g550 landing gear

LANDING GEAR

2A-32-10: GeneralThe Gulfstream G550 landing gear has dual wheels on both the main and steerable nosegear. The main gear retract laterally into wheel wells on either side of the fuselage keelat the wing root area. The auto-centered nose gear retracts forward into a wheel wellbeneath the cockpit. Integrated gear doors close upon completion of gear extension andretraction providing an aerodynamically conformal fuselage surface.

Landing gear operation normally uses 3000 psi pressure from the left hydraulic system.Extension and retraction selections are made with the cockpit landing gear handle thatelectrically signals the solenoid on the landing gear selector/dump valve. The selector/dump valve directs hydraulic pressure in a defined sequence to open the gear doors,extend or retract the gear and close the doors when the gear have reached the selectedposition. If left hydraulic system pressure is not available due to pump failure or engineshutdown, the Power Transfer Unit (PTU) can be employed to provide pressure forlanding gear operation, provided adequate fluid remains in the left system. In the eventthat left system fluid is lost, the dump function of the selector/dump valve will open lefthydraulic system return lines to allow emergency extension of the gear. A single-useemergency gear extension system operated with three thousand (3000) psi nitrogen gasbottles will extend the landing gear, but leave the gear doors open. The landing gearoperating system is shown in Figure 1.

Landing gear position is indicated on the gear control panel by a green light for each gearand a red light in the landing gear control handle. Illumination of the red landing gearhandle light indicates a disagreement between the selected and actual position of one ormore of the landing gear or gear doors. The individual gear green lights come on whenthe corresponding landing gear is down and locked.

In addition to the red landing gear position disagreement light in the gear handle, awarning horn will sound if the landing gear is not down and locked and the aircraftconfiguration, altitude and/or power setting approximates the landing configuration.

A proximity, or Weight-On-Wheels (WOW) switch is installed on each gear. As the weightof the aircraft compresses the shock struts of the landing gear, the WOW switches closecompleting circuits for relays to aircraft systems that only operate on the ground.Conversely, when airborne, the WOW switches open, providing an in-the-air signal tosystems that operate only in flight.

Main landing gear brakes are air-cooled multiple carbon-fiber discs with anti-skidprotection. Overheat protection for the tires/wheels is provided by fusible plugs in thewheels that melt, releasing tire pressure, if high temperature thresholds are exceeded.Brake temperatures are monitored and indicated on cockpit displays.

The left hydraulic system provides 3000 psi for normal brake operation. Hydraulic fusesare incorporated into the brake hydraulic lines and close to prevent fluid loss if ahydraulic line is damaged or cracked. If left system pressure is not available, but brakesystem hydraulic lines are intact, the PTU or Auxiliary (AUX) pump can be used for brakeoperation. If no hydraulic pressure is available from aircraft systems, accumulatorpressure from the Parking/Emergency brake system will provide approximately five to sixbrake applications without anti-skid protection.

Control of the brake system is through a hydro-mechanical linkage. Cockpit brake pedalapplication is mechanically linked to the system brake metering valves that applyhydraulic pressure proportional to pedal depression. Application of hydraulic pressure tothe brake assemblies on each wheel is moderated by the anti-skid system. The anti-skid

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PRODUCTION AIRCRAFT SYSTEMS 2A-32-00

July 15/04

system compares the speed of each wheel with aircraft speed as derived from theInertial Reference System (IRS) and interrupts application of hydraulic pressure to thebrake of any wheel that is not within three percent (3.0%) of aircraft speed.

The landing gear system is composed of the following components.

• 2A-32-20: Landing Gear and Doors

• 2A-32-30: Extension and Retraction

• 2A-32-40: Wheels and Brakes

• 2A-32-50: Nose Wheel Steering

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PRODUCTION AIRCRAFT SYSTEMS2A-32-00July 15/04

Simplified Landing GearHydraulic System Diagram

Figure 1

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2A-32-00 / 4July 15/04

2A-32-20: Landing Gear and Doors System1. General Description:

The major components of each main landing gear assembly are: a landing geardoor, a shock strut, a sidebrace actuator/downlock mechanism and an uplockmechanism. Each of the main landing gear assemblies has dual wheels withintegrated brakes mounted on an articulated trailing arm attached to the landinggear structural post. The pneumatic/hydraulic shock strut is attached between thewheel end of the articulated trailing arm and the top of the structural post. In flightwith the landing gear down, the shock strut extends and the articulated trailingarm pivots downward, positioning the wheels for ground contact and compressionof the shock strut. The major components of the main landing gear are shown inFigure 2 and Figure 3.

A side brace hydraulic actuator/downlock extends and retracts the main gear. Inthe extended position, hydraulic pressure positions mechanical downlocks thatprevent gear retraction. The downlocks include mechanical springs loads thatmaintain the locked position if hydraulic pressure is lost. In gear retraction,hydraulic pressure compresses the springs, displaces the downlocks and allowsgear retraction inward into the wheel well in the wing root next to the fuselagekeel. An uplock roller mounted on the dual wheel axle assembly engages anuplock hook in the wheel well and rotates to latch the main gear in the uplockedposition. The main landing gear door, hinged at the inboard side of the wheel well,closes after the gear retracts and mates with the fairing door panel mounted onthe outside of the main gear assembly to seal the wheel well.

The nose gear assembly includes steerable dual wheels mounted on a shock strutattached to a trunnion and a hinged two piece truss brace. A drag brace betweenthe truss brace and the shock strut provides additional structural support andlimits the extension of the nose gear shock strut. The truss brace and trunnionpivot axes allow the hydraulic actuator cylinder to retract the nose gear assemblyforward and up and extend the gear aft and down. See the illustration in Figure 4.

When the nose gear is fully extended, the forward and aft sections of the trussbrace move over center to lock the nose gear down. A downlock hydraulic cylinderextends an actuating arm holding the truss brace in the over center position.Springs are also incorporated in the over center downlock to maintain the overcenter lock in the event of hydraulic failure. When the nose gear retracts, thedownlock actuator arm moves aft, allowing the forward and aft portions of thetruss brace to fold at the hinge point. As the nose gear reaches the fully retractedposition, a roller attached to the dual wheel axle assembly pushes aside a springlatch and engages an uplock hook, with the spring latch then falling into placesecuring the roller into the uplock hook. A hydraulically actuated pushrod thenrotates the uplock hook upward to provide positive nose gear latching.

Two symmetrical nose wheel doors, hinged on the outside of the forward nosewheel well and mechanically linked together, operate in conjunction with the nosegear. Timing valves integrated into the extension/retraction cycle operate thehydraulic door actuating cylinder in the correct sequence. When the nose gear isretracted and seated in the uplock hook, the doors close to cover the forwardwheel well, with the fairing panel attached to the rear of the nose gear drag braceenclosing the aft portion of the wheel well.

Indication of landing gear position is provided to the flight crew through lights onthe landing gear control panel, an aural warning horn, and by messages on theCrew Alerting System (CAS) in the event of malfunction.

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2. Description of Subsystems, Units and Components:

A. Landing Gear Doors:

The landing gear doors are shaped to conform to the surrounding fuselageareas, providing an aerodynamically smooth covering of the wheel wellareas when closed. Each main landing gear wheel well is enclosed by asingle door in conjunction with a fairing panel attached to the outside of thegear strut. The nose gear wheel well is enclosed by two symmetricalcurved doors and a fairing panel attached to the aft side of the nose gear.When the landing gear is fully extended, the gear doors close to reducedrag and air noise. All gear door movement is sequenced by mechanicallinkages operating hydraulic actuators to coordinate actuation with landinggear extension and retraction.

Procedures exist to open and close the doors on the ground in order toaccomplish maintenance tasks. Hydraulic power from the Auxiliary (AUX)system and operation of a ground service valve, located adjacent to thenose wheel well, allow movement of the gear doors by pressurizing thelanding gear on the ground. The ground service valve is shown in Figure 5.

B. Shock Struts:

Standard oleo-pneumatic (gaseous nitrogen and hydraulic fluid) shockstruts are incorporated into each gear assembly. The shock struts consistof a movable stainless steel piston within a cylinder containing hydraulicfluid pressurized with gaseous nitrogen. O-ring seals maintain strutpressure and allow the movement of the steel piston. The struts absorb theshock of landing, and provide damping during taxi, takeoff and landingrollout. An air/oil filler valve is provided at the top of each strut for servicing.During normal operations, approximately three to five inches of innercylinder chrome is exposed at the bottom of each main landing gear strut,depending on aircraft gross weight and outside air temperature. A placardattached to upper portion of the landing gear strut indicates the correctextension for ambient conditions.

C. Extension / Downlock Mechanisms:

(1) Main Landing Gear Sidebrace Actuator / Downlock Mechanism:

The sidebrace hydraulic piston and actuator arm extends the mainlanding gear outboard and down from the wheel well and locks thelanding gear in the down position using internal locking keys locatedin the lower end of actuator cylinder. The keys slide into annularslots when the sidebrace actuator reaches full extension. A lock rampositioned by down hydraulic pressure and a spring maintain thekeys in the slots preventing any further actuator movement. Springpressure is necessary to maintain the keys in the locked position inthe event of hydraulic pressure failure. The movement of the keysinto the slots also provides the input for the downlock microswitchthat provides the cockpit indication that the gear is down and locked.When the gear is selected up, hydraulic pressure is ported to theother side of the lock ram, displacing the keys from the slots andoverriding spring pressure, allowing the actuator to move in theretract direction.

