Legacy 500 - Flight Controls

Front 1

Introduction FLT CONTROLS

Back 1

The Flight Control System (FCS) is a functionally distributed Fly-By- Wire (FBW) control system

The primary flight control system consists of:

- Ailerons and multi-function spoilers for lateral (roll axis) control;

- Elevators for longitudinal (pitch axis) control;

- Rudder for directional (yaw axis) control.

The secondary flight control system consists of:

- Horizontal stabilizer;

—Flaps;

- Multi-function spoilers (when used as speed brakes or ground spoilers).

The secondary flight control

system includes the Pitch Trim Actuation System (PTAS) and Flap Actuation System (FAS).

Front 2

Introduction FLT CONTROLS

Back 2

The Flight Control System (FCS) is a functionally distributed Fly-By- Wire (FBW) control system

The primary flight control system consists of:

- Ailerons and multi-function spoilers for lateral (roll axis) control;

- Elevators for longitudinal (pitch axis) control;

- Rudder for directional (yaw axis) control.

The secondary flight control system consists of:

- Horizontal stabilizer;

—Flaps;

- Multi-function spoilers (when used as speed brakes or ground spoilers).

The secondary flight control

system includes the Pitch Trim Actuation System (PTAS) and Flap Actuation System (FAS).

Front 3

SIDE STICK CONTROLLER

Back 3

AP QUICK DISCONNECT/PRIORITY SELECT (AP/PTY) BUTTON

– When the priority button is pressed, the autopilot is automatically disengaged if active. The priority annunciation is activated only after 1 second of having the priority button pressed, in order to differ from the AUTOPILOT quick disengagement action.

THERE IS A BREAKPUT FORCE THAT THE PILOT CANA COMMAND AND IT WILL DISCONNECT THE AUTOPILOT

- This button gives exclusive control to the sidestick while pressed, inhibiting the inputs from the cross-side sidestick to be transmitted to the controls. Depending on the time the button is held down, two results are possible as follows:

PRIORITY BUTTON:

If the priority button is held down for more than 20 seconds, the respective sidestick will have exclusive control

TO RELEASE PRIORITY,: MAINTAINS PRIORITY until the cross-side sidestick priority button is pressed or the priority button of the sidestick in exclusive control is released and pressed again;

If the priority button is held down for less than 20 seconds and released, the sidestick authority reverts to the algebraic sum of the two stick inputs.

- If dual sidestick input occurs, the output is the algebraic sum of both pilots’ sidestick inputs. The FCC limits this sum to the maximum command of a single sidestick.

– If dual sidestick input occurs and priority has not been requested by any pilot, a tactile warning in both sidesticks is activated, as well as a dual input aural warning and a visual announcement on PFD are triggered.
PRIORITY BUTTON FAILS

- If the priority button fails, the respective sidestick will not be able to request exclusive control;
If the priority button fails so that the pilot cannot press it, the respective sidestick will not be able to request exclusive control as well;
- If the priority button fails jammed pushed in, the priority button of the cross-side sidestick can be used to remove the exclusive control from the sidestick with the failed priority button in the same way as when the priority is latched.
- In either case, the sidestick commands will be computed normally when the cross-side sidestick is not in exclusive control and the sidestick itself is not failed.

NOTE: The sidestick controller NM priority logic described above is not available in DM.

Front 4

SIDE STICK CONTROLLER: TOUCH CONTROL STEERING (TCS) BUTTON

Back 4

– While pressed, with the AP engaged, it momentarily allows the pilots to manually control the airplane without AP disengagement or AP mode change. After button release, the AP returns to the previous engaged mode.

- During manual flight with flaps 3 or FULL and gear down, the pilots are capable of resetting the TCS trim speed bug, in order to re-trim the horizontal stabilizer in relation to the airplane speed, when the button is released.

- The TCS trim synchronizes the airplane current airspeed when the TCS button is pressed indicating the approach TCS trim control law (flaps 3 or FULL and with landing gear down) activation. The control law activation is indicated through a TCS TRIM green flag displayed beneath the speed tape (refer to AOM Section 9-09-30 AIRSPEED INDICATION).

Front 5

FLAP SELECTOR LEVER

Back 5

Lever Position/Flap Position

0/0 position

1/7 deg

2/ 21 deg

3/21 deg

Full/37 deg

– Intermediate positions are not enabled. If the lever is left at an intermediate position, flaps remain in the last selected position.

Front 6

SPEED BRAKE LEVER

Back 6

All multi-function spoilers’ panels are proportionally deployed, in a continuous response to the speed brake lever position. If the lever is left at an intermediate position, the speed brake panels will assume an angle between 0° (stow) and 30° (full).

Front 7

FLIGHT CONTROLS SYSTEM/TRIM PANELS

Back 7

PITCH TRIM SWITCH: For trim purposes there is a dual switch for pitch axis trim adjustment. Both switches must be simultaneously actuated in the same direction to provide command to the PTS.

– The PTS can be manually commanded only on ground or in DM.

Yaw Trim Knob: normal ops

Front 8

FLTCTRL NORMAL MODE BUTTON (GUARDED)

Back 8

A FLTCTRL NORMAL MODE button installed on center pedestal panel to recover NM or to revert to DM in case of NM misbehavior. Manual reversion to DM is inhibited while the following FCS envelope protections are engaged:

Low speed protection;
Angle Of Attack (AOA) limiter protection; High speed protection.

Front 9

FLIGHT CONTROLS SYNOPTIC PAGE

Back 9

The flight controls synoptic page on the Multi Function Window (MFW) provides flight crew with the control surfaces position, actuators status, surface awareness and electrical and hydraulic Power-up Built-in Test (PBIT) expire time. This synoptic page is selected through the Cursor Control Panel (CCP) which comprises the trackball.

The pictorial rear view of the airplane enables the crew to easily recognize the surface position with respect to the neutral and maximum positions.

A rectangle represents the surface position range and will be filled up proportionally to the deflection of the surface.

Each control surface has a surface range scale. When the surface is fully deployed, the control surface range scale and the filled rectangle will be at the same level.

A red X indication in the surface position area indicate the surface data is invalid.

Also, the system layout on the display provides a mean of easily identifying the complete system status, which allows the flight crew to quickly recognize system condition.

The surface positions displayed on the flight controls synoptic on MFW is available in Normal Mode (NM) and Direct Mode (DM).

Front 10

PITCH/YAW TRIM INDICATION

Back 10

– GREEN: Position is valid and available.

– RED (inverse video): Before takeoff, pitch trim position is outside of the green band.

