Flight Controls.docx

2A-27-10: General Flight Controls General Description: (See Figure 1. Flight Controls System: Control Surface Nomenclature, Figure 3. Flight Control Mode Comparison, and Figure 2. Flight Control Power Sources.) The flight control system consists of the primary Flight Control System (FCS), the...

2A-27-10: General Flight Controls General Description: (See Figure 1. Flight Controls System: Control Surface Nomenclature, Figure 3. Flight Control Mode Comparison, and Figure 2. Flight Control Power Sources.) The flight control system consists of the primary Flight Control System (FCS), the flap system and the horizontal stabilizer system. The primary FCS includes the components that control the two ailerons, six spoilers, two elevators and one rudder. These surfaces are hydraulically actuated and electrically controlled with fly-by-wire technology. The flap actuation system includes the Flap Electronic Control Unit (FECU) and the drive system. The horizontal stabilizer system includes a Horizontal Stabilizer Control Unit (HSCU) and a Horizontal Stabilizer Trim Actuator (HSTA). Nosewheel steering and brakes are not considered to be part of the FCS in this document. The pilot and copilot primary flight controls include control wheels, columns and rudder pedals. Each primary control is connected to position sensors and artificial feel mechanisms. The position sensors provide the position of the respective control device to the Flight Control Computers (FCCs) and Backup Flight Control Unit (BFCU). The artificial feel mechanisms provide tactile cues to the pilots. The pilot and copilot flight controls are mechanically linked and are designed to operate in unison. They have springs and dampers to provide appropriate "feel" forces and mechanical stops to limit their maximum displacement. The wheel and column (but not the rudder pedals) are linked by bungee rods to allow control in the event of a mechanical jam on the pilot or copilot side. The bungee rods act as a solid link unless the forces on the pilot and copilot side differ by a large amount due either to a jam or opposing inputs from the pilots. The pilot and copilot rudder pedals are mechanically connected with a hard link which cannot be overridden. In the event of a rudder pedal jam, the aircraft can be safely flown with rudder trim. Trim switches are provided for pitch, roll, and yaw axes. There are two dual-channel FCCs and a Backup Flight Control Unit (BFCU) which generate commands for the control surface actuators. The FCCs house the operational logic (also known as Control Laws) for control of the flight control surfaces. Any one of the four FCC channels is independently fully capable of safely operating the entire aircraft FCS. In the event that both FCCs (all four channels) are inoperative, the BFCU controls the elevators, ailerons, outboard spoilers (roll control only, no speed brakes), and rudder. The aircraft may be dispatched (with restrictions) with one FCC channel inoperative or the BFCU inoperative with restriction. See MMEL for dispatch requirements. The primary FCS actuators provide a damping feature to protect the primary control surfaces from potential damage caused by wind gusts (up to 65 knots) while the aircraft is on the ground with hydraulic systems depressurized. There is no "gust lock" mechanism. The primary and secondary flight controls are discussed in the following subsections: 2A-27-20: Aileron and Roll Spoiler Control System 2A-27-30: Rudder Control System 2A-27-40: Elevator Control System 2A-27-50: Horizontal Stabilizer System 2A-27-60: Wing Flaps System 2A-27-70: Speed Brake and Ground Spoilers System System Power On Self-Test (SPOST): The Cockpit Preflight Inspection checklist (AFM 02-01-10) and the Before Starting Engines checklist (AFM 02-03-20) both have distinct SPOST executions. When performing the 'Cockpit Preflight Inspection' checklist (AFM 02-01-10): During the first flight of the day, a separate FCC SPOST is performed with the MAIN BATTERIES selected ON without APU GEN power. FCC 1B and 2A are powered by the Main Batteries. When the EBHA and then the UPS switches are selected ON, the SPOST will exercise the system without having the FCC Batteries receiving power from the EBHA and UPS Battery chargers in order to accurately check battery health. The SPOST will exercise the powered systems, associated with flashing indications on the pedestal in the cockpit and 8 cycles of the pilot side stick shaker. Additionally, at the end of the cockpit preflight check initiated and the be present, the Flight Control synoptic will be stable (not flashing), and the FLT CTRL Battery voltage should be within acceptable voltage parameters as indicated in the preflight checklists. The CAS message is due to the fact that the Flight Control System is not receiving inputs from the IRUs, ADS, and other unpowered systems. If the SPOST is in progress and does not successfully finish (indicated by less than eight stick shaker momentary activations and the presence of the FCS Maintenance Required (caution) CAS) then the FCC reports the SPOST as failed and posts an FCS Test Fail (caution) CAS. The most likely reason for this failure is removing power to any FCC channels while the SPOST is in progress. Once an FCC channel reports the SPOST as failed, it is latched until power is cycled to the FCC channels twice. Thus, if during the Cockpit Preflight checks the crew were to remove power during the SPOST by prematurely deselecting FCS or Main battery power, then the SPOST will not run during the normal Before Starting Engines sequence and the CAS will post. To reset a latched FCC SPOST failure, power must be totally removed for one minute before the next aircraft power up attempt. When performing the normal 'Before Starting Engines' checklist (AFM 02-03-20): The SPOST is a feature of the FCC channels that runs when power is first applied by selecting the FLT CTRL BATTERIES (EBHA/UPS) switches to the ON position. During the SPOST, the FCC runs through a series of checks and the pilots will observe multiple flight control capsule lights on the center pedestal flashing as well as both stick shakers momentarily running for a total of eight cycles. A successful SPOST will terminate with the FCC Alternate Mode (caution) CAS and no FCS Maintenance Required (caution) CAS approximately 45-50 seconds after UPS power and about 10 seconds after the last stick shaker vibrates. An unsuccessful SPOST will be indicated by an FCS Maintenance Required (caution) CAS and this will require a complete removal of power for a minute to reset the FCCs The APU GEN powers both the Essential DC and the Main DC busses. When this happens, FCC channels 1B and 2A are powered along with all of the normal associated Flight Control System busses. When the EBHA switch is selected ON, the MCEs are powered. Subsequent selection of the UPS switch to ON powers the remaining two FCC channels (1A and 2B) and the SPOST will initiate within 10 seconds with all the FCS components electrically powered. At the conclusion of the SPOST, approximately 45 seconds after initiation, the FCC Alternate Mode (caution) CAS should be the only FCS CAS message. It will extinguish when the four 'ANTI ICE HTR AIR DATA PROBE' heaters are selected ON after Engine Start is complete. Primary Flight Control System Architecture: The aircraft