Abstract: The present disclosure provides methods and system for controlling the operation of a fly-by-wire aircraft. One or more yaw commands are received from an operator control, and one or more actual induced rolls rates are determined based on the yaw commands. A yaw signal and a roll-countering command are sent to flight control components of the aircraft, the yaw signal to cause a yaw motion in the aircraft, and the roll-countering command to counter the actual induced rolls. A standardized roll rate command is determined based on the yaw command, and the standardized roll rate command is sent to the flight control components to cause a roll motion in the aircraft.

Abstract: A trimmable horizontal stabiliser actuator (THSA) control system includes least one main THSA motor for driving the THSA in response to inputs from a flight control computer (FCC), a pilot trim control system for manual control of the THSA by a pilot overriding the FCC. The pilot trim control system comprises a trim control for receiving an input from the pilot, the trim control including a trim control sensor for detecting the input from the pilot and providing a pilot trim control signal; a THSA trim control motor for driving the THSA in response to the pilot trim control signal; and an electronic override system for giving inputs from the pilot priority over inputs from the FCC.

Abstract: A method for determining an aircraft angle-of-attack for aircraft stall protection includes providing an output signal from an angle-of-attack sensor and determining an initial angle-of-attack signal based on the output signal. The initial angle-of-attack signal is compensated to provide a pseudo angle-of-attack signal, and the pseudo angle-of-attack signal is mapped to a true angle-of-attack signal based on flight test data. The true angle-of-attack signal is compensated based on roll rate and sideslip or estimated sideslip to provide a compensated angle-of-attack. A complementary filter is applied that complements the compensated angle-of-attack signal with a higher frequency inertial angle-of-attack rate signal, calculated from aircraft inertial data, to provide an angle-of-attack complementary filter output. An angle-of-attack threshold for aircraft stall protection is determined based on one or more compensation parameters.

Abstract: Aircraft control systems to enable shared input controls. A single inceptor may provide outputs to control both land-based aircraft controls and air-based aircraft controls to reduce the number of inceptors required. In an example a side-stick provides control of both pitch and aircraft nose-wheel steering. Active inceptor technology may be utilised to provide feedback to the operator on the control system state.

Abstract: A computer-implemented method for determining a scheduled maintenance session for a proportional actuator includes transmitting a command signal to a controller of the proportional actuator and generating, using the controller, a drive signal for controlling the proportional actuator using the command signal. The position of a component actuator of the proportional actuator is changed, on the basis of the drive signal, and the position of the component actuator is measured, using a sensor, and a position feedback signal is generated on the basis of the measurement. A condition indicator is estimated, with a processor, using only one of the command signal, the drive signal, or the position feedback signal. A health indicator is estimated, with the processor, using the condition indicator, and a scheduled maintenance session of the proportional actuator is determined, with the processor, using the health indicator.

Abstract: Provided are airplane trim systems and methods of controlling such systems. These systems utilize smaller portions of the stabilizer total travel range for takeoff trims, in comparison to other trim systems. A trim system described includes stabilizer and elevator, and these components are used together to achieved a takeoff total tail pitching moment. The elevator or, at least a portion of the elevator operating range, is available for flight control. As such, takeoff trim settings include stabilizer and elevator orientation settings. Addition of the elevator to control the takeoff tail pitching moment allows reducing the stabilizer total travel. The elevator orientation can be changed much faster than that of the stabilizer providing pilot more control.

Abstract: An automatic pilot device for a rotary wing aircraft and rotary wing aircraft comprising such device are disclosed. In one aspect, the automatic pilot device comprises at least one automatic pilot assembly including at least two primary actuators, at least one or each of the primary actuators incorporating an electronic computation unit. The computation unit is configured to: communicate with a measuring system configured to generate measuring signals and/or a cockpit configured to generate control signals as a function of the actions by a crew, and compute, as a function of the measuring signals and/or control signals, a piloting setpoint for the primary actuator incorporating the computation unit and/or a piloting setpoint for at least one or each other primary actuator of the automatic pilot assembly, for the piloting of the aircraft by the automatic pilot device.

