Gulfstream G550 Operating Manual

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FLIGHT CONTROLS

2A-27-10: General

1. General Description:

The aircraft ﬂight controls allow the ﬂight crew to guide the aircraft in the longitudinal, vertical and horizontal axes (see Figure 2). The primary ﬂight controls are:

• Elevator to control aircraft pitch

• Rudder to control aircraft yaw

• Aileron to control aircraft roll

The primary ﬂight controls are positioned by moving the pilot and copilot control yokes and rudder pedals. Both yokes are mechanically linked together so that either crew position has full control authority and control inputs are transparent to both crew members since movement of one set of controls will move the corresponding set of controls. Each yoke has a dedicated cable connection to the elevator and aileron control on that respective side, but since the yokes are mechanically linked, moving one side elevator or aileron will move the ﬂight controls on both sides. For instance, the copilot yoke is cable-linked to the right elevator and the right aileron but any movement of the copilot yoke also moves the pilot yoke that is in turn cable-linked to the left aileron and left elevator. This system of split control authority and linked yoke movement provides a means to maintain some ﬂight control movement if one of the ﬂight controls or associated linkages becomes jammed. If a malfunction prevents movement of either the elevators or ailerons, the mechanical links joining the two pilot yokes can be severed, thereby allowing movement of the left or right elevator or aileron that remains operational by commands using the yoke connected to the side of the free control surface.

The two sets of rudder pedals are similarly mechanically linked together, but both are connected to the rudder by a single cable. For this reason, there is no provision for interrupting the linkage between the two sets of pedals, since each set does not have an independent route to the rudder.

The cable connections from the yokes and rudder pedals are continuous loop installations, providing feedback to the moveable controls. The control cables engage bell cranks that translate cable movement into displacement commands for hydraulic actuators that boost contol inputs to move the ﬂight controls. Each hydraulic actuator has a single shaft, but dual piston chambers in order that the actuator may be driven by both (or either) left and right hydraulic system. (The hydraulic system power sources for the ﬂight controls are shown in Figure 1.) The rudder actuator may also be powered by the Auxiliary (AUX) hydraulic system in the event of dual hydraulic power failure. Each of the hydraulic actuators is connected to the associated ﬂight control by pushrods and bell cranks to impart mechanical movement to the control surface. A bungee piston ﬁlled with hydraulic ﬂuid moderates the rate of actuation of the ﬂight control and provides an artiﬁcial feel input to the ﬂight crew through the closed loop cable system. The failure of a single hydraulic system does not degrade ﬂight control operation - the remaining system provides adequate power for ﬂight control movement, and the actuator chamber for the failed system bypasses ﬂuid so there is no resistance to ﬂow. If both (or all, in the case of the rudder) hydraulic systems fail, both actuator chambers bypass ﬂuid, and pilot input through the cable connection moves the

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internal shaft of the actuator and the associated pushrods and bellcranks without powered assistance. More pilot force is required, but full movement of the ﬂight controls is attainable.

The autopilot is interfaced with the primary ﬂight controls through electric servos to move parallel cable linkages to the hydraulic actuators. Each ﬂight control surface has a Linear Variable Displacement Transducer (LVDT) that provides an electrical signal proportional to ﬂight control surface displacement from neutral. The LVDT signal provides feedback to the autopilot for positioning the ﬂight controls, and also communicates control surface position for display on systems and synoptic windows through interface with the Modular Avionics Units (MAUs). The description of the Digital Automatic Flight Control Systems (autopilot) makes up the entire content of Section 2B of this manual, and for that reason is not covered in this Section. However, for the convenience of the reader, a tabulation of the Crew Alerting System (CAS) messages associated with the autopilot is included below.

The elevators and ailerons have trim tabs to position the ﬂight controls with aerodynamic forces to moderate the amount of physical effort to maintain the control surfaces in the desired steady-state condition. The rudder is not equipped with a trim tab, but instead has a mechanical trim input to reset the neutral position of the rudder. The mechanical trim uses the hydraulic actuator to project the rudder slightly into the windstream in the desired direction to compensate for induced yaw.

The secondary ﬂight controls and functions are:

• Wing ﬂaps - enhance wing lift characteristics

• Flight and Ground Spoilers - reduce wing lift and add overall drag

• Movable horizontal stabilizer - aligns the elevator with aircraft angle of

attack

• Yawdamper - uses the autopilot rudder servo to moderate aircraft heading oscillation

Some functions of the secondary ﬂight controls are integrated with the operation of the primary ﬂight controls and other functions mutually complement the operation of other secondary controls.

• As ﬂaps extend, moving the wing center of lift aft, a downward pitch moment is created - the moveable stabilizer automatically compensates for the pitch moment by trimming downward. The opposite movement occurs as ﬂaps are retracted and the nose of the aircraft pitches up.

• When ailerons are used, ﬂight spoilers on the downward wing activate to increase roll rate and provide yaw into the turn. Ground spoilers decrease ground roll distance during landings and aborted takeoffs.

• The yaw damper provides a degree of turn coordination for aileron roll commands, provided wing ﬂaps are not selected to more than thirty degrees (30°).

Safety features incorporated into the ﬂight controls system include:

• A stick shaker warning and a stick pusher stall prevention actuator.

• A gust lock that prevents damage to ﬂight controls while the aircraft is

secured on the ground.

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2. Primary and Secondary Flight Controls Subsections:

The primary and secondary ﬂight controls are discussed in the following subsections:

• 2A-27-20: Elevator Pitch Control

• 2A-27-30: Rudder Yaw Control

• 2A-27-40: Aileron Roll Control

• 2A-27-50: Horizontal Stabilizer and Wing Flaps

• 2A-27-60: Stall Warning and Prevention System

• 2A-27-70: Speed Brake and Ground Spoilers

• 2A-27-80: Flight Controls Gust Lock

3. Autopilot Crew Alerting System (CAS) Messages:

The following CAS messages are associated with the operation of the autopilot and integrated subsystems:

Area Monitored: CAS Message: Message Color:

Flight Guidance Computer Internal Monitor

Flight Guidance Panel AP Engage Switch

AP Engage Switches AP Engage Inhibit -Sw

AP 1-2 Fail Amber AP Control Switch

Stuck Active

Blue Blue

Air Data System AP Inhibit - ADS Blue Inertial Reference System AP Inhibit - IRS Blue Control Column Force AP Inhibit - Left Column Blue Control Column Force AP Inhibit - Right

Column

Blue

Control Wheel Force AP Inhibit - Left Wheel Blue Control Wheel Force AP Inhibit - Right Wheel Blue Manual Trim Wheel AP Inhibit - Man Trim

Autopilot Quick Disconnect Switch

Active AP Inhibit - QD Blue

Blue

Stall Shaker AP Inhibit - Stall Blue Autopilot Touch Control

Steering Switch Control Wheel Electric Trim

Switch Weight On Wheels (WOW)

System

AP Inhibit - TCS Blue AP Inhibit -Trim Cmd Blue AP Inhibit - WOW Blue

Autopilot Power Source AP 1-2 Power Fail Blue Autopilot Elevator Trim Servo AP / Trim Fail Blue Flight Guidance Panel Speed

Window Flight Guidance Computer /

WOW System Take Off and Go Around

(TOGA) Engage Switch / Flight Guidance Panel Manual Speed

Check Speed Target Blue FGC - WOW Fault Blue Go Around Pitch Blue

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Flight Controls System:

Simpliﬁed Fluid Power

Diagram Figure 1

2A-27-00

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Flight Controls System Components

Figure 2

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2A-27-20: Elevator Pitch Control

1. General Description:

The aircraft has a dual elevator installation to control aircraft pitch attitude. The elevators are composed of a baked graphite-epoxy material. Each of the cockpit yokes is connected to one of the aircraft elevators. The pilot yoke is connected to the left elevator,the copilot yoke to the right elevator.Eachyokeisalsoconnected to the other by a mechanical torque tube beneath the cockpit ﬂoor. Since both yokes are interconnected, moving one yoke moves both elevators.

Braided steel cables run from each yoke to hydraulic assist actuators in the tail of the aircraft. The cables are routed beneath the aircraft ﬂoor using pulley connections to clear other installed equipment. The cables mate with the hydraulic assist actuators via bellcranks that translate pulley rotational motion into forward and aft motion. The actuators each have a single shaft powered by two piston chambers, one chamber for each (left and right) hydraulic system. Both hydraulic systems normally power the actuators, but one system is sufficient for full elevator movement. The actuators are connected to the respective elevator by linkages and bellcranks, moving the elevator up or down about the pivot points on the aft of the horizontal stabilizer. The deﬂection range of the elevators is twenty-four degrees (24°) up and thirteen degrees (13°) down.

Each connection of yoke to elevator is a continuous loop. Incorporated into the loops adjacent to the actuators is a bungee cylinder ﬁlled with viscous ﬂuid to resist yoke / elevator movement in order to provide artiﬁcial feel to each yoke. Each elevator also has a stability spring incorporated into the cable linkage to provide a forward pull to the control yoke and to contribute additional feel input.

Both sides of the hydraulic actuators are monitored to assure correct operation. The cockpit cable input motion must result in a corresponding actuator output motion, and similarly the output side of the actuator should not move without cockpit input. If input and output do not correspond, actuator hydraulic pressure is bypassed to prevent movement of the elevator.

Anytime hydraulic pressure to the actuators is bypassed or lost (in the instance of dual hydraulic system failure) the elevators remain operable with manual yoke movement that positions the actuator shaft and connecting linkages to the elevator. Control forces will be higher, since normal hydraulic assist provides a six (6) to one (1) boost advantage to move the elevator surfaces.

Each elevator is equipped with a trim tab that uses aerodynamic pressure to aid in positioning the control surface. The trim tabs are controlled manually by rotating a wheel on the cockpit pedestal or electrically using switches on the control yokes. Manual trim uses a dedicated braided wire connection from the cockpit to a mechanical linkage in the tail. Electrical switch trim movement commands an electric servo to move the same linkage.

Both the elevators and elevator trim incorporate Rotary Variable Differential Transducers (RVDTs) to feed back position information to the autopilot for elevator control and trim and to the ModularAvionics Units (MAUs) for formulation of control surface position display on the Flight Controls 2/3 synoptic page. RVDTs measure the angle of the elevators and trim tabs and transmit an electrical signal proportional to displacement from a neutral position.

When the autopilot is engaged, the elevators are positioned by electric servos that move parallel cable connections to the hydraulic actuators. The autopilot also uses the electric trim servo to move the trim tabs, minimizing hydraulic actuator force.

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If a malfunction or failure in any portion of the loopbetweenacockpityokeandthe corresponding elevator prevents control surface movement, the mechanical torque tube connection between the two control yokes can be separated to allow control of the aircraft with the free (unjammed) elevator. The autopilot may be used in single elevator operation.

2. Description of Subsystems, Units and Components: A. Elevator Hard Over Prevention System (HOPS):

Movements of the elevators contrary to the commanded position are limited by a Hard Over Prevention System (HOPS), illustrated in Figure 3. The system incorporates eight switches for each elevator to monitor mechanical and hydraulic elevator operation. Four external mechanical switches are integrated into the elevator control linkage to provide a comparison reference for four switches mounted internally within the hydraulic actuator. Of the four external switches, two are for left hydraulic system reference and two are for right hydraulic system reference. Of the two switches for each hydraulic system, one provides a up elevator command reference and the other provides a down elevator command reference. The switches are plunger-type contact switches, and are installed on each side of a bracket attached to the command input side of the elevator actuator. On one side of the bracket are the up elevator command input switches for the left and right hydraulic systems. On the other side of the bracket are the down elevator command input switches for the left and right hydraulic systems. Inserted between the switches in the bracket is a cam-type arm mated to the elevator hydraulic actuator output linkage. The cam arm is positioned with a deﬁned amount of clearance between the plunger-type switches. Under normal conditions, the bracket holding the switches moves with elevator command input and the cam arm moves with the elevator hydraulic actuator output, so the clearance between the switches and the arm is maintained.

If a malfunction occurs and the elevator moves opposite to or further from the commanded direction, the cam arm that is attached to the output linkage of the elevator actuator will move to close the clearance gap between the cam arm and the plunger-type switches, making contact with the switches on the side of the bracket. When the plungers of the switches are depressed, a relay closes and an electrical signal is sent to a corresponding set of switches mounted internally within the hydraulic actuator.

