FLIGHT CONTROLS GENERAL Pitch Control Mechanically actuated Elevator via cable and pulleys Mechanically actuated Elevator Trim Tabs Electrically actuated Movable Horizontal Stabilizer Roll Control Hydraulically actuated Ailerons with mechanical backup via cable and pulleys Hydraulic neutral point adjustment Yaw Control Hydraulically actuated with mechanical backup via cable and pulleys Hydraulic neutral point adjustment Lift Devices Electrically actuated Fowler Flaps Spoilers Electrically controlled and hydraulically actuated Speed Brakes and Ground Spoilers ELEVATORS Each control yoke operates a set of cables, pulleys and bell cranks so as to operate its respective Elevator. The systems are linked together with a torque tube and a disconnect device. The disconnect device is used if there is a jam in the system. When the Elevators are disconnected: An ELEV DISC Light illuminates on the center pedestal. Stick Pusher is available only on the LEFT side. If disconnected do NOT use the Autopilot. The system can be reconnected only on the ground by maintenance. The Autopilot and Stick Pusher use the LEFT Elevator system ELEVATOR TRIM TABS Two types of Elevator Trim Tabs automatically function throughout the airspeed envelope to alleviate air loads required to control aircraft pitch. They are not controllable from the cockpit. Inboard Spring Tabs Activated by a torsion rod, bellcrank, torque tube assembly. Deflection is proportional to control column force and aerodynamic loads. Deflection is opposite that of Elevator travel. Used to assist smaller Elevator travel at high airspeed and high air loads. Remains neutral at low speeds. Outboard Servo Tabs Deflection is proportional to Elevator travel. Deflection is opposite that of Elevator travel. Used to assist larger Elevator travel at low airspeed and low air loads. PITCH TRIM The Pitch Trim System uses electric motors and a jackscrew to move the leading edge of the stabilizer and change its angle of incidence. The position of the Horizontal Stabilizer can be confirmed visually with markings on the Vertical Stabilizer. The thick markings on the tail indicate 4° Nose DOWN, NEUTRAL, and 10° Nose UP. The thin markings indicate 2° intervals. The system is commanded electrically by Trim Switches located on the control yokes and center pedestal. The system can also be commanded automatically by the Autopilot. In both cases the pitch trim signal is processed in the Horizontal Stabilizer Control Unit (HSCU). The HSCU commands the Horizontal Stabilizer Actuator (HSA). The HSA uses two electric motors to move the Horizontal Stabilizer. One motor operates as Main Trim and the other as the Backup Trim. Each system sends commands to its respective motor. When one Motor is commanded it will simultaneously drive the other motor. The non-commanded Motor will run deenergized. Both control systems have a cutout relay. The MAIN SYS CUTOUT button deactivates the Main system, and the BACKUP SYS CUTOUT button deactivates the Backup system. Trim inputs can be made using: Trim switches on the control yokes (Main System) Auto Trim by Autopilot or Speed brakes (Main System) Backup Switches on the center pedestal (Backup System) The trim rate is based on airspeed received from the Air Data Computers (ADC). Below 160 KCAS the trim rate is 2° per second Between 160 KCAS and 250 KCAS the trim rate decreases as airspeed increases Above 250 KCAS the trim rate is 1°per second. During Stick Shaker activation, nose-up trim actuation is INHIBITED. PITCH TRIM PROTECTIONS A “TAKEOFF – TRIM” warning is made if the stabilizer trim position is NOT in the GREEN when the TO CONFIG button is pressed or the Thrust Levers are advanced prior to takeoff. Both halves of a trim switch must be operated simultaneously to function. This prevents a pitch trim runaway due to a faulty switch. A trim switch will be inhibited if operated in a split condition for more than 7 seconds. The aircraft must be powered down and repowered to regain control of the switch. Excessive air loads could potentially stall the electric motor if the aircraft is allowed to accelerate in an untrimmed condition after takeoff. This would result in a temporary loss of pitch trim command. In the case of a stalled electric motor, and a cumulative unsatisfied pitch trim command of 16 seconds will result in the associated trim system to be de-activated. To minimize the effects of an inadvertent trim command, each trim switch is only effective for 3 seconds at a time. The pilot must release the switches every 3 seconds and then use them again. Should a pitch trim runaway ever occur, the Quick Disconnect button will override any trim activation. This button may be PRESSED and HELD until a full disengagement has been accomplished through the Main Cutout Button. AILERONS The ailerons are positioned by the control wheels, which are linked together by a torque tube and cables to supply mechanical inputs to two separate hydraulic actuators. Each aileron actuator is supplied by both hydraulic systems. Either hydraulic system is capable of providing full power control. Each hydraulic system supply can be shut off if necessary via the overhead panel Aileron Shutoff Buttons. If both hydraulic systems are lost, rotating the control yoke mechanically positions the ailerons. The Artificial Feel