Abstract: An actuator system for controlling a flight surface of an aircraft includes a first actuator having a first actuator input and a first linear translation element that moves based on rotational motion received at the first actuator input and a first sensor coupled to the first linear translation element that generates a first output based on a displacement of the first linear translation element. The system also includes a second actuator having a second actuator input and a second linear translation element that moves based on rotational motion received at the second actuator input and a second sensor coupled to the second linear translation element that generates a second output based on a displacement of the second linear translation element. The system also includes a control unit that receives the first and second outputs and determines if an error condition exists for the system based on first and second output.

Abstract: An actuator system for controlling a flight surface of an aircraft includes a first actuator having a first actuator input and a first linear translation element that moves based on rotational motion received at the first actuator input and a first laser distance sensor disposed inside the first actuator that generates a first output based on a displacement of the first linear translation element. The system also includes a second actuator having a second actuator input and a second linear translation element that moves based on rotational motion received at the second actuator input and a second laser distance sensor disposed inside the second actuator that generates a second output based on a displacement of the second linear translation element. The system also includes a control unit that receives the first and second outputs and determines if an error condition exists for the system based on first and second output.

Abstract: Actuator systems and methods for controlling aerodynamic control surfaces of aircraft including a first actuator receiving a first input and a first linear translation element that moves based thereon, the first linear translation element operably connected to a first portion of the control surface. A first sensor assembly is disposed relative to the first actuator that generates an output based on a displacement of the first translation element. A second actuator receives a second input and a second linear translation element moves based on the second input, the second linear translation element operably connected to a second portion of the control surface. A second sensor assembly is disposed relative to the second actuator that generates a second sensor output based on a displacement of the second translation element. A controller generates the inputs and receives the sensor outputs to determine if an error condition exists for the system.

Abstract: Method for predicting an operational malfunction in an aircraft equipment, the parameters of the equipment being monitored and recorded during flights by measurements or signals, the aircraft further including an equipment breakdown detection unit, the breakdown also being recorded in the aircraft memory. During maintenance phases between flights, measurements and breakdowns are retrieved on a programmable computer to form a database, and (a) a data analysis program is executed, a first time, to determine a set of pairs of parameter pairs whose signals develop positively over time in a correlated manner and in the absence of a breakdown, and (b) once the set of parameters pairs has been determined, a detection program is executed to calculate correlations for the pairs in the determined set of pairs and, when the calculated correlation value of one pair for a given flight falls below a predetermined positive detection threshold, a malfunction is reported.

Abstract: A control unit for controlling movement of a stabilizer and an elevator on an aircraft, the control unit comprising a processor which is configured to receive a pitch reference command from a flight control input device of the aircraft, determine, based on the pitch reference command, an elevator command configured to command the elevator in the aircraft nose-up or nose-down direction, determine a difference between the elevator command and a desired elevator value corresponding to the pitch reference command, determine a threshold value as a desired maximum elevator deflection at a given pitch reference command, compare the difference with the threshold value and in response thereto generate a command signal configured to retain a current position of the stabilizer when the difference is smaller than or equal to the threshold value, and to command the elevator in the aircraft nose-up or nose-down direction based on the pitch reference command.

Abstract: In accordance with an embodiment of the present invention, a method of operating a rotorcraft includes operating the rotorcraft in a heading control mode that includes activating a yaw channel path of a heading controller and deactivating a roll channel path of the heading controller when a speed of the rotorcraft is less than a first speed threshold or a heading error is less than a heading error threshold, and activating the roll channel path of the heading controller and deactivating the yaw channel path of the heading controller when the speed of the rotorcraft is greater than a second speed threshold and the heading error is not less than the heading error threshold.

Abstract: A wafer edge protection apparatus, including an electrical control module having a servomotor, a vertical motion mechanism and a protection mechanism. The protection mechanism includes: a shaft coupler, in fixed connection with a shaft of the servomotor and having a plurality of first connecting components; and a transmission sleeve, in fixed connection with the vertical motion mechanism and having a plurality of second connecting components in movable connection with the plurality of first connecting components.