The sidebrace actuators have provisions for inserting ground lockpins to maintain the main landing gear in the extended position

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during ground operations. See the illustration in Figure 6.

(2) Nose Landing Gear Extend / Retract Actuator:

Unlike the main landing gear sidebrace actuator that pushes themain gear into the extended position, the nose gear actuator pullsthe nose gear aft and down to the extended position. One end of theactuator is attached to the aft structural wall of the nose wheel well.The other end is attached to the hinged nose gear trunnion thatpivots aft into the extended position.

(3) Nose Landing Gear Downlock:

Nose gear downlock protection is provided by the over centerposition of the hinged two part truss brace with integrated downlockactuator and spring. The forward and aft sections of the truss braceunfold as the nose gear extends. As the truss brace unfolds, it forcesopen the “C” shaped suitcase springs. When the nose gear is fullyextended, a hydraulic downlock actuator and linkage pulls the trussbrace slightly (approximately 0.050 inch) over center at the hingepoint. The truss brace hinge is also held in the over center positionby the energy of the “C” shaped suitcase springs that close slightlyin the over center position. The springs are necessary to maintainthe nose gear in the downlocked position in the event of hydraulicpressure failure. Once locked into the over center position, hydraulicpressure on the retract side of the downlock actuator is required topush the truss brace past the over center position and overcomespring pressure.

The truss brace hinge has a provision for inserting a down lock pinto secure the nose gear in the extended position. See the illustrationin Figure 6.

(4) Downlock Switches:

Microswitches installed on the main and nose landing gear providedownlock electrical signals for the landing gear position indicatorson the cockpit control panel. On the main landing gear, when theinternal locking keys move into the annular groves of the actuator inthe fully extended position, a plunger with an integrated cam moves,depressing the contact of the downlock microswitch.

The downlock microswitch on the nose gear is located on the trussbrace. As the truss brace unfolds to the over center position thecontact on the nose gear microswitch is depressed sending adownlocked signal to the landing gear control panel.

D. Main and Nose Gear Uplock Mechanisms:

The uplock mechanisms of the main and nose landing gear operate in anidentical manner, although the components are shaped somewhatdifferently due to the direction of movement of the nose and main landinggear in the extension/retraction cycle. Operation of the main gear uplocksis described as typical of all landing gear installations. The landing gearuplocks are shown in Figure 2 and Figure 4.

When the main landing gear is fully retracted into the wheel well, the gearis locked into position by a hook and latch that engage an uplock rollermounted on the front of the dual wheel axle housing. The uplock hook andlatch mechanism is mounted on the interior forward side of the wheel well.

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The uplock hook is mechanically positioned by the uplock roller on thelanding gear and the uplock latch is positioned by dual springs. (The uplockhook also has dual springs that will maintain the hook in the locked positionin event of hydraulic failure.) The integrated latch assembly mechanicallyblocks movement of the uplock hook until the latch is moved aside by thegear uplock roller as the gear reaches the retracted position. When thelatch is moved by the gear uplock roller from the position blockingmovement of the uplock hook, the uplock roller enters the concave portionof the uplock hook as the hook pivots into a cradling position. Springtension on the latch drops the latch into position behind the uplock roller,securing the uplock roller in the uplock hook. The action of the spring heldlatch counteracts any tendency for the uplock roller to rebound out of theuplock hook and retains the uplock roller in the grasp of the uplock hookafter system hydraulic pressure is released.

CAUTION

WHEN THE LANDING GEAR IS EXTENDED, THEGEAR UPLOCK HOOKS ARE OPEN ANDHYDRAULIC PRESSURE IS REMOVED FROM THEUPLOCK ACTUATORS. IT IS POSSIBLE TO MANU-ALLY PUSH UP THE UPLOCK LATCH AND ROTATETHE UPLOCK HOOKS TO THE LOCKED POSITION.SUBSEQUENT RETRACTION OF THE GEAR WILLRESULT IN THE LANDING GEAR DOORS CLOSINGPREMATURELY, AND THE LANDING GEAR WILLIMPACT THE DOORS, FAILING TO RETRACT. ALLUPLOCKS SHOULD BE VERIFIED OPEN DURINGPREFLIGHT WALK AROUND INSPECTION.

3. Limitations:

A. Nose Strut Servicing:

Recommended nose strut servicing is shown in the following table.

Service Temp:Cold Day

Service (0°F / -18° C)

Std DayService

(70° F / 21°C)

Hot Day Service (110° F /43° C)

Piston Extension (inches) Strut Pressure (psig)15.51 207 248 27311.75 300 358 3948.75 460 548 6027.25 620 738 8096.50 750 892 9786.00 870 1034 11345.88 900 1070 11735.75 940 1117 12255.63 980 1164 12775.50 1030 1224 13425.38 1075 1277 1400

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Service Temp:Cold Day

Service (0°F / -18° C)

Std DayService

(70° F / 21°C)

Hot Day Service (110° F /43° C)

Piston Extension (inches) Strut Pressure (psig)5.25 1130 1342 14725.13 1190 1413 15505.00 1260 1496 16404.75 1415 1680 18424.50 1610 1911 20954.25 1870 2220 24333.36 3860 4579 5017

NOTE:

(1) Piston extension is measured from bottom surfaceof gland nut to black circumferential line on piston /axle.

(2) This table is used only for checking and adjustingstrut service pressure and not for full strut servicing.

(3) Interpolate between temperatures for properservice pressure.

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Main Landing Gear ComponentsFigure 2

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Main Landing Gear StructureFigure 3

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Nose Landing Gear ComponentsFigure 4

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Ground Service ValveFigure 5

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Landing Gear Safety PinsFigure 6

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2A-32-30: Extension and Retraction System1. General Description:

Landing gear extension and retraction is controlled with the landing gear lever onthe copilot side of the center instrument panel. The landing gear lever iselectrically connected to the extend/retract solenoid on the landing gear selector/dump valve mounted in the main wheel well. Moving the landing gear handle intothe extend or retract detent electrically signals the extend/retract solenoid toposition the selector/dump valve to route left system hydraulic pressure to theextend or retract side of actuators for the main and nose gear and doors. If leftsystem pressure is not available, but hydraulic lines and fluid quantity are intact,the Power Transfer Unit (PTU) can provide pressure to operate the landing gear.If no hydraulic pressure is available, but system lines remain intact, the landinggear may be lowered by pressurizing the lines with nitrogen gas stored inemergency bottles located in the nose wheel well. The emergency extensionsystem may only be used once and cannot retract the landing gear.

A Weight-On-Wheels (WOW) switch is installed on each landing gear. The WOWswitches are closed when the aircraft is on the ground and the pneumatic-oleostruts are compressed with the aircraft weight. The switches provide sensinginformation to aircraft systems and sub-systems that operate only in the air or onlyon the ground.

To facilitate maintenance on the landing gear or on components located in thewheel wells, the aircraft batteries can power the Auxiliary (AUX) pump topressurize left hydraulic system fluid and operate the landing gear system and/orthe landing gear doors. A switch in the forward external switch panel, located onthe left side of the aircraft just aft of the radome, controls the aircraft batteries. Aground service valve, located inside a hinged panel on the right side of the aircraftjust aft of the nose wheel well, is used to power the AUX pump and provide leftsystem pressure to the landing gear and doors. The valve is spring-loaded off,and must be held to the on position to make contact with a plunger switch thatengergizes the AUX pump. To open or close the gear doors, a manual handle oneach door selector valve is rotated and pinned in position to open the doors - theselectors are spring-loaded to the door closed position. The nose gear doorselector valve is within a panel outboard and aft of the left nose wheel door, andthe selector valves for the main gear doors are mounted on the forward bulkheadof the respective gear well. If the aircraft is raised on jacks, AUX pump pressureand left system fluid may also be used to raise and lower the landing gear with thecockpit gear control handle.

2. Description of Subsystems, Units and Components:

A. Landing Gear Control Panel:

The landing gear control panel, located on the copilot side of the centerinstrument panel, contains the gear lever, position and warning indicatorsand safety features. See Figure 7. Landing gear position and warning lightson the panel are powered by the Emergency DC bus and may be dimmedand tested with switches on the cockpit side consoles.

(1) Landing Gear Lever:

The landing gear lever is selected into either the UP or DOWNdetents on the control panel. To move the lever from one selection tothe other, the lever must first be moved to the left to clear the detent.When the lever is moved to a detent an electrical contact is made

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with the detent switch and transmitted to the extend/retract solenoidon the landing gear hydraulic selector/dump valve in the main wheelwell. The solenoid commands the selector valve to route left systempressure to the extend or retract sides of actuators for the landinggear and doors. Lever position has no effect on operation of theemergency landing gear extension system, but the position of thelever is integral to a correct landing gear position indication.

The transparent wheel shaped handle of the landing gear lever hasa red bulb for annunciation of disagreement between gear controllever position and the actual position of the landing gear or landinggear doors as determined by the uplock or downlock switches.When the gear lever is selected up, the red light illuminates until alllanding gear are up and the gear doors closed. With the gear leverselected down, the red light illuminates until all landing gear aredown and locked and doors closed.

