Abstract: The present disclosure provides methods and systems for fault detection for a speed sensing system of a multi-engine rotorcraft. A shaft speed for a first engine and a rotor speed for at least one rotor of the multi-engine rotorcraft are obtained. The shaft speed is compared to the rotor speed. When the shaft speed is greater than the rotor speed, a first fault in the speed sensing system is detected and a first speed sensing system fault signal is issued. When the shaft speed is less than the rotor speed, a determination is made regarding whether the first engine is coupled the at least one rotor based on a fuel flow to the first engine. A second fault in the speed sensing system is detected and a second speed sensing system fault signal is issued responsive to determining that the first engine is coupled to the at least one rotor.

Abstract: An electrically distributed yaw control system for a helicopter having a tailboom includes a plurality of tail rotors rotatably coupled to the tailboom and a flight control computer implementing a tail rotor balancing module. The tail rotor balancing module includes a tail rotor balancing monitoring module configured to monitor one or more parameters of the helicopter and identify a first set of one or more tail rotors in the plurality of tail rotors based on the one or more parameters. The tail rotor balancing module also includes a tail rotor balancing command module configured to modify one or more operating parameters of the first set of tail rotors.

Abstract: A system for inspecting an acoustic treatment applied to the hull of a maritime vessel includes an underwater inspection system including at least one signal emitter configured to emit a signal, at least one receiver configured to receive the signal after the signal has come into contact with the acoustic treatment applied to the hull of the maritime vessel, wherein the signal provides information about the integrity of the acoustic treatment and a processor configured to perform signal processing on the received signal and generate an output comprising information about one or more parameters of the acoustic treatment. The inspection system thereby allows for in-water inspection of acoustic treatments applied to a maritime vessel thus avoiding the need for a costly and time-consuming inspection of such a vessel while in dry-dock.

Abstract: A control system configured to control a deceleration process of an air vehicle which comprises at least one tiltable propulsion unit, each of the at least one tiltable propulsion units is tiltable to provide a thrust whose direction is variable at least between a general vertical thrust vector direction and a general longitudinal thrust vector direction with respect to the air vehicle.

Abstract: Methods and systems for generating flight plans for aerial vehicles based at least in part on communication performance are described. First information about a set of access points may be accessed. Second information about a flight path for an aerial vehicle to move from a first point to a second point may be accessed. A set of directions for the aerial vehicle to move between the first point and the second point may be generated. The set of directions may be generated based at least in part on a communication criterion and the second information about the flight path. The flight path may be updated based at least in part on the set of directions.

Abstract: A vehicle system includes a control unit with a weather module configured to receive weather data and to identify a wind shear zone at a location based on the weather data, the weather module further configured to generate wind shear coordinate data and wind shear characteristic data based on the weather data. The control unit further includes a display module configured to generate display commands based on the wind shear coordinate data and wind shear characteristic data from the weather module. The vehicle system further includes a display device coupled to receive the display commands from the control unit and configured to display a three-dimensional forward perspective view corresponding to a vehicle environment. The display device is further configured to display first wind shear symbology within the view at a position that indicates the location of the wind shear zone.

Abstract: A method of controlling an active aerodynamic element for a vehicle includes determining a target position for the active aerodynamic element from a target aerodynamic force, which may be a given value that is provided based on dynamic conditions of the vehicle. The method actuates the active aerodynamic element to the target position and senses an aerodynamic response characteristic of the active aerodynamic element while actuated to the target position. An estimated applied aerodynamic force is determined from the aerodynamic response characteristic, and is compared to the target aerodynamic force. A force error is determined from the comparison of the estimated applied aerodynamic force and the target aerodynamic force, and a modified position for the active aerodynamic element is determined from the force error and the target aerodynamic force. The active aerodynamic element is actuated to the modified position.

Abstract: A method for determining a state of credibility of measurements of an angle-of-attack sensor of an aircraft is provided. This method includes at least one coherence test between angle-of-attack measurements from said angle-of-attack sensor, and the measurements of a flight characteristic of the aircraft, distinct from the angle-of-attack. The coherence test includes determining an angle-of-attack value from said angle-of-attack sensor, determining said flight characteristic of the aircraft, determining a value of at least one indicator of the coherence of the angle-of-attack value with the value of said flight characteristic, and activating a low state of credibility, in which the measurements of said angle-of-attack sensor are deemed unreliable, or an intermediate state of credibility, in which the measurements from said angle-of-attack sensor are deemed coherent with said flight characteristic, based on the value of said coherence indicator.