Four pressure switches monitor left and right hydraulic system pressures within the pistons of the elevator actuator. In normal conditions, all four switches sense stabilized pressures since hydraulic outputs positioning the elevator are balanced by air load pressures on the elevator surface acting against actuator pressures. When a malfunction occurs and the elevator moves contrary to the commanded direction, the hydraulic actuator shaft moves in the contrary direction, causing an increase in hydraulic pressure on the opposite sides of the pistons within the actuator. The left and right system opposite side pressure switches close, completing the circuit initiated by closure of the bracket plunger switches, and an electrical signal is sent to a timing relay. If the contrary elevator movement persists for longer than one tenth (1/10) of a second, hydraulic pressure from both left an right systems is shut off to the elevator actuator.

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Left Hydraulic System Right Hydraulic System

Elevator

←

Down

Elevator

→

Elevator

←

Hydraulic Actuator Shaft

Pressure

Switch

Pressure

Switch

Pressure

Switch

In a similar manner, if a hydraulic malfunction causes the shaft of the actuator to move in a direction opposite the commanded elevator direction, the increased pressures in the wrong direction sides of the pistons would close the monitoring switches, and movement of the actuator shaft would cause the cam arm to contact the plunger switches on the external bracket, and hydraulic pressure would be shut off to the elevator actuator after a one tenth (1/10) of a second delay.

If a hydraulic malfunction in a single system (left or right) side of the actuator moves the actuator shaft in a wrong direction, only the hydraulic pressure of the malfunctioning system is shut off. For instance in a stabilized condition, if the left hydraulic system piston attempts to move the actuator shaft in the up direction, the increased pressure in the down elevator side of the left piston will close the monitor switch and actuator shaft displacement will close both the left and right hydraulic system up elevator bracket plunger switches, completing the shut off circuit for the left hydraulic system after a one tenth (1/10) second delay. (In this case hydraulic pressure in the up elevator side of the right hydraulic system piston will decrease due to the increase in area caused by actuator shaft movement.) The elevator hydraulic actuator will continue to function using the remaining hydraulic system.

The operation of the hydraulic shut off valves by the HOPS is signaled to the MAUs (left elevator to MAU #1, right elevator to MAU #2) over ARINC429 connections. The shut off condition is monitored by the MWS, and a CAS message corresponding to the condition is displayed on the CAS window. If either or both hydraulic systems are shut off, an amber caution message of “L (or) R Elevator Hydraulics Off” is displayed. If only a single hydraulic system has been shut off, the remaining hydraulic system will provide full elevator operation. If both hydraulic systems have been shut off, manual elevator control may remain possible, depending upon the cause of the hardover condition. If the cause of the condition is thought to be momentary, and the use of the elevator is deemed necessary for continued safe ﬂight and landing, the hydraulic shut off valve(s) may be reset by cycling the RIGHT ELEV HYD S/O and/or LEFT ELEV HYD S/O circuit breaker. If the cause has not been rectiﬁed, the shut off valve(s) will close and hydraulic boost for the elevator will be unavailable if both hydraulic systems have been shut off. (HOPS is powered by the left essential DC bus for the left elevator and the right essential DC bus for the right elevator.)

To prevent the HOPS from shutting off hydraulic system pressure to the elevator during normal ﬂight maneuvers that may involve rapid changes in elevator direction, HOPS activation is buffered by three elements:

• The clearance between bracket plunger switches and the actuator

Down

Elevator

→

Pressure

Switch

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cam arm

• Inertia in the build up of hydraulic pressures in the elevator actuator

• The one tenth (1/10) second delay in electrically latching the shut off

of hydraulic pressure

B. Elevator Disconnect Handle:

(See Figure 4.) The mechanical torque tube beneath the cockpit connecting the pilot and

copilot yokes may be disconnected if a malfunction in one of the cable connections, hydraulic actuators or elevators renders the respective elevator inoperative. Disconnecting the torque tube prevents both elevators from being disabled by a malfunction in one elevator linkage.

An elevator disconnect handle is located on the pilot side of the center pedestal beneath a protective cover. The handle is connected by a cable to a pin securing the two halves of torque tube together. Pulling out on the disconnect handle removes the pin and allows each yoke to move independently. If the elevator linkage malfunction has resulted in opposite movement between the two yokes, making retraction of the mating pin difficult, a power assist gas-spring cartridge may be activated to provide additional force to remove the pin. The power assisted disconnect is activated by pulling a trigger beneath the disconnect handle, and provides an upward lifting moment of thirty-three feet per second (33 ft./sec).

After the yokes have been separated, the malfunctioning elevator is isolated and the operable elevator may be used to control the aircraft. with manual or autopilot inputs. If it is discovered that separating the yokes freed the previously malfunctioning elevator linkage, the yokes may be reconnected by pushing in on the disconnect handle when the yokes are aligned, provided that the power assist disconnect was not used.Areset of the yoke torque tube coupling is not possible without special maintenance tools after activation of the power assist disconnect.

C. Pitch Trim System:

(See Figure 4 through Figure 6.) Each elevator has a trim tab installed on the trailing edge. The trim tabs are

manufactured from the same graphite-epoxy material as the elevators, but incorporate a ceramic heating element that is continuously electrically powered to maintain a temperature of one hundred seventy-ﬁve degrees Fahrenheit, plus or minus twenty degrees (175°F±20) around the tab actuator linkage. Elevator trim heat is powered by 115V AC from the right main bus. The trim tabs have a range of movement of twenty-two degrees (22°) trailing edge down (aircraft nose up) to eight degrees (8°) trailing edge up (aircraft nose down). Limit switches are installed at the travel limits that will prompt the display of Crew Alerting System (CAS) messages notifying the crew that the elevator trim tabs are at maximum displacement. If a trim limit message is displayed while the autopilot is engaged, extreme care should be taken prior to disengaging the autopilot. An abrupt attitude change will occur if trim displacement is not moderated prior to disengaging the autopilot.

Operation of the trim tabs employs aerodynamic force to maintain the elevator in the desired position. As the trim tab is moved from the neutral position (faired with the elevator) into the airstream, the air impinging on

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the tab forces the hinged elevator into the opposite direction. As the elevator is moved from the neutral position (faired with the horizontal stabilizer), it also encounters pressure forces from the airstream, thus the amount of elevator movement from neutral is determined by a balance of airstream forces acting on both the trim tab and the elevator. Since the surface area of the trim tab exposed to the airstream is less than the surface area of the elevator, the elevator is deﬂected in much smaller increments than trim tab displacement (excluding other factors, trim tab effectiveness is a ratio of tab surface area to elevator surface area).

The ﬂight crew manually controls the amount of trim tab deﬂection by moving control wheels on either side of the center console. The wheels are hubs connected to a common axial shaft, so that moving one wheel moves the other. The shaft is connected to a continuous loop of wire cables that connect through a series of pulleys and bellcranks to the elevator trim tabs. Rotating a trim wheel forward positions the trim tab up forcing the elevator down resulting in an aircraft nose down moment. Trim wheel rotation aft results in an aircraft nose up moment.

The ﬂight crew has the option of electrically moving the elevator trim tabs. A pushbutton, labelled PITCH TRIM ENG / DISENG, located to the left of the standby ﬂight instruments on the lower instrument panel enables electrical operation of trim switches mounted on the outboard side of the control yokes. Electric pitch trim is normally engaged. The amber DISENG legend in the pushbutton will illuminate if the button is not pushed in to engage electric trim. The yoke trim switches are composed of split halves. Both halves must be moved in the same direction to move the elevator trim. The split switch design helps to prevent accidental trim input. The switches are wired to an electric servomotor that is located in the tail of the aircraft and incorporated into the cable linkage to the trim tabs. An electric signal from a cockpit trim switch results in the servomotor rotating an attached pulley, moving the elevator trim cable loop in the desired direction. The manual trim wheels on the pedestal will rotate with electric trim inputs, since the control cabling is a continuous loop.

When the autopilot is engaged, autopilot servomotors move the elevators and also control elevator trim with the same servomotor employed by the yoke electric trim switches. The PITCH TRIM ENG / DISENG switch is automatically engaged whenever the autopilot is engaged (to enable autopilot trim). However the reverse is not true - electric pitch trim will not disengage when the autopilot is disengaged, but must be selected off with the switch.

Elevator trim must be within a deﬁned range for takeoff, with the speciﬁc setting within the range determined by aircraft Center of Gravity (COG) and takeoff gross weight. The limits of the acceptable pitch trim range are eight degrees nose up to nineteen degrees nose up (8° - 19° up). The takeoff elevator trim range is marked in green on the trim setting indices of the manual trim wheels at each side of the center pedestal. The same limits are shown in a green band on the trim scale of the Flight Controls 2/3 synoptic page. Failure to set pitch trim within the deﬁned range will result in a warning annunciation and CAS message as the power levers are advanced for takeoff.

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D. Mach Trim:

The autopilot employs elevator control to provide the aircraft with Mach trim. Mach trim is necessary because at high speed ﬂight the center of lift on the wing transits aft with increases in speed, producing a nose down pitch moment termed Mach tuck. The autopilot electrically repositions the elevator trim to neutralize the nose down force. Mach trim is an automatic function of the Flight Guidance Computers.

NOTE:

The autopilot does not have to be engaged to provide Mach trim. Automatic Mach trim is available whenever the PITCH TRIM ENG/DISENG switch is engaged.

3. Controls and Indications:

(See Figure 4 through Figure 6.)

NOTE:

A full description of the Flight Controls 2/3 synoptic page appears in section 2B-07-00.

A. Circuit Breakers (CBs):

The following CBs protect elevator pitch control:

Circuit Breaker Name: CB Panel: Location: Power Source:

ELEV SERVO #1 POP D-4 L ESS DC Bus ELEV SERVO #2 CPOP D-4 R ESS DC Bus PITCH TRIM SERVO #1 POP E-2 L ESS DC Bus PITCH TRIM SERVO #2 CPOP E-1 R ESS DC Bus L ELEV TRIM HEAT REER E-16 R MAIN AC Bus R ELEV TRIM HEAT REER F-16 R MAIN AC Bus LEFT ELEV HYD S/O POP C-5 L ESS DC Bus RIGHT ELEV HYD S/O CPOP C-5 R ESS DC Bus

B. Crew Alerting System (CAS) Messages:

The following CAS messages are associated with the elevator pitch controls:

Area Monitored: CAS Message: Message Color:

Elevator Trim Tab RVDTs vrs takeoff range

Elevator and Trim Tab RVDTs and MAUs

L/R Elevator HOPS and MAUs L/R Elevator Hydraulics

Aircraft Conﬁguration Red Elevator Mistrim Nose

Down

Amber

Elevator Mistrim Nose Up Amber

Off

Amber

Electric Pitch Trim Servos Elevator Trim 1-2 Fail Amber

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Area Monitored: CAS Message: Message Color:

Electric Pitch Trim Servos / Trim ENG/DISENG Switch

Electric Pitch Trim Servos and MAUs

Mach Trim Off (inhibited below 0.82

Mach)

NOTE

Amber

Pitch Trim 1-2 Power Fail Blue Elevator Trim Down Limit Blue Elevator Trim Up Limit Blue

4. Limitations: A. Mach Trim / Electric Elevator Trim Functions:

(1) Use of Mach Trim / Electric Elevator Trim Functions:

Mach trim / electric elevator trim must be ON during all ﬂight operations except as provided for in Section 05-02-40: Mach Trim Failure.

(2) With both Mach Trim / Electric Elevator Trim Inoperative:

is reduced to 0.80MT.

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Hard Over Prevention

System (HOPS)

Figure 3 (Sheet 1 of 2)

2A-27-00

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Hard Over Prevention

System (HOPS)

Figure 3 (Sheet 2 of 2)

2A-27-00

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Pitch Flight Controls

System Controls and

Indications (Cockpit

Center Pedestal)

Figure 4

2A-27-00

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Electric Pitch Trim Engage / Disengage Switch Controls and Indications

Figure 5

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Electric Pitch Trim / Stall Barrier System Controls and Indications

Figure 6

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2A-27-30: Rudder Yaw Control

1. General Description:

The ﬂight crew controls the direction of the aircraft around the vertical axis by movement of the rudder. The rudder panel is composed of graphite-epoxy fabric baked at high temperatures to form a strong but light weight structure. The rudder is positioned by inputs from the pilot or copilot rudder pedals, or by autopilot electro-servos. The pilot and copilot rudder pedals are connected by a common torque tube so that either may control rudder movement. (The individual pedal pairs are adjustable to accommodate differences in pilot leg length.) The common torque tube is connected by a bellcrank to a single stranded wire cable loop located on the right side of the aircraft beneath the cockpit and cabin ﬂoor.(Since there is only one rudder ﬂight control surface, there is no need for dual cable linkages from both yokes to the rudder, nor is there a need for a system to separate pilot and copilot rudder pedal inputs to the rudder.) The cable loop incorporates pulleys and bellcranks to route the cable around other installations beneath the aircraft ﬂoor. At the end of the loop is a bellcrank connected to a dual piston hydraulic actuator. The bellcrank translates movement of the control cable into lateral displacement inputs to the actuator. The rudder control cable linkage is illustrated in Figure 7 and the termination of the cable linkage at the rudder actuator is shown in Figure 8. The actuator has a single shaft with a piston chamber for each (left and right) hydraulic system. Both hydraulic systems provide up to three thousand (3,000) psi pressure to assist in moving the rudder surface. Internal regulator valves limit the pressure output of the two pistons within the hydraulic actuator to a maximum of ﬁfteen hundred (1,500) psi each to prevent structural damage to the rudder and vertical stabilizer. The output end of the hydraulic actuator shaft is connected to linkages that move the rudder around the pivot point connections on the vertical stabilizer. If one hydraulic system fails, the regulator valve of the remaining system shifts to provide up to three thousand (3,000) psi to move the rudder. Additionally, in the event of dual hydraulic system failures, the auxiliary (AUX) hydraulic pump will power the rudder using the left system piston in the actuator provided sufficient ﬂuid remains in the auxiliary reservoir of the left hydraulic system. AUX hydraulic system power for the rudder is activated by the selection of the STBY RUD (Standby Rudder) switch on the lower portion of the pilot instrument panel to the left of the standby ﬂight instruments.