(2) Ground Safety Lock and LOCK RELEASE Button:

On the ground with weight on the aircraft wheels (WOW switches onboth main landing gear compressed), the landing gear lever isblocked from movement out of the DOWN detent by a solenoidactuated pin. The pin prevents inadvertent retraction of the landinggear due to unintentional movement of the landing gear lever. If theblocking solenoid or WOW switches malfunction during takeoff inthe transition to the weight off wheels mode, preventing movementof the landing gear lever to the UP position, pressing the LOCKRELEASE button will manually move the blocking pin clear of thelanding gear lever.

(3) Down and Locked Lights:

Three green light modules, one for each landing gear, are installedabove the landing gear lever. Each light is an independent circuit,illuminating only when the associated gear downlock switch contactis closed. The green lights do not annunciate landing gear dooruplock engagement, and are also separate from the red warninglight in the landing gear lever handle. Thus, if a landing gear doesnot fully extend and engage the downlock switch, a red light willilluminate in the gear lever handle and the green light associatedwith the malfunctioning landing gear will not illuminate, identifyingthe malfunctioning gear. However, if a landing gear does not fullyretract (downlock not engaged, but gear door uplock not engaged)the red light in the landing gear lever handle will illuminate, but noneof the green lights will illuminate.

(4) Gear Warning Horn and HORN SILENCE Switchlight:

If the landing gear are not down and locked during flight conditionsassociated with landing, a warning horn will sound, alerting the crewthat the landing gear is not in the correct position. The gear warninghorn tone is generated by the Monitor and Warning System (MWS)and annunciated over the cockpit overhead speakers. See Figure 9.In some conditions, the gear warning horn may be silenced bypressing the HORN SILENCE switchlight on the landing gear controlpanel to the left of the landing gear lever. In other conditions, thewarning horn can be silenced only by configuring the aircraft

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correctly for landing. If the HORN SILENCE switchlight has beenpressed, the HORN SILENCE switchlight will illuminate to remindthe crew that the condition prompting the horn still exists.

The following conditions will sound the landing gear warning horn:

• The aircraft at or below three hundred forty-five (345) feetradio altitude and both power levers retarded to five degrees(5°) or less and all landing gear not down and locked. Thewarning horn may be silenced with the switchlight in thiscondition. If the power lever(s) is subsequently advanced tomore than five degrees (5°), the aircraft climbs above threehundred forty-five (345) feet, or if Calibrated Air Speed (CAS)drops below sixty (60) knots (a speed only associated withground operation in the instance of WOW system failure), theHORN SILENCE switchlight will extinguish and the hornwarning circuit will rearm.

• The flap handle selected to more than twenty-five degrees(25°) or flap position twenty-two degrees (22°) or more andall landing gear not down and locked regardless of altitude.The warning horn cannot be silenced and will sound until theaircraft configuration is corrected.

• The aircraft is on the ground (WOW switches compressed),the Air Data System (ADS) indicates less than sixty knots (60kts), and the gear handle is in the retract position. Thewarning horn cannot be silenced and will sound until thelanding gear handle is selected down (or WOW switches arein the air mode or ADS indicates more than 60 kts).

(5) The landing gear control and position switches are connected to theModular Avionics Units (MAUs) through Input / Output (I/O)modules. The modules report switch status to the MWS thatgenerates Crew Alerting System (CAS) messages that accompanythe illumination of the red light in the gear control lever. See Figure8 and the following table that lists the monitored switches and thecorresponding MAUs.

Switch Element MAU #1 MAU #2Gear Control Handle Up* XRight Main Downlock XNose Gear Downlock XRight Main Uplock XRight Main Door Up Contact XNose Gear Door Up Contact XLeft Main Downlock XNose Gear Uplock XLeft Main Uplock XLeft Main Door Up Contact XHorn Silence Switchlight X XHorn Muted (silenced) X XGear Emergency ExtensionBottle Pressure

X

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Switch Element MAU #1 MAU #2*Since the gear control lever has only two positions, only one position switch ismonitored (if the handle is not in the retract position it must be in the extendposition).

B. Normal Retraction and Extension System:

(1) Retraction Sequencing

(a) Landing gear lever is placed in UP position

(b) Red light in landing gear lever handle is illuminated

(c) Landing gear selector/dump valve is moved to RETRACTposition

(d) Hydraulic pressure is routed to OPEN side of landing geardoor actuators

(e) Landing gear doors open

(f) The timer valves route hydraulic pressure to the RETRACTside of the main landing gear sidebrace actuators, unlockingthe internal downlock keys, and to the nose gear actuator andrelease the nose gear downlock actuator

(g) The nose gear downlock actuator unlocks

(h) Three green DOWN AND LOCKED lights on landing gearcontrol panel are extinguished

(i) Landing gear retracts and locks into uplock hooks andcompress uplock switches

(j) Hydraulic pressure is routed to CLOSE side of landing geardoor actuators

(k) Landing gear doors close and compress door close switches

(l) Red light in landing gear lever handle is extinguished

(2) Extension Sequencing

(a) Landing gear lever is placed in DOWN position


(c) Landing gear selector/dump valve is moved to EXTENDposition

(d) Hydraulic pressure is routed to OPEN side of landing geardoor actuators

(e) Landing gear doors open

(f) Hydraulic pressure is routed to UNLOCK side of landing gearuplock actuators

(g) Uplock hooks unlock

(h) Hydraulic pressure is routed to EXTEND side of main landinggear sidebrace actuators and nose landing gear extend /retract actuator

(i) Landing gear extends down and downlocks engage andcompress downlock switches

(j) Three green DOWN AND LOCKED lights on landing gearcontrol panel are illuminated

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(k) Red light in landing gear lever handle is extinguished

(l) Hydraulic pressure is routed to CLOSE side of landing geardoor actuators

(m) Landing gear doors close and compress door closedswitches

C. Emergency Extension System:

The landing gear emergency extension system will operate only if thehydraulic lines for gear extension are intact. The system substitutesnitrogen gas pressure for hydraulic pressure to move the actuators of thelanding gear components. Emergency gear extension takes approximatelysix seconds to extend and lock the landing gear. Emergency extension gaspressure will not close the landing gear doors - after the gear are down andlocked, the gear doors remain open. Once activated, emergency systempressure remains in the three actuators of each landing gear (uplockrelease, door open and gear extend) until the pressure is vented overboardwhen the emergency extension T-handle is stowed. Controls for theemergency extension system are shown in Figure 10.

(1) Units and Components

(a) Nitrogen Bottles

Pressurized nitrogen used for emergency landing gearextension is stored in two bottles located in the nose wheelwell. A bottle containing one hundred fifty (150) cubic inchesof nitrogen is installed on the left side of the nose wheel welland a bottle containing four hundred twelve (412) cubicinches is located on the right side. Both bottles arepressurized to three thousand plus or minus fifty (3,000 ±50)psi at seventy degrees Fahrenheit (70°F / 21.1°C). Thebottles have a pressure relief valve set at 3,750 psi to preventdamage to aircraft components should gas pressure exceedthe structural limits of the bottle. A service panel in the nosewheel well contains a direct reading pressure gage for bothbottles and a nitrogen filler valve. The pressure gagetransmits pressure information to MAU #1 that forwards thedata to the MWS to format for display on the CAS Summarywindow.

(b) Gear Emergency Extension Cable and T-Handle

The emergency extension T-handle is located in right cockpitside wall near the copilot’s right knee. The handle is labeledEMER LDG GEAR. The T-handle is connected by a flexiblecable to the pressurized nitrogen bottles. Pulling the T-handleto full extension opens the bottle outlet valves, releasingcompressed nitrogen to the dump function of the landing gearselector/dump valve, and on through separate lines to thegear and door actuators. If the emergency T-handle issubsequently returned to the fully stowed position, a nitrogenbottle vent port opens, and nitrogen pressure in the actuatorlines is discharged through an overboard vent.

(c) Dump Valve

Nitrogen bottle pressure enters the dump function of the

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selector/dump valve and shuts off normal hydraulic pressureto the selector/dump valve and opens a return path forhydraulic fluid in the actuator lines back to the left systemreservoir. The action prevents a hydraulic lock in the gear anddoor extension lines. Nitrogen then pressurizes the unlockport of the three gear uplock actuators, the open port of thethree door actuators and the extend port of the three landinggear actuators.

When the dump function is activated, an electrical signal issent to the cockpit overhead panel to a pushbutton labelledLDG GR DUMP V. The DUMP legend within the pushbuttonwill illuminate blue confirming that normal system hydraulicpressure has been bypassed.

If emergency extension of the landing gear with thepressurized nitrogen bottle has been satisfactorily completed,the landing gear would normally be left extended until alanding could be made. If however, the landing gear must beretracted and hydraulic fluid and left system pressure isavailable, resetting the dump valve will allow retraction of thelanding gear. The emergency T-handle must first be stowedvent the pressurized nitrogen in the actuator lines.Depressing the illuminated LDG GR DUMP V pushbuttonresets the selector/dump valve to the normal configuration(the DUMP legend within the pushbutton will extinguish),allowing left system hydraulic pressure to the retract ports ofthe gear actuators and the close side of the door actuatorswhen the gear handle is selected up to the retract detent.