System (AFS) provides tactile feedback on the aerodynamic loading of the Ailerons. The Artificial Feel System is installed in the RIGHT Aileron system. The systems are linked together with a torque tube and a disconnect device. The disconnect device is used if there is a jam in the system. When disconnected: An AIL DISC Light illuminates on the center pedestal. If disconnected do NOT use the Autopilot. The system can be reconnected only on the ground by maintenance. The Autopilot is installed in the LEFT Aileron system. ROLL TRIM SYSTEM An electro-mechanical Roll Trim Actuator is installed on the RIGHT side of the Torque Tube which is linked to the AFS. Roll Trim is accomplished by repositioning of the neutral point of the ailerons. The ROLL TRIM switch on the Trim Control Panel will energize the actuator in either direction. ROLL TRIM PROTECTIONS Both halves of a trim switch must be operated simultaneously to function. This prevents a pitch trim runaway due to a faulty switch. A trim switch will be inhibited if operated in a split condition for more than 7 seconds. The aircraft must be powered down and repowered to regain control of the switch. To minimize the effects of an inadvertent trim command, each trim switch is only effective for 3 seconds at a time. The pilot must release the switches every 3 seconds and then use them again. The Quick Disconnect button will override any trim activation. RUDDERS Yaw control is provided through two Rudders (Forward and Aft) which are mechanically linked together. The Rudder Pedals are linked together with a torque tube. However, each set of Rudder Pedals operates its own set of cables, pulleys and bell cranks which provide mechanical inputs to the Rudder Power Control Unit (PCU). The PCU hydraulically powers the Forward Rudder. There is an Artificial Feel System within the PCU which gives pilots aerodynamic loads feedback. RUDDER SYSTEM 1 AND 2 The PCU uses both hydraulic systems to provide pressure to two separate Rudder Actuators which divide Rudder control into Rudder System 1 and Rudder System 2. Normally both rudder systems are powered at speeds less than 135 knots. Above 135 knots, Rudder system 1 automatically shuts OFF. If the automatic shut off fails to shut off a system above 135 knots, a RUDDER OVERBOOST caution message is presented. It is then necessary to manually shut off Rudder System 1 or 2. RUDDER SYSTEM SHUTOFF Each Rudder system can be shut off manually by the associated RUDDER SHUTOFF button on the FLIGHT CONTROL Panel. Manual reversion is provided by way of the conventional cables, pulleys and bell cranks. Greater control forces are to be expected since the pilot will deal directly with aerodynamic loads on the Rudder. The Artificial Feel System is not available in the mechanical reversion mode. If Rudder System 2 fails, the Rudder System 1 automatically takes over Rudder control. A caution message is presented until Rudder System 2 normal operation is re-established. The Autopilot is installed in the LEFT Rudder Pedal cable system YAW TRIM Yaw trim is accomplished by the PCU in repositioning of the neutral point. Pressing the ROLL TRIM switch on the Trim Control Panel will adjust the PCU in either direction. Yaw Trim is not available in the manual reversion mode. RUDDER HARDOVER PROTECTION The Rudder Hardover Protection system monitors Rudder Pedal forces applied by the pilots. Each Rudder Pedal is equipped with a 130 lb Spring Loaded Cartridge and a microswitch. If the system recognizes 130 Lbs. of Rudder Pedal force applied by the pilots, and at least a 5° Rudder deflection CONTRARY to the direction commanded by the Rudder Pedals, the Rudder Hardover Protection system disables both sides of the PCU and generates an EICAS Message. The pilots must then shut off each Rudder System manually. After a proper automatic shutoff during a Hardover occurrence, do not reset rudder systems. Rudder Hardover Protection is INHIBITED during single engine operations. FLAP SYSTEM Each wing has two double-slotted Fowler-type panels. The pilot chooses flap position with the Flap Selector Lever (FSL). There are four detent positions to set the Flaps at: 0, 9, 22, and 45. The Flaps 18 detent is blocked. Flap position, velocity and command signals are processed in the Flap Electronic Control Unit (FECU). The FECU compares FSL position and Flap positions and commands the FPDU to move the Flaps. Two electric motors are housed in the Flap Power Drive Unit (FPDU). The motors drive flexible drive shafts which actuate Ball Screw Actuators. If one electric motor fails, the other motor can operate the flaps at half speed. EICAS will present a FLAP LOW SPEED caution message. If both motors or control channels fail, an EICAS FLAP FAIL caution message is presented to indicate that the system is inoperative. FLAP SYSTEM PROTECTIONS If there is any disagreement on flap position or velocity between both systems, the FECU disables the system and informs EICAS and other related Systems. If the FSL Switch fails, the position(s) can NOT be selected. A protection circuit INHIBITS flap movement if the flap position and FSL disagree upon initial aircraft power-up. The FSL must lifted UP to release the protection. Flap actuators are torque limited to safeguard the structure against excessive loading should the flaps or the