Mechanical stops are incorporated into the rudder mounting structure to physically limit rudder displacement to a maximum of twenty-two degrees (22°) either side of neutral, although full displacement is available only at low airspeeds. As airspeed increases, the airload on the rudder surface increases proportionally. When the airload on the rudder surface equals the available hydraulic pressure output of the rudder actuator, no further rudder displacement is possible. The Monitor and Warning System (MWS) software monitors aircraft speed, angle of attack and rudder displacement to formulate an advisory message informing the ﬂight crew when maximum rudder displacement has been reached.

2. Description of Subsystems, Units and Components: A. Yaw Damper and Autopilot Rudder Operation:

Automatic rudder compensation for aircraft yaw produced by dutch roll inherent to swept-wing aircraft is provided by the yaw damper function of the autopilot. The yaw damper function is engaged with the YAW DAMP ENG / DISENG switch located on the lower pilot side instrument panel

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adjacent to the STBY RUD switch, shown in Figure 9. The yaw damper is normally engaged even if the autopilot is not operating. (The DISENG legend in the switch will be illuminated amber if the yaw damper is not engaged.) If the autopilot is engaged, the yaw damper must be engaged, since autopilot rudder commands use the yaw damper circuits to displace the rudder.

The yaw damper function is a redundant dual-channel installation. Yaw damper channel #1 is controlled by the autopilot function hosted in processor modules in Modular Avionics Unit (MAU) #1, channel #2 by MAU #2. Each channel has a dedicated Electro-Hydraulic Servo Valve (EHSV) internal to the rudder actuator. The autopilot processor detects an uncommanded yaw displacement by monitoring data from the Inertial Reference Units (IRUs). The autopilot processor signals a rudder displacement to counter the aircraft yaw through an Actuator Input/Output Processor (AIOP) module via ARINC-429 bus connection to the EHSV on the rudder actuator. The amount of rudder displacement necessary is a function of airspeed / Mach number, and the AIOP uses information from the Air Data Application (ADA) in the MAU to determine the amount of rudder to apply. A feedback loop from a Rotary Variable Displacement Transducer to the AIOP conﬁrms the rudder position. The maximum amount of rudder displacement available to the yaw damper is ﬁve degrees (5°).

Since only one yaw damper channel is necessary for rudder control, the active channel alternates on each ﬂight segment (a function of weight-onwheels) to prolong system life. If the active channel fails, the standby channel will automatically assume yaw damper control.

The yaw damper function also provides a rudder input for aircraft turn coordination provided the ﬂaps are not set to thirty degrees (30°) or more. The yaw damper will add up to ﬁve degrees (5°) of rudder in the direction of turn without pilot rudder input.

Autopilot control of the rudder is the same functional process as the yaw damper, but the amount of rudder displacement available to the autopilot is greater (up to the 22° limit). Larger rudder inputs are necessary for the lower airspeeds associated with coupled approaches or during single engine operations.

B. Rudder Trim:

The rudder is trimmed by manual inputs from the trim wheel mounted on the cockpit aft center pedestal (see Figure 10). There is no trim tab on the rudder, rather the whole rudder panel moves in response to trim input. The rudder may be displaced up seven and a half degrees (7½°) left or right with trim commands. Rotation of the trim wheel moves a cable linkage to a drum mounted adjacent to the rudder hydraulic actuator. Movement of the trim wheel rotates the drum and a linkage attached to the drum moves the shaft of the rudder actuator. The rudder actuator hydraulically positions the rudder by the commanded amount of deﬂection, and the linkage from the drum establishes the trimmed rudder position as a new neutral setting for the actuator. Manual or yaw damper / autopilot rudder deﬂections are then summed to the existing trimmed rudder displacement. (The autopilot does not have a separate trim input to the rudder.)

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C. Rudder Limiting:

The G550 does not have a separate load limiter unit, but relies instead upon MWS software to compute maximum rudder deﬂection for a given airspeed, matching airloads on the rudder surface with the hydraulic function of the rudder actuator.Although at low airspeeds full rudder travel of twenty-two degrees (22°) is available, at higher airspeeds less rudder travel is necessary to achieve the desired amount of aircraft heading control. To avoid excessive loads, the rudder hydraulic actuator uses internal pressure switches to signal the MWS when full hydraulic pressure output of the actuator has been reached. The MWS formulates a CAS blue advisory message text of “Rudder Limit” for display on the CAS window indicating the maximum rudder hydraulic power assist condition.

The amount of rudder deﬂection at which maximum rudder hydraulic assist occurs is dependent upon airspeed. As speed increases, air loads increase as the rudder is displaced. Since the rudder surface is linked mechanically to the rudder hydraulic actuator shaft, the force of the rudder airload opposes the force of the hydraulic system(s) moving the actuator. Whenever the airload force equals hydraulic system force (1,500 psi with both left and right systems operating or 3,000 psi with a single hydraulic system) no further rudder displacement is possible, and the “Rudder Limit” advisory message is displayed on the CAS window.

D. Rudder Hard Over Prevention System (HOPS):

Movements of the rudder contrary to the commanded position are limited by a Hard Over Prevention System (HOPS) that incorporates eight switches to monitor mechanical and hydraulic rudder operation (see Figure

3. Four external mechanical switches are integrated into the rudder control linkage to provide a comparison reference for four switches mounted internally within the hydraulic actuator. The external HOPS switches are noted on Figure 8. Of the four external switches, two are for left hydraulic system reference and two are for right hydraulic system reference. Of the two switches for each hydraulic system, one provides a left rudder command reference and the other provides a right rudder command reference. The switches are plunger-type contact switches, and are installed on each side of a bracket attached to the command input side of the rudder actuator. On one side of the bracket are the left rudder command input switches for the left and right hydraulic systems. On the other side of the bracket are the right rudder command input switches for the left and right rudder systems. Inserted between the switches in the bracket is a cam-type arm mated to the rudder hydraulic actuator output linkage. The cam arm is positioned with a deﬁned amount of clearance between the plunger-type switches. Under normal conditions, the bracket holding the switches moves with rudder command input and the cam arm moves with the rudder hydraulic actuator output, so the clearance between the switches and the arm is maintained.

If a malfunction occurs and the rudder moves opposite to or further from the commanded direction, the cam arm that is attached to the output linkage of the rudder actuator will move to close the clearance gap between the cam arm and the plunger-type switches, making contact with the switches on the side of the bracket. When the plungers of the switches are depressed, a relay closes and an electrical signal is sent to a corresponding set of switches mounted internally within the hydraulic

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actuator. Four pressure switches monitor left and right hydraulic system pressures

within the pistons of the rudder actuator. In normal conditions, all four switches sense stabilized pressures since hydraulic outputs positioning the rudder are balanced by air load pressures on the rudder surface acting against actuator pressures. When a malfunction occurs and the rudder moves contrary to the commanded direction, the hydraulic actuator shaft moves in the contrary direction, causing an increase in hydraulic pressure on the opposite sides of the pistons within the actuator. The left and right system opposite side pressure switches close, completing the circuit initiated by closure of the bracket plunger switches, and an electrical signal is sent to a timing relay. If the contrary rudder movement persists for longer than one half (½) second, hydraulic pressure from both left an right systems is shut off from the rudder actuator.

Left Hydraulic System Right Hydraulic System

Left

Rudder ←

Right

Rudder →

Left

Rudder ←

Hydraulic Actuator Shaft

Pressure

Switch

Pressure

Switch

Pressure

Switch

In a similar manner, if a hydraulic malfunction causes the shaft of the actuator to move in a direction opposite the pilot or yaw damper / autopilot commanded direction, the increased pressures in the wrong direction sides of the pistons would close the monitoring switches, and movement of the actuator shaft would cause the cam arm to contact the plunger switches on the external bracket, and hydraulic pressure would be shut off to the rudder actuator after a one half (½) second delay.

If a hydraulic malfunction in a single system (left or right) side of the actuator moves the actuator shaft in a wrong direction, only the hydraulic pressure of the malfunctioning system is shut off. For instance in a stabilized rudder condition, if the left hydraulic system piston attempts to move the actuator shaft to the right, the increased pressure in the right side of the left piston will close the monitor switch and actuator shaft displacement will close both the left and right hydraulic system bracket plunger switches, completing the shut off circuit for the left hydraulic system after a one half (½) second delay. (In this case hydraulic pressure in the right rudder side of the right hydraulic system piston will decrease due to the increase in area caused by actuator shaft movement.) The rudder hydraulic actuator will continue to function using the remaining hydraulic system. During single hydraulic system operation the pressure regulator valve will open, allowing the remaining hydraulic system pressure output to increase up to three thousand (3,000) psi for rudder actuation.

The operation of the hydraulic shut off valves by the HOPS is signaled to MAU #2 over an ARINC-429 connection. The shut off condition is monitored by the MWS, and a CAS message corresponding to the condition is displayed on the CAS window. If both hydraulic systems are shut off, an amber caution message of “Rudder Hydraulics Off” is displayed. Manual rudder control may remain possible, depending upon

Right

Rudder →

Pressure

Switch

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the cause of the hardover condition. If the cause of the condition is thought to be momentary, and the use of the rudder is deemed necessary for continued safe ﬂight and landing, the hydraulic shut off valves may be reset by cycling the RUDDER HYD S/O circuit breaker. If the cause has not been rectiﬁed, the shut off valves will close and hydraulic boost for the rudder will be unavailable. Loss of rudder hydraulic pressure will also prevent yaw damper (and autopilot rudder) operation.

If only one hydraulic system is shut off to the rudder, the remaining system will provide full boost to the rudder and yaw damper / autopilot rudder operation. The amber caution CAS message of ”Rudder Hydraulics Off will be accompanied by a blue advisory “Single Rudder” message.

To prevent the HOPS from shutting off hydraulic system pressure to the rudder during normal ﬂight maneuvers that may involve rapid changes in rudder direction, HOPS activation is buffered by three elements:

• The clearance between bracket plunger switches and the actuator cam arm

• Inertia in the build up of hydraulic pressures in the rudder actuator

• The one half (½) second delay in electrically latching the shut off of

hydraulic pressure

E. Standby Rudder System:

In the event that both hydraulic systems fail during ﬂight, the Auxiliary (AUX) hydraulic pump can be used to pressurize the left hydraulic system to actuate the rudder and yaw damper provided that ﬂuid remains in the left system reservoir and there is no leak in the left hydraulic system lines connecting to the rudder actuator. The AUX hydraulic pump, located in the right main landing gear wheel well, is powered by left essential DC bus and can provide three thousand (3,000) psi at a ﬂow rate of two (2) gallons per minute. Normally the AUX is used to actuate the ﬂaps, ground spoiler servos, brakes and nose wheel steering in the event of a dual hydraulic system failure. However a valve in the left hydraulic system plumbing, controlled by the STBY RUD switch on the lower portion of the pilot instrument panel, can be used to divert the total output of the AUX hydraulic pump to the left hydraulic system piston of the rudder actuator. Powering the rudder and yaw damper, especially at higher altitudes, provides dutch roll compensation and avoids the airspeed restrictions necessary without an operating yaw damper. As the aircraft descends to lower altitudes and airspeeds during the approach phase, airloads on the rudder are reduced and manual rudder control is adequate for steering commands. The STBY RUD switch can then be selected off to enable the AUX hydraulic pump to provide the pressure for ﬂap extension, pressurize ground spoiler servos, wheel brakes and nose wheel steering during landing. If the STBY RUD switch is not selected off prior to landing, the standby rudder valve will automatically close when the nose gear weighton-wheels (WOW) switch is compressed on landing, enabling AUX pump power for stopping and steering the aircraft.