PITCH TRIM GREEN BAND
The pitch trim green takeoff band extends from 26 to 68 corresponding to the allowable pitch trim position for takeoff.

Front 11

FLIGHT CONTROLS INDICATIONS ON PFD

Back 11

The following flight controls indications are displayed on PFD:

– Sidestick priority;
– Sideslip target;
– Low Speed Awareness (LSA);
– Pitch Limit Indicator (PLI);
– Touch Control Steering (TCS) trim.

Front 12

FLIGHT CONTROLS INDICATIONS ON PFD :

SIDESTICK PRIORITY

Back 12

The Sidestick priority indication indicates which sidestick is on command and warns the crew of dual sidestick input, combined with an aural and tactile warning.
SCENARIO:

BOTH SIDESTICKS COMMANDED AT THE SAME TIME (NO PRIORITY REQUEST)

WARNINGS Indicated on PFD 1 and PFD 2 DISPLAY , AURAL WARNING: DUAL INPUT TACTILE WARNING: ON

PRIORITY REQUESTED BY THE PILOT (Captain)

- STICK PTY STICK PTY shows on PFD 1 and 2 and has an arrow pointing to the Captain Position

- AURAL WARNING: PRIORITY LEFT,

TACTILE WARNING: OFF

PRIORITY REQUESTED BY THE COPILOT

STICK PTY STICK PTY indicated on PFD 1 and 2 with an arrow pointing to the copilot

AURAL WARNING: PRIORITY RIGHT,

TACTILE WARNING: OFF

NORMAL OPERATION (NO PRIORITY REQUEST)

AURAL WARNING: − TACTILE WARNING: OFF

Front 13

FLIGHT CONTROLS INDICATIONS ON PFD:

SIDESLIP TARGET

Back 13

The sideslip target is a required yaw command, presented by a blue slip-skid indicator, in a One Engine Inoperative (OEI) condition during the takeoff, climb and go-around flight phase.

Front 14

FLIGHT CONTROLS INDICATIONS ON PFD: LOW SPEED AWARENESS AND PITCH LIMIT INDICATOR

Back 14

The LSA, adjacent to the speed tape, and the PLI, always parallel to the attitude reference airplane symbol on the Attitude Direction Indicator (ADI), alerts the flight crew of a low speed condition. In Normal Mode (NM), the AOA limiter function is operational and the LSA indicates:

- VSS: Represented by the top of the white tape and indicates the low speed protection function and should be close to 1.13 VS1G. In landing flap configurations, it indicates the minimum airspeed the aircraft can be trimmed at. In the other flap. configurations, it indicates the addition of speed stability;

– VAOA: Represented by the top of the amber tape and indicates the AOA limiter function activation (1.08 VS for 1 g conditions). This is the AOA limiter protection;

– VLIM: Represented by the top of the red tape and indicates the speed associated with the maximum allowable AOA airspeed (1.03 VS for 1 g conditions).

Since VSS is defined in terms of speed and VAOA/VLIM are defined in terms of angle of attack, those indications have different behavior in curves and other maneuvers. The white tape is not affected by load factor variation, so that it does not change when the aircraft is maneuvered into or out of a curve. On the other hand, VAOA/VLIM will increase if the aircraft is maneuvered into a curve and the yellow tape may cover the white tape in this condition.

- In Direct Mode (DM), the AOA limiter function is not operational and the LSA is modified to maintain correspondence with the stall aural alert. There are two airspeed indication: Where:

– VSA: Represented by the top of the red tape and indicates the airspeed at which the stall aural alert message is triggered (1.10 VS);

– VMS: Represented by the top of the amber tape and indicates an airspeed with appropriate margin to VSA (1.15 VS).

The LSA, adjacent to the speed tape, and the PLI, always parallel to the attitude reference airplane symbol on the Attitude Direction Indicator (ADI), alerts the flight crew of a low speed condition. In Normal Mode (NM), the AOA limiter function is operational and the LSA indicates:

- VSS: Represented by the top of the white tape and indicates the low speed protection function and should be close to 1.13 VS1G. In landing flap configurations, it indicates the minimum airspeed the aircraft can be trimmed at. In the other flap. configurations, it indicates the addition of speed stability;

– VAOA: Represented by the top of the amber tape and indicates the AOA limiter function activation (1.08 VS for 1 g conditions). This is the AOA limiter protection;

– VLIM: Represented by the top of the red tape and indicates the speed associated with the maximum allowable AOA airspeed (1.03 VS for 1 g conditions).

Since VSS is defined in terms of speed and VAOA/VLIM are defined in terms of angle of attack, those indications have different behavior in curves and other maneuvers. The white tape is not affected by load factor variation, so that it does not change when the aircraft is maneuvered into or out of a curve. On the other hand, VAOA/VLIM will increase if the aircraft is maneuvered into a curve and the yellow tape may cover the white tape in this condition.

- In Direct Mode (DM), the AOA limiter function is not operational and the LSA is modified to maintain correspondence with the stall aural alert. There are two airspeed indication: Where:

– VSA: Represented by the top of the red tape and indicates the airspeed at which the stall aural alert message is triggered (1.10 VS);

– VMS: Represented by the top of the amber tape and indicates an airspeed with appropriate margin to VSA (1.15 VS)

Front 15

FLIGHT CONTROLS INDICATIONS ON PFD: PITCH LIMIT INDICATOR

Back 15

- The PLI, displayed on the pitch angle scale, indicates the margin to VAOA and VLIM when in NM.
- At high margins the PLI is not displayed. However, as the margin reduces, the PLI color varies from amber to red.
- If the airspeed drops below VAOA (1.08 VS), the PLI becomes amber. Else if the airspeed drops below VLIM (1.03 VS), the PLI becomes red.
- The PLI is not displayed in DM.

Front 16

FLIGHT CONTROLS AURAL WARNINGS

Back 16

AURAL WARNING MESSAGE

DUAL INPUT: Both sidesticks commanded at the same time.

PRIORITY LEFT: Priority requested by the left seat pilot.

PRIORITY RIGHT: Priority requested by the right seat pilot.

Front 17

Flight Controls: memory Items

Back 17

Flight Control switch( off)

Front 18

FLY-BY-WIRE (FBW)

Back 18

FBW is an electronic system designed to operate the flight controls replacing the control cables of a conventional airplane.

The Legacy 500 FBW system provides closed-loop control and monitoring of all primary and secondary flight control surfaces within the system.

