through a transistor is controlled by the NPN signal applied to the base. The control current (3 to 5% of total current) flows between base and emitter. The main flow of E current (95 to 97%) is between the collector and emitter. Transistors may be drawn without the circle. C PNP E

COLLECTOR BASE

EMITTER

CONTROL SIGNAL

MAIN CURRENT FLOW

C B

DARLINGTON

E

APP-4

FOR TRAINING PURPOSES ONLY

HIGH-TEMPERATURE WIRE