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3. Controls and Indications:

(See Figure 9 and Figure 10.)

NOTE:

A full description of the Flight Controls 2/3 and Hydraulics 2/3 synoptic pages appears in section 2B-07-00.

A. Circuit Breakers (CBs):

The following CBs protect the rudder ﬂight controls:

Circuit Breaker Name: CB Panel: Location: Power Source:

RUDDER HYD S/O CPOP C-3 R ESS DC Bus YAW DAMP SERVO #1 POP D-6 L ESS DC Bus YAW DAMP SERVO #2 CPOP D-6 R ESS DC Bus AUX HYD PUMP LEER C-16 L ESS DC Bus

B. Crew Alerting System (CAS) Messages:

The following CAS messages are associated with the yaw ﬂight controls system:

Area Monitored: CAS Message: Message Color:

Rudder Hydraulic Shutoff Valves Rudder Hydraulics Off Amber Yaw Damper Servo Valves Yaw Damper 1-2 Fail Amber YAW DAMP ENG/DISENG Switch Yaw Damper Off Amber Autopilot, YAW DAMP ENG/

DISENG Switch MWS Software, Aircraft Speed,

Rudder Position (RVDT)

No YD Turn Coordination Blue Rudder Limit Blue

Rudder Hydraulic Shutoff Valves Single Rudder Blue Standby Rudder Switch / AUX

Pump

Standby Rudder Hyd On Blue

Yaw Damper Servo Valves YD 1-2 Power Fail Blue

4. Limitations:

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A. Yaw Damper Inoperative Speeds:

Maximum Speeds:

• Above 10,000 Feet: 260 KTS / 0.80 MT

• Below 10,000 Feet: 250 KCAS

Minimum Speeds:

• Above 20,000 Feet: 210 KTS

• Below 20,000 Feet:

The minimum speed is in accordance with the following schedule until ready to conﬁgure for approach and landing. V the airspeed tape of the PFD, is the approach speed for landing for the current ﬂap setting.

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REF, as shown on

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Flaps 0, 10,

Fuel - lb

23,000 V 24,000 V 25,000 V 26,000 V 27,000 V 28,000 V 29,000 V 30,000 V 31,000 V 32,000 V 33,000 V 34,000 V 35,000 V 36,000 V 37,000 V 38,000 V 39,000 V 40,000 V 41,000 V 41,300 V

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Sea Level to 5000 ft 5000 to 20,000 ft

REF VREF VREF 135 REF VREF VREF 141 REF VREF VREF 147 REF VREF VREF 153 REF VREF VREF 159 REF VREF 147 160 REF VREF 153 160 REF VREF 158 160 REF VREF 163 160 REF VREF 168 160 REF VREF 174 160 REF VREF 179 160 REF VREF 184 160 REF VREF 189 160 REF VREF 195 160 REF VREF 200 160 REF VREF 205 160 REF VREF 211 160 REF VREF 216 160 REF VREF 217 160

Flaps 39

Flaps 0, 10,

Flaps 39

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THIS PAGE IS INTENTIONALLY LEFT BLANK.

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Rudder Control Cable

Linkage

Figure 7

2A-27-00

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Control Cable Termination at Rudder Actuator

Figure 8

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Yaw Flight Controls System Controls and Indications

Figure 9

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Rudder Trim Controls

Figure 10

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2A-27-40: Aileron Roll Control

1. General Description:

The ﬂight crew controls the motion of the aircraft around the longitudinal axis by displacement of the ailerons. The ailerons, composed of baked epoxy graphite fabric, are mechanically linked to the pilot and copilot control yokes by a dual installation of multi-strand braided cables and a system of pulleys and bellcranks. A hydraulic actuator is incorporated in the cable linkage of each aileron to provide a powered assist in deﬂecting the ailerons. The hydraulic actuators provide a boost advantage of six to one (6 : 1) to pilot manual yoke inputs. Rotation of the control yokes left or right produce a roll in the corresponding direction. Each control yoke is linked to one aileron: the pilot yoke is cable-linked to the left aileron and the copilot yoke is cable-linked to the right aileron. Although the ailerons are independently linked, the control yokes are coupled together by a mechanical arm beneath the cockpit ﬂoor so that rotation of one yoke will rotate the other yoke, producing motion in both ailerons. Like the elevator system previously described, the link between the control yokes may be separated if a malfunction causes one aileron to become inoperative. A disconnect handle on the copilot side of the center pedestal can be pulled out to separate the link between the yokes, allowing independent operation of the free (unjammed) aileron.

Operation of the ailerons is symmetrical but opposite. If a control yoke is rotated to the left to command a left roll, the aileron on the trailing edge of the right wing deﬂects downward, producing a small change in the chord of a section of the wing, and creating lift. The right wing then rises due to the increased lift component. At the same time the aileron on the left wing trailing edge deﬂects upward, disrupting airﬂow over the wing and decreasing lift. The left wing falls due to the decrease in lift on a portion of the wing. A roll in the opposite direction is accomplished by reversing the direction of the aileron deﬂection, creating the same circumstances on opposite wings. The maximum aileron deﬂection is eleven degrees (11°) up or down.

To increase the response time for roll commands, two of the three spoiler panels on each wing are mechanically coupled to the aileron control cable linkage. The two outboard spoilers on the falling wing (with the aileron deﬂected up) deploy, creating additional disruption of airﬂow over the wing to increase roll rates. The spoiler panels on the upward moving wing (with aileron deﬂected downward) do not move, remaining faired with the upper wing surface.

A single trim tab on the left aileron is manually positioned to reduce control yoke forces in maintaining a stabilized condition around the longitudinal axis. An aileron trim wheel on the aft section of the cockpit center pedestal is linked by a cable directly to the trim tab. The trim wheel is rotated left and right to produce the corresponding trim input to position the left aileron. Since the ailerons are linked through the mechanical connection between the control yokes, moving the left aileron will move the right aileron in the opposite direction for a equal amount of deﬂection. The trim tab has a ceramic surface surrounding the hinge points that is heated to prevent ice formation that might interfere with tab operation.

The autopilot is integrated into the cable linkage to the ailerons through electric servos. The aileron cables are wrapped around a drum driven by the servos. The autopilot servos rotate the drum, producing cable inputs to position the ailerons.

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NOTE:

The autopilot has no trim input to the ailerons

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The ailerons, like the other primary ﬂight controls, are monitored for correct operation by a Hard Over Prevention System (HOPS). The HOPS will shut off hydraulic pressure to both aileron actuators if aileron movement does not correspond to ﬂight crew of autopilot commands.

If no hydraulic pressure is available to the aileron actuators, the hydraulic control valves of the actuators will bypass any residual pressure and full aileron deﬂection will be available using manual control inputs from the cockpit yokes. A slower response time to aileron manual operation can be expected due to the lack of boost authority and the absence of assist from the ﬂight spoilers (if the spoilers cannot be hydraulically powered).

2. Description of Subsystems, Units and Components: A. Aileron Control Linkage:

Rotation of each control yoke produces a parallel rotation of the other yoke through the mechanical arm beneath the cockpit ﬂoor. As both yokes rotate bellcranks translate rotational movement into linear displacement of the control cables leading to the hydraulic actuators that move the ailerons. Integrated into the cable linkage are the autopilot electric servos and a series of bellcranks and linkages to route the cables around existing equipment beneath the cabin ﬂoor. The aileron cable links transit aft along the fuselage until reaching the aft bulkheads of the main landing gear wheel wells. At this point the cables pass though pulleys that translate the forward and aft motion of the cables ninety degrees (90°) and out the aft face of each main wing beam. The cable linkages connect to the ﬂight spoiler mixing and summing links and continue out to the aileron hydraulic actuators. The hydraulic actuators are located on the aft wing beam between the ﬂaps and the ailerons. The actuators boost manual or autopilot control inputs and position the ailerons in the commanded direction.

Each hydraulic actuator is powered by both hydraulic systems. The actuators have a central shaft surrounded by a dual piston chamber. Each chamber is powered by a dedicated left or right hydraulic system. The pistons are moved by hydraulic pressure ported to the extend or retract sides of the chamber by control valves moved by the aileron cable linkage. The pistons extend or retract the central shaft of the actuator that in turn moves a bellcrank and linkage to position the aileron. A mechanical linkage, termed a force link, between the control cable input to the actuator and the output of the actuator to the aileron parallels actuator operation and provides the sensing element for operation of the HOPS provision. Loss of a single hydraulic system will not degrade aileron activation since the remaining system is capable of providing sufficient power for a full range of aileron displacement. See the illustration of the linkages and actuators in Figure 11.

B. Aileron Control Disconnect:

If a malfunction causes one of the ailerons to become inoperative, the mechanical link between the two cockpit control yokes can be separated to allow independent operation of the normally functioning aileron. A handle on the right side of the cockpit center pedestal, labelled AIL DISC is connected to a pin in the mechanical linkage. (See Figure 12). Pulling up on the handle retracts the pin mating the two yokes and allows full control of the operative aileron. If the ﬂight crew is unable to exert sufficient force

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to retract the pin, a trigger mechanism beneath the handle can be activated to supply additional gas cartridge pressure to assist in retracting the pin. If the cartridge assist is used, the aileron channel linkage cannot be reconnected in ﬂight since it requires special maintenance tools. However, if only manual force is used to disconnect the aileron channels, the pin may be reinserted in the mechanical arm link by pushing in on the handle if it is determined that the malfunctioning aileron has returned to normal operation.

C. Aileron Hard Over Prevention System (HOPS):

The conformity of the motion of aileron hydraulic actuators to aileron control inputs from cockpit yokes or the autopilot is monitored by force link mechanisms. The force links are the sensing elements that initiate aileron HOPS activation. (See the ﬂight controls HOPS system diagram in Figure

3). The mechanisms are essentially telescoping tubes with two sections one section that travels within the other. The end of one section of the tube is connected to the cable linkage control inputs to the hydraulic actuator and the end of the other tube section of the is connected to the actuator shaft output to the aileron. The dual tube mechanism thus parallels the action of the aileron actuator. Each tube section is spring loaded to resist the push and pull motion transmitted through the actuator to move the aileron. If a control cable input to the actuator is in the push direction and the actuator shaft does not transmit a push assist or the aileron does not move, the spring at the push end of the force link tube is compressed, and an electrical contact within the spring initiates a signal to the hydraulic control valves of both left and right aileron hydraulic actuators to bypass system pressure. (If a resistance to control link input is present in the pull direction, the spring and electrical contact at the end of the other section would operate in the same way to activate HOPS bypass of hydraulic pressure.)

The electrical signal incorporates a one half (½) second delay before activating the hydraulic pressure bypass. Once activated, the control valves are latched to the bypass position. If, in the judgement of the ﬂight crew, the condition causing the malfunction is transient, and the use of the ailerons is deemed essential to continued safe ﬂight and landing, the HOPS initiated bypass can be reset by cycling both the LEFTAIL HYD S/O and RIGHT AIL HYD S/O circuit breakers. If the malfunction is still present the control valves will re-latch to the bypass state after the one half (½) second delay.The left aileron shut off valve is powered by the left essential DC bus and the right aileron shut off valve by the right essential DC bus.

When the electrical signal bypasses hydraulic pressure to the actuators, signals are also sent to Input / Output (I/O) modules in the Modular Avionics Units (MAUs). The left aileron state is communicated to MAU #1 and the right aileron state to MAU #2. The MAUs communicate with the Monitor and Warning System (MWS) that formats the appropriate messages for presentation on the Crew Alerting System (CAS) window.

To prevent inadvertent HOPS bypass of the aileron actuators during ﬂight conditions requiring quick reversals of the ailerons, the actuation of the HOPS is ﬁltered by:

• The compression of the springs within the force link sections

• Inertia in hydraulic piston action

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• The one half (½) second electrical delay

D. Aileron Position Sensing:

The position of each aileron is sensed by a Rotary Variable Displacement Transducer (RVDT). The RVDTs are installed at the hinge point of the ailerons, and transmit an electrical signal proportional to the displacement of the aileron from the neutral position. The signals are communicated to the autopilot processor modules in the MAUs where, in conjunction with the Actuator Input / Output Processors (AIOPs), autopilot commands are formulated for the electric servos in the cable control linkage for the ailerons. The MAUs also provide aileron position information to the MWS for generation of the graphic display of control surface position on the Flight Controls 2/3 synoptic page.