Abstract: The invention relates to a flap system for an aircraft high lift system or an engine actuation with a rotary shaft system, one or more drive stations as well as elements for transmitting the drive energy from the rotary shaft system to the one or more drive stations, wherein at least one drive station includes at least two independent load paths with at least one rotational transmission each for actuating the flap kinematics, and per load path at least one mechanically coupling-free synchronization unit is provided for compensating regular load fluctuations between the load paths. The invention furthermore relates to a method for monitoring a flap system with at least two redundant load paths which each comprise at least one rotational transmission, wherein it is cyclically checked whether the difference of the output-side torques of the at least two load paths exceeds a defined threshold value and/or lies within a defined limit range.

Abstract: A method of assisting a pilot in order to minimize the risks of an aircraft rolling over on the ground, the aircraft having a main rotor and a yaw movement control rotor (4), together with lateral cyclic pitch control means for the main rotor and yaw movement control means for the yaw movement control rotor. At least one measurement is taken relating to left forces exerted on a left side undercarriage and to right forces exerted on a right side undercarriage in order to evaluate asymmetry, if any, between the left and right forces, and then a recommended position (26) is determined for at least one of the yaw and lateral cyclic pitch control means in order to make the left and right forces more symmetrical. Each recommended position is displayed on a display unit.

Abstract: A method of controlling a deflection angle of pitching stabilizer means for an aircraft. The method includes a preliminary stage for preparing at least one control relationship delivering a setpoint angle that is to be reached by the deflection angle as a function of at least one flight parameter of the aircraft and of an incidence parameter of the pitching stabilizer means. During an operational stage an operational stage (STP2), a “flight” value is determined for each of the flight parameters, a “setpoint” value is determined for the incidence parameter, a setpoint angle is determined by inputting the flight values and setpoint value into the control relationship, and at least one actuator is actuated in order to cause the deflection angle of the pitching stabilizer means to reach the setpoint angle.

Abstract: A multi-axis serially redundant, single channel, multi-path fly-by-wire control system comprising: serially redundant flight control computers in a single channel where only one “primary” flight control computer is active and controlling at any given time; a matrix of parallel flight control surface controllers including stabilizer motor control units (SMCU) and actuator electronics control modules (AECM) define multiple control paths within the single channel, each implemented with dissimilar hardware and which each control the movement of a distributed set of flight control surfaces on the aircraft in response to flight control surface commands of the primary flight control computer; and a set of (pilot and co-pilot) controls and aircraft surface/reference/navigation sensors and systems which provide input to a primary flight control computer and are used to generate the flight control surface commands to control the aircraft in flight in accordance with the control law algorithms implemented in the flight

Abstract: A flight control system for an aircraft directs a primary surface that is pivotable with respect to the aircraft with a moveable tab located at the end of the primary surface. A primary surface actuator maintains a position of the primary surface. A tab actuator moves the tab with respect to the primary flight control surface. A controller configured to coordinate the movement of the primary flight control surface with the movement of the tab by deflecting the tab with respect to the primary flight control surface, thereby allowing aerodynamic forces acting upon the deflected tab to position the primary flight control surface at a desired position. The primary flight control surface can then be maintained at the desired position by the primary surface actuator.

Abstract: Embodiments of a flight-control system and methods of reducing the probability of a roll-control reversal in a canard-controlled flight vehicle are generally described herein. In some embodiments, the flight-control system may monitor angular velocities of the flight vehicle to detect the onset of instability and adaptively control an acceleration limit of the flight vehicle based on the detected instability to reduce the probability of a roll-control reversal. The onset of instability may be detected by persistently high angular velocities. The acceleration limit may be further adaptively controlled based on an approach of a vehicle trim limit.

Abstract: An automatic trim system and method is disclosed for automatically trimming a flight control surface of an aircraft. A force sensor measures a force applied by a pilot to a flight control system actuator. The length of time that the force is applied by the pilot is then timed by a timer. A trim system to reduce the applied force is included on the flight control surfaces. A processor determines if trim is required if a predetermined amount of time is exceeded based on the force sensor measurement. The processor can set the trim system to the trim required therein. An airspeed sensor is used to verify that the aircraft has sufficient speed for flight. A force sensor can be utilized to measure the input force being applied by the pilot. If a pilot input force is applied to the controls and the aircraft is in a steady state, a timer can be activated.