(2) Landing Gear Emergency Extension Sequence

(a) Landing gear lever is placed in DOWN position


(c) EMER LDG GEAR T-handle is pulled to full limit of travel

(d) Compressed nitrogen is released to landing gear selector/dump valve

(e) Dump valve isolates landing gear extend lines fromremainder of hydraulic system and opens returns forhydraulic fluid in actuators to left system reservoir

(f) Pressurized nitrogen is routed to OPEN side of landing geardoor actuators, UNLOCK side of landing gear uplockactuators, EXTEND side of main landing gear sidebraceactuators and nose landing gear extend / retract actuator

(g) Landing gear doors open

(h) Uplock actuators unlock

(i) Landing gear extends down and locks

(j) Three green DOWN AND LOCKED lights on landing gearcontrol panel illuminate

(k) Red light in landing gear lever handle extinguishes

(l) Landing gear doors remain open

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D. Weight On Wheels (WOW) Switch System:

The Weight-On-Wheels (WOW) switch system provides in-air or on-groundsensing for aircraft components that operate only in specific regimes.WOW signal distribution is illustrated in Figure 11 and Figure 12. Sensing isprovided by three switches installed on the aircraft landing gear: one oneach main gear and one on the nose gear. The switches are hermeticallysealed and are equipped with gold contacts for reliability. The main landinggear switches are positioned on the side of the shock struts and are closedwhen the shock strut torque arm rotates when the aircraft weightcompresses the switch arms. The nose gear switch is wired differently, withthe switch arms are closed in the air and open when the aircraft is on theground - wiring logic is reversed to provide a signal compatible with theswitches on the main landing gear.

NOTE:

For purposes of clarity in this section and throughoutthis manual, the nose gear WOW switch is discussedas if the operation were identical to the main landinggear switches - i.e., that the nose landing gear WOWswitch is closed or compressed with the aircraft weighton wheels to give an on-the-ground signal. Thiseditorial convention is necessary to avoid mentioningthe difference in switch operation whenever a systemthat requires WOW switch input is discussed.

Each WOW switch has dedicated control authority over designated systemoperations. Control is exercised by a relay associated with each switch.The relays have multiple internal pole dual air / ground contacts to provideair / ground sensing to the designated systems. In addition, the sum of themain landing gear WOW switch signals is used for control authority forother systems that require additional confirmation of aircraft ground / airstate for operation. The summed switch signal is termed Combined WOW.

WOW relays provide sense signals powered by twenty-eight volt directcurrent (28v DC) from either the right and left essential DC buses and/orthe left emergency DC bus or Right Battery bus. The following tableprovides the power sources and signal outputs of the WOW switch relays:

Power Source WOW Relay Signal OutputRight Ess DC bus Right Main WOW Brake Control and

IndicationRight Ess DC bus Right Main WOW Engine Thrust

ReversersRight Ess DC bus Right Main WOW Ground SpoilersLeft Emerg DC bus Right Main WOW Ground Service Bus

ground only feed toEngine Oiler and FQSCFuel Quantity Gages

Left Emerg DC bus Right Main WOW Combined WOW signalto MAUs #1 and #3

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Power Source WOW Relay Signal OutputRight Battery bus Right Main WOW Ground Service Bus

power switching fromMain DC / External DCto Right Battery

Left Ess DC bus Left Main WOW Brake Control andIndication

Left Ess DC bus Left Main WOW Engine ThrustReversers

Left Ess DC bus Left Main WOW Ground SpoilersLeft Emerg DC bus Left Main WOW Combined WOW signal

to MAUs #1 and #3Right Ess DC bus Left Main WOW Left Bus Power Contol

Unit for Power Up andGround Self Test

Left Emerg DC bus Nose Wheel WOW Nose Wheel SteeringLeft Emerg DC bus Nose Wheel WOW AUX Hydraulic Pump

Output to StandbyRudder

Left Emerg DC bus Nose Wheel WOW Nose Wheel WOWsignal to MAUs #1 and#3

R Essential DC bus Ground Service WOW Fuel QuantityEngine Oiler

The summed left and right main landing gear Combined WOW signal isdistributed to a wide range of aircraft system relays to provide air / groundsense for correct operational performance. The Combined WOW signal isprovided by a single wire connection that furnishes a twenty-eight volt (28v)essential DC current whenever both main landing gear WOW switchesindicate that the aircraft is on the ground. Only a positive indication of theon-the-ground state is necessary, since the using system relays arenormally loaded to the in-the-air position - an electrical signal is required toposition the relays to the on-the-ground state. The following table lists therelays and associated systems using the Combined WOW signal:

Combined WOW Signal DistributionRelay Number Associated Systems

Combined Relay #1 Bleed Air ControlCockpit Clocks

Combined Relay #2 Galley PowerCockpit Voice RecorderCockpit Clocks

Combined Relay #3 Reserved for future useCombined Relay #4 Probe Heat

Water SystemCombined Relay #5 Nose Wheel Steering

Probe HeatCombined Relay #6 Cabin Window Heat

TCASBleed Air Control

Combined Relay #7 APU ControlLanding Gear Control and Indication

Combined Relay #8 (Not Installed)

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Combined WOW Signal DistributionRelay Number Associated Systems

Combined Relay #9 Windshield Rain Removal BlowerAir Flow Control

Combined Relay #10 (Not Installed)Combined Relay #11 Display Units and Pedestal Cooling

Air Flow ControlCombined Relay #12 (Not Installed)Combined Relay #13 Electronic Display System

Weather RadarCombined Relay #14 Electronic Display System

Weather RadarTransponder

Combined Relay #15 (Not Installed)Combined Relay #16 (Not Installed)Combined Relay #17 (Not Installed)Combined Relay #18 Cabin Pressure ControlCombined Relay #19 Flight Data RecorderCombined Relay #20 Transponder

Flight Management SystemEnhanced Vision System

Combined Relay #21 Crew Alerting SystemFlight Management System

Combined Relay #22 (Not Installed)Combined Relay #23 Cabin Window Heat

Landing Gear Control and IndicationCombined Relay #24 Crew Alerting System

Flight Management System

The WOW system is designed to protect the aircraft by wiring the WOWrelays and contingent systems to the in-the-air state, requiring positiveflight crew intervention prior to system operation if WOW relay inputbecomes inoperative. For example, the relays controlling the engine thrustreversers require a WOW on-the-ground signal for deployment. If theWOW switch system is inoperative, thrust reversers will also be inoperativeon the ground, but the consequence of the loss of thrust reversers forlanding is less severe than if the thrust reverser WOW relays were wired inthe opposite configuration that would allow the thrust reversers to deploy inflight if the WOW system relays failed.

NOTE:

Failure of the WOW system will result in inoperativeand/or abnormal operation of aircraft systems,depending upon whether the WOW system fails to theground or flight mode. Consult the appropriateabnormal procedure based upon system indicationsand CAS messages.

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3. Controls and Indications:

A. Landing Gear Indications on Cockpit Display Windows:

In addition to the green lights on the control panel and red position warninglight in the gear handle, landing gear position and system operation arepresented graphically on the Flight Controls Synoptic 2/3 window display.When the landing gear are down and locked, the wheels and strut of eachgear are depicted in green. When the gear are in transition from either thedown or up position, each strut and corresponding wheels are shown inmagenta, a hue related to the red light in the landing gear control handle.When the gear are up and locked, they are not depicted on the display. Ifany (or all) of the gear are in disagreement with the position of the landinggear control handle, the malfunctioning gear(s) is shown in amber.

The status of the WOW switch system is also displayed on the FlightControls Synoptic window. Beneath each landing gear is a display spacefor the letter A or G indicating that the corresponding WOW switch is in theair or ground position. The Combined WOW signal status is shown belowthe individual gear WOW switch status letters. If the individual gear andCombined WOW switch signals are valid, the letters A or G are shown inwhite. If any of the WOW switch signals are invalid, the A or G is shown inamber and an amber crosshatch is displayed on the corresponding wheeland tire outline on the display (and invalid Combined WOW signal causesan amber crosshatch display on the nosewheel outline).

A digital readout of the nitrogen pressure in the emergency gear extensionbottle is presented on the Summary synoptic 2/3 window display. Thepressure readout is shown in a range from zero (0) to four thousand(4,000) psi, with any reading below two hundred (200) psi, shown as zero(0). The numerals of the readout are shown in white when the pressure isbetween two thousand four hundred and fifty and three thousand fivehundred and twenty (2,450 - 3,520) psi. When the readout is outside of thisrange, the numerals are depicted in amber. If the bottle pressuretransmitter is faulty, the digital readout is replaced with amber dashes.

NOTE:

Pressure within the nitrogen bottle will decrease to aslow as two thousand (2,000) psi during extendedflights at high altitude - pressure will return to thenormal range when the aircraft descends to altitudeswith warmer temperatures.

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B. Circuit Breakers (CBs):

The following circuit breakers protect the operation and indication of thelanding gear and associated WOW system elements.

NOTE:

The nose wheel WOW switch shares a circuit breakerwith the Combined WOW system since both arepowered by emergency DC and neither requiressignificant amperage (the circuit breaker is rated atfive amps).