Fly-By-Wire (FBW) primary FCS is powered by the aircraft electrical and hydraulic systems. The primary FCS utilizes the following equipment: Flight deck controls and displays Flight Control Computers Air Data sensors Inertial sensors Flight Control Actuators Flight control surfaces Analog and digital communications to link the above components The left and right hydraulic systems are connected to the actuators in such a way that if one hydraulic system is lost, all primary flight control surfaces (other than one mid or inboard spoiler pair) remain operational. Components: The flight deck control devices include: Conventional control wheel, control column and rudder pedal assembly for each pilot Speed brake handle Trim control switches Flap handle Flight Control Reset Switch Each of the above controls have multiple position sensors that are read by the FCCs The BFCU receives position sensor information from the control wheel, control column and rudder pedal assembly for each pilot and trim control switches (Backup Pitch trim ONLY) The primary FCS includes the following components: Two dual-channel FCCs One BFCU Nine Hydraulic Actuators (HA) - one rudder, two ailerons, two elevators, two mid spoilers and two inboard spoilers Seven Electric Backup Hydraulic Actuators (EBHA) - one rudder, two ailerons, two elevators and two outboard spoilers The Flap Actuation System includes the following components: FECU in the Right Electronics Equipment Rack (REER) Flap Hydraulic Control Module (HCM) in the main wheel well Flap Power Drive Unit (PDU) in the main wheel well Inboard flap actuators (one per wing) Outboard flap actuators (one per wing) Torque shafts (eight per wing) Pillow blocks (six per wing) Flight Control Computers (FCCs): (See Figure 4. LEER Flight Control Computer (FCC) Location and Figure 5. REER Flight Control Computer (FCC) Location.) Two dual-channel FCCs are installed in the aircraft (located in the Left and Right Electronics Equipment Racks [LEER and REER]). Each of the four channels is comprised of a command lane and a monitor lane. The two lanes provide integrity to the system by operating with different software and as a self-checking pair. The monitor lane continuously checks its own computations with the commands transmitted by the command lane. Significant differences between the two lanes cause that channel to be shut down. Any single channel is capable of controlling all flight control surfaces. This configuration provides four redundant dissimilar and independent channels of operation. Each FCC channel contains the operation logic for each flight control surface, previously identified as Control Laws. The FCCs provide excitation to and receive inputs from the flight deck control device sensors (and other aircraft sensors and systems). The FCCs use this data to calculate position command signals to the 16 flight control surface actuators. Position sensors at the control surfaces feed the position of the control surface back to the FCC. The FCCs monitor and report the health status of the computing channels, flight deck sensors, air data sensors, inertial data sensors, horizontal stab system, BFCU, and the HA and EBHA actuators. Any abnormal condition is reported through the CAS, FCS Synoptic page, and/or the Central Maintenance Computer (CMC). Functions such as system rigging and system configuration checks can be also be accomplished through the CMC. Communication between the FCCs and the BFCU, HAs, EBHAs, HSTS, FECU and other aircraft systems is primarily through ARINC 429 digital communication buses. The buses from the FCCs to the actuators have redundancy that allows system operation after damage from birdstrikes, tire bursts, engine rotor bursts, etc. The FCS has extensive built-in test capabilities. Some tests are run automatically with each power cycle and others are initiated by maintenance personnel. Each FCC provides the following functions: Flight control surfaces commands as a function of flight conditions and aircraft configuration for pilot and autopilot inputs based on the control laws Secondary command inputs to the FECU based on flaps handle microswitch status Command inputs to the HSTS (based on pitch trim inputs and Control Law commands) Yaw damper Computation of low speed awareness information such as Normalized AOA, pitch limit indicator, and low speed limits for the airspeed tape on the PFD Interface with the surface actuators, FECU and the HSCU by data buses FCS Built-In-Test initiation and fault reporting Backup Flight Control Unit (BFCU): (See Figure 6. Backup Flight Control Unit Architecture.) The BFCU is designed to maintain emergency flight and landing capability in the event all four channels of the FCCs become unavailable. The BFCU is located under the cabin floor aft of the main entrance door. The BFCU interfaces with the following components: Provides excitation to and processes signals from flight deck control sensors (one for each wheel, one for each column, and one for rudder pedals) Provides BFCU health status and flight deck sensor position data to each FCC (which provides system monitoring of the BFCU to report a failure) Provides surface position commands directly to each of the seven EBHA REUs to allow operation during either normal hydraulic or Electrical Backup operation. The backup mode is not selectable by the flight crew. The BFCU monitors all four FCC channels and will assume control of the FCS within one half of a second if all four channels of the FCCs become unavailable. An FCC channel is unavailable if it is unpowered or is unable to compute in either normal, alternate, or direct modes. The BFCU provides control surface inputs to the EBHAs on the following control surfaces: Both Ailerons Outboard Spoilers Both Elevators Rudder When in the BFCU mode of operation, these flight control surfaces have limited travel. This helps to avoid over-controlling or over-stressing the aircraft when operating at high speed. Hydraulic Actuators (HAs) and Electric Backup Hydraulic Actuators (EBHAs): (See Figure 7. Elevator Electrical Backup Hydraulic Actuator (EBHA) and Electro-Hydraulic Servo Actuator (EHSA) Components through Figure 10. Inboard, Mid and Outboard Spoiler Hydraulic Components.) The flight control surfaces are electrically-controlled and hydraulically-powered. There are two types of primary flight control surface actuators: the HA and the EBHA. The EBHAs have the ability to operate normally using the engine-driven-pump hydraulic system supply or in Electric Backup (EB) operation using a built-in motor-pump in the EB manifold powered by 28V DC from the EBHA battery bus. HA: Each HA is controlled by a Remote Electronics Unit (REU) mounted on the actuator manifold. The REU sends commands to a hydraulic valve in the manifold which ports pressure to either the "extend" or "retract" side of the actuator until the desired actuator position is achieved. Each HA REU is powered by a single electrical source on one of the Essential DC buses. The midboard spoiler REUs each have two electrical sources from the Right Essential DC bus with separate wing-leading-edge and trailing-edge routings. The HA has two states - Hydraulically Active or Damped Bypass: Hydraulically Active - HA in normal operation. The aircraft hydraulic system is providing pressurized fluid to position the flight control surface. To be