FLIGHT CONTROLS -

FLIGHT CONTROLS -
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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