Circuit Breaker Name CB Panel Location Power SourceLDG GEAR IND REER B-9 Right Emergency DC busLDG GEAR CONT POP E-4 Essential DC busLEFT WOW POP C-1 Left Essential DC busCOMBINED WOW(and nose wheel WOW-see note above)

POP C-2 Left Emergency DC bus

RIGHT WOW CPOP C-1 Right Essential DC bus

C. Crew Alerting System (CAS) Messages:

The following CAS messages are associated with the landing gear andWOW systems:

Area Monitored: CAS Message: Message Color:Landing Gear Control Handleand Aircraft Configuration Aircraft Configuration Red

Main Gear Door Uplocks Main Gear Door Open,L-R* Amber

Main Gear Downlocks Main Gear Not Down,L-R* Amber

Main Gear Uplocks Main Gear Not Up, L-R* AmberNose Gear Door Uplock Nose Gear Door Open* AmberNose Gear Downlock Nose Gear Not Down* AmberNose Gear Uplock Nose Gear Not Up* AmberGear WOW Switch SignalsLogic / Switch Comparison WOW Fault Amber

Gear WOW switch status andlogic WOW Fault Blue

No WOW signal to MAU(s) andan Air Data Module Miscompare WOW Power Fail Blue

* CAS messages associated with landing gear and/or door position are notpresented until fifteen (15) seconds after the landing gear control handle is placed ina detent - the delay allows for completion of the extension or retraction cycle.

D. Other Annunciations:

Annunciation Cause or MeaningAural warning horn sounds (gear positionnot corresponding to aircraft landingconfiguration)

Landing gear unsafe

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4. Limitations:

A. Landing Gear Extended Speed (V LE / MLE):

Do not exceed 250 KCAS / 0.70 MT with landing gear extended (geardoors open or closed).

B. Landing Gear Operation Speeds (V LO / MLO):

(1) Normal Operation:

Do not lower or raise landing gear at speeds in excess of 225 KCAS/ 0.70 MT.

(2) Alternate Operation:

Do not lower landing gear using alternate system at speeds inexcess of 175 KCAS.

C. Landing Gear Extended / Operating Altitude:

Maximum operating altitude for extending landing gear or flying withlanding gear extended is 20,000 ft. MSL.

D. Speed Brakes With Landing Gear Extended:

Speed brakes are not approved for extension with flaps set at 39° (DOWN)or with the landing gear extended in flight.

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Landing Gear ControlPanel

Figure 7

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2A-32-00 / 28

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Landing Gear Control andPosition Interface

Figure 8

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2A-32-00 / 30

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Gear Warning AudioInterfaceFigure 9

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2A-32-00 / 32

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2A-32-40: Wheels and Brakes System1. General Description:

The G550 landing gear incorporates dual wheels on each main and nose strut.The nose gear wheels are free to rotate, with no braking system installed. Themain gear wheels are fitted with a hydro-mechanical analog braking system withthe cockpit brake pedals mechanically connected to valves that apply hydraulicpressure to the brakes in response to brake (rudder) pedal displacement. Thebraking system incorporates an electronic sub-system that offers full anti-skidprotection with additional features to prevent any brake pressure application to thewheels prior to touchdown or locked wheels during landing rollout. The brakesystem uses pressure from the left hydraulic system and, if the left engine or

Emergency Gear Extension ControlsFigure 10

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THIS PAGE IS INTENTIONALLY LEFT BLANK.


Combined Weight-On-Wheels (WOW) System

Figure 11

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2A-32-00 / 36

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Weight-On-Wheels(WOW) System

Figure 12 (Sheet 1 of 2)

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2A-32-00 / 38

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Weight-On-Wheels(WOW) System

Figure 12 (Sheet 2 of 2)

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2A-32-00 / 40

July 15/04

hydraulic pump is not operating, the Power Transfer Unit (PTU) or Auxiliary (AUX)electric pump can pressurize left system fluid to provide braking. If left systempressure or fluid is not available due to a combination of malfunctions, the brakingsystem is supplemented with an independent emergency / parking brake systemthat uses stored accumulator pressure for manually modulated brake applicationwithout anti-skid protection.

2. Nose Landing Gear Wheels And Tires:

The nose landing gear has dual wheels. Each wheel is twenty one inches indiameter and seven and a quarter inches thick (21 x 7.25 in) and made of twoforged aluminum alloy halves. The two wheel halves are mated together by tenbolts, with the inner joint fitted with an O-ring seal, forming an airtight structure formounting tubeless tires. The tires are 21 x 7.25 -10 12 ply rated at 225 mph. Eachwheel weighs sixteen pounds (16 lbs) and each tire twenty pounds (20 lbs).

3. Main Landing Gear Wheels and Tires:

The main landing gear wheels are forged aluminum and consist of a hub with aremovable flange to facilitate tire servicing. The flange is mated to the wheel withan O-ring that provides and airtight seal for the tubeless tires. Each wheel ismounted to the landing gear axle with two tapered roller bearings. A disc brakeassembly is integrated into the wheel, fitting into the space between the bearinghousing and inside wheel rim. Each wheel assembly weighs eighty three pounds(83 lbs) and each brake assembly weighs one hundred seven pounds (107 lbs).The main gear wheels are fitted with H35x11.0-18 20 ply tires rated at 225 mph.Each tire weighs eighty two pounds (82 lbs).

Since wheel braking friction generates considerable heat, the wheels are fittedwith four (4) fusible plugs, spaced equally around the wheel circumference, thatmelt at 390°F (199°C) to release tire pressure if the wheel overheats. An increasein tire temperature produces an increase in tire pressure (Boyle’s law), so eachwheel also incorporates a safety plug that will deflate the tire if internal tirepressure exceeds 412.5 ±37.5 psi. (Both type plugs are installed to compensatefor tire under or over-inflation.)

4. Brake Assemblies:

Brakes have four rotating discs (rotors), three stationary discs (stators), an endplate and a pressure plate. All of these elements are composed of carbon-metallicalloy, and are referred to as a whole as the disc stack or heat pack. The statorsare attached to a torque tube that in turn is bolted to the aircraft landing gearassembly and holds the wheel bearings. The rotors are attached to the wheel,with notches in the rotors fitting keys on the interior of the wheel hub. The wheelwith attached rotors turns on the wheel bearing, with the rotors spinning betweenthe brake stators. Brake bolts attached at the brake housing and fastened to theoutside of the end plate maintain sufficient clearance between the rotors and thestators to permit the wheels to turn freely. Five hydraulic actuating pistons are builtin the brake housing. When the brakes are activated, hydraulic pressure isincreased within the pistons, moving them outward from the housing compressingthe pressure plate, and reducing the clearance between the rotors and stators.The surfaces of the rotors and stators are pressed against each other, producingfriction that slows the rotating wheel. Brake wear indicators are installed on eachwheel. The indicators are pins that extend from the inner stator through thechassis and outboard, protruding from the brake assembly. The amount of pinprotrusion is indicative of the combined thickness of the rotor and stator discs.

Main wheel brakes are equipped with temperature monitoring and anti-skid

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protection, and also automatically apply hydraulic pressure at a reduced level(400 ± 50 psi) to stop wheel spin as the landing gear are retracted into the wheelwell.

5. Brake Operation:

The brake system applies up to 3000 psi hydraulic pressure from the left hydraulicsystem to the main landing gear brakes in response to cockpit brake pedalcommands. (See the diagram of the brake operation in Figure 13.) The left andright pilot and copilot brake pedals are mechanically linked to the left and rightbrake metering valves in the nose wheel well. The metering valves open inresponse to brake pedal deflection, porting increased pressure with greater brakepedal application. The left and right brake metering valves are plumbed to boththe left and AUX hydraulic systems. Immediately downstream of the meteringvalves in the nose wheel well are two shuttle valves (one for each pilot and copilotmetering valve) that shift between either left or AUX system pressure. The valvesare installed to avoid the necessity of plumbing lines for both left and AUX systempressures aft to the wheel wells.

If left system pressure drops below 1500 psi during braking, internal stops in themetering valves shift, allowing AUX system pressure though the shuttle valves tooperate the brakes with a slight increase in pedal effort required. The AUXhydraulic pump will pressurize automatically (if armed) with the loss of left systempressure below fifteen hundred (1,500) psi and subsequent deflection of the brake(rudder) pedals ten degrees (10°) or more, closing limit switches. The limitswitches in combination with the left system low pressure switch complete thecircuit to power the AUX pump. When AUX pressure is in use, the augmentervalve is bypassed, and pressure is supplied to the brakes through the anti-skidcontrol valves and the shuttle valves.

During operation with normal left system hydraulic pressure, hydraulic fluid isrouted to the augmenter valve in the main wheel well. The augmenter valveincreases hydraulic fluid flow rates and intensifies hydraulic pressure tocompensate for any system response loss caused by the distance between thebrake metering valves in the nose wheel well and the main landing gear. From theaugmenter valve, hydraulic pressure then passes through anti-skid control valvemodules, hydraulic line fuses, pressure transmitters and shuttle valves to eachbrake piston assembly.