In a closed-loop control system, a sensor monitors the system output (e.g. the actuator’s position) and feeds the data back to a controller which adjusts the control (e.g. the actuator) as necessary to maintain the desired system output (e.g. match the actuator’s position to the reference position).

Front 19

FLIGHT-BY-WIRE COMPONENTS

Back 19

The Legacy 500 FBW architecture is structured around four separated stick-to-surface control paths. Each control path consists of:

• – One Inceptor Interface Modules (IIM);

• – Two Remote Electronic Units (REU);

• – Two Flight Control Computer (FCC).

  
  

  Each one of the four control paths incorporates one dual channel data bus that interface to both FCCs and two REUs.

• 1. The FCS operates in two distinct modes: Direct Mode (DM) and Normal Mode (NM). In DM the surface displacement is directly related to the inceptor command. In NM the A/C is commanded by the closed loop FBW Control Law (CLAW) based on pilot inceptor inputs.

  
  

Front 20

INCEPTOR INTERFACE MODULE (IIM)

Back 20

The basic IIM function is to transmit digital data from the cockpit interfaces listed below to the FCC and REU.

The cockpit interfaces are as follows:

• – Left and right seat pilot pedals;

• – Left and right seat pilot sidestick controller:
  AP Quick Disconnect/Priority Select (AP/PTY) button; Touch Control Steering (TCS) button.

• – Flap selector lever;

• – Speed brake lever;

• – Pitch Trim switch;

• – YAW TRIM knob;

• – FLTCTRL NORMAL MODE button.

  
  

  The IIM also perform data acquisition, in-line monitoring of the sidestick controller, pedals and levers sensors, and handling of digital data exchange between other devices, such as airplane avionics.

Front 21

1. REMOTE ELECTRONIC UNIT (REU)

Back 21

1. he REU perform the following function:

   
   

   • – Controls the actuator position based on FCC commands, and limits the primary surfaces deflection rate in NM and DM by interfacing and controlling two EHA and one EMA. Each surface has its own rate limiters;

   • – The two EHA belonging to one REU are dedicated to one specific hydraulic system;

   • – There are two REU types to prevent failure due to common cause REU failure propagation;

   • – Determines flap commands based on FSL position and performs closed-loop position control of the EHA and FAS;

   • – Controls and monitors MCE for flap and HSTAB actuation;

   • – Monitor the actuators health;

   • – Performs closed-loop velocity control of the PTAS;

   • A Cross-Channel Data Link (CCDL) provides communication between the two REUs controlling each pair of actuators on each primary surface.

Front 22

FLIGHT CONTROL COMPUTER (FCC)

Back 22

The FCC hosts the following functions:

• – Fly-By-Wire Control Laws (CLAWS) in DM and NM;

• – One FCC is enough to provide NM functionality to all control channels;

• – FCS failure management (OMS);

• – Sidestick controller, pedals, levers, button and switches input signals for selection and voting purposes, system monitoring, built-in-test processing and others;

• – Coordinates PBIT execution within FCS;

• – Determines actuators commands (except for flaps) based on control laws and AP.

  
  

  There are two FCC to provide a fail-operational capability. One FCC is enough to provide NM functionality to all control channels.

Front 23

1. ELECTRO-HYDRAULIC ACTUATORS (EHA)

Back 23

Each primary surface is controlled by two EHA in active-active configuration.

The primary surface hydraulic actuators have two modes of operation as follows:

– Active;

– Damped bypass.
Each primary surface is commanded by both REU types, such that no

primary surface is lost due to a common cause REU failure.

1. There are EHA for primary and spoiler surfaces:

   
   

   • – Four aileron EHA;

   • – Four elevator EHA;

   • – Two rudder EHA;

   • – Six MFS EHA.

Front 24

ELECTRO-MECHANICAL ACTUATORS (EMA)

Back 24

There are EMA for the following systems:

• – Pitch Trim Actuation System (PTAS);

• – Flap Actuation System (FAS).

  
  

  The PTAS and FAS utilize EMA powered from airplane 28 VDC power buses.

  
  

  The electrical power module for REU to handle EMA is the Motor Control Electronics (MCE) unit.

  
  

  The REU commands the EMA indirectly through a separate MCE channel.

  
  

  There are two MCE in the system: one to control both flap channels and another to control both pitch trim channels.

Front 25

1. PITCH TRIM ACTUATION SYSTEM (PTAS)

Back 25

1. The PTAS is implemented via dual-channel control architecture, composed by two different types of REU and two independent MCE channels.

   
   

   Each PTAS channel consists of a FCC, IIM, REU, a MCE channel and the PTA channels, paired together to control one of two PTA motors.

   
   

   The active channel selection is automatically controlled by the FCC, REU and MCE.

   
   

   The PTA, installed in the vertical tail, utilizes two electrical motors in an active-standby configuration that commands the actuators via a slip clutch.

   
   

   The PTA can be manually commanded from the Pitch Trim switch, only on ground or in DM, or automatically commanded by the CLAWS hosted in the FCC.

Front 26

FLAP ACTUATOR SYSTEM (FAS)

Back 26

The FAS is composed by two flap surfaces per wing, named inboard (IB) and outboard (OB) flap surfaces.

There are three flap linear actuators per wing. The OB flap surfaces are driven by two flap actuators each, while the IB flap surfaces are driven by one flap actuator each, and they are interconnected by a driveline composed by eight flap flexible shafts in the airplane. A Power Drive Unit (PDU) drives the drive line and the actuators.

Electrically, the FAS is a dual channel control architecture commanded independently by two different REU types and respective MCE channels to provide a high integrity REU/MCE command path for each flap PDU motor.

Front 27

FLIGHT CONTROL SYSTEM (FCS) INTERFACES

Back 27

The FCS also interfaces through data buses with other airplane systems as follows:

Avionics system:

• – Two DMC Cabinet;

• – One Radar Altimeter (RAD ALT) plus an optional;

• – Four Attitude Heading Reference Systems (AHRS);

• – Four Air Data Smart Probes (ADSP);

• – Four Display Units (DU);

• – Flight Guidance (FG) hosted in the DMC Cabinet.

  
  

  Propulsion system:

– Two Full Authority Digital Engine Control (FADEC).

Brake control system:
– Two Brake Control Unit (BCU).

Landing gear system:
– Two Landing Gear and Hydraulic Control Unit (LGHCU).

Air management system:
– Pneumatic System Controller (PSC).

Electric system:

• – Two main DC buses 1 and 2;

• – Two emergency DC buses 1 and 2, that includes A/C ram air turbine and batteries;

• Black Airplane batteries (BA) dedicated for FCS.