in this state, hydraulic pressure must be available and the actuator REU must electrically hold a solenoid. The REU can force Damped Bypass operation by removing power from the solenoid. Damped Bypass - The actuator is in a passive condition that resists movement proportional to the rate at which the surface is moved by either airflow over the surface or by the other actuator. Hydraulic fluid is trapped within the actuator to prevent control surface flutter in flight. During ground operation, the trapped fluid in the actuator provides gust protection. Pilots may directly experience the behavior of damped bypass actuators by rotating an aileron surface upward or downward with the palms of their hands during preflight when hydraulic pressure is not available to either the left or right system. EBHA: The EBHA has three modes of operation: Hydraulically Active Electric Backup Damped Bypass During normal hydraulic operation, the EBHA has the same inputs and functions as the HA. However, in the event that a surface cannot be powered by an engine-driven hydraulic source, an Electrical Backup (EB) hydraulic pump will provide the EBHA actuator with hydraulic pressure to operate the flight control surface. The EBHA REUs each receive electrical power from two separate circuit breakers. One is located on the UPS bus and the other on one of the Main DC buses except for the rudder which uses two UPS electrical supplies routed through separate rotor burst zones. The Motor Control Electronics (MCE) is used to control the EBHA electric motor-pump when the actuator is in the Electric Backup state due to hydraulic or REU failures. If an EBHA is in the normal hydraulically active state, the MCE is in a standby condition wherein it still processes commands and sends them back to the FCC through its REU to monitor its availability. Normally, the MCE receives commands from its associated REU. As a backup, each FCC has a single MCE bus for direct MCE control. In the event of failure of all REUs on a surface (except inboard and mid spoilers), the MCE bus allows an FCC to retain control of the surface. FCC 1 connects to both aileron MCEs and the right elevator MCE. FCC 2 connects to both outboard spoiler MCEs, the left elevator MCE and the rudder MCE. MCEs each receive electrical power from a single 65 Amp circuit breaker on the 28V DC EBHA bus located in the tail compartment at floor level forward of the ladder. The MCE uses the power for its internal electronics and to power the backup electrical motor. Operation: The primary FCS has five modes of operation: Normal Alternate Direct Backup Maintenance The FCCs will not allow mixed mode operation. All of the flight controls will be in Normal, Alternate, Direct, or Backup mode. Normal Mode: Under normal conditions, the Flight Control System (FCS) will be in Normal mode. Normal mode has four sub-modes: On-ground, Takeoff and Landing, Cruise, and Angle of Attack (AOA) Limiting. The FCS will be in Normal mode when the following conditions are satisfied: Two or more air data probes are heated and providing consistent data AND At least one IRU is operating and without conflicting data from another IRU or all operating AHRS (standby Attitude Heading Reference Systems) AND The horizontal stabilizer is not reporting to the FCC that it is in Backup Mode (wherein the stabilizer is directly controlled by the center console backup trim switch) On Ground sub-mode: The FCS will be in the On-ground sub-mode when Normal mode conditions are satisfied and at least one of the four conditions is satisfied: WOW from both the left and right main gear indicate weight on wheels OR WOW from the left main gear indicates weight on wheels and the wheel speed from the left wheel is 47 knots or greater OR WOW from the right main gear indicates weight on wheels and the wheel speed from the right wheel is 47 knots or greater OR Both the left and right main gear wheel speeds are 47 knots or greater and Radar altitude is less than 10 feet During a normal takeoff, this last condition will be the last one to cancel as the aircraft climbs through 10 feet AGL allowing transition from On-Ground sub-mode to the Takeoff and Landing sub-mode. During a normal landing, the normal sub-mode will not transition to On-ground until after touchdown. Indications of On-ground sub-mode: The pitch, roll and yaw trim indicators will include a green band when the FCS is in the On-Ground sub-mode While On-ground, AOA limiting is unavailable. Weight off wheels AOA limiting is available Pitch Control - Column deflection translates directly to elevator command. At speeds of approximately 135 KCAS and below, full column will result in full elevator except for a small flap bias described below. As speed increases above approximately 135 KCAS, the amount of elevator deflection per column deflection will decrease so that the pitch response will not be too sensitive. The elevator is biased one degree nose up for each "notch" of flap handle deflection - i.e. one degree for flaps 10, two degrees for flaps 20, and three degrees for flaps 39. Roll Control - Wheel deflection translates directly to aileron and roll spoiler deflection. The amount of roll spoiler deflection will vary with flap position and air speed, and could be as much as full deflection (55°). The outboard and midboard roll spoilers are commanded to move together. Higher speeds and higher flap deflections will tend to a smaller spoiler deflection schedule. Yaw Control - Below 60 KCAS, rudder deflection is only a function of rudder pedal deflection; full pedal will result in full rudder displacement. Between 60 KCAS and 90 KCAS the yaw damper will smoothly transition from 0% to 100% effectiveness. The yaw damper will then be active using simple yaw rate feedback (from the IRS sensors) to the rudder in combination with the rudder pedal input. Pitch Trim - Actuation of any pitch trim switch commands the stabilizer to move at its maximum rate of 0.4 degrees per second in the direction commanded. Takeoff and Landing sub-mode: The FCS will be in the Takeoff and Landing (TL) sub-mode when Normal mode conditions are satisfied and the following conditions are all satisfied: The On-ground logic is not satisfied AND AOA limiting logic is not satisfied AND Autopilot is not engaged AND Landing gear handle is down OR flap handle is greater than 5° Indications of Takeoff and Landing sub-mode: The pitch trim indicator will display the numeric value of stabilizer position without a green band indication Pitch Control - Identical to On-ground sub-mode and repeated here for convenience. Column deflection translates directly to elevator command. At speeds of approximately 135 KCAS and below, full column will result in full elevator except for a small flap bias described below. As speed increases above approximately 135 KCAS the amount of elevator deflection per column deflection will decrease so that the pitch response will not be too sensitive. The elevator is biased one degree nose up for each "notch" of flap handle deflection - i.e. 1 degree for flaps 10, 2 degrees for flaps 20, and 3 degrees for flaps 39. Roll Control - Wheel deflection will command aileron and roll-spoiler deflection. The roll augmentation function may increase or decrease the roll control surface deflections depending on the magnitude of roll rate being realized compared to the