E. Aileron Trim System:

(See Figure 12.) Only one trim tab is installed for both ailerons, since both the left and right

ailerons work in symmetrically opposite directions through the cable linkage and mechanical arm connecting both control yokes. The trim tab is located on the left aileron and is made of carbon epoxy fabric. The aileron trim tab is a manually operated by cable linkage from the aileron trim wheel on the aft section of the cockpit center pedestal to the trim tab bellcrank actuator. Moving the trim wheel to the left positions the trim tab down into the airstream, deﬂecting the left aileron up to reduce lift on the left wing, moving the wing downward. At the same time the connecting linkage moves the right aileron down to increase lift on the right wing and move the wing up. The range of travel for the aileron trim tab is 15 degrees up or down.

The aileron trim tab actuator has ceramic element that is heated to prevent the actuator from freezing. The actuator is maintained at a constant temperature of 175°±20°F. The heater is controlled by the AIL TRIM HEAT circuit breaker on the Right Electronic Equipment Rack (REER) and receives power from the right main AC bus.

The trim tab is also equipped with a RVDT to provide information to MAU #1 for the autopilot processor and the MWS for graphic display on synoptic and system windows.

F. Flight Spoiler Roll Augmentation:

The aircraft is equipped with three spoiler panels on the top surface of each wing, just forward of the ﬂaps. The spoiler panels are rectangular in shape and are hinged at the forward edge to allow the panels to pivot upward, into the airﬂow over the wing to disrupt lift. The spoiler panels are hydraulically powered by actuators supplied with three thousand (3,000) psi pressure from both hydraulic systems. Loss of one hydraulic system will not effect spoiler operation since a single system supplies sufficient pressure for full spoiler activation. The spoilers serve three functions:

• As speed brakes, all six spoilers (three on each wing) are extended to disrupt airﬂow over the wing and decrease lift in order to slow the aircraft and/or increase descent rates

• On landing, with weight on the wheels, all six panels (three on each wing) extend to provide aerodynamic drag to aid in slowing the aircraft

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• As ﬂight spoilers, only the two outboard spoilers on each wing (a total of four spoiler panels) can be used in conjunction with the ailerons to increase the rate of roll of the aircraft

Flight spoilers deploy upward on the side of the upward displaced aileron to contribute to the decrease in lift on the downward rolling wing. By interrupting additional airﬂow over the wing on the inside of a turn, roll rate into the turn is increased. Spoilers are activated by the same cable linkage that controls aileron deployment. A bellcrank in the aileron cables on the aft wing beam provides an input to a mixing and summing series of pushrods and hinges to provide a mechanical command to the hydraulic actuator for the two outboard spoiler panels.

The amount of ﬂight spoiler extension is proportional to aileron displacement. Maximum aileron displacement is eleven degrees (11°) from neutral and maximum ﬂight spoiler extension is forty-seven degrees (47°) up from the faired position. The following table shows examples of the aileron / ﬂight spoiler displacement ratio:

Aileron Displacement Flight Spoiler Extension

Up one half degree (+½°) Up one half degree (+½°) Up one degree (+1°) Up ﬁve and one half degrees (+5.5°) Up seven degrees (+7°) Up twenty-eight degrees (+28°) Up eleven degrees (+11°) Up forty-seven degrees (+47°)

If the spoilers are in use as speed brakes, all three spoiler panels on each wing can be extended up to thirty degrees (30°) with commands from the speed brake control handle on the cockpit center pedestal. If a turn is initiated with speed brakes extended, the two outboard spoiler panels on the downward rolling wing will extend an additional amount up to ﬁfty-ﬁve degrees (55°) to increase the roll rate into the turn.

Malfunctions in the spoiler system are controlled by selecting the spoilers off. There is no manual extension provision for spoilers. A SPOILER CONTROL pushbutton is installed in the cockpit center pedestal just forward of the speedbrake handle, as shown in Figure 12. Pushing the button in will bypass all hydraulic pressure to the spoilers and illuminate the amber OFF legend on the face of the button.

3. Controls and Indications:

(See Figure 12.)

NOTE:

A full description of the Flight Controls 2/3 synoptic page appears in section 2B-07-00.

A. Circuit Breakers (CBs):

The following circuit breakers protect the aileron and ﬂight spoiler roll controls:

Circuit Breaker Name: CB Panel: Location: Power Source:

LEFT AIL HYD S/O POP C-4 L ESS DC Bus RIGHT AIL HYD S/O CPOP C-4 R ESS DC Bus

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Circuit Breaker Name: CB Panel: Location: Power Source:

SPLR FLT PWR S/O CPOP E-3 R ESS DC Bus AIL TRIM HEAT REER D-16 R MAIN AC Bus

B. Crew Alerting System (CAS) Messages:

The following CAS messages are associated with the roll ﬂight controls system:

Area Monitored: CAS Message: Message Color:

L / R Aileron Hydraulic Actuators L - R Aileron Hydraulics

Off

Amber

Spoiler Control Switch Spoilers Hydraulic Off Amber Aileron Trim Tab RVDT and

Autopilot Aileron Servos Aileron Trim Tab RVDT and

Autopilot Aileron Servos

Retrim Left Wing Down Amber Retrim Right Wing Down Amber

4. Limitations:

There are no limitations established for the aileron roll control system at the time of this writing.

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Aileron Linkage and

Actuators Figure 11

2A-27-00

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Aileron and Spoiler

Controls and Indications

Figure 12

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2A-27-50: Horizontal Stabilizer and Wing Flaps

1. General Description:

The aircraft horizontal stabilizer is mounted at the top of the vertical stabilizer.The horizontal stabilizer is movable in order to provide a ﬂexible aircraft neutral stability position in the pitch axis as ﬂaps are extended or retracted. As the ﬂaps extend they effectively produce a large change in the wing airfoil. The airfoil change moves the center of lift aft, causing a nose down moment in the pitch axis. The horizontal stabilizer will automatically compensate for pitch changes as the ﬂaps extend by moving leading edge down, moderating elevator trim change and conserving elevator deﬂection for landing. As ﬂaps are retracted, the stabilizer moves leading edge up to reduce elevator forces as the wing center of lift transits forward.

The Fowler-type ﬂaps installed on the trailing edge of each wing provide the additional lift required to ﬂy at slower airspeeds. By extending aft and down, the ﬂaps change airﬂow over the wing, effectively lengthening wing chord and increasing wing thickness. By emulating a larger wing surface, the ﬂaps decrease takeoff and landing speeds.

2. Description of Subsystems, Units and Components: A. Horizontal Stabilizer and Actuator

The horizontal stabilizer is an all-metal structure mounted to the top of the vertical stabilizer at two points. The aft point consists of a shaft installed through an opening allowing the stabilizer to pivot. The front attachment is an electrically powered jackscrew that moves the leading edge of the stabilizer up and down. The range of movement available for the horizontal stabilizer is 1.25°up to 4.6°down. The jackscrew is driven by an electric actuator with two motor windings, and incorporates a locking ratchet mechanism that prevents movement of the stabilizer if there is no electrical input to the AC drive motor. The two AC motor windings are independently powered by the left and right standby AC buses that can be energized by the Hydraulic Motor Generator (HMG) in the event of electrical failure. Each AC motor winding has an independent essential DC bus powered control channel for increased redundancy. (The horizontal stabilizer is fully operational using a single motor and control channel.) A Rotary Variable Displacement Transducer (RVDT) is mounted on the actuator to provide feedback to control signals and to furnish position information for display on the Flight Controls 2/3 synoptic page and on the Horizontal Situation Indicator (HSI) portion of the Primary Flight Director (PFD). Stabilizer position data is communicated to MAU #1 and #2 through ARINC-429 bus connections from the Flap Control Unit (FCU) where it is shared with the Monitor and Warning System (MWS).

NOTE:

The CrewAlerting System (CAS) messages generated by the MWS use the following convention to denote the redundant power and control sources for the operation of the horizontal stabilizer:

• Stabilizer channel A = left standby AC power and left essential DC control

• Stabilizer channel B = right standby AC power and

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right essential DC control

Control inputs to drive the actuator come from two (2) sources:

(1) The Flap / Horizontal Stabilizer Control Unit (FCU) to compensate

for ﬂap extension.

(2) Pilot or Copilot yoke trim switches in the Emergency Stabilizer mode

to provide aircraft pitch control if the elevators malfunction or jam.

All control inputs to move the horizontal stabilizer are routed through the Flap Control Unit (FCU) that coordinates stabilizer position with aircraft conﬁguration. The FCU prioritizes control inputs to the stabilizer AC motors so that:

• The stabilizer position schedule is matched to ﬂap position

• Arming the Emergency Stabilizer cancels all other stabilizer

operating modes

Other control and coordination functions of the FCU are discussed in the following wing ﬂap section - the position of the stabilizer is scheduled by wing ﬂap extension settings to allow the greatest range of elevator effectiveness for takeoffs and landings.

If the elevators become jammed or otherwise malfunction, selecting the EMERG STAB pushbutton switch (shown in Figure 14) to ARM (the pushbutton legend will illuminate amber) enables the trim switches on the pilot and copilot yokes to control the position of the stabilizer. As soon as the yoke trim switch is moved in either direction, the FCU isolates the stabilizer from all other command inputs. The yoke trim switches can then be used for a reduced amount of pitch control during ﬂight and landing.

The emergency stabilizer mode is communicated from the FCU to the Flight Management System (FMS) Actuator Input / Output Processor (AIOP) modules in MAU #1 and #2 and the MWS processors in MAU #1, #2 and #3. The MWS initiates the display of a CAS message indicating emergency stabilizer operation and provides the stabilizer position display on the Flight Controls 2/3 synoptic page and HSI on the PFD.

B. Flap Control Unit (FCU) and Horizontal Stabilizer Operation:

The Flap and Stabilizer Controller, most often termed the Flap Control Unit (FCU) is located in the Baggage Compartment Electronic Equipment Rack. The FCU is a processor that receives external position commands for the ﬂaps and horizontal stabilizer and also internally computes position commands for the horizontal stabilizer to coordinate stabilizer position with ﬂap position. External commands for the ﬂaps are initiated by the ﬂap lever in the cockpit, and external stabilizer commands are sourced from the electric trim switches when the emergency stabilizer mode is selected. The FCU responds to position commands for the ﬂaps and stabilizer by providing dual channel DC control signals to position the ﬂaps and dual channel DC control and AC actuator power to move the stabilizer to commanded positions. A block diagram of the Flap / Stabilizer Controller (or FCU) is shown in Figure 13.

Flap position commands are obtained from two RVDTs on the cockpit ﬂap control lever.The FCU compares the commanded position with the existing ﬂap position derived from RVDTs on each wing ﬂap. If there is a difference in the two positions, the FCU signals operation of the Power Drive Unit (PDU) that hydraulically actuates the ﬂaps to the commanded position.

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(The operation of the aircraft ﬂaps is described in the following sections.) The FCU moves the horizontal stabilizer to compensate for the change in the center of lift on the wing as ﬂaps are extended. The FCU processor schedules stabilizer position according to the following table:

Flap Position Horizontal Stab Position

0° (ﬂaps up) −1.5° ±.1°

10° −2.7° ±.1° 20° −3.6° ±.1°

39° (ﬂaps down) −4.6° ±.1°

If a malfunction such as a ﬂap asymmetry prevents the ﬂaps from moving, the FCU cancels movement of the horizontal stabilizer. The horizontal stabilizer remains operative in the EMER STAB mode, and can be positioned with the cockpit yoke trim switches.

If the EMER STAB mode is used to control the horizontal stabilizer, and subsequently control of the horizontal stabilizer is returned to normal operation (by deselecting theARM state with the EMER STABpushbutton), the FCU will move the stabilizer to align with the existing ﬂap setting according to the processor internal schedule. During this process, an advisory (blue) CAS message of “Stabilizer Syncing A - B” will be displayed.

C. Wing Flaps:

A Fowler-type ﬂap is installed on the trailing edge of each wing. Fowler ﬂaps, when extended, move aft and down from the wing trailing edge, increasing wing area to generate additional lift for low speed ﬂight. Fowler ﬂaps have a cross-section similar to wing airfoils, and by deploying aft of the wing trailing edge, open up a gap between the wing and ﬂap allowing additional airﬂow over the ﬂap. RVDTs are installed on the outer section of the ﬂaps to provide position feedback information to the FCU.