Abstract: An alternative system for damping the dutch roll mode in an aircraft is provided using roll control surfaces. Classical yaw dampers for the dutch roll mode utilize the yaw control surfaces such as a rudder to dampen the dutch roll mode oscillations. An alternative damper is described that utilizes roll control surfaces such as spoilers or ailerons to dampen the dutch roll mode.

Abstract: A longitudinal control law is designed to optimize the flying qualities when aircraft is set to approach configuration, i.e. when the flap lever is set to the landing position and landing gears are locked down. Under such circumstances, the effort of trimming the aircraft speed can be extremely reduced by the usage of a momentary on-off switch or other control in the sidestick, instead of or in addition to a conventional trim up-down switch, making easier the task of airspeed selection by the pilot. This control law provides excellent handling qualities during approach and landing, with the benefit of not needing or using radio altimeter information in safety-critical applications.

Abstract: The invention relates to a method for actuating an adjusting drive for adjusting an elevator (12) and an adjusting drive for adjusting a moveable tail (23) provided with the steps: Generation of an elevator command to actuate the elevator adjusting drive; Calculating a moveable tail command (IHC1) for actuating the moveable tail adjusting drive in such a manner that the moveable tail (23) is tracked to the elevator input signal (10); Depending on the adjusting states of the elevator (12) and/or the moveable tail (23) or flight states, retaining the adjusting state of the moveable tail adjusting drive or actuating the moveable tail adjusting drive with a moveable tail command (IHCMD) for changing the adjusting state of the moveable tail (23), during actuation of the elevator adjusting drive with an elevator command for changing the adjusting state of the elevator (23) and in the event of a deviation from the calculated moveable tail command (IHC1) and the commanded moveable tail command (IHCMD), acting upon th

Abstract: Methods and apparatus for driving rotational elements of a vehicle are provided. The apparatus includes a first driving member configured to rotate a first rotational element of the vehicle relative to a body of the vehicle. The apparatus also includes a second driving member. The second driving member includes a drive structure rotatably coupled to the body. The second driving member also includes a gimbal element coupled to the drive structure via a first joint pin. The gimbal element is configured to couple to a second rotational element of the vehicle such that when the drive structure is rotated relative to the body, the second rotational element rotates relative to the first rotational element. The gimbal element is configured to pivot about the first joint pin relative to the drive structure based on the rotation of the first rotational element and the rotation of the second rotational element.

Abstract: A device for controlling vehicles having a manual control unit configured to influence the direction of movement of a vehicle. The manual control unit provides, in a neutral position of the manual control unit, a trim point to determine a preferred direction of movement. The device further includes a force generating device, generating at least one force acting in the direction of the neutral position of the manual control unit; a trim coupling operable to reduce the at least one force acting on the manual control unit; and a trim control unit configured to store and retain the trim point existing prior to an operation of the trim coupling.

Abstract: The present invention provides systems and methods for controlling the speed of flap retraction on aircraft, and alerts to the pilot of potentially unsafe flap position. The invention accepts direction from a pilot and senses operation of the aircraft to determine appropriate flap position and flap retraction speed.

Abstract: The invention relates to an actuator having an (auto)synchronous rotary electric motor and a reversible speed-reducing gearbox coupled to the motor to be driven in rotation thereby An outlet shaft is coupled to the speed-reducing gearbox to be driven in rotation thereby A first angular position sensor and a control circuit are connected to the motor. The circuit delivers a motor power supply signal that varies as a function of a position setpoint signal applied to the control circuit and as a function of signals delivered by the angular position sensor. The gearbox has a plurality of speed-reducing each having a pair of gears mounted to rotate about two parallel axes of rotation.