• The BA indication calls for actuators and equipment that operates under complete loss of normal and emergency electrical system, considering all main batteries, bus bars, RAT, etc or a typical Black Airplane (BA) scenario. This battery is able to withstand the operation of a minimum of 15 minutes in order to allow crew to recover the proper airplane electrical system operation.

  
  

  NOTE: The SFIS will also be powered the backup batteries under this scenario.

Front 28

MODES OF OPERATION

Back 28

1. NORMAL MODE (NM)

   
   

   The NM provides improved flying handling qualities, reduced crew workload, increased passenger comfort and enhanced flight safety through the CLAWS implemented in the FCC.

   
   

   This NM is supported by two FCC units configured as a dual active-standby channel, but only one FCC is capable to provide NM functionality to all control channels. The FCC unit in control will be determined based on priority logic which selects the FCC that has the most resources available to maintain the highest operational mode. If both FCCs are equally capable, the FCC in control will alternate during airplane power up based on the last FCC in command.

2. 1. In case of a FCC failure, the remaining FCC will maintain the FCS in NM, and the FLTCTRL NO DISPATCH CAS message will be displayed in-flight (cruise flight phase). However, if the remaining FCC fails, the FCS will revert to DM.

      
      

      The crew is able to recover NM or revert to DM through the FLTCTRL NORMAL MODE button, located on the central pedestal, in case of NM misbehavior.

   
   

   DIRECT MODE (DM)

   
   

   The FCS reverts automatically to DM operation when it is not possible to keep the NM operation due to multiple system failures, which affected the availability of the flap position, ADSP or AHRS data or the internal FCC integrity.

   
   

   In DM, the airplane behaves like a conventional airplane, and the sidestick controller and pedal deflections are directly related to the surfaces position, limited by the airplane airspeed and flap position or, in case of loss of air d

   
   

   The FCS reverts to DM whenever:

   
   
   
   

   – The inputs to both FCC necessary to maintain the NM are not available, then the system reverts to FCC DM or;

   
   
   
   
   
   

   – Not enough integrity of applicable sensors (ADSPs, flap position and AHRS);

   
   
   
   
   
   

   – Loss of both FCC. Both FCC internal integrity check has detected an error to maintain NM, and the system reverts to REU DM;

   
   
   
   
   
   

   – When one primary surface is not capable to be controlled in NM, the FCS reverts to DM, as long as that surface is controllable in DM;

   
   
   
   
   
   

   – Loss of both databuses of one primary surface;

   
   
   
   
   
   

   – Manual reversion performed by the crew through the FLTCTRL NORMAL MODE button, if the envelope protection is not activated.

   
   
   
   

Front 29

NORMAL MODE VERSUS DIRECT MODE SELECTION

Back 29

NORMAL MODE VERSUS DIRECT MODE SELECTION

The system is designed in a way that all the primary surfaces are always in the same FCS operational mode. In case both actuators on a primary surface are not able to operate in NM, the FCS maintains NM only if it identifies that the surface is also not recoverable in DM (not DM capable).

The FLTCTRL NORMAL MODE button installed on the central console allows NM recovery when FCS is NM capable. In case of NM misbehavior, this button may be used to revert the FCS to DM, as long as no FCS protection function is active.

NOTE: The flight under DM should be avoided as some airplane protection functions are only available in NM.

he FCS reverts automatically to DM operation when it is not possible to keep the NM operation due to multiple system failures, which affected the availability of the essential sensors (flap position, ADSP or AHRS data) or the internal FCC integrity.

Front 30

FCS BUILT-IN TEST

Back 30

STARTUP BUILT-IN TEST

The SBIT is a basic core processing capability test, applicable to IIMs, REUs, MCEs and FCCs to ensure that the unit has been correctly started-up and to determine the airplane configuration and unit position.

The SBIT duration time does not exceed 10 seconds, and is automatically performed at every power-up, no matter the airplane is in the air or on the ground. Also, it does not prevent the in-air power-up time requirements from being achieved.

POWER-UP BUILT-IN TEST

The PBIT automatically tests the FCS at each power-up under FCC control in order to ensure the FCS is correctly configured and to reduce the FCS component’s exposition to latent faults.

Each PBIT operation takes less then 1 minute, and if conditions are not satisfied, the FCS automatically perform another PBIT operation, up to three times, before declaring a failure, extending the PBIT to approximately 4 minutes. During the PBIT operation, the FLTCTRL TEST IN PROG CAS message is displayed.

If the PBIT operation does not pass or a failure is detected, the FLTCTRL TEST FAIL CAS message will be displayed. In this case, the Onboard Maintenance System (OMS) shall be checked.

If the PBIT operation is interrupted, the PBIT TIME EXPIRED CAS message will be displayed and the airplane must be powered down and then powered up to run the PBIT again.

The system keeps track of the hours interval between successful PBIT completions and prompts PBIT execution through the PBIT TIME EXPIRED CAS message when such interval exceeds the limit set by safety analysis, which corresponds to 20 hours starting at each airplane power up. When this CAS message is displayed, a new successful PBIT is required to clear this message.

The PBIT comprises the following tests:

• – Electrical PBIT;

• – Hydraulic PBIT.

Front 31

FCS BUILT-IN TEST

The PBIT comprises the following tests:

• – Electrical PBIT;

• – Hydraulic PBIT.

Back 31

ELETRICAL POWER-UP BUILT-IN TEST

The FCC initiates the electrical PBIT after airplane power up provided that:

• – Airplane is on the ground;

• – Engines are not running (OFF);

• – FCC power is valid for 5 consecutive seconds.

  
  

  During PBIT, no surface movement should occur under normal conditions. However, some small flap movement due to PBIT failure may occur before it is detected and arrested.

  
  

  HYDRAULIC POWER-UP BUILT-IN TEST

  
  

  The hydraulic PBIT is used to perform tests that either requires the hydraulics to be pressurized or result in a small control surfaces movement.

  
  

  The FCC initiates the hydraulic PBIT after engine start provided that:

• – Airplane is on the ground;

• – All tree hydraulic systems are powered on;

• – REUs are available to start the test.

Control laws Flight controls 9-08-40 page 20

Front 32

FLY-BY-WIRE CONTROL LAWS

Back 32

The CLAWS are a set of functions, named closed-loop functions, that computes the control surfaces commands based on airplane sensors data, airplane configuration and pilots’ inputs through the sidestick controllers and the rudder pedals.

Due to these closed-loop functions (CLAWS), pilots’ commands are interpreted as demands for airplane response rather than surface deflection. In this sense, the CLAWS command the surfaces so that the airplane achieves the desired response.