Flap extension and retraction is controlled with the ﬂap handle on the copilot side of the center console, shown in Figure 16. Each ﬂap extends back from the wing along four tracks with rollers attached to the ﬂaps moving within the tracks. The two middle track installations on each wing also contain the jackscrews that extend or retract the ﬂaps. The jackscrews on each wing are driven by torque tubes installed on the aft section of the wing running inboard to a hydraulic Power Drive Unit (PDU) located in the main landing gear wheel well. The PDU is powered by the left hydraulic system and uses a gearbox to transmit hydraulic pump rotation to the torque tubes. If the left hydraulic system is not available due to a malfunction, the auxiliary (AUX) hydraulic pump or the Power Transfer Unit (PTU) can be used to pressurize left hydraulic system ﬂuid to operate the PDU. A schematic of the wing ﬂaps system operation is shown in Figure

15. The direction of PDU rotation is controlled by two solenoid valves, one to

extend the ﬂaps and the other to retract the ﬂaps. Each solenoid valve has dual control channels (A and B) from the FCU. The FCU compares the position of the RVDTs on the ﬂap handle in the cockpit to the position indicated by the ﬂap RVDTs, and signals the extend or retract solenoid valve on the PDU, powering the PDU in the required direction to extend or

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retract the ﬂaps. Incorporated into the ﬂap jackscrew actuators are force limiters that

prevent structural damage to the ﬂaps from high airspeeds. If the ﬂaps are selected to a setting greater than that allowed by airspeed limits, the force limiters will brake the motion of the jackscrews, preventing further extension of the ﬂaps (the jackscrews will continue to attempt to turn, but are unable to overcome the force limiters). During the overspeed condition, the ﬂaps will operate in the opposite direction if retracted to a lower setting. If the ﬂaps are not retracted, the force limiters will release once airspeed drops below the limit for the selected ﬂap setting.

NOTE:

If airspeed exceeds the limit for the current ﬂap setting, the overspeed condition will be signalled by the change in color of the airspeed tape on the PFD and the OVERSPEED aural clacker will sound over cockpit speakers.

D. FCU and Flap Operation:

The FCU uses RVDT data from the ﬂap handle and ﬂap position to drive the PDU in the correct direction to extend or retract the ﬂaps. The FCU also uses ﬂap handle RVDT data to position the horizontal stabilizer in the correct position according to the schedule corresponding to ﬂap position. The FCU processor software also monitors performance of the ﬂaps, horizontal stabilizer and ﬂap / stabilizer scheduled positions. If a fault is detected, the operation of the malfunctioning unit (ﬂaps or stabilizer) is interrupted and through ARINC-429 connections to the MAUs, the MWS generates the display of the appropriate CAS message(s). For instance, if the FCU detects an asymmetry during the extension of the two wing ﬂaps, the FCU stops operation of the PDU and signals the condition to the MWS through the MAUs. The MWS generates an amber “Flap Asymmetry” caution message on the CAS display.(The complete list of CAS messages pertaining to ﬂap and horizontal stabilizer operation are shown at the end of this section.)

Flap RVDT reported position is supplied by the FCU to open or close a relay at a ﬂap setting of twenty-two degrees (22°). The relay provides the ﬂap position input for the conﬁguration warning horn of the landing gear control system. The landing gear warning horn will sound if:

• Power levers at idle, ﬂaps extended less than 22°, landing gear not down and altitude less than three hundred ﬁfty feet (350 ft.)

• Power levers at any setting, ﬂaps extended more than 22° and landing gear not down

The FCU is used to perform maintenance functions involving recalibration of wing ﬂap position to ﬂap handle position and horizontal stabilizer position to ﬂap position. The maintenance process uses selector switches on the face of the FCU to perform the recalibration. Various CAS messages will be displayed during the maintenance procedure, and are included in the list at the end of this section for information only.

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3. Controls and Indications:

(See Figure 14.)

NOTE:

A full description of the Primary Flight Display appears in section 2B-05-00 and a description of the Flight Controls 2/3 synoptic page appears in section 2B-07-

00.

A. Circuit Breakers (CBs):

The following CBs protect the horizontal stabilizer system:

Circuit Breaker Name: CB Panel: Location: Power Source:

FLAP/STAB LEFT DC POP C-6 L Essential DC Bus FLAP/STAB RIGHT DC CPOP C-6 R Essential DC Bus FLAP/STAB L STBY AC POP D-2 L Standby AC Bus FLAP/STAB R STBY AC CPOP D-2 R Standby AC Bus

B. Crew Alerting System (CAS) Messages:

The following CAS messages are associated with the horizontal stabilizer system:

Area Monitored: CAS Message: Message Color:

EMER STAB Switch / Stabilizer

Motor Channels

Emergency Stabilizer On A-B

Amber

Flap Control Unit Flap Asymmetry Amber Flap Control Unit Flap Command Invalid Amber Flap Control Unit Flaps Failed Amber

Flap Control Unit

Flap/Stab IndependentOpAmber

Flap Control Unit Flap/Stab Miscompare Amber Flap Control Unit Flap/Stab Sync Fail Amber Flap Control Unit Stabilizer Failed Amber Flap Control Unit Uncommanded Flaps Amber Flap Control Unit Uncommanded Stabilizer Amber Flap Control Unit Flap/Stab Maint Reqd A-B Blue

Flap Control Unit

Flap/Stab System Fail A-B

Blue

Flap Control Unit Flap/Stab Rig Complete Blue Flap Control Unit Stabilizer Syncing A-B Blue

4. Limitations: Flap Altitude and Airspeed Limits:

• Maximum operating altitude for extending landing ﬂaps (39° DOWN) or ﬂying with landing ﬂaps extended is 20,000 ft MSL

• Maximum operating altitude for extending ﬂaps to 10° or 20° or ﬂying with the ﬂaps extended to 10° or 20° is 25,000 ft MSL

• Maximum airspeed for ﬂaps extended to 10° is 250 KCAS / 0.60 M

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• Maximum airspeed for ﬂaps extended to 20° is 220 KCAS / 0.60 M

• Maximum airspeed for ﬂaps extended to 39° is 170 KCAS / 0.60 M

Use of Flaps:

Use of ﬂaps in icing conditions is restricted to takeoff, approach and landing only. Holding in icing conditions is limited to 0° ﬂaps only.

Emergency Stabilizer Trim:

Maximum airspeed with the emergency stabilizer armed with the autopilot engaged and a jammed elevator is 270 KCAS / M.80.

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Flap / Horizontal Stabilizer

System Block Diagram

Figure 13

2A-27-00

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Horizontal Stabilizer System Controls and Indications

Figure 14

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Wing Flaps System

Figure 15

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Flap Handle

Figure 16

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2A-27-60: Stall Warning and Prevention System (SWPS)

1. FAA Certiﬁed Airplanes Versus JAA Certiﬁed Airplanes:

In some instances, the Stall Warning and Prevention System (SWPS) differs between FAAcertiﬁed airplanes and JAA certiﬁed airplanes. This section presents a description of the system installed in FAA certiﬁed airplanes. A synopsis of the differences for JAA certiﬁed airplanes is presented in the Limitations portion of this description.

2. General Description:

The SWPS provides the ﬂight crew with visual indications of deteriorating airspeed and high angles of attack, a physical warning of an impending stall by activation of a yoke stick shaker and automatic elevator ﬂight control movement by a stick pusher to decrease angle of attack.

The SWPS functions are hosted in the autopilot (AFCS) processor modules of Modular Avionics Units (MAUs) #1 and #2. Since each MAU autopilot module is a fully-redundant independent processor, the SWPS is also a redundant system, with each MAU using separate data inputs from Angle of Attack (AOA) sensors, Flap Control Unit (FCU) channels, Inertial Reference Systems (IRSs), Air Data Systems (ADS) and Display Controllers (DCs). Each MAU also has independent outputs through Actuator Input/Output Processors (AIOPs) to separate channels for the yoke stick shakers and hydraulic actuator valves of the yoke stick pusher. A schematic of the system is contained in Figure 17.

Operation of the SWPS is controlled by the STALL BARR switch on the cockpit center console. The SWPS must be operative prior to commencing ﬂight, and the switch is required to be selected on for all ﬂight operations unless a malfunction occurs during ﬂight. (See Figure 18.)

3. Description of Subsystems, Units and Components: A. Angle of Attack (AOA) Sensors:

Because an aircraft stall is an aerodynamic condition caused by separation of the boundary layer of airﬂow over the wing at high angles of attack, AOA data is more accurate than airspeed in determining the onset of a stall. Airspeed is closely related to AOA, but is subject to indication and compressibility errors - however, since airspeed is the standard reference used by ﬂight crews to control angle of attack, the airspeed indication displayed on the Primary Flight Display (PFD) provides the best cues to the onset of a stall condition. Angle of attack is included on the PFD, but in an abbreviated format.

Two AOA sensors are installed on the G550, one on each side of the aircraft positioned below and aft of the pilot and copilot windows. The AOA sensors are powered and heated by the respective side essential DC bus (left AOA by left essential DC bus, etc.). The AOA sensors are vanes that measure the relative wind generated by the aircraft ﬂight path. Each AOA vane is mounted on a shaft that allows the vane to rotate, aligning the vane into the relative wind. The amount of vane rotation is measured internally by the AOA sensor and transmitted over ARINC-429 connections to the MAUs (the left AOA sensor to MAU #1, right AOA sensor to MAU #2). The SWPS function of the AFCS modules in the MAUs compare the relative wind data with effective chord of the aircraft wing to determine the aircraft angle of attack.

The effective chord of the wing changes with ﬂap extension and is sensed

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by the SWPS function through the dual channels of the Flap Control Unit (FCU). Position of each wing ﬂap is reported by Rotary Variable Displacement Transducers(RVDTs)to the respective FCU control channel. Each of the FCU channels is connected over ARINC-429 links to the MAUs. The SWPS software is programmed with a set of AOA values for each ﬂap setting, and triggers stall warning and prevention measures if the values are exceeded.

To ensure that the AOA values used by the SWPS are the most accurate possible, each MAU processor ﬁlters the input AOA data using time averaging and compares inputs with the AOA data from the other MAU processor and also with IRS pitch, roll, acceleration, change rates and True Air Speed (TAS) a ﬁnal reﬁned AOA value.

Since the display of actual AOA would not be meaningful to the ﬂight crew due to the change in signiﬁcance of the values with ﬂap setting, the software in the SWPS computes a normalized value for AOA. The normalized value uses a format of zero (0) to one (1) with the indication shown to two (2) decimal places, where 0.00 indicates zero lift on the wing and 1.00 indicates the point of maximum wing lift. A normalized AOA value of 1.00 is also the threshold of stick pusher activation, since any further increase in AOA above the maximum lift angle results in a stall. The normalized AOA value is displayed on the PFD below the airspeed tape.

B. Low Airspeed Indications:

The airspeed and AOA indications on the PFD provide the ﬂight crew with visual indications of deteriorating airspeed and high angles of attack. As airspeed slows, the stick shaker threshold point is shown beside the airspeed tape as a red bar descending with airspeed from the shaker initiation point. If airspeed is slowing at a rate that will soon reach the shaker initiation speed, the airspeed trend vector indicator will change to amber. If airspeed reaches the stick shaker threshold, the digital airspeed readout and AOA display will change to amber. Further airspeed loss that initiates activation of the stick pusher is denoted by the digital airspeed and AOA displays changing to red.

C. Yoke Stick Shakers:

Each cockpit yoke is equipped with an electrically powered stick shaker installed under the ﬂight deck ﬂoor. The motor for the pilot stick shaker is powered by the left essential DC bus; the copilot shaker is powered by the right main DC bus. The pilot yoke stick shaker is controlled by the SWPS function in MAU #1 and the copilot stick shaker by MAU #2. The SWPS software function activates the stick shakers at a threshold that provides adequate margin for pilot action to avoid a stall. The stick shaker activation point is set at:

• A normalized AOA value of 0.90 for the ﬁrst six (6) seconds after takeoff (weight off wheels) to prevent nuisance activations

• A normalized AOA value of 0.85 thereafter for all ﬂap settings

These stick shaker thresholds provide at least a ten percent (10%) airspeed margin prior to stick pusher activation and stall onset. Once the stick shakers have been triggered, only a decrease in AOA below the activation point will turn off the stick shakers. If a stick shaker malfunctions and activates below the normalized AOA trigger threshold, the circuit

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breaker for the faulty stick shaker can be pulled (SHAKER #1 on the pilot overhead panel, SHAKER #2 on the copilot overhead panel).

The MAUs monitor their respective data input paths and SWPS activation paths for operational capabilities. If faults or failures are found, the status is reported to the Monitor and Warning System (MWS) that will in turn generate appropriate CAS messages for ﬂight crew notiﬁcation.