Abstract: A system for determining a region of interest for an imaging device based on instrument landing system (ILS) is provided herein. The system may include an imaging device attached to an aircraft; an ILS detector; a computer processor configured in to calculate in a line of sight between said aircraft and a planned touch down point, based on the received ILS signals; a touchdown positioning module executed by the computer processor and configured to calculate a position in a field of view (FOV) of said imaging device which represents the planned touchdown point, based on said line of sight; and a region of interest (ROI) module executed by the computer processor and configured to define a region of interest (ROI) of the imaging device based on said position in said FOV, wherein said computer processor is further configured to apply an image processing operation only to data within said ROI.

Abstract: A control system configured to control a deceleration process of an air vehicle which comprises at least one tiltable propulsion unit, each of the at least one tiltable propulsion units is tiltable to provide a thrust whose direction is variable at least between a general vertical thrust vector direction and a general longitudinal thrust vector direction with respect to the air vehicle.

Abstract: Ground detection of a touch sensitive device is disclosed. The device can detect its grounded state so that poor grounding can be selectively compensated for in touch signals outputted by the device. The device can include one or more components to monitor certain conditions of the device. The device can analyze the monitored conditions to determine the grounding condition of the device. The device can apply a function to compensate its touch signal outputs if the device determines that it is poorly grounded. Conversely, the device can omit the function if the device determines that it is well grounded.

Abstract: A method and apparatus for carrying out a symmetric deflection of the spoilers so as to reduce the lift of the aircraft during a takeoff roll thereby improving lateral and directional control on the ground.

Abstract: A method for reducing yawing motions of an aircraft in-flight, wherein a spoiler adjustment drive of a spoiler and a regulating flap adjustment drive of a regulating flap of the same respective airfoil are adjusted in a time segment in such a way that the motion of the spoiler being adjusted and the motion of the regulating flap of the same airfoil deflect in mutually opposite directions in the time segment. The spoiler and the regulating flap are adjusted on the airfoil, on which the adjusted deflections counteract the respectively occurring yawing motion. Also provided are a computer program product for carrying out this method and an aircraft with a directional stabilization device for carrying out this method.

Abstract: The invention relates to a method of controlling the steering of a steerable portion of an aircraft undercarriage that is fitted both with a steering member for steering the steerable portion and also with at least two angular position sensors for sensing the angular position of the steerable portion in order to deliver respective signals representative of the angular position of the steerable portion, wherein the steering member is controlled by means of servo-control using information that is representative of the angular position of the steerable portion. According to the invention, the information representative of the angular position that is used is a mean of the angular positions sensed by the at least two angular position sensors.

Abstract: The invention relates to a method of managing movement of an aircraft on the ground, the aircraft including at least one left main undercarriage and at least one right main undercarriage, each comprising wheels associated with torque application members for applying torque to the wheels in response to a general setpoint, the general setpoint comprising a longitudinal acceleration setpoint and an angular speed setpoint, the method including the successive steps of braking down the general setpoint into general torque setpoints for generating by the torque application members associated with each of the wheels.

Abstract: Some embodiments relate to a method of controlling a flight of a flight vehicle according to a first mode of operation and changing the mode of operation to a second mode of operation having a different bandwidth than the first mode of operation. Other embodiments relate to a flight-control system for a flight vehicle configured to control a flight of a flight vehicle according to a first mode of operation and to control the flight of the flight vehicle according to a second mode of operation to use less energy than the first mode of operation. Other embodiments relate to a control actuation system configured to control positions of aerodynamic elements in a flight vehicle in response to commands from a guidance system according to a first mode of operation and to change the mode of operation to a second mode of operation having a different bandwidth than the first mode of operation.

Abstract: A method of controlling the yaw attitude of a hybrid helicopter including a fuselage and an additional lift surface provided with first and second half-wings extending from either side of the fuselage, each half-wing being provided with a respective first or second propeller. The hybrid helicopter has a thrust control suitable for modifying the first pitch of the first blades of the first propeller and the second pitch of the second blades of the second propeller by the same amount. The hybrid helicopter includes yaw control elements for generating an original order for modifying the yaw attitude of the hybrid helicopter by increasing the pitch of the blades of one propeller and decreasing the pitch of the blades of the other propeller, the original order is optimized as a function of the position of the thrust control to obtain an optimized yaw control order that is applied to the first and second blades.

Abstract: A method and device control the thrust of a multi-engine aircraft. At least one engine command is determined by a command determining unit, with the determined command commanding each engine that has not failed to deliver a thrust substantially equal to a reduced thrust value. A command application unit is configured to apply the determined command to each aircraft engine that has not failed. The reduced thrust value is determined, by a weight determining device, according to the current weight of the aircraft and according to a reduced thrust value FOEI, which is calculated by a thrust calculation unit.