Abstract: Actuator systems and associated methods are disclosed herein. One aspect of the disclosure is directed toward an actuator system that includes a first structure and a second structure movable relative to the first structure. The system further includes an actuator apparatus and an actuator device coupled in series between the first structure and the second structure. The system can still further include a controller operably coupled to the actuator apparatus and the actuator device. The controller can be programmed with instructions to automatically actuate the actuator apparatus and the actuator device so that a position of the first structure relative to the second structure after the actuator apparatus and the actuator device have been actuated is at least approximately the same as a position of the first structure relative to the second structure before the actuator apparatus and the actuator device have been actuated.

Abstract: A device for controlling vehicles having a manual control unit configured to influence the direction of movement of a vehicle. The manual control unit provides, in a neutral position of the manual control unit, a trim point to determine a preferred direction of movement. The device further includes a force generating device, generating at least one force acting in the direction of the neutral position of the manual control unit; a trim coupling operable to reduce the at least one force acting on the manual control unit; and a trim control unit configured to store and retain the trim point existing prior to an operation of the trim coupling.

Abstract: The invention relates to an electric flight control system for aircraft elevators. According to the invention, the flight control system can be controlled in terms of load factor or rate of pitch. The inventive system comprises built-in protections in relation to load factor, incidence and pitch attitude.

Abstract: A breakable coupling device for coupling together first and second main transmission shafts stationary in translation along a longitudinal axis of rotation of the device comprises a blocking device having a discontinuous first housing and a continuous second housing forming a closed loop; a compression device; at least one drive device connecting the blocking device and the compression device together in rotation about the longitudinal axis below a predetermined torque, wherein the discontinuous first housing is configured to receive the at least one drive device below the predetermined torque; and a shifting device configured to shift the at least one drive device non-reversibly from the discontinuous first housing towards the continuous second housing when the torque exerted on the at least one drive device is greater than the predetermined torque.

Abstract: A locking mechanism, which permits fixation to a piston rod of a linear actuator relatively to a cylinder of the actuator, as well as a linear actuator equipped with such a locking mechanism as well as to an aircraft with at least one linear activator that is equipped with the locking mechanism. The locking mechanism includes a plurality of first positive locking elements which are arranged lengthwise with the outside circumference of the cylinder and in certain distances to one another and at least one second positive locking element, which is moved together with the piston rod upon activation of the linear actuator and thus passes a segment of the cylinder. The at least one second positive locking element is positively engageable with one of the plurality of first locking elements at discrete positions of the cylinder lengthwise, by which the piston rod is fixed relatively to the cylinder.

Abstract: The invention relates to an actuator (10) comprising: an (auto)synchronous rotary electric motor (11); a reversible speed-reducing gearbox (13) coupled to the motor to be driven in rotation thereby; an outlet shaft (16) coupled to the speed-reducing gearbox to be driven in rotation thereby; a first angular position sensor responsive to the angular position of the outlet shaft; and a control circuit connected to the angular position sensor and to the motor, said the circuit delivering a motor power supply signal that varies as a function of a position setpoint signal applied to the control circuit and as a function of signals delivered by the angular position sensor. The gearbox comprises a plurality of speed-reducing modules or stages (131 to 134), each comprising a pair of gears mounted to rotate about two parallel axes of rotation (17, 160).

Abstract: A piloting device may include a section for generating control orders for the control surfaces acting on the yaw movement of an aircraft. A central unit determines an instruction to rotate the aircraft, based on control orders, and determines a global moment to be applied to the aircraft about the yaw axis so that the aircraft performs the rotation instruction. The central unit divides the global moment into a sum of elementary moments and computes, for each control surface, the instruction to be applied to its actuator so that the latter generates the associated elementary moment.

Abstract: A system for providing automatic trim control associated with a control surface in an aircraft that utilizes at least two sensors that must agree in direction before trim adjustment is made. Each sensor is provided with a separate an independent controller channel to further enhance fail-safe operation. A trim sensor is placed in the coupling link between a servo and aircraft primary control linkage leading to the associated control surface. A trim sensor is provided that utilizes a spring with a portion disposed laterally with respect to the direction of the force to be measured. An arm is attached to the lateral portion of the spring to effect motion that can be sensed by various sensors including optical, mechanical switch and magnetic sensors.