Front 33

FLY-BY-WIRE CONTROL LAWS:

NORMAL MODE FUNCTIONS

Back 33

NORMAL MODE FUNCTIONS

The NM provides automatic compensation for thrust variations, landing gear and different flap configurations, and provides the same flight characteristics for different weights, CG positions and speeds.

There is damped airplane response during turbulence, minimizing airplane oscillations; reduced transient in case of a system failure such as engine shutdown or surface hardover, so that the airplane controllability is easier, letting the flight crew have clear indications/cues of the failure.

In-flight, there is no manual command to PTAS, only automatic pitch trim compensation controlled by the CLAWS. On ground, the manual command to PTAS is available.

Front 34

NORMAL MODE FUNCTIONS; AXIS CONTROL FUNCTIONS AND PROTECTION FUNCTIONS

Back 34

Pitch Axis- CONTROL FUNCTIONS AND FEATURES

– Longitudinal neutral stability command;

– Pitch damper;
– Autotrim;
– Lift compensation in turns.

PITCH AXIS- PROTECTION FUNCTIONS

• Angle-of-attack limiter;

• – High speed protection;

• – Load factor protection;

• – Pitch protection (soft limit);

• – Elevator surface limiter.

Roll Axis-CONTROL FUNCTIONS AND FEATURES

- Roll rate command;

– Neutral spiral stability, positive spiral stability for large bank angle;

– Automatic lateral compensation with sideslid.

ROLL AXIS- PROTECTION FUNCTIONS

• – Bank angle limiting when in high speed protection.

Yaw Axis- CONTROL FUNCTIONS AND FEATURES

Sideslid command;

– Yaw damper (on ground);
– Turn coordination;

– Yaw trim (sideslip command);
– Sideslip target.

YAW AXIS - PROTECTION FUNCTIONS

• – Yaw input limiter;

• – Thrust asymmetry transient reduction.

Lift Dump - CONTROL FUNCTIONS AND FEATURES

- Speed brake;

– Ground spoilers.

Front 35

LONGITUDINAL CONTROL LAW TRANSITION LOGIC DIAGRAM

Back 35

Front 36

LONGITUDINAL CONTROL LAW TRANSITION LOGIC; STICK TO SURFACE CONTROL LAW

Back 36

The stick to surface control law is activated when the airplane is on ground and up to 65 kt on the takeoff roll. In this control law, there is a stick to surface direct link with no augmentation (no airspeed gain schedule) and without autotrim, which means that there is a direct relationship between the sidestick controller and elevator deflection irrespective of the airplane airspeed.

The takeoff trim is manually set via the pitch trim switch located on the center pedestal. The reverse process occurs after touchdown.

Front 37

LONGITUDINAL CONTROL LAW TRANSITION LOGIC: TAKEOFF CONTROL LAW

Back 37

The takeoff control law as well as the pitch dumper function is active when the airspeed is greater than 65 kt during takeoff roll. The control law receives a pitch rate demand as a function of the aft sidestick deflection, providing precise capture of pitch angle after lift off and standardizing the airplane response for different weights and CGs. At full aft position, the sidestick provides 15° per second of pitch rate demand without auto trim.

Front 38

LONGITUDINAL CONTROL LAW TRANSITION LOGIC: LONGITUDINAL NEUTRAL STABILITY CONTROL LAW

Back 38

The longitudinal neutral stability control law is a flight path demand mode which is active 5 seconds after airplane lift off with the pitch rate less than 2° per second. In this mode, the airplane becomes flight path stable rather than speed stable, and the control law response remains almost unaffected by the weight or CG shift. The longitudinal neutral stability control is active after takeoff until landing, where the de-rotation control law is active.

Front 39

LONGITUDINAL CONTROL LAW TRANSITION LOGIC:

APPROACH TCS TRIM CONTROL LAW

Back 39

The approach TCS trim control law adds speed stability to the longitudinal neutral stability control law, becoming active when the landing gear is down and the flap is set to 3 or FULL position with the AP disengaged. The control law activation is indicated through a TCS TRIM green flag below the airspeed tape.

NOTE: The speed stability control law will also engage in case of flap and/or landing gear failure.

Upon control law activation, the auto trim function is not available and the airplane must be manually trimmed via the TCS button. Refer to AOM Section 9-09-30 AIRSPEED INDICATION. When the TCS button is pressed, the airplane is trimmed to the current airspeed and the TCS trim bug is displayed. The airplane will maintain the load factor and stall protection, restoring the trimmed airspeed in case of atmospheric turbulence and sidestick controller deflection until a new airspeed is selected through the TCS button. The TCS trim reference speed is limited at the low speed awareness activation.

The airplane flare is accomplished by pulling the sidestick controller back as required in order to reduce the airplane airspeed, providing conventional flare characteristics without using the radio altimeter data.

The approach TCS trim control law is active until airplane touchdown.

Front 40

LONGITUDINAL CONTROL LAW TRANSITION LOGIC: DE-ROTATION CONTROL LAW

Back 40

The de-rotation control law is active when the airplane touchdown, providing a smooth nose down movement by commanding the elevator to a nose down pitch rate of 2° per second when the sidestick controller is set to the neutral position.

- The airplane will transition back to takeoff control law when the airplane pitch is less than 2° and then to stick to surface control law when the airplane airspeed is less than 65 kt during the landing roll. If a go around is required, the crew must press the TO/GA button on the power quadrant.

In case of bounced landing, the airplane will transition from the de-rotation law back to approach TCS trim control law, with the elevator trimmed for the last approach speed set prior to the flare. If a go around maneuver is required, as the landing gear and flaps are retracted, the airplane will transition to the takeoff control law and then to the longitudinal neutral stability control law.

Front 41

LATERAL AND DIRECTIONAL CONTROL LAW – ROLL AND YAW AXIS

Back 41

Front 42

ON GROUND LATERAL AND DIRECTIONAL CONTROL LAW -YAW DAMPER

Back 42

The lateral and directional ground mode is activated when the airplane is on ground or 2 seconds after touchdown. The lateral control command is generated through a direct link between the sidestick controller and the ailerons and spoilers surfaces, as well as the directional control command is generated through a direct link between the pedals and the rudder surface.

YAW DAMPER-

1. The yaw damper computes a rudder deflection based on the airplane yaw rate response to improve the airplane dutch roll characteristics. In NM, the YD is available with the airplane airspeed greater than 60 kt during the takeoff roll, and has the objective to improve airplane response under engine failure scenario on ground. The YD command is summed to the pilot command but is limited to 20% of the maximum available deflection. The pilot remains with full authority over the directional axis.