Abstract: A control system (12) for controlling velocity and attitude of a dynamic system (20) includes a velocity controller configured to receive a desired velocity command and output a velocity error in the form of Euler angle commands to an attitude controller system. The attitude controller system includes a converter (14) configured to receive the Euler angle commands and output a desired quaternion based on the Euler angle commands. The attitude controller system further includes an attitude error generator (16) configured to receive the desired quaternion and an estimated quaternion indicative of an estimated velocity and attitude of the dynamic system (20) and to output attitude errors associated with the dynamic system (20). The attitude controller system also includes an attitude controller (18) configured to receive the attitude errors and to output error commands to the dynamic system (20) based on the attitude errors.

Abstract: An altitude profile representative of the terrain overflown by an aircraft is established. Thereafter, an altitude limit curve which comprises an intersection with the altitude profile upon the engagement of a terrain avoidance maneuver is determined. As soon as there is no longer any intersection of the limit curve with the altitude profile, the terrain avoidance maneuver is interrupted.

Abstract: The fuzzy logic-based control method for helicopters carrying suspended loads utilizes a controller based on fuzzy logic membership distributions of sets of load swing angles. The anti-swing controller is fuzzy-based and has controller outputs that include additional displacements added to the helicopter trajectory in the longitudinal and lateral directions. This simple implementation requires only a small modification to the software of the helicopter position controller. The membership functions govern control parameters that are optimized using a particle swarm algorithm. The rules of the anti-swing controller are derived to mimic the performance of a time-delayed feedback controller. A tracking controller stabilizes the helicopter and tracks the trajectory generated by the anti-swing controller.

Abstract: A device comprises means for computing the air-craft (A) current position, means for determining at least one maximum permitted deviation (E1) around a set position of the flight path of the flight plan according to accuracy and integrity performances of said current position computation and to the restriction of a flight range authorized in a flight corridor (6A, 6B), and a display system (7) for displaying at least one a distance scale (9) on a viewing screen (8), at least one a fixed symbol (10) displaying the current position and two movable pointers (13, 14) displaying the limits of said maximum permitted deviation (E1).

Abstract: A redundant electromechanical actuator is provided that includes first and second electric motors, a common output shaft, first and second drive cable drums, first and second driven cable drums, a first cable, and a second cable. The first and second electric motors are each adapted to be controllably energized to supply a drive torque. The common output shaft is coupled to receive the drive torque supplied from one or both of the motors. The first and second drive cable drums are coupled to the common output shaft to receive the drive torque from one or both of the motors. The first and second driven cable drums are each configured, upon being driven, to rotate. The first cable is coupled between the first drive cable drum and the first driven cable drum, and the second cable is coupled between the second drive cable drum and the second driven cable drum.

Abstract: Methods and systems for a position indicating display system for an aircraft are provided. The system includes a map display unit configured to display a map representative of an area being traversed by the aircraft, and an overlay comprising an own ship depiction, said overlay displayed on the map for a period of time in response to an input from at least one of a user and an aircraft sensor.

Abstract: The invention relates to a method for managing a steering control of an aircraft landing gear provided with at least one wheel steerable by the steering control, comprising the following steps: monitoring one or several taxi parameter(s) (51, 54, 64, 59, 61) of the aircraft to determine whether said aircraft enters a towing situation, setting the steering control in a free steering mode of the steerable wheel, if the aircraft enters the towing situation; activating the steering control so that the steerable wheel has an angular controlled steering by the steering control if the aircraft leaves said towing situation.

Abstract: An objective of the invention, focusing on these issues involved in the use of a small, hobby-type, unmanned helicopter, is to develop an autonomous control system comprising autonomous control systems for a small unmanned helicopter, to be mounted on said small unmanned helicopter; a servo pulse mixing/switching unit; a radio-controlled pulse generator; and autonomous control algorithms that are appropriate for the autonomous control of the aforementioned small unmanned helicopter, thereby providing an autonomous control system that provides autonomous control on the helicopter toward target values.

Abstract: Aerospace vehicle yaw generating systems and associated methods are disclosed herein. One aspect of the invention is directed toward a yaw generating system that can include an aerospace vehicle having a fuselage with a first portion and a second portion. The system can further include a movable control surface coupled to the fuselage and extending generally in a horizontal plane. The control surface can be movable to a deflected position in which the control surface can be positioned to create a flow pattern proximate to the fuselage when the aerospace vehicle is located in a flow field. The flow pattern can be positioned to create a pressure differential between the first portion of the fuselage and the second portion of the fuselage. The first and second portions can be located so that the pressure differential produces a yawing moment on the aerospace vehicle.