amount of roll-rate expected based on the wheel and pedal inputs and normal aircraft response. The roll augmentation function is a safety feature which significantly reduces any roll rate which is not commanded by the pilot (e.g. roll rate caused by wake turbulence or a spoiler panel malfunction). In the event of a spoiler panel malfunction, the roll augmentation feature will deactivate after five seconds and remain so for the duration of the flight. The amount of roll spoiler deflection will vary with flap position and air speed, and could be as much as full deflection (55°). The outboard and midboard roll spoilers are commanded to move together. Higher speeds and higher flap deflections will tend to a smaller spoiler deflection schedule. Yaw Control - The yaw damper will be active using sensor feedback (sideslip rate calculation) to the rudder in combination with the rudder pedal input in order to damp dutch roll oscillations. Pitch Trim - Actuation of any pitch trim switch commands the stabilizer to move at its maximum rate of 0.4 degrees per second in the direction commanded. Cruise sub-mode: The FCS will be in the Cruise sub-mode when Normal mode conditions are satisfied and the following conditions are all satisfied: The On-ground sub-mode logic is not satisfied AND AOA limiting sub-mode logic is not satisfied AND Autopilot engaged OR both the landing gear and flap handles are up Indications of Cruise sub-mode: When the pitch trim indicator does not display the numeric value of the stabilizer position (indicates NOT in On-ground or Takeoff and Landing sub-modes) and there is no (advisory) CAS message, the FCS is in the Cruise sub-mode Pitch Control - Pitch response is designed to be conventional, but with additional damping (especially at high altitude). Elevator movement is augmented with sensor feedback to provide a more consistent response to column input with variations in weight and center-of-gravity. Overspeed protection is provided by limiting column authority in the nose down direction if the aircraft exceeds the maximum operating speed. There is no protection against exceeding the maximum mach number. Roll Control - Identical to Takeoff and Landing sub-mode which is repeated here for convenience. Wheel deflection will command aileron and roll-spoiler deflection. The roll augmentation function may increase or decrease the roll control surface deflections depending on the magnitude of roll rate being realized compared to the amount of roll-rate expected based on the wheel and pedal inputs and normal aircraft response. The roll augmentation function is a safety feature which significantly reduces any roll rate which is not commanded by the pilot (e.g. roll rate caused by wake turbulence or a spoiler panel malfunction). In the event of a spoiler panel malfunction, the roll augmentation feature will deactivate after five seconds and remain so for the duration of the flight. The amount of roll spoiler deflection will vary with flap position and air speed, and could be as much as full deflection (55°). The outboard and midboard roll spoilers are commanded to move together. Higher speeds and higher flap deflections will tend to a smaller spoiler deflection schedule. Yaw Control - Identical to Takeoff and Landing sub-mode which is repeated here for convenience. The yaw damper will be active using sensor feedback (sideslip rate calculation) to the rudder in combination with the rudder pedal input in order to damp dutch roll oscillations. Pitch Trim - Pitch trim will initially offset the elevator. In the long term, any persistent elevator offset greater than one degree will be offloaded to the horizontal stabilizer. AOA Limiting: The FCS will be in the AOA limiting sub-mode when Normal mode conditions are satisfied and the following conditions are all satisfied: WOW from both main landing gear indicates weight off wheels AND AOA limiting is available as indicated by the absence of the Stall Protection Unavail (caution) CAS message AND Normalized AOA is greater than the AOA limiting sub-mode entry threshold. During normal maneuvering, this threshold is approximately 0.88-0.93 normalized AOA depending on airspeed deceleration rate. In cases of significant pitch rate and rapid/large aft column input, this entry threshold can be lowered substantially. Entry into the AOA limiting sub-mode can occur at any normalized angle of attack with abrupt nose up control input. Indications of AOA Limiting sub-mode: The FCC AOA Limiting (advisory) CAS message will be displayed. Additionally, the (caution) CAS message will be displayed if the AOA is near its limit value of 0.96. Pitch Control - Column commands angle of attack, with full aft column resulting in maximum allowed angle of attack (alpha limit of approximately 0.96 normalized AOA). Reducing angle of attack will exit AOA limiting and return typically to either Cruise or Takeoff-and-Landing sub-mode depending on landing gear and flap handle positions. On-ground conditions have priority and a transition to on-ground sub-mode will occur if the aircraft should touch down while in AOA limiting mode. To protect against unnecessary entry into AOA limiting during landing and flare, approach speed should be increased during strong or gusty wind conditions. In nominal wind conditions, cross the threshold at VREF. In strong wind conditions, cross the threshold at an increased approach speed of VREF plus % of the steady state wind plus the full gust increment, not to exceed an approach speed of VREF + 20 KCAS. Landing distance shall be calculated utilizing planned speed at the threshold. Roll Control - Identical to On-ground sub-mode which is repeated here for convenience. Wheel deflection translates directly to aileron and roll spoiler deflection. The amount of roll spoiler deflection will vary with flap position and airspeed, and could be as much as full deflection (55°). The outboard and midboard roll spoilers are commanded to move together. Higher speeds and higher flap deflections will tend to a smaller spoiler deflection schedule. Yaw Control - The yaw damper will be active using sensor feedback to the rudder in combination with the rudder pedal input in order to damp dutch roll oscillations. Pitch Trim - Nose down trim will command the stabilizer in a nose-down direction. Nose up trim has no effect. Alternate Mode: The FCS will be in Alternate mode when one or more of the following conditions are satisfied: Less than two air data probes are heated and providing consistent data OR IRU data is inconsistent, is conflicted by all operating AHRS, or is invalid (all off or malfunctioned) OR The horizontal stabilizer is reporting to the FCC that it is in Backup Mode Indications of Alternate mode: Alternate mode is primarily indicated by an FCC Alternate Mode (caution) CAS message. Since airspeed is not used in Alternate mode, the amount of surface deflection per pilot input (pitch, roll, and yaw) will be greatest when either the flap or gear handle is not up (assumes 250 KCAS); less when both handles are up (assumes 340 KCAS). Pitch Control - Column deflection translates directly to elevator command. Roll Control - Wheel deflection translates directly to aileron and spoiler deflection. The outboard and mid roll spoilers are commanded to move together. Yaw Control - Rudder pedal deflection commands rudder deflection. If IRU #3 data is available, a simplified yaw damper will damp dutch roll oscillations. Pitch Trim - Actuation of any pitch trim switch commands the stabilizer to move at its maximum rate of 0.4 degrees per second. Direct Mode: The FCS will be in Direct mode after all four FCC channels become invalid. The Direct mode behavior is the same as Alternate mode but some of the software and hardware is different and/or simplified to help guarantee its availability. Indications of Direct mode: Direct mode is primarily indicated by an (caution) CAS message. While in Direct mode, AOA limiting is unavailable as annunciated by the message. (caution) CAS The following applies to both Alternate and Direct mode and is repeated here for convenience: Since airspeed is not used in Alternate or Direct mode, the amount of surface deflection per pilot input (pitch, roll, and yaw) will be greatest when either the flap or gear handle is not up (assumes 250 KCAS); less when both handles are up (assumes 340 KCAS). Pitch Control - Column deflection translates directly to elevator command. Roll Control - Wheel deflection translates directly to aileron and spoiler deflection. The outboard and mid roll spoilers are commanded to move together. Yaw Control - Rudder pedal deflection commands rudder deflection. If IRU #3 data is available, a simplified yaw damper will damp dutch roll oscillations. Pitch Trim - Actuation of any pitch trim switch commands the stabilizer to move at its maximum rate of 0.4 degrees per second. Backup Mode: The FCS will be in Backup mode after all four FCC channels have failed (not capable of Normal, Alternate, or Direct Modes). Backup mode behavior is similar to Alternate mode except the only spoiler capability is outboard roll control spoilers. Speed brakes and ground spoilers are not available. Indications of Backup mode: Backup mode is primarily indicated by a BFCU Active (caution) CAS message. While in Backup mode, AOA limiting is unavailable as annunciated by the Stall Protection Unavail (caution) CAS message. Since airspeed is not used in Alternate, Direct, or Backup mode, the amount of surface deflection per pilot input (pitch, roll, and yaw) will be greatest when either the flap or gear handle is not up (assumes 250 KCAS); less when both handles are up (assumes 340 KCAS). Pitch Control - Column deflection translates directly to elevator command. Roll Control - Wheel deflection translates directly to aileron and outboard spoiler deflection. The mid roll spoilers will remain retracted while in Backup mode. Yaw Control - Rudder pedal deflection commands rudder deflection. If IRU #3 data is available, a simplified and reduced authority yaw damper will provide a small amount of additional dutch roll damping. Pitch Trim - The pedestal pitch trim switch commands the stabilizer to move at its Backup rate of 0.15 degrees per second in the direction commanded. The wheel-mounted trim switches will not have any effect. Maintenance Mode: The Maintenance mode is used to rig the FCC and REUs via the CMC. The FCC can only enter Maintenance mode if the aircraft is in the "on-ground stationary" condition (Both MLG WOW = true AND wheel speed is <47 knots). The Maintenance switch is located on the maintenance test panel above the observer Audio Control Panel (ACP). The Maintenance mode allows the following tasks to be accomplished: System Rigging Initiated Built-In Test (IBIT) Controls and Indications: Crew Alerting System (CAS) Messages: The following CAS messages are associated with the Primary Flight Control System (PFCS): CAS Message Possible Cause(s) FCC Alternate Mode (caution) FCCs are in Alternate mode. CAS Message Possible Cause(s) FCC Direct Mode (caution) FCCs are in Direct mode. BFCU Active (caution) BFCU is in Backup mode. Limitations: Environmental Conditions with Degraded Flight Control Law Mode: Flight into known icing conditions is prohibited when operating in a flight control law mode other than Normal (i.e., Alternate, Direct, or Backup). If the flight control law mode degrades from Normal while in icing conditions, exit icing conditions. Figure 1. Flight Controls System: Control Surface Nomenclature RIGHT AILERON RIGHT OUTBOARD SPOILER RIGHT MIDBOARD SPOILER RIGHT INBOARD SPOILER RIGHT ELEVATOR HORIZONTAL STABILIZER LEFT ELEVATOR LEFT INBOARD SPOILER LEFT MIDBOARD SPOILER LEFT OUTBOARD SPOILER LEFT AILERON RUDDER TIL-000208 Figure 2. Flight Control Power Sources TIL-005311 Figure 3. Flight Control Mode Comparison FLIGHT CONTROL MODE COMPARISON NORMAL ALTERNATE DIRECT BACKUP Gain Control deflection scheduled with airspeed Two fixed gains Two fixed gains Two fixed gains Yaw damper Normal Simple Simple Limited Pitch Trim Switches Normal (all three) Normal (all three) Normal (all three) Controlled with BACKUP PITCH switch only Speed brake and ground spoilers available √ √ Speed brakes only Autopilot, AOA limiting, high speed protection, turn coordination, dynamic rudder limiting, and maneuver load alleviation √ TIL-033850 Figure 4. LEER Flight Control Computer (FCC) Location FLIGHT CONTROL COMPUTERS TIL-006386 Figure 5. REER Flight Control Computer (FCC) Location FLIGHT CONTROL COMPUTERS TIL-006387 Figure 6. Backup Flight Control Unit Architecture PILOT INCEPTOR SIGNALS PILOT WHEEL RVDTS TO FCC 2 TO FCC 1 TO FCC 2 TO FCC 1 COPILOT WHEEL RVDTS COPILOT INCEPTOR SIGNALS BACKUP FLIGHT CONTROL UNIT ARCHITECTURE Pilot Column RVDTs Copilot Column RVDTs TO FCC 2 TO FCC 1 UPS BFCU TO FCC 1 TO FCC 2 BCFU BFCU R AIL A429 TX 28 VDC 2A BFCU R ELEV A429 TX BFCU R OB SPLR A429 TX BFCU RUDDER A429 TX BFCU L OB SPLR A429 TX BFCU L ELEV A429 TX BFCU L AIL A429 TX REU 4 EBHA REU 6 EBHA REU 5 EBHA REU 3 EBHA REU 14 EBHA REU 16 EBHA REU 13 EBHA TO FCC 1 TO FCC 2 Pilot Rudder Pedals Rudder Position RVDTs Copilot Rudder Pedals TIL-000643 Figure 7. Elevator Electrical Backup Hydraulic Actuator (EBHA) and Electro-Hydraulic Servo Actuator (EHSA) Components SURFACE LVDT DETAIL A TIL-007165 Figure 8. EBHA and EHSA Rudder Actuators and Components HA MANIFOLD REU FWD ACTUATOR EB MANIFOLD REU MEC TIL-006388 Figure 9. Aileron Hydraulic Components SEE DETAIL A EHSA ACTUATOR EHSA MANIFOLD EBHA ACTUATOR HA MANIFOLD EHSA REU EB MANIFOLD EBHA REU MCE DETAIL A TIL-013901 Figure 10. Inboard, Mid and Outboard Spoiler Hydraulic Components EHSA MANIFOLD MCE (EBHA) REU (EBHA) EHSA ACTUATOR EB MANIFOLD EBHA ACTUATOR HA MANIFOLD DETAIL A TIL-003010 2A-27-20: Aileron and Roll Spoiler Control System General Description: (See Figure 11. Aileron Control System Schematic.) Aircraft roll control is accomplished by movement of the ailerons and mid/outboard spoilers located on the wing. In addition to providing roll control, spoilers also provide speed brake and ground spoiler functions described in a separate section of this document. The left and right inboard aileron actuators are EBHAs and the left and right outboard aileron actuators are EHSAs. Each mid and inboard spoiler panel is controlled by one EHSA and each outboard spoiler panel by one EBHA. The inboard spoilers are not used for roll control. Based on the roll control wheel displacements produced either by the pilot, copilot or autopilot servos (when engaged), RVDT position sensors attached to the pilot and copilot wheels (via cables inside the column and under the floor) generate electrical commands to FCCs 1 and 2. The FCCs send digital commands via