Front 43

1. IN-FLIGHT LATERAL AND DIRECTIONAL CONTROL LAW

Back 43

- The in-flight lateral and directional control is provided by the neutral spiral stability control law, based on the lateral sidestick controller and pedal deflection and activated after airplane lift off.

- The amount of aileron, spoiler and rudder control surface deflection that results from the sidestick controller and pedals deflection depends upon:

1. – Current airplane flap configuration.

   
   

   When the sidestick controller is laterally deflected, a roll rate demand is generated and the control law combines the aileron, spoiler and rudder control surfaces movement to achieve the roll rate demanded, coordinates the turns, and damps the dutch roll.

   
   

   When the pedals are deflected, a sideslip angle demand as well as a bank angle demand is generated proportional to the pedals deflection. The bank angle can be compensated by commanding, but without holding, the sidestick controller to level the wings. Furthermore, commanding and holding a pedal deflection results in a stabilized bank angle (zero roll rate).

2. – Airspeed from the ADSP;

3. – Inertial data from flight control AHRS;

Front 44

NEUTRAL SPIRAL STABILITY CONTROL LAW : to maintain the flight level during turns up to 33° of bank angle.

Back 44

After a sidestick controller roll rate command, the neutral spiral stability control law automatically trims the elevator when the sidestick controller is set back to the neutral position, to maintain the flight level during turns up to 33° of bank angle.

Front 45

POSITIVE SPIRAL STABILITY CONTROL LAW : Bank angles > 33dg with sidestick neutral rolls aircraft back to 33 deg bank angle

Back 45

The positive spiral stability control law is provided for bank angles beyond 33° with the sidestick controller at neutral position. The control law starts to add an amount of spiral stability proportional to the difference between the actual bank angle and the 33° reference bank angle, rolling the airplane back to 33° when the sidestick controller is set to the neutral position, providing a soft limit envelope protection. At 33° bank angle, the control law transition back to the neutral spiral stability control law.

- The amount of spiral stability applied is proportional to the bank angle beyond 33°, and inversely proportional to the sidestick deflection. Increasing the sidestick controller lateral deflection towards the lateral stop, the amount of spiral stability is proportionally decreased and, at the stop, the amount of spiral stability added is null. The maximum roll rate capability is achieved by setting the sidestick controller at the lateral stop.

- commanding a bank angle > 33 deg will require a constant lateral pressure from the sidestick and in order to maintain altitude, you will need to move the sidestick rearward commanding the elevator to pitch up

- This feature combined with automatic pitch compensation up to 33° of bank allows safe recovery from roll upset by returning the stick to the neutral position.

Front 46

SIDESLIP TARGET

Back 46

The sideslip target is activated in a OEI condition during the takeoff, climb and go-around flight phase and changes the slip-skid indicator color from white to blue. The control law generates a sideslip indication through the slid-skid indicator proportional to the residual sideslip.

- The residual yaw and bank angle can be compensated by applying rudder and sidestick controller in order to maintain airplane heading and improve climb gradients. The required amount of rudder or pedal deflection to counteract the residual OEI yaw moment is provided by centering the roll scale and the slip-skid indicator.

- NOTE: At full range, the sideslip indicator color will change from blue to yellow.

Front 47

LIFT DUMP FUNCTION

Back 47

The lift dump function of lift dump, airplane drag increase, and steeper angle of descent creation, is implemented in the FCC.

The lift dump function increases the airplane drag and provides a steeper angle of descent. The lift dump functions are as follows:

• – Speed brake;

• – Ground spoiler;

• – Flap 3 with spoiler function.

Front 48

LIFT DUMP FUNCTION:

SPEED BRAKE (test question)

Back 48

The speed brake control is provided by the speed brake lever. The three pair of multifunction spoiler panels are symmetrically deployed at the same angle, varying from 0° (stow) to 30° (full), according to the speed brake lever position.

The speed brake function is enabled provided that:

• All consumed data are available and valid;
• Flap lever and position is less or equal to 2;
• Airspeed is greater than 160 kt;
• TLA position is less than 55° (no takeoff/go-around condition);
• The AOA limiter function is not engage

-If any of the above conditions are removed with the speed brake deployed, the control law generates a command to retract the speed brake with its lever in the deployed position, and displays the SPDBRK LEVER DISAG CAS message. In this case, the speed brake lever must be set back to the CLOSE position, to clear the CAS message and allowing the speed brake reactivation.

Front 49

LIFT DUMP FUNCTION:

GROUND SPOILERS

Back 49

When the ground spoiler function is activated, all multifunction spoiler panels are commanded to the full deployed position (40°).

The ground spoiler function operates in four different modes:

• – ARMED MODE: When one of the following conditions are met: Reject takeoff condition: When the wheel speed is greater than 30 kt;
  Touchdown condition: When the WOW sensors indicate air for more than 10 seconds.

• – ENGAGED MODE: After the ARMED mode, the ground spoiler engaged when:

  
  

  Both TLA indicates idle position (less than 26°); WOW sensor indicates ground.

• – DISARMED MODE: When the wheel speed and airspeed is less then 30 kt and 40 kt respectively.

• – FAILED MODE: When all of the following conditions are met: When the airplane airspeed is greater than 130 kt for more than 60 seconds and the WOW sensors still indicates ground; Loss of required data to compute the ground spoiler function.

  
  

Front 50

LIFT DUMP FUNCTION:

FLAP 3 WITH SPOILER FUNCTION

Back 50

This function provides lift dump and increases the airplane drag in order to assist the pilot to maintain the airspeed and trajectory while performing landing with flap 3 configuration. When this function is activated, the three pair of spoilers is commanded to 15°.

The flap 3 with spoiler function is activated when the following conditions are met:

• Flap position is equal to 3;
• Both TLA position less than 50° (the takeoff/go-around condition is not met);
• The AOA limiter function is not engaged;

• All spoiler panels are engaged.

If any of the above conditions are removed when the spoilers are already deployed for the flap 3 with spoiler function, the control law commands the multifunction spoilers back to the retracted position.

Front 51

FLIGHT CONTROL NORMAL MODE MISBEHAVIOR

Back 51

A flight controls NM misbehavior is any airplane behavior that the flight crew realize it is behaving unexpectedly, not responding adequately to the commands or is presenting lack of response.