Abstract: The present invention is structured such that lift control apparatus such as flaps and flaperons operate in association with elevator operation performed by a pilot. Lift control conforming to elevator operation by the pilot is performed.

Abstract: A system for automatically controlling lift augmentation devices of an aircraft, during a phase of take-off by the aircraft, includes a controllable actuator that shifts the lift-augmentation devices, and a control unit for generating control demands to control the actuator to bring the lift-augmentation devices into a defined position. The control unit further includes a detector that detects actual take-off by the air craft and, if appropriate, signals such detection to the control unit. At start of the take-off phase, the lift-augmentation device are brought into a first position, in which they are deployed. The control unit generates, at least when the detector signals the actual take-off, a control demand making it possible to bring the lift-augmentation devices into a second position, in which the lift-augmentation devices are retracted by comparison with the first position.

Abstract: A control system utilizes hydraulic power only to control a fixed-wing aircraft horizontal stabilizer trim control surface with no electrical control devices except for a cockpit pilot/co-pilot initiated 3-position spring-centered toggle trim switch, a backup electrical motor and a standard electronic logic pilot interface. The toggle trim switch controls two solenoid valves which control the operation of a directional control valve. The control system also includes a hydraulic motor, a rate control valve, a blocking-bypass valve, a shutoff valve with an integrated position sensor and a gear set driving an acme threaded output shaft for elevating or lowering the horizontal stabilizer trim control surface on command by the pilot. The control system controls the horizontal stabilizer trim control surface angular displacement rate as a function of angular position by means of a mechanical feedback linkage between the spool of the rate control valve and the horizontal stabilizer trim control surface.

Abstract: The present invention relates to an arrangement and a method in a mechanical control system of an aircraft. From the cockpit of the aircraft control surfaces (1) of the aircraft can be acted upon via the control system by means of at least one control element (3), for example a wheel, a control column or a pair of pedals operatively connected to the control system. A servomotor (5) together with sensors is also connected to the control system. The sensors are designed, when acted upon by the control element (3), to detect the torque exerted on the control element (3) by a force applied thereto, the trim angle of the control element and the angular velocity with which the action occurs. The arrangement comprises a control device (9) which, on the basis of the conditions detected by the sensors, is designed to control the servomotor (5) so that a ratio between the force applied to the control element (3) and the trim angle of the control element assumes a desired, essentially constant value.

Abstract: An actuator system (10, 100, 200) for use in an aircraft control or operating system, comprising a controller (12) operable in response to an input for generating a control signal, and an electrical actuator (20) responsive to the control signal for operating an aircraft flight control surface or other aircraft apparatus. A tab (24; 124) for aerodynamically assisting the electrical actuator is also provided in order to reduce the load on the electrical actuator in use.

Abstract: The present invention relates to a flight control device for an aircraft comprising a control (2) and a means for actuating a controlled member (RP, RQ) to which a command is applied. According to the invention, said device (1) additionally includes a sensor (E) for determining a second value that is representative of the control executed by the aircraft (He) with respect to the control axis, and second means (M2) which determine: as long as the second value is lower than or equal to a reference value, a first trim command that is proportional to the actuation of the control (2); and when the second value is higher than the reference value, in addition to a second trim command, a speed command that is proportional to the additional actuation over and above said second value.

Abstract: An improvement to a method of automatic flight using a flight management system is provided. The flight management system includes a lateral navigation (LNAV) control mode in which the flight management system provides roll commands via a flight director system to the pilot during manual flight or to an autopilot during autoflight to effectuate lateral flight guidance. The improvement includes evaluating the energy state of the airplane during the LNAV control mode, calculating a Thrust Based Roll Limit (TBRL), and providing an appropriately limited bank angle command signal via the flight director system for use by the pilot during manual flight or to the autopilot during autoflight while flying in a thrust-limited condition. The TBRL is calculated as a function of the energy state of the airplane. The use of the TBRL to limit the bank angle command signal avoids an uncommanded change in altitude and/or airspeed.