data bus to each REU. Roll control wheel forces are provided by two feel units connected to the pilot and copilot wheels. The roll feel-force is proportional to the displacement of the control wheel. Roll control wheel damping is provided by two dampers connected to the pilot and copilot roll control wheels. In the event of either an aileron or spoiler surface jam, only normal pilot action is required, since pilot and copilot control wheels both remain available to control the non-jammed surfaces. The Autopilot may be disengaged by depressing the AIP DISC button on either control wheel, selecting the AIP switch on the guidance panel to OFF, or by actuation of a pitch trim switch. Additionally, the autopilot will disconnect if overpowered by the flightcrew applying approximately 25 pounds of force input in pitch or 10 pounds of force in roll upon the control wheel. Roll trim, yaw trim, and rudder pedal displacement does not disconnect the autopilot. Roll Trim: The FCS has both a normal and a backup roll trim capability (determined by the position of the ROLL MOTOR CONTROL switch on the center console TRIM panel). A split roll trim switch to the left of the motor disconnect switch is used to activate roll trim during either normal or backup operation. During normal operation, the roll trim switch commands the roll trim motor connected to the pilot's side wheel under the floor. The motor moves both control wheels by offsetting the neutral position of the centering springs. The offset wheel provides a command to the FCCs as if the pilot moved the control wheel manually. Electrical power to the roll trim motor is provided through the Solid State Power Controller (SSPC). In the event of a runaway roll trim motor, holding the Autopilot Disconnect (A/P DISC) switch located on either control wheel removes power from the roll trim motor. Continue to hold the A/P DISC switch until the ROLL MOTOR CONTROL switch is depressed which removes the motor's electrical power and causes roll trim commands to be sent directly to the FCCs for backup roll trim. Autopilot Roll Servo Motors: An autopilot roll servo electric motor is mechanically connected to each wheel underneath the floor. The pilot side motor is controlled by Automatic Flight Control System (AFCS) 1 located in MAU 1 and the copilot side motor is controlled by AFCS 2 located in MAU 2. The servo motors are used to directly move the control wheels. The servos control the aircraft by moving the control wheels in the same manner as the flight crew in manual control. This method has the benefit of providing the flight crew visual feedback of what the AFCS system is doing to control the aircraft. Either servo drives the movement of both control wheels, due to the wheel mechanical interconnect. Aileron and Roll Spoiler Surface Activation: Each aileron surface is actuated by an inboard EBHA and an outboard HA. During normal operation, both actuators are hydraulically pressurized and actively moving the control surface. The outboard spoiler panels each have a single EBHA and the midboard panels each have a single HA. Aileron Droop during Cruise Flight: To improve cruise performance, the FCC commands the ailerons to symmetrically displace trailing-edge-down ("droop") at long-range cruise speeds to minimize cruise drag. The symmetric aileron droop is relatively small (a maximum of 0.6°) and has no effect on airplane roll control capability. The aileron droop command is proportional with speed: activating at 0.80M; fully applied at 0.84M; and faded out between 0.88M and 0.89M, or between 300 and 320 KCAS. Controls and Indications: A full description of the Flight Controls 2/3 synoptic page appears in section 2B-07-00. 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 HA LEER A-9 L ESS 28V DC Bus L AIL EBHA PRI REER A-3 UPS 28V DC Bus L AIL EBHA SEC LEER B-9 L MAIN 28V DC L AIL EBHA PWR EBHA PDB (Tail Cmpt) EBHA Bus 28V DC R AIL HA REER A-6 R ESS 28V DC Bus R AIL EBHA PRI REER A-4 UPS 28V DC Bus R AIL EBHA SEC REER C-7 R MAIN 28V DC Bus R AIL EBHA PWR EBHA PDB (Tail Cmpt) EBHA Bus 28V DC ROLL TRIM MCDU SSPC (#2702) R ESS 28V DC Crew Alerting System (CAS) Messages: The following CAS messages are associated with the roll flight controls system: CAS Message Possible Cause(s) L-R Aileron Fail (caution) NOTE: Not available in FCC Direct mode. Both REUs failed. OR Loss of aileron command to both ailerons. Aileron Single Actuator (caution) Single aileron actuator failed. Roll Authority Limit (caution) NOTE: Not available in FCC Direct mode. Roll axis flight control surfaces approaching maximum displacement. Roll Control Miscompare (caution) NOTE: Not available in FCC Direct mode. Pilot and copilot control wheel positions are different. Roll Trim Motor Off (advisory) ROLL MOTOR CONTROL switch has been selected OFF. Roll Trim LWD Limit (advisory) Roll trim at left wing down limit. Roll Trim RWD Limit (advisory) Roll trim at right wing down limit. Limitations: There are no limitations established for aileron and roll spoiler control system at the time of this writing. Figure 11. Aileron Control System Schematic TIL-000640 2A-27-30: Rudder Control System General Description: (See Figure 12. Rudder Control System Block Diagram.) The rudder control system provides electrical and hydraulic control of the rudder, as well as fault monitoring and annunciation. Aircraft yaw control is accomplished through a single rudder surface in the vertical tail. Two electrically-controlled hydraulic actuators provide power for rudder surface movement. Conventional pilot and copilot rudder pedals provide the control inputs through Rotary Variable Differential Transformer (RVDT) position sensors. The commanded position information is provided to the two Flight Control Computers (FCCs) and the Backup Flight Control Unit (BFCU), which then transmit rudder surface commands on digital data buses to the Remote Electronic Units (REUs) located at each actuator. The REUs control the hydraulic actuators which move the rudder surface. Description of Subsystems, Units and Components: Rudder Pedals: (See Figure 13. Cockpit Rudder Control Components.) Standard rudder pedals are used by the flight crew to control aircraft yaw. The pilot and copilot rudder pedals are part of the assembly that also includes the RVDT position sensors, applied force transducers, a feel and centering unit, and rudder pedal dampers. The pilot and copilot pedals are mechanically linked with no disconnect provided between the two pedal sets. The control module is located above the cockpit floor under the flight deck instrument panel. RVDT Postion Sensors: Rudder pedal position is measured by five separate RVDT position sensors that are contained within the same mechanical enclosure. This enclosure, called the RVDT cluster, contains a common shaft along the center line which simultaneously drives all five sensors. The sensors are located concentrically around the shaft. The pilot and copilot rudder pedals drive the same RVDT cluster through the mechanical linkage and gears. Two RVDTs provide rudder pedal position information to FCC 1, two for FCC 2, and one for the BFCU. The FCCs and the BFCU provide the excitation signal to their designated RVDTs and then receive voltage signals back that are proportional to the rudder pedal position. The FCCs or BFCU then command the rudder to move to the