- A manual reversion to DM is available in case the flight crew identifies a flight control NM misbehavior. Under this scenario, the pilot shall press the FLTCTRL NORMAL MODE button in order to revert the FCS to DM (refer to AOM Section 4-02-01 FLTCTRL N-MODE

MISBEHAVIOR Non Annunciated).

- The reversion to DM is indicated through a > FLTCTRL N-MODE FAIL CAS message followed by a triple chime aural warning. The FCS synoptic page presents the control surfaces in yellow and the airplane looses the auto trim capability, besides the different piloting characteristics (control to surface).

- However, if the airplane is flying inside the low speed, AOA limiter or high speed envelope protection, the FLTCTRL NORMAL MODE button is inhibited, precluding the DM engagement. This will be immediately noticed by the absence of > FLTCTRL N-MODE FAIL CAS message. The PFD will also keep displaying the white tape and PLI (if inside AOA limiter) in case of a low speed and high AOA scenario.

- The flight crew must wings level, trim the airplane and assure the airplane is in a safe altitude before attempt to recover NM by pressing the FLTCTRL NORMAL MODE button.

Front 52

FLIGHT ENVELOPE AND PROTECTION FUNCTIONS

Back 52

The CLAWS also provides flight envelope protections to prevent the airplane from exceeding the flight envelope limitation, by calculating the appropriate control surface deflections in order to maintain the airplane inside the limits. There are two flight envelopes for the airplane:

• – Normal flight envelope;

• – Limit flight envelope.

  
  

  The longitudinal and lateral/directional envelope protection functions are the AOA limiter and the high speed protection.

Front 53

FLIGHT ENVELOPE AND PROTECTION FUNCTIONS : NORMAL FLIGHT ENVELOPE

Back 53

1. In the normal flight envelope, airplane control requires less pilot workload, with significant improvement to passengers comfort. The CLAWS maintain the airplane trimmed with the sidestick controller at neutral position in any condition inside this envelope. The boundaries of this envelope are reached with less than full control application.

   
   

   As illustrated in the following figure, the limits of the normal flight envelope are:

• – 33° bank angle protection;

• – High speed protection (VMO);

• – Low energy avoidance protection (1.13 VS);

• – Pitch Attitude Protection.

Front 54

FLIGHT ENVELOPE AND PROTECTION FUNCTIONS : HIGH SPEED PROTECTION

Back 54

The control law automatically recovers by commanding an additional load factor proportional to the difference between the current speed and the maximum operating speed, limited to 1 g, towards the VMO when the airspeed limit of VMO + 6 kt or MMO + 0.01 is exceeded.

- The pilot sidestick controller nose down authority is reduced during the recover maneuver, but if a full forward sidestick controller deflection is commanded, the airspeed is limited to VMO + 22 kt. Positioning the sidestick controller to the neutral position, the airplane returns to the normal flight envelope, maintaining the load factor and AOA (stall) protection.

- A visual warning comprised by an emphasized VMAX marker (barber pole) shall be already in view before the airplane reaches the VMO, and the airspeed digits (knots and Mach) become red inverse video indicates that the current airspeed becomes greater than VMO (refer to AOM Section 9-09-30 AIRSPEED INDICATION). A HIGHSPEED aural warning continually sounds when the airspeed is greater than VMO + 4 kt.

Front 55

FLIGHT ENVELOPE AND PROTECTION FUNCTIONS : LOW ENERGY AVOIDANCE PROTECTION

Back 55

The low energy avoidance provides adequate low energy awareness to the pilot, ensuring that the airplane will not inadvertently reach speeds significantly below the normal operating speeds. The system is comprised of Low Speed Protection and Low Speed aural Warning.

- The low speed protection function is activated when the speed becomes lower than a value of approximately 1.13 times the stall speed, represented by the top of the white tape (VSS). In landing flap configurations, it indicates the minimum speed the airplane can be trimmed at. In the other flap configurations, it indicates the addition of

speed stability to the longitudinal control law.

- When the autopilot is engaged, the autothrottle (if installed) will automatically engage slightly above VSS in order to prevent the airspeed to reduce below the top of the white tape. A stabilized flight within the white tape range will only be possible with an aft sidestick controller deflection, commanding the airplane to a pitch up attitude and decreasing the airspeed to a value smaller than VSS.

- Positioning the sidestick controller back to the neutral position, the airplane will return to VSS. The TCS trim reference speed is also limited at VSS.

A LOW SPEED aural warning is also provided to get the pilots attention to a low energy condition. The aural warning is activated when the airspeed is reduced to VSS and is interrupted when it is increased above this value. Below 2000 ft AGL, the low speed aural warning may be anticipated or activated with the airplane airspeed above VSS based on the following condition:

– If a high rate of energy loss is detected represented by the airspeed trend vector;

– If an altitude loss is detected based in a glide slope deviation.

NOTE: A barometric altitude reference of 16000 ft and below is considered, instead of 2000 ft AGL, in case of radar altimeter failure.

- The aural warning is inhibited in case the throttles are close to TO/GA and/or when radar altimeter is below 100 ft AGL.

Front 56

FLIGHT ENVELOPE AND PROTECTION FUNCTIONS : PITCH ATTITUDE PROTECTION

Back 56

The pitch attitude protection is an additional function to increase the effectiveness of the stall and high speed protections, and activates at pitch angles of:

• – 30° of pitch up;

• – 15° of pitch down.
  When this protection is activated, the control law commands an additional positive or negative load factor in low or high pitch angle scenarios respectively. Additionally, the pilot authority is restricted making it difficult to lower or increase the airplane pitch attitude. If the sidestick controller is positioned to the neutral position, the elevator is commanded to achieve activation pitch angles of 30° nose up or 15° nose down.

  
  

Front 57

FLIGHT ENVELOPE AND PROTECTION FUNCTIONS: LIMIT FLIGHT ENVELOPE

Back 57

- Outside the normal flight envelope, there is the limit flight envelope. The limit flight envelope represents the condition in which the airplane is close to the envelope hard limits, thus requiring more pilot effort to maintain the airplane in this condition.

- The pilot is not able to override the hard limits, and if the sidestick controller is released, the airplane is brought back to the normal flight envelope.

- The external boundaries of the limit flight envelope are reached at full control application. The hard limits of this flight envelope are:

• – Maximum angle of attack - AOA limiter;

• – Airplane structural limits - load factor protection;

• – Maximum sideslip protection;

• – Dive speed protection.