Abstract: A device for re-trimming a pilot control of an aircraft to keep a flight parameter at a reference value includes a first detection means for detecting the actual position of the pilot control, a calculation unit for determining the direction and magnitude of displacement of the control in order to re-trim the control based on the actual position and reference position, and a display structure for displaying to the pilot a characteristic sign indicating the direction and magnitude of the displacement needed to re-trim the control.

Abstract: An autopilot controller, for controlling a vehicle traveling in a fluid medium, designed for controlling the vehicle in any one of pre stall, stall and post stall regions. The autopilot controller provides an approach for one-to-one mapping between a target command which is received and control surface deflections for effecting the command based on vehicle normal force and vehicle moment response characteristics. In one embodiment, the autopilot controller determines linear functions representative of the vehicle dynamic response for the current operating conditions in order to determine a direction of change of fin deflection for effecting the command. In another embodiment, the autopilot controller controls transverse vehicle accelerations by controlling a body pitch moment generated by a control surface.

Abstract: The apparatus of this invention includes a pilot-induced oscillation (PIO) detector, a PIO compensator and a pilot input modifier. The PIO detector is coupled to receive aircraft state signal including the aircraft's pitch, roll and yaw attitudes. The PIO detector is also coupled to receive pilot control signal generated by the aircraft's pilot by manipulation of flight control instruments. Preferably, the PIO detector includes a feature calculator and a discriminator. Based on the aircraft state signal and the pilot control signal, the feature calculator generates at least one feature signal indicative of whether a PIO or non-PIO condition exists in the aircraft. The feature calculator supplies the feature signal to the discriminator, that uses the feature signal to determine whether or not a PIO condition exists.

Abstract: A cyclic stick system for a helicopter or other aircraft includes memories for holding minimum and maximum speed values, a comparator for comparing an air speed V1 with the minimum and maximum speed values, a router for selecting either the air speed V1 or a quiescent speed, a subtracter for obtaining a difference between the output of the router and a reference speed, another router for selecting between the difference and a signal representing an orientation of the stick, and a switch for selecting times when the air speed V1 is to be stored as the reference speed. On the basis of the air speed V1 and under the control of the device for retrimming said cyclic stick, the system acts on the motorization of stick in order to confer on the helicopter apparent longitudinal static stability in terms of forces on the cyclic stick so that to accelerate (or decelerate) and maintain a new higher (or lower) speed.

Abstract: An underspeed protection system for an aircraft under autopilot control selects a target speed based upon the greater of a minimum maneuver speed and a stick shaker speed. The system then compares a monitored speed to the target speed to produce an error signal. The system also monitors vertical speed to determine if tie aircraft begins to descend. If the error signal due to the underspeed condition causes the aircraft to descend, the system provides a hold zero vertical speed signal in place of the error signal such that the aircraft seeks to maintain its altitude. The hold zero vertical speed signal overrides the underspeed error signal such that the aircraft does not pitch forward to seek an increased speed.

Abstract: In an aircraft under autopilot control, a stall protection system overrides established autopilot parameters in response to a monitored angle of attack for the aircraft exceeding a trigger angle of attack. The trigger angle of attack is established with reference to a stick shaker angle of attack for the aircraft. When the aircraft exceeds the trigger angle of attack, the stall protection system produces an error signal causing the autopilot to seek an angle of attack 1 degree below the stick shaker angle of attack. Initially, the stall protection system provides a boost command to the error signal to accelerate the response of the aircraft to the error signal. The boost command is produced by a driving signal generator that selects one of four driving signals in response to monitored flap angle and impact pressure.

Abstract: An underspeed protection system for an aircraft under autopilot control selects a target speed based upon the greater of a minimum maneuver speed and a stick shaker speed. The system then compares a monitored speed to the target speed to produce an error signal. The system also monitors vertical speed to determine if the aircraft begins to descend. If the error signal due to the underspeed condition causes the aircraft to descend, the system provides a hold zero vertical speed signal in place of the error signal such that the aircraft seeks to maintain its altitude. The hold zero vertical speed signal overrides the underspeed error signal such that the aircraft does not pitch forward to seek an increased speed.