position established by the rudder pedals, yaw damper, and turn coordinator. Force Transducers: Flight crew applied rudder pedal force is measured by two separate force transducers. One transducer is located in the pilot rudder pedal linkage, and the other in the copilot linkage. Both transducers are single channel strain gauge type, with the pilot side connected to FCC 1 and the copilot side connected to FCC 2. The rudder pedal force information is used only by the Flight Data Recorder (FDR). The FCCs provide the information on a digital data bus to the Honeywell Modular Avionics Units (MAUs), which then transmit it to the FDR. The FCCs provide an excitation signal to the force transducers and then receive a voltage signal back that is proportional to the applied force on the rudder pedals. Feel Force Units: Rudder pedal feel forces are provided by two feel force units connected to the pilot and copilot rudder pedals. The rudder pedal force is proportional to the displacement of the rudder pedals. Full pedal displacement nominally requires approximately 65 to 80 pounds of pedal force. Pedal Dampers: Rudder pedal damping is provided by two dampers connected to the pilot and copilot pedal assemblies. Trim Control: (See Figure 14. Rudder Trim Controls.) Rudder Trim Switches: The rudder trim switch is located on the center console. Actuating both halves of the switch in a left or right direction from the center neutral position allows the rudder to trim left or right. The FCCs command the rudder surface to move in the direction requested by the trim switch signals. As long as both switches are held actuated, the rudder surface moves until the maximum allowed trim angle is achieved. When the switches are released, the rudder holds the commanded trim position. Trim authority is limited to % of the available rudder deflection. Trimming does not move the rudder pedals. Rudder Trim Auto Center Switch: A push button trim centering switch provides a simple and quick method for the flight crew to command rudder trim to neutral. Once pressed (holding the button is not necessary) the FCCs will neutralize the rudder trim at the standard rudder trim rate. Automatic Rudder Trim with Autopilot Engaged: The rudder auto-trim function objective is to minimize cruise drag by applying a rudder command in response to sensed lateral acceleration during wings-level auto-flight. With the autopilot engaged, the rudder is constantly and automatically trimmed to keep the slip/skid trapezoid centered, which may result in the rudder trim display not appearing centered. This is normal. The flight control system, through the Flight Control Computers (FCC), effectively produces zero sideslip during non-maneuvering flight in order to compensate for the small variations encountered during normal operation. The automatic rudder trim function has programmed limited authority, with the FCC only being capable of commanding ±1° rudder deflection. The rudder trim display depicts the FCC command to the rudder for zero sideslip. When the autopilot is disengaged, the automatically trimmed location of the rudder is held constant. Pressing the rudder trim AUTO CENTER button will reset the rudder to zero (neutralized). At cruise speeds near 0.90M with the autopilot engaged, an automatic rudder trim input may result in the autopilot and automatic rudder trim system inducing a continuous airplane yawing/rolling motion up to approximately X slip-skid trapezoid displacement and 5° bank angle, with a period of approximately 17 seconds. Disengaging the autopilot stops the yawing/rolling motion. Changing speed and/or altitude prior to re-engaging the autopilot will reduce the possibility of this condition recurring. If changing the speed and/or altitude is not viable, this condition can be eliminated with the autopilot engaged by manually adjusting rudder trim in one direction and using small increments until the automatic rudder trim authority is exceeded. Exceedance of the automatic rudder trim authority is indicated on the Rudder Trim display by the manually adjusted rudder location not automatically returning to the previous setting within a few seconds. The manual rudder trim capability is not affected when the automatic rudder trim authority is exceeded. The automatic rudder trim function can be reinstated with the autopilot engaged by pressing the rudder trim AUTO CENTER button. Improved Turn Coordination During Single-Engine Operations: In FCC software versions prior to V6.2 (ASC 037), turn coordination faded out as a function of total yaw input (rudder pedal displacement plus rudder displacement as a result of rudder trim). During single engine operation, when a large amount of yaw trim might be present, turn coordination was effectively disabled. With FCC V6.2 and subsequent installed, the rudder pedal input has priority over the turn coordination rudder command such that turn coordination is nullified with greater than approximately 50% rudder pedal displacement; however, the turn coordination command is not affected by rudder trim input. Rudder Control System Fault Monitoring and Annunciation: The FCCs receive status, health and validity information from the rudder actuation system The FCCs transmit the data received from the REUs to the Honeywell MAUs. In addition to actuation data, the FCCs also pass information on the rudder pedal position, rudder pedal applied force, trim switch position and other required data. Controls and Indications: A full description of the Flight Controls 2/3 and Hydraulics 2/3 synoptic pages appears in section 2B-07-00. Circuit Breakers (CBs): The following CBs protect the rudder flight controls: Circuit Breaker Name CB Panel Location Power Source RUD EBHA PRI REER B-1 UPS 28V DC RUD EBHA PWR EBHA PDB (Tail Cmpt) EBHA Bus 28V DC Circuit Breaker Name CB Panel Location Power Source RUD EBHA SEC REER B-2 UPS 28V DC RUD HA REER B-7 R ESS 28V DC Crew Alerting System (CAS) Messages: The following CAS messages are associated with the yaw flight controls system: CAS Message Possible Cause(s) Yaw Authority Limit (caution) Rudder approaching maximum displacement. Yaw Damper Fail (caution) Multiple IRU failure. Rudder Single Actuator (caution) Single rudder actuator failed. Rudder Fail (caution) Both REUs failed. OR Loss of rudder command. Pedal Steering Fail (caution) Rudder pedal steering failed. Pedal Steering Off (advisory) Rudder pedal steering is off. Yaw Trim Auto Center (advisory) Yaw trim automatic centering feature is active. Limitations: There are no limitations established for the rudder control system at the time of this writing. Figure 12. Rudder Control System Block Diagram TIL-000214 Figure 13. Cockpit Rudder Control Components FORCE SENSOR PUSHROD ASSEMBLIES RUDDER PEDALS VIEW LOOKING FWD DETAIL A TIL-002953 Figure 14. Rudder Trim Controls TIL-002955 2A-27-40: Elevator Control System General Description: (See Figure 15. Pitch Control System Block Diagram and Figure 16. Elevator Control System Diagram.) The Elevator Control System provides control of the left and right elevators, as well as fault monitoring and annunciation. Aircraft pitch control is accomplished by movement of the elevator surfaces located on the horizontal stabilizer, as well as by trim system movement of the entire horizontal stabilizer. Electrically-controlled hydraulic actuators provide power