Front 58

FLIGHT ENVELOPE AND PROTECTION FUNCTIONS: ANGLE OF ATTACK (AOA) LIMITER PROTECTION

Back 58

The AOA limiter protects the airplane from low speed stall and high Mach buffeting, and is activated when the angle of attack becomes greater than a value which corresponds to approximately 1.08 times the stall speed, represented by the top of the yellow tape (VAOA), engaging the AOA limiter control law.

The aft sidestick controller deflection corresponds to an angle of attack demand instead of load factor demand, and the control law will command the elevators to achieve the demanded angle of attack. Full aft sidestick controller deflection corresponds to the maximum commanded angle of attack (VLIM), which is approximately 1.03 times the stall speed represented by the top of the red tape.

The control law allows the pilot to maintain the angle of attack limit at full aft sidestick controller deflection, providing adequate roll maneuverability, good flight characteristics, and enough margins to the stall (CLMAX).

If the pilot sets the sidestick controller to the neutral position, the control law will command the airplane back to the VAOA, preventing the airplane from stalling.

Front 59

FLIGHT ENVELOPE AND PROTECTION FUNCTIONS: LOAD FACTOR PROTECTION

Back 59

The control law commands the elevator to maintain the load factor proportional to the sidestick deflection, airplane airspeed and flap configuration. The load factor is limited as follows:

– Flap up:-1g to 2.5g;

NOTE: Full forward stick command equal to -0.2 g, becoming -1 g in 3 seconds if full forward sidestick is maintained.

– Flap down:0g to 2g for flaps down.

Front 60

FLIGHT ENVELOPE AND PROTECTION FUNCTIONS: MAXIMUM SIDESLIP

Back 60

MAXIMUM SIDESLIP: As the pedal position corresponds to a sideslip command, the control law provides directly a sideslip limiter function. The sideslip angle is limited with the pedals at the stops.

Front 61

FLIGHT ENVELOPE AND PROTECTION FUNCTIONS: DIVE SPEED (VD)

Back 61

The control law does not allow the airplane to exceed the maximum dive speed of Vd/Md equal to 360 kt/.90 M.

Front 62

THRUST ASYMMETRY TRANSIENT REDUCTION

Back 62

THRUST ASYMMETRY TRANSIENT REDUCTION

In OEI condition, the control law automatically compensates almost all the yaw moment generated by the operating engine, providing only a residual sideslip and bank angle for the pilot to recognize the OEI condition and which engine is inoperative.

NOTE: If no pilot action is taken, the airplane will slowly diverge the heading on a stabilized sideslip angle and bank angle condition.

The OEI condition is detected if all the following conditions are satisfied:

• – There is a N1 asymmetry equal or greater than 20%;

• – Both engines are not accelerating or decelerating up to 3 seconds after the N1 asymmetry detection or the asymmetry has been detected for longer than 3 seconds.

  
  

  The OEI condition is latched if there is an announced engine failures in any of the engines.

Front 63

DIRECT MODE FUNCTIONS

normal mode

direct mode

Ultimate backup

Back 63

- The FCS automatically reverts to DM operation when data availability from the flap position, ADSP or AHRS to both FCC is affected due to system failure or the FCC integrity required to maintain NM CLAW is also affected.

- The DM functions are classified as control functions/features and protection functions, as presented in the following table:

Pitch Axis: CONTROL FUNCTIONS AND FEATURES

- Surface command

- Manual stabilizer trim

Pitch Axis: PROTECTION FUNCTIONS

- Stall aural alert elevator gain schedule

Roll Axis: CONTROL FUNCTIONS AND FEATURES

– Surface command

Roll Axis: PROTECTION FUNCTIONS

- Aileron and spoilers gain schedule

Yaw Axis: CONTROL FUNCTIONS AND FEATURES

- Surface Command

- Yaw Trim (rudder command)

- Yaw Damper

Yaw Axis: PROTECTION FUNCTIONS

- Rudder gain schedule

Lift Dump: CONTROL FUNCTIONS AND FEATURES

- Speedbrake (if FCC is available)

- Ground spoilers (if FCC is available)

Lift Dump: PROTECTION FUNCTIONS

- none

In DM, the airplane behaves like a conventional airplane, and the sidestick controller and pedal deflections are directly related to the surfaces position, limited by the airplane airspeed and flap position or, in case of loss of air data, only flap position information (ultimate backup).

- Furthermore, the automatic pitch trim is not available in DM, and the airplane has to be manually trimmed in-flight and on ground via the PTAS. The yaw damper is present for both in air and on ground.

NOTE: The AP is not operational in DM.

The DM can be hosted in either FCC or REU. When the FCS DM is running in the FCC, it is referred to as FCC DM. When the FCS DM is running in the REU, it is referred to as REU DM.

- In FCC DM, the FCC continues to process the available flap position, ADSP or AHRS sensors, the sidestick controller lateral deflection (aileron) and speed brake control lever inputs through the same data path as in NM. Thus, the MFS control is available for all three roll, speed brake and ground spoiler functions.

In REU DM, the MFS control for speed brake and ground spoiler functions are not available, however MFS roll command is kept operational.

Front 64

DIRECT MODE PROTECTION FUNCTION

Back 64

DIRECT MODE PROTECTION FUNCTION

The following protections are provided in DM:

• – Primary control surface gain schedule;

• – Stall aural alert.

  
  

  PRIMARY CONTROL SURFACE GAIN SCHEDULE

  
  

  The elevator, aileron and rudder deflection are limited by the airplane airspeed and flap position.

1. STALL AURAL ALERT

   
   

   The AOA limiter is not available in DM. A stall aural function is implemented to provide pilot awareness and an aural alert when the stall aural AOA is exceeded indicating the proximity with a stall condition.

   
   

   The stall aural in DM is computed in the DMC and is a function of:

   
   

   • – Flap position;

   • – Ice condition;

   • – Angle-of-attack (ADS);

   • – Mach (ADS).

     
     

     The stall aural output is then routed through RIU and ACP, and finally is sent to the loudspeakers to provide the aural alert when appropriate. The aural consists of a continuous Stall... stall... stall..., non-cancelable.

     
     

     When flap position or ice detector status is not available, the computation assumes a conservative value for the stall aural angle of attack and, in case of a invalid angle of attack information from the ADS, the AOA can not be computed.

Front 65

CAS MESSAGES:

WARNING: 2:

- FLTCCTRL N-MODE FAIL

- SIDESTICK LH(RH) FAIL

Back 65

WARNING:

- > FLTCTRL N-MODE FAIL:Loss of NM, including the following main functions: required yaw command, AOA limiter, longitudinal neutral stability, auto trim, autopilot.

- SIDESTICK LH (RH) FAIL: Loss of LH (RH) sidestick