D. Yoke Stick Pusher:

If the visual display of low airspeed and high AOA in conjunction with activation of the cockpit yoke stick shakers does not prompt a decrease in AOA and airspeed gain, the SWPS function of the autopilot processors will command the aircraft elevators to a nose down position in an attempt to avoid stalling the aircraft wing. If the aircraft approaches a normalized AOA of 1.00, each SWPS channel will separately signal activation of a dedicated hydraulic solenoid valve on the stick pusher mechanism in the tail compartment. The stick pusher mechanism is installed just aft of the dual elevator hydraulic actuators, and employs a linkage connected by a cam arm to provide a nose down movement to the input side of the elevator actuators. Activation of the stick pusher will not interrupt the action of the stick shakers.

The SWPS software provides an anticipatory function for the stick pusher, so that at lower airspeeds, stick pusher activation will respond to the rate of change in aircraft body angle (as sensed by the IRS) to prevent the aircraft from overshooting the 1.00AOA stall point. As airspeed increases from .76 Mach to .80 Mach the anticipatory function is biased out of the activation logic, since large control inputs are not necessary at high airspeeds.

The stick pusher will deactivate the nose down input to the elevators when aircraft acceleration drops below 0.5 G or the actual wing AOA decreases

3.6° below the stick pusher activation threshold. If a malfunction causes a spurious activation of the stick pusher, a manual

force on the control column of approximately seventy-ﬁve pounds (75 lbs) will overcome the nose down control input. The stick pusher may be electrically disconnected by depressing either of the pilot or copilot yoke A/P DISC BARR DISC switches or by selecting the STALLBARR switch on the cockpit center pedestal to off (the OFF legend in the pushbutton switch will illuminate amber). Normally the yoke switches are used if the crew is hand ﬂying the aircraft and the pedestal pushbutton is used during autoﬂight to avoid disconnecting the autopilot.

The MAUs monitor the performance of the data inputs used to activate the stick pusher and the health of the stick pusher hydraulic actuation installation. Any degraded performance is annunciated to the crew through MWS initiated CAS messages.

E. SWPS Ground Test:

Although the stall warning and prevention system does not operate with the aircraft on the ground (weight on wheels), a preﬂight test of the system may be conducted by selecting the system on the TEST menu of the Display Controllers (DCs). Each independent SWPS channel has a dedicated Line Select Key (LSK) on the DC TEST menu. Depressing the LSK next to the STALL 1 or STALL 2 legend will ﬁrst activate the stick shaker for one second, then activate the stick pusher to the nose down position. The test

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function of the SWPS is operational only with the aircraft on the ground to preclude inadvertent activation of the system during ﬂight.

4. Controls and Indications:

(See Figure 0.)

NOTE:

A full description of the Primary Flight Display appears in section 2B-02-00: Electronic Display System Description.

A. Circuit Breakers (CBs):

The following CBs provide power to SWPS components:

Circuit Breaker Name: CB Panel: Location: Power Source:

AOA PROBE #1 POWER LEER F-2 Left Essential DC Bus #1 AOA HTR POP A-2 Left Essential DC Bus AOA PROBE #2 POWER REER F-13 Right Essential DC Bus #2 AOA HTR CPOP A-2 Right Essential DC Bus SHAKER #1 POP E-6 Left Essential DC Bus SHAKER #2 CPOP E-6 Right Main DC Bus STALL BARR VALVE #1 POP E-5 Left Essential DC Bus STALL BARR VALVE #2 CPOP E-5 Right Essential DC Bus

B. Crew Alerting System (CAS) Messages:

The following CAS messages are associated with the SWPS system:

Area Monitored: CAS Message: Message Color:

AOA Probe AOA Probe 1-2 Fail Amber SWPS function AOA Miscompare Amber SWPS function Stall Barrier 1-2 Amber

SWPS function / STALL BARR

SWPS function / hydraulic SWPS function / hydraulic SWPS function / hydraulic

Switch actuator actuator

actuators

Stall Barrier Off Amber Stick Push 1-2 Fault Amber Stick Push 1-2 Fail Amber Stick Push Unavailable Amber

SWPS function / electric motor Stick Shake 1-2 Fail Blue

5. Limitations: A. AFM Limitations for Stall Warning / Stall Barrier System (FAA Certiﬁed

Airplanes):

(1) Takeoff Requirements:

Both stall warning / stall barrier systems must be operable for takeoff.

(2) Use of System:

Operative stall barrier systems must be ON during all ﬂight

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operations, unless required to be selected OFF for procedural reasons. See Section 05-13-50, Stall Barrier Malfunction for additional information.

B. Synopsis of Stall Warning / Stall Barrier System Differences for JAA

Certiﬁed Airplanes:

(1) Stall Warning System Operational Logic:

The stall warning system operational logic has been modiﬁed to allow the system to continue to operate until Angle of Attack (AOA) is reduced below the initiation threshold.

(2) Stick Shaker System Inhibition Logic:

The stick shaker system has been modiﬁed to no longer be inhibited during control stick push.

(3) Stick Pusher System Inhibition Logic:

The stick pusher system has been modiﬁed to no longer be inhibited when the autopilot disconnect switch is held in the ″IN″ position.

(4) Stick Pusher System Activation Logic:

The stick pusher system has been modiﬁed to not activate until both AOA vanes are at or above the push AOA.

(5) Stick Shaker / Stick Pusher Threshold Ratio:

• [00b7] The AOA stick shaker/stick pusher threshold ratio has been modiﬁed to 0.9 when the Weight-On-Wheels (WOW) system is in the GROUND mode and for the ﬁrst six (6) seconds after transition to the AIR mode

• [00b7] For all other phases of ﬂight, the AOA stick shaker/ stick pusher threshold ratio will equal 0.85

(6) Stall Barrier Disconnect (BARR DISC) Buttons:

The control yokes have been modiﬁed removing the BARR DISC legend from the autopilot disconnect button.

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Stall Warning / Stick

Pusher Block Diagram

Figure 17

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STALL BARR Switch

Figure 18

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2A-27-70: Speed Brake and Ground Spoilers

1. General Description:

The speed brake and ground spoilers use six panels on the upper wing surface to increase drag and reduce lift allowing the ﬂight crew to increase descent rates and, on landing, to shorten landing roll and increase braking effectiveness by quickly transferring aircraft weight to the main landing gear wheels. The three panels on each wing are hydraulically powered and activated by either manual command through the SPEED BRAKE handle on the pilot side of the cockpit center console, or electrically actuated as an automatic selectable feature during landing. The automatic ground spoilers will also deploy during an aborted takeoff to reduce aircraft stopping distance.

Hydraulic power for speed brake and ground spoiler actuation is provided by both the left and right hydraulic systems, with the pressure of each system reduced from three thousand (3,000) psi to ﬁfteen hundred (1,500) psi by a pressure control module to prevent structural damage to the aircraft from speed brake deployment at high airspeeds. Two hydraulic actuators on each wing operate the spoiler panels in response to manual or electrical control inputs. Each actuator has tandem pistons, one for each hydraulic system, that move a single actuator shaft. Failure of one hydraulic system will not effect the actuators - the remaining system pressure reducer will allow a full three thousand (3,000) psi to the operational piston to power actuator shaft movement, and the condition will by annunciated by a blue “Single Speed Brake” advisory Crew Alerting System (CAS) message. Diagrams of the spoiler system is included in Figure 19and Figure 20.

The two hydraulic actuators are necessary to provide the multi-purpose operation of the spoiler system. The two outer spoiler panels on each wing operate in three (3) modes:

• To assist the ailerons in aircraft roll control

• To actuate as speed brakes

• To actuate as ground spoilers

The single inner spoiler panel on each wing operates as:

• A speed brake

• A ground spoiler

The two hydraulic actuators and the three spoiler panels on each wing are mechanically linked together; however, the linkage is unidirectional, operating from the inboard ground spoiler actuator outboard to the ﬂight spoiler actuator. All speed brake or ground spoiler commands, either manual or electric, are directed to the ground spoiler actuator that mechanically relays the command to the ﬂight spoiler actuator. The single direction of the mechanical linkage is necessary in order for the two outboard spoiler panels of each wing to assist the ailerons without operating the inboard spoiler panel. The aileron control linkage to the outer panels is a separate command input that is cumulative to the speed brake deployment command. For more information regarding the roll assist function of the outer spoiler panels, see section 2A-27-40: Aileron Roll Control.

Both actuator linkages on each wing have dedicated Rotary Variable Displacement Transducers (RVDTs) that furnish position information to the Modular Avionics Units (MAUs). The left ground and ﬂight spoiler actuator linkages are monitored by RVDTs linked to MAU #1, and both right wing spoiler RVDTs are monitored by MAU #2. The MAUs share position information with the

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Monitor and Warning System (MWS) in order to generate CAS messages in the event of a malfunction, and to format the graphic display of spoiler position on the Flight Controls 2/3 synoptic page.

Two additional contact switches, one on each wing, are closed when the spoilers are ﬂush with the wing surface. The switches, located on the aft wing beam provide position signals to the MAUs (left spoiler to MAU #1, right spoiler to MAU #2) conﬁrming that the spoilers are stowed.

If a malfunction of the spoiler system occurs, a single pushbutton, labelled SPOILER CONTROL, located just forward of the speed brake handle can be used to deactivate the spoilers. Depressing the pushbutton will illuminate the amber OFF legend within the button and close two (left and right) spoiler control valves located in the aft wing to fuselage fairing, bypassing left and right system hydraulic pressure to the spoiler actuators.

2. Description of Subsystems, Units and Components: A. Speed Brake:

All six spoiler panels will deploy upward from the wing surface in response to manual commands from the speed brake handle on the cockpit center console, shown in Figure 16. An internal light, powered by the right essential DC bus, will illuminate the speed brake handle as it is moved aft from the stowed position. The speed brake handle is connected by wire cables and linkages to the ground spoiler hydraulic actuators that position the inboard spoiler panels on each wing. The ground spoiler hydraulic actuators are in turn mechanically linked to the ﬂight spoiler actuators on each wing that position the middle and outboard spoiler panels. The amount of spoiler panel deﬂection is dependent upon the position of the speed brake handle. The handle may be moved to any range from the retract position to fully extended position with no detents in the handle range. At the fully extended position, the two inboard spoilers will deploy to thirty degrees, plus or minus one degree (30° ±1°) and the two outboard spoilers on each wing will deploy to thirty degrees, plus or minus four degrees (30° ±4°). The difference in tolerances from the thirty degree (30°) full setting is due to the rigging of the linkages between the ground and ﬂight spoiler actuators.

The position of the SPEED BRAKE handle is monitored by an Input / Output (I/O) module in MAU #2. When the handle is out of the retract position, the I/O module communicates with the MWS to generate a blue “Speed Brake Extended” advisory CAS message. If the power levers are positioned forward of the idle position with the speed brakes extended, an amber “Speed Brake Extended” caution CAS message is displayed as a reminder to the ﬂight crew.

NOTE:

Rapid descents using the speed brake with the power levers at idle produce high closure rates with the desired or assigned altitude. If not recognized early, the recovery response is often to add power to slow the descent rate without ﬁrst retracting the speed brakes.

Additionally, if the ﬂight crew commences a roll command to the ailerons

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with the yoke or autopilot while the speed brake handle is extended, the two outboard spoiler panels on each wing will increase displacement over the speed brake commanded spoiler position to aid in aileron roll control. The maximum displacement of the outboard spoiler panels is ﬁfty-ﬁve degrees, plus or minus four degrees (55° ±4°) with full speed brake and aileron commands.

B. Ground Spoilers:

The automatic operation of all six spoiler panels to reduce landing roll distance or decrease stopping distance in the event of an aborted takeoff is an electrically controlled function. Ground spoilers use no manual inputs via control cables or linkages but instead use electro-hydraulic servo valves to control the ﬂow of hydraulic system pressure within the ground spoiler hydraulic actuators to extend the spoiler panels. The servo valves are electrically operated and use left hydraulic system pressure to control the actuators. If left system pressure is not available, Power Transfer Unit (PTU) pressure or Auxiliary (AUX) system pressure will operate the servo valves using left system ﬂuid.

To provide protection against spurious activation of the ground spoilers, two levels of servo control safeguards are incorporated. The servo valves are plumbed in series and both must be open to command actuation of ground spoilers, and each servo valve has an independent electrical power circuit for operation.

When the servo valves operate the ground spoiler actuators, system operation is the same as that in speed brake operation: the hydraulic actuators use both left and right hydraulic system pressure, reduced to ﬁfteen hundred (1,500) psi. to operate the inboard spoiler panels and the mechanical linkage to the outboard ﬂight spoiler actuators commands operation of the two outboard panels on each wing. All six wing panel spoilers extend to ﬁfty-ﬁve degrees plus or minus four degrees (55° ±4°).

For automatic operation, ground spoilers must ﬁrst be armed using the GND SPLR pushbutton on the cockpit center console (the SPOILER CONTROL switch must also be in the ON position). The ground spoiler pushbutton has two positions, OFF or ARMED, and the corresponding legend within the pushbutton will illuminate according to the selection made. See Figure 22. With the pushbutton ARMED, the ground spoilers will activate whenever all the required elements in the operational logic are satisﬁed.