for elevator surface movement. Conventional pilot and copilot control columns move the Rotary Variable Differential Transformer (RVDT) column position sensors that provide the electrical control inputs. The commanded position information is read by the two Flight Control Computers (FCCs) and the Backup Flight Control Unit (BFCU), which then transmit elevator commands on digital data buses to the Remote Electronic Units (REUs) located at each actuator. There are two separate and independent hydraulic actuators at each elevator surface. The outboard actuator is a conventional Hydraulic Actuator (HA) and the inboard is an Electric Backup Hydraulic Actuator (EBHA). The REUs normally control the elevator surface actuators. BFCU commands are ignored by the REUs unless there has been a total failure of all four channels of the FCCs. Both control columns are mechanically connected under the flight deck floor. Control column movement is limited by mechanical stops. Since the columns are not mechanically connected to the elevators, artificial feel is provided by springs and dampers located under the flight deck floor. The springs provide the pilot with column forces proportional to column displacement. The dampers smooth the column motion and prevent overshoot. Column force sensors provide electrical input to the FCCs, avionics system, and Flight Data Recorder (FDR). The force sensor signal is used to disengage the autopilot system if either pilot provides greater than 25 lbs of force input to the column. In the event of a control column jam, the column bungee-interconnect-rod serves as a force-based override mechanism. In this case, the other control column can override the jammed control column to continue providing control inputs to the FCCs. The FCCs always command the elevator positions to the average of both the pilot and copilot RVDT inputs. Therefore, with one column jammed, elevator response to a single column displacement will be half the normal amount. The force required to override the spring in the override rod is approximately 44 lbs between the columns. The override mechanism in the rod never disconnects and returns to normal operation whenever the jam condition clears. Description of Subsystems, Units and Components: Elevator Data Bus Control: (See Figure 16. Elevator Control System Diagram.) Each FCC transmits digital elevator commands on two dedicated busses and one split bus to the four elevator REUs. Three buses are required to provide for separate bus routings in each of the three rotor burst zones. Signals are separated in the three zones in such a way that any two zones can be cut and elevator control can still be maintained. Under normal operation, each REU averages the two FCC commands. The BFCU provides a separate bus for each elevator EBHA. The HAs do not receive any input from the BFCU. Flight Deck Control Columns: (See Figure 17. Longitudinal Control Components (1 of 2).) Elevator control is provided through control columns installed at the pilot and copilot stations. Each column is connected by a pushrod to an assembly located under the flight deck floor. These assemblies include feel and centering units, an override mechanism, and other components described below. The pilot and copilot control columns are mechanically connected and during normal operation always move together. Either pilot has the ability to override the other by exceeding the break away force and maintaining that force. This allows control to be maintained if one of the columns becomes jammed. Pilots can familiarize themselves with the override feature while parked on the ground as follows (test will not be affected by the state of electric, hydraulic, engine, APU etc.): One pilot holds their column fixed in any desired position to simulate a jam while the other pilot applies force as necessary to override the interconnect and move their own column full travel. Pilots may then switch roles and repeat. No reset is required and the system returns to normal operation on its own when the opposing forces are released. Other components located under the flight deck floor include: RVDT Position Sensors Applied Force Transducers Stick Shaker Motors Autopilot Pitch Servo Motors RVDT Position Sensors: Each control column position is measured by five separate RVDT sensors that are contained within the same mechanical enclosure. This RVDT "cluster" contains a common shaft along the center line which simultaneously drives all five sensors. A centering spring positions the RVDT cluster in the neutral position in the unlikely event that the common shaft fails. The sensors are located concentrically around the shaft. The control columns drive their respective RVDT clusters through direct mechanical linkages. On each RVDT cluster, two RVDTs provide column position information to FCC 1, two for FCC 2, and one for the BFCU. The FCCs and the BFCU provide the excitation signal to their designated RVDTs, and then receive 2 voltage signals back that are related to the column positions in a way that has a built-in integrity check (against shorted or broken wires and other failures). The FCCs or BFCU then command the elevators to move to the position established by the average of the two control columns. Under normal operation, each elevator actuator command results from voted and averaged signals originating from eight column position sensors. Force Transducers: The applied force on each control column is measured by its own force transducer. The transducers are located at the ends of their associated pushrods at the column attachment points. Each transducer is a two channel strain gauge type, with one channel connected to FCC 1, and the other channel connected to FCC 2. The FCCs provide an excitation signal to the transducers, and then receive a voltage signal back that is proportional to the applied force. Control column force information is used by the Autopilot System and the FDR. The FCCs provide the information to the Honeywell Modular Avionics Units (MAUs), where the autopilot software resides. The autopilot uses the force information to automatically disconnect if excessive force is applied opposing the servo. The MAUs transmit the force information on a digital data bus to the FDR. Stick Shaker Motors: A stick shaker motor is mounted to each control column. The motors are located at the lower end of each column below the floor mounts and are used to shake the control column during an approaching stall, before the maximum Angle-of-Attack (AOA) is exceeded. The stick shaker does not normally activate when the stall protection system (AOA limiter) is active. When the stall protection system is unavailable the (caution) CAS message is posted and the stick shaker threshold is set to a lower AOA value to give the pilots earlier warning of impending stall. The FCCs control activation of the shakers directly by sending 28V DC power to the shaker motors. The column shakers may be tested on the ground by selecting the test page of the Standby Multifunction Controller (SMC) and pressing and holding the stall warning button. The pilot's column shaker will activate first for 3 seconds and after a short delay the copilot's column shaker will activate for 3 seconds. Autopilot Pitch Servos: Autopilot pitch servos are mounted on the artificial feel units for both columns (one on each side of the col