The anti-skid control valves electronically regulate hydraulic pressure to the brakeassemblies to maintain wheel rotation during braking. Hydraulic fuses detect abroken or damaged hydraulic line as indicated by increased fluid flow and close toprevent system fluid loss. Pressure transmitters installed in the brake lineselectronically provide readings to the MAUs. The transmitters in the lines to theleft outboard and right inboard wheels report either left hydraulic or emergencybottle pressure applied to the brakes. The left inboard and right pressuretransmitters report only left system pressure application. (This arrangementavoids unnecessary duplication of hydraulic fittings, since emergency brakepressure application is uniform to inner and outer wheels.) The left brakes aremonitored by MAU #1, the right brakes by MAU #2. The MAUs communicateapplied pressures to the MWS for display on the Brakes system page.

The shuttle valves on each wheel brake installation open in response to either lefthydraulic system or emergency / parking brake accumulator pressure. In normaloperations, left system pressure displaces the valves and operates the aircraftbrakes. In the absence of left system or AUX pressure, the shuttle valve is

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displaced by emergency brake accumulator pressure when the emergency /parking brake handle in the cockpit is activated.

If left or AUX hydraulic pressure are not available, a parking / emergency brakesystem with stored accumulator pressure may be used to stop the aircraft. Thebrake accumulator is in the nose wheel well, and has a nitrogen gas pre-charge of1200 psi. The accumulator is shown in Figure 15. The accumulator is fullypressurized to three thousand (3,000) psi. by the AUX pump. Accumulatorpressure may be read on the accumulator in the nose wheel well, on the indicatoron the copilot lower instrument panel or on the Summary 2/3 synoptic windowsystem display. The accumulator, when fully charged, contains sufficientpressurized hydraulic fluid for approximately 5 to 6 brake applications. The PARK/ EMER BRAKE handle (see Figure 16) on the cockpit center console isconnected by cable to the modulating valve on the right side of the nose wheelwell. Emergency brake pressure is applied simultaneously to all four (4) wheelbrakes - some modulation can be accomplished by pulling out the handle between¼ and ½ inch. Handle travel beyond approximately ½ inch results in theapplication of full accumulator pressure. No anti-skid is available with emergencybraking. To apply accumulator pressure for parking the aircraft, the PARK / EMERBRAKE handle is pulled up and then rotated clockwise, latching the handle in theapplied position. To release the PARK / EMER BRAKE, pull the handle up androtate counterclockwise allowing the handle to retract to the stowed position.Parking / emergency brake operation is monitored by a pressure switchdownstream of the accumulator valve. When the PARK / EMER BRAKE handle isextended to open the valve at the accumulator, the pressure switch will trip whenpressure in the parking / emergency brake line exceeds one hundred sixty-five(165) psi. The switch reports system activation to MAU #1 that in turn forwardssystem operation notification to the MWS for generating a CAS message oncockpit window displays. The pressure switch will reset when line pressure isreduced to less than one hundred thirty (130) psi, and the MWS will remove theCAS message from the display. (MWS will also generate an aircraft configurationCAS message if the power levers are advanced for takeoff with the parking /emergency brake on.)

An anti-rotation feature is incorporated in the braking system to stop main wheelspin prior to gear retraction into the main wheel wells. When the landing gearcontrol handle is selected to the retract position, a microswitch in the gear controlpanel sends a signal to the anti-skid Electronic Control Unit (ECU) to open theanti-skid control valves, allowing hydraulic pressure to the main wheel brakes.Hydraulic pressure from the nose landing gear retract line is routed through a portin the brake modulating valve and applied to the wheel brakes for three (3)seconds. The pressure applied to the brakes is reduced by the increasedhydraulic flow requirements during gear retraction so that normal system pressureof three thousand (3,000) psi is lowered to provide a moderated braking effect.After three (3) seconds, the anti-skid control valves close, preventing any furtherapplication of pressure to the main wheel brakes.

6. Brake Anti-Skid Protection:

Anti-skid protection for the main landing gear brakes operates by regulatinghydraulic pressure applied to the brakes to maintain wheel rotation. An overviewof the anti-skid system is shown in Figure 14. Anti-skid is effective at taxi speedsabove ten (10) kts. (Below 10 knots, the main wheel brakes may be locked bybrake pedal commands in order to turn the aircraft within a minimum radius.)Hydraulic pressure is metered by the anti-skid control valve modules in response

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July 15/04

to commands from the anti-skid ECU. The ECU, located in the left avionicsequipment rack aft of the cockpit, monitors wheel speeds using data shared withthe Wheelspeed Monitor Unit (WMU). Each main landing gear wheel is equippedwith a transducer mounted on the wheel axle that is spun by the rotation of thewheel. The transducers act as generators, producing an alternating current (AC)signal at a frequency proportional to wheel revolution rate. The AC frequencies ofeach wheel are transmitted to the WMU and the ECU The WMU provides anoversight function for the transducers on each wheel, generating a fault messageif both transducers on each main gear do not report congruent information.

The WMU also uses wheel transducer speed data to supply a supplementaldiscrete signal to enable the activation of the engine thrust reversers and groundspoilers should there be a failure in the Weight-On-Wheels (WOW) sensingsystem. The discrete signal is provided to reversers and spoilers whenever atleast one wheel on each strut spins up to fifty-three knots (53 kts). The signal isinterrupted when wheels slow to less than forty-two knots (42 kts) - if reverser andspoiler activation was initiated by only the WMU discrete, both will stowautomatically upon slowing to this speed.

A. Anti-Skid Electronic Control Unit (ECU):

The ECU is a digital electronic component that receives multiple inputsfrom aircraft systems to determine the effectiveness of wheel braking onaircraft deceleration. Information is derived from:

• Landing gear control handle position

• Left and right main gear downlocks

• Left and right main gear weight-on-wheels (WOW) switches

• Wheel rotation speeds from each inboard and outboard wheel onboth main gear struts

• Aircraft speed from Micro Inertial Reference Unit (IRU) #1

Landing gear control handle position, gear downlock switch signals and air/ ground (WOW) sensing are used to determine the control mode for anti-skid valve modulation. IRU sensed aircraft speed is compared to wheeltransducer speeds to detect wheel skid during ground braking.

The ECU has three circuit boards for monitoring wheel speeds andcontrolling brake application: one board for the inboard wheels on eachstrut, and one board for the outboard wheels on each strut and a thirdcircuit board used for fault detection and Built-In-Test (BIT) functions. Eachboard is separately powered by the essential DC busses with a dedicatedcircuit breaker in order to provide system redundancy. Failure of one boardwill not effect symmetric anti-skid protection by the remaining board (if theboards were each to control the inboard and outboard brakes on a gear, aboard failure would cause directional control difficulties with only theremaining good board releasing both wheel brakes on a gear during askid). The board for the outboard wheels is powered by the left essentialbus, the inboard wheel circuit board is powered by the right essential busand the monitor board is powered by the left essential bus.

The ECU compares the frequency of the AC signals produced by therotation of the wheels to the speed of the aircraft over the ground asdetermined by the IRU to detect a skidding wheel (wheel speeddeceleration faster than overall aircraft deceleration rate). When the ECUdetects a rotational speed difference it signals the anti-skid control valve for

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that wheel to reduce hydraulic pressure to the brake to eliminate the wheelskid. The amount of hydraulic pressure reduction is proportional to thedeceleration rate difference. Anti-skid protection endeavors to keep wheelrotation velocity within two percent (2%) of IRU sensed ground speed. Anti-skid does not operate below ten (10) knots - this allows the flight crew tolock both main wheel brakes on one side to accomplish a tight radius turnif necessary.

B. Anti-Skid Control Valves:

Hydraulic pressure to each wheel brake is modulated by an anti-skidcontrol valve. The valves are mounted in pairs on a common module, onemodule installed in each respective main landing gear wheel well. Thevalves reduce the hydraulic pressure commanded by flight crew actuationof the rudder pedal brakes to prevent tire damage and preserve aircraftdirectional control during braking. Each inboard or outboard valve iselectrically connected to the respective circuit board in the ECU for valvemodulation commands. The control circuitry is designed such that thefailure of a ECU board would preserve symmetrical anti-skid protection forthe valves and brakes controlled by the remaining circuit board.

C. Locked Wheel and Hydroplaning Protection:

Additional protective features are provided by the anti-skid system to avoidflat-spotted and/or blown tires: touchdown protection and locked wheelprotection. Touchdown protection prevents hydraulic pressure fromreaching a wheel brake until the aircraft is established on the ground andthe wheel has begun to rotate and achieve a speed sufficient to confirmthat a malfunction has not occurred that prevents the wheel from turningfreely. Touchdown protection is initiated when the landing gear controlhandle is selected to the extend position. The anti-skid ECU compares thecontrol handle position with the gear downlock switch position to confirmthe position of the landing gear, and also monitors wheel rotation speed. (Itis normal for the wheels to spin slightly when the gear is extended in flightdue to aerodynamic forces.) The ECU determines that the aircraftconfiguration indicates that a landing is in progress, and establishes thetouchdown protection mode. The mode prevents the anti-skid controlvalves from opening to port hydraulic pressure to a wheel brake until thewheel spins up to a velocity within fifty (50) knots of the present aircraftground speed as determined by the IRU input to the ECU. This protectionmode prevents landing with brake pressure inadvertently applied to thewheel brakes. Note that this mode does not require any confirmation of anon the ground state by the WOW switches.