actuators jam. Velocity sensors installed at the end of the flexible shafts detect panel asymmetry. The Flap System is disabled if the panels are asymmetric. Flap Transmission Brakes stop flap movement, locking them in place. It can only be reset on the ground by maintenance. If the flaps are not in the appropriate takeoff position (9° or 22° for the EMB-145, and 9° or 18° for the EMB-135 and 140) when the Takeoff Configuration Check Button is pressed or when either thrust lever angle is advanced for takeoff, the EICAS flap digits and box will turn red, and an aural warning TAKEOFF FLAPS and an EICAS NO TAKEOFF CONFIG warning message will be displayed in the cockpit. Two switches on the Flap Selector Lever send signals to the Landing Gear Warning System to alert pilots any time the aircraft is in a landing configuration and the landing gear are not down and locked. Flap position is displayed on Engine EICAS. Marks on the wing also indicate the 9 and 22 positions. GUST LOCK A Gust Lock system locks the elevator to avoid wind damage on the ground. It is either a mechanical lock which mechanically inhibits control column movement or an electro-mechanical lock which activates solenoid locking pins which engage the Elevator. The electro-mechanical type is identified by the yellow and black striped paint and the ELEC GUST LOCK inscription. Ailerons and Rudders do not require a lock as hydraulic pressure dampens any control surface movement. The GUST LOCK light on the glareshield illuminates during the unlocking cycle on the ground. It also indicates a system failure when in flight. LOCKING Pull the control column backwards to any position from neutral to full nose up. Lift the safety lock device Move the gust lock lever from the unlocked FREE to the LOCKED position Push the control column fully forward until the control column movement is restricted. UNLOCKING Lift the safety lock device and move the gust lock lever to its intermediate position. The locking pins are commanded to open when the gust lock lever is in the intermediate position. The elevators will be unlocked after approximately eight seconds. The indication lights will illuminate during the unlocking cycle, and then should remain off after unlocking. Lift the safety lock device Pull the gust lock lever from the intermediate position to its full forward FREE position GROUND SPOILERS AND SPEEDBRAKES Each wing is equipped with two spoiler panels, inboard and outboard. Speedbrake function uses outboard spoilers only. Ground Spoiler function uses both inboard and outboard spoilers. The Speedbrakes and Ground Spoilers are used as separate systems. The EICAS displays spoiler position as OPN or CLSD. Spoilers are controlled by a common Spoiler Control Unit (SCU). The SCU processes spoiler position, air/ground logic, wheel speed, Speedbrake Lever position, and Thrust Lever Angle signals, and commands hydraulic deployment and retraction of the panels. Hydraulic System 1 operates through the Ground Spoiler Valve. Hydraulic System 2 operates through the Spoiler/Speedbrake Valve. GROUND SPOILERS SYSTEM The SCU automatically commands DEPLOYMENT when the following conditions are met: Aircraft on the ground Main Wheels > 25 knots N2 < 56.4% Thrust Lever Angle (TLA) of both Engines are less than 30° SPEEDBRAKE SYSTEM The SCU commands DEPLOYMENT when the following conditions are met: Flap position 0° or 9° Speedbrake Lever selected OPEN Thrust Lever Angle (TLA) of both engines less than 50° If any of the conditions are NOT met the deployment does NOT occur and EICAS SPBK LVR DISAGREE Message is displayed. If the speed brake panels are open and either of the speed brake open conditions are not met, the speed brake panels automatically close and an EICAS SPBK LVR DISAGREE is presented. The speed brake lever must then be moved to the CLOSE position to remove the EICAS message in both cases. Spoiler Aural Warning (Takeoff Spoilers) If any spoiler/speed brake panel is deployed when the TO CONFIG Check Button is pressed or when either thrust lever is advanced for takeoff, the EICAS spoiler OPN label will turn red, and an aural warning TAKEOFF SPOILERS and an EICAS NO TAKEOFF CONFIG warning message will be displayed in the cockpit. PIT TRIM 1 (2) INOP: Respective Pitch Trim System is inoperative: 1 -- is the Main Trim System, 2 -- is the Backup System AIL SYS 1 (2) INOP: Respective Aileron System Actuator Hydraulic Power is inoperative. RUDDER SYS 1 INOP: Rudder System 1 is inoperative. Message is presented if: 1. Less than 135 kts 2. Greater than 135 kts if both ADC's airspeed information is invalid. RUDDER SYS 2 INOP: Rudder System 2 is inoperative. RUDDER SYS 1-2 INOP: BOTH Rudder Systems are inoperative. RUDDER OVERBOOST: BOTH Rudder Systems Hydraulic Actuators are pressurized above 135 kts. RUDDER HDOV PROT FAIL: 1. Disagreement between both FADEC of the same Engine. 2. Rudder Position Micro-switches indicate the Rudder is both LEFT and RIGHT at the same time. FLAP FAIL: BOTH Flap Channels are inoperative. SPOILER FAIL: Any Spoiler panel is: 1. Inadvertently OPEN, 2. FAILED to OPEN 3. Any failure in the input signals. SPBK LVR DISAGREE: Speedbrake Lever is at OPEN but the deployment logic is NOT satisfied. FLAP LOW SPEED: One Flap Channel is inoperative.
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