Normal Operational Logic:

Because the simultaneous deployment of all six spoiler panels to ﬁfty-ﬁve degrees (55°) results in such an abrupt change in the wing airfoil, the logical steps necessary for activation are designed to ensure that automatic operation occurs only on the ground and with the concurrence of the ﬂight crew. The following conditions must be met to automatically deploy ground spoilers during normal landing operations:

(1) The SPOILER CONTROL pushbutton selected ON. (2) The GND SPLR pushbutton selected to ARM. (3) Wing ﬂaps selected to greater than twenty-two degrees (22°). (4) Both power levers retarded to ﬂight idle.

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(5) Both main landing gear weight on wheels (WOW) switches in the

ground position OR both main landing gear wheel speeds (sensed by the anti-skid sensors) greater than forty-seven knots (>47 kts) OR one main landing gear WOW switch in the ground position and the other main landing gear wheel speed greater than forty-seven knots (>47 kts).

NOTE:

The relationship between the WOW switches and wheel speed is necessary to accommodate the wide range of conditions encountered during landing. An extremely gentle touchdown would spin up the wheels but not immediately compress the WOW switches. A very ﬁrm landing may compress the WOW switches but not immediately spin up the wheels. Strong crosswinds may result in one WOW (upwind) switch being compressed and the opposite wheel (downwind) spinning up prior to the WOW switch being compressed.

CAUTION

A HARD LANDING THAT RESULTS IN A SIGNIFICANT BOUNCE WILL DEPLOY THE GROUND SPOILERS IF THE POWER LEVERS ARE MAINTAINED AT IDLE, RESULTING IN A RAPID LOSS OF LIFT AND AIRSPEED DURING THE SUBSEQUENT TOUCHDOWN. ADVANCING THE POWER LEVERS WILL RETRACT THE GROUND SPOILERS AND DECREASE THE SINK RATE IF THE AMOUNT OF RUNWAY REMAINING ALLOWS SAFE COMPLETION OF THE LANDING.

For automatic ground spoiler deployment in conditions that require landing with less than full (39°) ﬂaps, the following logic parameters are necessary for abnormal landing operations:

(6) The SPOILER CONTROL pushbutton selected ON. (7) The GND SPLR pushbutton selected to ARM. (8) Wing ﬂaps selected to less than twenty-two degrees (22°). (9) The GPWS / GND SPLR FLAP ORIDE pushbutton on the center

console forward of the ﬂap handle selected to ON (amber ON legend in the pushbutton illuminated) - see Figure 21.

(10) Both power levers retarded to idle.

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(11) Both main landing gear weight on wheels (WOW) switches in the

NOTE:

The GPWS / GND SPLR FLAP ORIDE pushbutton selection furnishes an alternate relay path bypassing the ﬂaps greater than twenty-two degrees (>22°) requirement and also prevents nuisance annunciations from the GPWS module of the MAUs due to ground proximity without the normal landing ﬂap conﬁguration.

To enable automatic deployment of ground spoilers for reducing stopping distance in event of an aborted takeoff, the following conditions must be met for normal takeoffs.

(12) The SPOILER CONTROL pushbutton selected ON. (13) Wing ﬂaps selected to takeoff setting - ten or twenty degrees (10° or

20°). (14) Power levers advanced greater than idle. (15) GND SPLR pushbutton selected to ARMED. (16) WOW switches in the ground mode (compressed). (17) Power levers subsequently retarded to idle.

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NOTE:

It is not necessary for wheel speed to increase over forty-seven knots (47 kts) for the ground spoilers to deploy during an aborted takeoff.

C. Malfunction Monitoring - Landing and Takeoff:

When the automatic ground spoilers are armed for landing or for aborted takeoff protection, the system is monitored by the MWS for correct performance. The MWS veriﬁes all of the control logic parameters, senses uncommanded pressurization of the ground spoiler control servos, and circuit integrity to the servos and power lever position switches. If all of the control logic parameters for the speciﬁc aircraft conﬁguration as previously described have been satisﬁed, but either or both of the ground spoilers do not deploy, both red NO GND SPLRS lights located on either side of the forward windshield center post, will illuminate, a red CAS warning message of “Ground Spoiler” is displayed accompanied by illumination of the Master Warning lights on the cockpit glareshield and annunciation of the warning chime signal over the cockpit speakers. (Ground spoiler operation is sensed by the stow position contact switches on the inboard spoiler panels on each wing - the inboard ground spoilers are position monitored since the mechanical linkage from the inboard actuators operates the actuators for

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the two outboard ﬂight spoilers.) The malfunction annunciations are depicted in Figure 22.

NOTE:

Failure of the automatic deployment of ground spoilers does not preclude operation of the spoiler panels manually. Extending the SPEED BRAKE handle will deploy all spoiler panels to thirty degrees (30°) - less than the full ﬁfty-ﬁve degrees (55°) with automatic operation, but still effective in slowing the aircraft.

At any time the left main landing gear is locked in the down position (inﬂight or on the ground) and the GND SPLR pushbutton switch is not in the ARMED position (amber OFF legend within the switch illuminated), a blue advisory CAS message reading “Ground Spoiler Unarm” is displayed. The same message will be displayed if a malfunction in the logic arming circuitry prevents operation of automatic ground spoilers with the GND SPLR switch selected toARMED (amber ARMED legend illuminated in the switch).

D. Malfunction Monitoring Inﬁght:

The uncommanded deployment of a spoiler panel during ﬂight may cause structural damage or seriously degrade the stability of the aircraft. If the SPEED BRAKE handle is in the forward stowed position and the GND SPLR switch is not armed, or if the switch is armed in preparation for landing but the aircraft is not yet on the runway, and one or both of the ground spoiler panels is not in the stowed position (sensed by the contact switches), the red “Ground Spoiler” CAS message will be displayed along with illumination of the Master Warningglareshield lights and aural warning tones.

The MWS will also generate a red “Ground Spoiler” warning if conditions are detected that could cause an uncommanded spoiler activation. Areas sampled are the same as those described during takeoff and landing with the addition of fault monitoring of the WOW switches and wheel speed circuits.

Removing hydraulic power to the spoilers by pulling the SPLR FLT PWR S/O circuit breaker on the copilot overhead panel will bypass all hydraulic pressure in the ground and ﬂight spoiler actuators and allow aerodynamic pressures to stow the extended spoiler panel. See 05-13-60, Ground Spoiler Failure Inﬂight..

E. Ground Spoiler Test Switch:

A GND SPLR TEST switch is located on the cockpit center console above the GND SPLR ARMED / OFF pushbutton as shown in Figure 22. Depressing and holding in the switch will illuminate the amber IN TEST legend in the switch and test the following indications:

• Both NO GND SPLRS red lights on the windshield center post will illuminate

• A red “Ground Spoiler” warning message is displayed on the CAS window

• The red Master Warning glareshield light and aural warning tones

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are activated

Prior to the ﬁrst ﬂight of the day, a complete check of the automatic ground spoiler system is usually conducted using the test switch and other controls to verify the operation of the system activation logic. See the Before Takeoff: 03-04-00 section of this manual for a complete description of the entire test.

3. Controls and Indications:

(See Figure 21 and Figure 22.)

NOTE:

Since the spoiler panels on the upper wing surfaces are not visible from the cockpit inﬂight, the position of the spoilers is monitored using the Flight Controls 2/3 synoptic page. A full description of the Flight Controls 2/3 synoptic page appears in section 2B-07-00.

A. Circuit Breakers (CBs):

The following CBs protect the spoiler system:

Circuit Breaker Name: CB Panel: Location: Power Source:

SPDBRK ALARM REER A-26 Right Essential DC Bus SPLR FLT PWR S/O CPOP E-3 Right Essential DC Bus GND SPLR CPOP E-2 Right Essential DC Bus LEFT WOW POP C-1 Left Essential DC Bus RIGHT WOW CPOP C-1 Right Essential DC Bus WHEEL SPEED POP C-3 Left Essential DC Bus THROT QUAD CPOP C-2 Right Main DC Bus

B. Crew Alerting System (CAS) Messages:

The following CAS messages indicate the status of the speed brake / ground spoiler system:

Area Monitored: CAS Message: Message Color:

MAUs #1 and #2, Speed Brake handle, spoiler panel RVDTs / stowed contact switches

MAUs #1 and #2, ground spoiler arming logic circuit, stowed contact switches

Spoiler Control pushbutton, spoiler control valves

MAU #2, Speed Brake handle switch and power lever position

Left main landing gear downlock, GND SPLR ARMED pushbutton

MAUs #1 and #2, pressure reducing module

Aircraft Conﬁguration Red

Ground Spoiler Red

Spoilers Hydraulic Off Amber Speed Brake Extended Amber Ground Spoiler Unarm Blue Single Speed Brake Blue

Speed Brake handle switch Speed Brake Extended Blue

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A. Flight Manual Limitations:

(1) Use of Speed Brakes:

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

(2) Automatic Ground Spoilers:

Takeoff is permitted with automatic ground spoilers inoperative provided the Anti-Skid is operative, 20° ﬂaps are used and the cowl and wing anti-ice systems are not used. Dispatch with reference to MMEL.

If a touch-and-go landing is to be performed, the GND SPLR must be OFF and manual spoiler landing distances must be taken into account.

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Spoiler System

Figure 19 (Sheet 1 of 3)

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Spoiler System

Figure 19 (Sheet 2 of 3)

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Spoiler System

Figure 19 (Sheet 3 of 3)

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Ground Spoiler Circuits

Figure 20

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Spoiler System Controls

Figure 21

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Spoiler System Indications

Figure 22

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2A-27-80: Flight Controls Gust Lock

1. General Description:

The ﬂight controls gust lock secures the aircraft moveable surfaces in stabilized positions to prevent damage to the surfaces and attached control linkages from high winds or jet blast. A GUST LOCK handle on the copilot side of the cockpit center console (see Figure 23) is connected through a system of cables, pulleys and rods to hooks or latches that engage the mechanical linkages of ﬂight controls to prevent movement. The control surfaces are locked in positions that offer the best protection to the aircraft. The ailerons and rudder are locked in neutral faired positions, and the elevators are locked in a slightly trailing edge down position so that any wind force would tend to maintain weight on the nosewheel to preserve aircraft stability.

NOTE:

The gust lock is effective in preventing ﬂight control movement during wind speeds up to sixty knots (60 kts). If weather conditions are forecast to include stronger winds, consideration should be given to securing the aircraft within a hangar or other suitable shelter.

2. Description of Subsystems, Units and Components: A. GUST LOCK Handle:

The aircraft gust lock is engaged with the T-shaped handle on the cockpit center console. In order to move the handle from either the engaged or disengaged position, a spring loaded latch on the back of the handle must be depressed. The spring latch prevents inadvertent engagement and also maintains engagement of the gust lock once selected. The handle rotates aft from the horizontal disengaged position up to a ninety degree (90°) vertical position when engaged in order provide a strong visual cue that the aircraft ﬂight controls are locked.

B. Flight Control Latches:

As the GUST LOCK handle is rotated aft and up, the motion is translated by a cable, rods and pulleys to the locking mechanisms on the ﬂight control linkages. The lateral motion of the single cable rotates pulleys that in turn move latching arms into position to engage the command input links of the ﬂight controls. The latches will not lock in the engaged position until the ﬂight control is placed in the speciﬁed position necessary to minimize the effect of wind forces on the aircraft. Section 03-06-20: After Landing / Shutdown contains the following CAUTION notiﬁcation:

ENSURE HYDRAULIC PRESSURE IS DEPLETED PRIOR TO ENGAGING THE GUST LOCK. IF IT IS NOT POSSIBLE TO READ HYDRAULIC PRESSURES BECAUSE THE AIRCRAFT IS WITHOUT ELECTRICAL POWER, CYCLE THE CONTROLS WITH THE CONTROL COLUMN, YOKE AND RUDDER PEDALS TO DEPLETE RESIDUAL HYDRAULIC PRESSURE.

C. Mechanical Power Lever Interlock:

A mechanical interlock within the cockpit center pedestal prevents advancing the power levers more than six percent (6%) forward of ground

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idle if the GUST LOCK handle is engaged. The blocking action of the interlock cannot be overcome with manual force to advance the power levers.

3. Controls and Indications:

(See Figure 23.) There are no circuit breakers or CAS messages related to the gust lock.

4. Limitations: A. Flight Manual Limitations:

There are no limitations for the gust lock system at the time of this writing.

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Gust Lock System Controls

Figure 23

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