A backup touchdown protection mode is available should IRU groundspeeddata or wheelspeed data not be available to the ECU. If IRU groundspeedis not available, the ECU will block hydraulic pressure to the brakes untilthe WOW switches confirm that the aircraft is on the ground, and continuesto withhold hydraulic pressure for an additional seven (7) seconds afterWOW switches indicate the on-the-ground state to protect the tires in caseof an aircraft bounce during landing. If IRU groundspeed is available, butWOW switch failure prevents an on-the-ground signal to the ECU,hydraulic pressure is blocked to a wheel until the wheel rotation indicates aspeed within fifty (50) knots of the IRU reported groundspeed. (Rotation ofthe wheel to within the groundspeed parameter could only be achieved ifthe aircraft were on the ground.)

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Once the aircraft is established on the ground in a normal brakingcondition, a similar protection acts as a backup to normal anti-skidprotection to avoid locked wheels on the ground during roll out. If arotational speed difference of approximately thirty percent (30%), asmeasured between inboard wheels or between outboard wheels isdetected, hydraulic pressure is removed from the slow wheel brake untilwheel rotation recovers to match the speed of the corresponding wheel onthe opposite landing gear. This feature is useful in circumstances of severehydroplaning or icing conditions where wheel rotation is minimal ornonexistent. Removing all hydraulic brake pressure to the locked or non-rotating wheel prevents damage / blow outs when the aircraft exitsconditions of very low friction (standing water or ice patches) during rollout.

Touchdown and locked wheel / hydroplaning protection is prevented oncethe aircraft has slowed to twenty-five (25) knots or less.

7. Brake Temperature Monitoring:

Braking effectiveness is greatly reduced when the temperature within the rotorsand stators of the brakes is elevated. Brake temperatures of each wheel aremonitored to detect loss of efficiency and as an indication of a malfunction such asa dragging brake. A temperature sensor, termed a Resistance TemperatureDevice (RTD), composed of a platinum wire is installed on each brake. As thetemperature of the brake rises, the wire loses resistance. A constant three (3) mAreference current is provided to each wire by the MAUs (left inboard and outboardwheels by MAU #1, right wheels by MAU #2) and a return voltage monitored bythe furnishing MAU. Since resistance changes linearly with temperature, theMAUs are able to determine the temperature of each brake by comparing thereference current with the return current. Data is provided by the MAUs to theMWS that formats temperature information for display on the Brakes system 1/6window display. The MWS also stores a reading of the highest brake temperatureindication during each flight cycle for display on the Brakes system page for useby the flight crew in determining brake cooling time requirements. The MWS willdelete the highest previously recorded reading and record a new peaktemperature whenever aircraft groundspeed exceeds sixty (60) knots or when thelanding gear is selected to the extend position. The sixty (60) knot groundspeedparameter is to enable recording the highest brake temperature during an abortedtakeoff.


A. Anti-Skid Control:

Although the anti-skid is normally selected on for all operations, apushbutton switch on the cockpit center console may be used to select thesystem off if a malfunction occurs. The switch is located immediatelyforward of the trim controls and is labelled ANTI SKID (see Figure 16).Selecting the system off will illuminate the OFF legend within the switch,prompt a text message display on the Brake system 1/6 window and initiatea Crew Alerting System (CAS) message.

B. Brake System Cockpit Displays:

The emergency / parking brake accumulator pressure is indicated on adirect reading gage mounted on the copilot side of the lower forwardinstrument panel. The accumulator pressure may also be noted duringpreflight inspections by viewing the gage mounted on the accumulatorservice bracket in the nosewheel well. Additional readings of the

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accumulator pressure are shown on the Summary synoptic 2/3 window andBrake system 1/6 window displays. The System synoptic windowaccumulator display has a range of zero to four thousand (0 - 4,000) psiwith a resolution of ten (10) psi. The display will read zero (0) ifaccumulator pressure is less than two hundred (200) psi. The numericpresentation is shown in white when the pressure is within the normalrange of two thousand four hundred fifty to three thousand five hundredtwenty (2,450 - 3,520) psi - otherwise the digits are displayed in ambercolor. If a malfunction renders the pressure information invalid, thenumerals are replaced by amber dashes. (During prolonged flights at highaltitudes, cold-soaking of the accumulator will often result in pressurereadings as low as 2,000 psi - readings return to normal after a descent towarmer altitudes.) The accumulator pressure indication on the Brakesystem window display is almost identical to that on the Summary synopticwindow with the exception that the normal pressure range shown in whitedigits is changed to twelve hundred to thirty-five hundred (1,200 - 3,500)psi. The lower range parameter is to account for normal accumulatornitrogen pre-charge pressure that exists when the accumulator has beendepleted. The accumulator should return a the normal pressure ofapproximately three thousand (3,000) psi (adjusted for ambienttemperature) once the AUX pump has been activated.

The full operation of the braking system can be monitored on the Brakesystem 1/6 window display. The system display incorporates bar graphs toindicate the hydraulic pressure applied to each brake and the temperatureof each brake. Applied brake pressure is shown at the top of the windowwith a centrally located scale calibrated in psi. Green bar graphs are shownon either side of the scale representing each brake, and arranged inpositions corresponding to brake / wheel positions on the aircraft. The bargraphs rise upward to indicate applied brake pressure with the value of thetop of the graph measured against the central scale calibrations. Anypressure below two hundred (200) psi is shown as zero (0). The centralscale ranges from zero to three thousand (0 - 3,000) psi with a resolution offifty (50) psi below one thousand (1,000) psi and a resolution of twohundred (200) psi above one thousand (1,000). If a brake pressuretransmitter has failed, a red “X” will replace the green bar graph of thecorresponding brake.

If the anti-skid system is turned off or has failed, an expanded brakeapplied pressure scale is available. The expanded scale is shown as an aidto the flight crew in monitoring the precise amount of hydraulic pressurecommanded with brake (rudder) pedal depression. The expanded scalehas a resolution of forty (40) psi and is depicted whenever applied brakepressure for all brakes is less than six hundred (600) psi. If moreaggressive braking is commanded, the scale will revert to the normal range(0 - 3,000 psi) when applied brake pressure exceeds eight hundred (800)psi.

Below the applied brake pressure graphics are provisions for text display toannunciate anti-skid system status. If the anti-skid system is selected off orhas failed, an amber text message will appear to annunciate the change ofsystem status. Adjacent to the text message space is the display of brakeaccumulator pressure shown in white text and digits when the readings arein the normal range.

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Below the text line is a bar graph display of brake temperatures. The formatis identical with the applied brake pressure graphics except that the centralscale is calibrated in degrees centigrade (°C) and has an expandablerange of zero to nine hundred ninety-nine degrees (0 - 999°C) with a onedegree (1°C) resolution. Each bar graph is depicted in green attemperatures below six hundred fifty degrees (650°C) - when a braketemperature exceeds that threshold, the bar graph will be shown in amber,and will remain amber colored until the associated brake cools to sixhundred degrees (600°C) or less. The central reference scale will alsodisplay amber cross-hatching when a brake temperature exceeds sixhundred fifty degrees (650°C).

The highest brake temperature recorded during the current flight cycle isshown below the temperature bar graphs. The data includes theidentification of the brake (e.g. “LO” = left outboard), the highesttemperature for that brake and the time that the high temperature wasrecorded in universal standard time (GMT).

CAUTION

ELEVATED TEMPERATURES SEVERELY LIMIT THESTOPPING CAPACITY OF THE WHEEL BRAKES.SINCE BRAKING IS ACCOMPLISHED BY FRICTIONBETWEEN THE BRAKE ROTORS AND STATORS,HEAT ISADIRECT INDICATION OF THEAMOUNT OFFRICTION AVAILABLE WITHIN THE MATERIALCOMPONENTS OF THE BRAKE INSTALLATION. THEMAXIMUM CAPACITY OF THE G550 BRAKINGSYSTEM IS 142.3 MILLION FOOT POUNDS (MFP),WHICH IS EXCEEDED AT APPROXIMATELY 893°C,AND A COMPLETE TEARDOWN AND INSPECTIONIS REQUIRED AT 121 MFP OR APPROXIMATELY760°C.

ALTHOUGH THESE TEMPERATURE / ENERGYLEVELS SEEM EXCESSIVE AND UNLIKELY INNORMAL OPERATIONS, THEY CAN BE REACHEDVERY QUICKLY IN SCENARIOS OF ABORTING AHEAVY WEIGHT TAKEOFF AFTER A SHORTTURNAROUND PERIOD. INADEQUATE BRAKECOOLING MAY RENDER AN ABORTED TAKEOFFUNSUCCESSFUL. OPERATORS ARE STRONGLYENCOURAGED TO REVIEW THE DISCUSSION ANDCONSULT THE TABLES REGARDING BRAKE COOL-ING TIME REQUIREMENTS IN APPENDIX C OF THEG550 FLIGHT MANUAL

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C. Circuit Breakers (CBs)

The wheels and brakes are protected by the following circuit breakers:

Circuit BreakerName

CB Panel Location Power Source

OUTBD ANTISKID LEER B-18 L ESS DC busINBD ANTISKID REER B-6 R ESS DC busANTISKIDMONITOR

REER B-7 L ESS DC bus

WHEELSPEED POP C-3 R ESS DC busWHEEL BRKACCUM PRESS

REER B-8 R ESS DC bus

APPLIED BRKPRESS

LEER B-16 L ESS DC bus

D. Crew Alerting System (CAS) Messages:

The following CAS messages are associated with wheels and brakes:

Area Monitored CAS Message Message ColorEmergency / Parking Brake PressureSwitch

AircraftConfiguration

Red

Anti-skid essential DC power, circuitryor AUX pump auto operation failure

Antiskid Fail Amber

ANTI SKID pushbutton switch Antiskid Off AmberBrake Temperature(s) over 650°C Brake Overheat AmberAnti-skid system integrity Touchdown Protect

UnavailAmber

Main Wheelspeed Monitors senserotation after gear retraction

Wheel Despin Fail Amber

Wheelspin Monitor Unit / transducers WheelspeedMonitor

Amber

Anti-skid / MAUs Brake MaintenanceReqd

Blue

Emergency / Parking Brake PressureSwitch

Parking Brake On Blue

Wheelspeed Monitor Unit power orwheelspeed transducer disagreementbetween both wheels on one gear

WheelspeedMonitor

Blue

9. Limitations:

Anti-Skid System:

Takeoff is permitted with Anti-Skid system inoperative, provided the runway is dry,Ground Spoilers are operative, 20° flaps are used, and the Cowl and Wing Anti-Icing systems are not used. Dispatch with reference to the MMEL.

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Brake SystemFigure 13 (Sheet 1 of 2)

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Brake SystemFigure 13 (Sheet 2 of 2)

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Anti-Skid SystemFigure 14

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Emergency Brake Accumulator and Dump ValveFigure 15

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Brake Controls and IndicatorFigure 16

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2A-32-50: Nose Wheel Steering System1. General Description:

The G550 has a steerable nose wheel that is controlled electrically and operatedhydraulically and mechanically. The steer-by-wire system uses inputs from a tillerhandle mounted on the pilot side console or signals from rudder pedal position.The nose wheel may be turned 80° left or right of center with the cockpit controltiller, or 7° left or right using only rudder pedal input. Control tiller and rudder pedalinputs are additive, but total nose wheel displacement is limited to 80° bymechanical stops. Nose wheel steering must be selected ON with the guardedsystem switch next to the control tiller. The system switch and other systemcomponents and controls are shown in Figure 17. Nose wheel steering useshydraulic pressure from the left hydraulic system for actuation. If left systempressure is not available the Power Transfer Unit (PTU) can provide systempressure, or the Auxiliary (AUX) electrical pump can pressurize the system oncethe aircraft is on the ground with weight-on-wheels.

The tiller handle is mechanically connected to system control elements mountedunderneath the handle within the side console. When the system is selected on,moving the control tiller turns a shaft connected to a Rotary Variable DisplacementTransducer (RVDT). The RVDT provides an electrical signal of the direction andamount of displacement of the tiller handle from the centered position. Theelectrical signal is received by a microprocessor called the Electronic ControlModule (ECM). The ECM modulates the turn command signal by compensatingthe turn command for groundspeed. The ECM receives groundspeed velocitiesfrom micro Inertial Reference Units (IRUs) #1 and #2 and reduces the amount ofnose wheel displacement from center with increased speed.

The tiller turn command is transmitted to the Electro-Hydraulic Servo Valve(EHSV) in the nose wheel well. Left system hydraulic pressure is routed to theEHSV through a shutoff valve and two solenoid valves. The EHSV directs leftsystem hydraulic pressure to the required intake ports on the hydraulic motor ofthe steering unit to turn the nose wheel. The hydraulic motor drives a reductiongearbox that engages a large ring gear that encircles the top of the nosewheelsteering unit. The direction of rotation of the gearbox, controlled by the EHSV,displaces the ring gear to the left or right, steering the nose wheel.

A RVDT is mounted on the gearbox and sends a feedback signal to the ECU. TheECU compares the moderated command signal from the cockpit tiller handleRVDT to the actual nosewheel position as reported by the gearbox. When thenosewheel corresponds to the commanded position, the ECU signals the EHSVto neutralize steering pressures to the hydraulic motor.

When not in active steering mode, the tiller control handle is returned to theneutral, or straight ahead position by a centering spring and cam when the handleis released. If the system is selected off, the nose wheel is free to caster in apassive mode, responding to the tracking forces of differential braking or enginepower settings. The trapped hydraulic fluid in the system will also act as a shimmydamper.

Steering the nose wheel with rudder pedal commands is functionally identical tooperation of the tiller handle. A RVDT is located on a connector rod beneath thecockpit floor on the copilot side. Displacement of the rudder pedals left or rightprompts the RVDT to signal the steering command to the ECU. The ECU limitsthe rudder steering command such that full pedal deflection will result in amaximum nose wheel displacement of seven degrees (7°) left or right. If the tiller

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control handle is displaced from neutral, the ECU will algebraically sum the tillerand rudder pedal commands up to the maximum nose wheel displacement ofeighty degrees plus or minus one point six (80° ±1.6). If the tiller handle RVDTmalfunctions, the nose wheel steering authority of the rudder pedals is increasedto sixteen degrees (16°) left and right below twenty knots (20 kts), with theauthority linearly decreased to ten degrees (10°) at sixty knots (60 kts) and limitedto the normal seven degrees (7°) at ninety knots (90 kts) and above.

To ensure that the nose wheel remains centered prior to nose wheel retractioninto the nose wheel well, internal centering cams are incorporated in the shockstrut, and the steering system is restricted to ground operation only by severalsafeguards. The landing gear lever handle must be in the down position tocomplete the electrical circuits for steering. Hydraulic pressure for steering isprovided through the nose landing gear extend lines. The hydraulic pressure mustpass through two (2) solenoid controlled shutoff valves (SOVs) before reachingthe EHSV and steering actuator. See the schematic in Figure 18. The solenoidsfor the SOVs are powered by 28V DC through circuit breaker STEER BY WIRE(position C-8) on the Right Electronics Equipment Rack (REER). For the SOVs toopen, the guarded PWR STEER switch next to the tiller handle must be on, andthe nose gear down lock switch must be engaged. When these conditions aremet, the ECM initiates a Built In Test (BIT) of the steering system and also opensSOV #1 to allow hydraulic fluid to circulate within the system to warm thecomponents prior to active steering command. The second SOV, #2, requires thatthe nose gear WOW switch be compressed in the ground (weight-on-wheels)position for the valve to admit hydraulic pressure for the steering motor. When thenose gear WOW switch is compressed, the ECM initiates a one second delaybefore responding to steering commands from the rudder pedals or the controlwheel, maintaining the nose wheel in the centered or straight ahead direction. Thedelay avoids abrupt steering commands caused by large rudder pedal inputsduring crosswind landings.

The nose wheel steering linkage may be disconnected for aircraft towing byremoving the connecting pin on the torque link between the actuator and the nosewheel. A tow bar is then connected to the nose wheel axle. Care should be takento ensure that the nose wheel is not pivoted in excess of the eighty degrees plusor minus one point six (80° ±1.6) limit. Exceeding the displacement limit willdamage the internal mechanical stops. A pop-up indicator is installed on theforward section of the top of the steering unit to warn of an overtravel condition. Ifthe indicator is visible during preflight inspection, the nose wheel steering collarshould be removed and inspected for damage.


A. Control Panel Switches:

Nose wheel steering is selected with the guarded POWER toggle switchforward of the tiller handle on the pilot side console. Rudder pedal steeringis selected with the PEDAL DISC switch to the right of the POWER switch.The POWER toggle switch must be selected ON for the rudder pedalswitch to operate. The rudder pedal switch is not illuminated in the normalon position, but a blue OFF legend within the switch will alert the crew thatthe system is selected off. Below the OFF legend are two internalmaintenance indicators for use in performing system checks and aligningthe rigging between the RVDTs in the tiller handle and on the steeringgearbox.

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B. System Monitoring:

The nose wheel steering system is monitored by Modular Avionics Units(MAUs) #1 and #2 through ARINC-429 discrete connections. The MAUscommunicate with the Monitor and Warning System (MWS) that willformulate Crew Alerting System (CAS) messages to alert the flight crew ofany system malfunction.

C. Circuit Breakers (CBs):

Nosewheel steering is protected by the following circuit breakers:

Circuit BreakerName

CB Panel Location Power Source

STEER BY WIRE REER C-8 R ESS C Bus

D. Crew Alerting System (CAS) Messages:

CAS messages associated with nosewheel steering are:

Area Monitored CAS Message Message ColorSystem ECU, tiller RVDT, rudder pedalRVDT

Steer By Wire Fail Amber

Tiller handle RVDT (rudder pedalauthority increased)

Tiller Steering Fail Amber

Rudder pedal RVDT Rudder SteeringFail

Amber

Rudder Pedal Steering Switch Rudder Steering Off BlueNose Wheel Steering Power Switch Steer By Wire Off BlueECU and/or IRU Interface forGroundspeed Modulation of SteeringAuthority

NWS Fixed Gain Blue

E. Limitations:

There are no limitations regarding the nose wheel steering system as ofthis writing.

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Nose Wheel Steering ControlsFigure 17

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Nose Wheel SteeringSystem

Figure 18

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g550 landing gear

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