Abstract: Method and system for displaying information on one or more aircraft flights, where at least one flight is determined to have at least one atypical flight phase according to specified criteria. A flight parameter trace for an atypical phase is displayed and compared graphically with a group of traces, for the corresponding flight phase and corresponding flight parameter, for flights that do not manifest atypicality in that phase.

Abstract: Aerospace vehicle yaw generating systems and associated methods are disclosed herein. One aspect of the invention is directed toward a yaw generating system that can include an aerospace vehicle having a fuselage with a first portion and a second portion. The system can further include a movable control surface coupled to the fuselage and extending generally in a horizontal plane. The control surface can be movable to a deflected position in which the control surface can be positioned to create a flow pattern proximate to the fuselage when the aerospace vehicle is located in a flow field. The flow pattern can be positioned to create a pressure differential between the first portion of the fuselage and the second portion of the fuselage. The first and second portions can be located so that the pressure differential produces a yawing moment on the aerospace vehicle.

Abstract: Method and system for analyzing aircraft data, including multiple selected flight parameters for a selected phase of a selected flight, and for determining when the selected phase of the selected flight is atypical, when compared with corresponding data for the same phase for other similar flights. A flight signature is computed using continuous-valued and discrete-valued flight parameters for the selected flight parameters and is optionally compared with a statistical distribution of other observed flight signatures, yielding atypicality scores for the same phase for other similar flights. A cluster analysis is optionally applied to the flight signatures to define an optimal collection of clusters. A level of atypicality for a selected flight is estimated, based upon an index associated with the cluster analysis.

Abstract: An airborne glide slope tracking system includes a radio altimeter for producing a signal indicative of the instantaneous altitude of the aircraft and a glide slope error indicator for producing a signal indicative of the angular deviation from glide slope. The two signals are multiplied to produce a glide slope error signal in feet and fed to a summing device directly and through a lead filter to generate a signal for a throttle servo to increase or decrease the thrust of the aircraft. A longitudinal accelerometer signal is then added to a signal indicative of a difference between an aircraft's reference angle of attack and actual angle of attack to produce a signal to a pitch command to provide an angle which will be sustained by the power of the aircraft.

Abstract: The present method and apparatus consists of storing past values of estimated IRU error and using these past values to update the coasting filter when switching from GPS to inertial mode. Through the storage of past IRU error estimates, it is possible to avoid misdirected guidance from an erroneous GPS signal. The MMR and ground station can require up to 6 seconds to identify a failed GLS signal.

Abstract: A measurement-based method to control and optimize the performance of an airborne vehicle. The stability and control of the vehicle is modified to induce a response in the airborne vehicle as reflected by a plurality of response signals. Excitations signals having multi-term sinusoidal waveforms are generated and applied to control signals controlling one or more control effectors. A time domain response of each of the state variables, response signals, and control signals arising from the application of the excitation inputs to the control signals is then measured. These time domain responses are then transformed into frequency domain models. The effectiveness and vehicle stability and control derivatives may then be identified from the frequency domain models of the state variables, response signals, and control signals.

Abstract: An emergency flight control system is disclosed for lateral control using only differential engine thrust modulation of multiengine aircraft having at least two engines laterally displaced to the left and right from the axis of the aircraft in response to a heading angle command .psi..sub.c to be tracked. By continually sensing the heading angle .psi. of the aircraft and computing a heading error signal .psi..sub.e as a function of the difference between the heading angle command .psi..sub.c and the sensed heading angle .psi., a track control signal is developed with compensation as a function of sensed bank angle .PHI., bank angle rate .phi., or roll rate p, yaw rate .tau., and true velocity to produce an aircraft thrust control signal ATC.sub..psi.(L,R).

Abstract: The invention is embodied in a method of integrating kinematics equations for updating a set of vehicle attitude angles of a vehicle using 3-dimensional angular velocities of the vehicle, which includes computing an integrating factor matrix from quantities corresponding to the 3-dimensional angular velocities, computing a total integrated angular rate from the quantities corresponding to a 3-dimensional angular velocities, computing a state transition matrix as a sum of (a) a first complementary function of the total integrated angular rate and (b) the integrating factor matrix multiplied by a second complementary function of the total integrated angular rate, and updating the set of vehicle attitude angles using the state transition matrix.

Abstract: An aircraft landing determination apparatus and method for determining landing status of an aircraft by tracking arriving aircrafts and providing a unique runway assignment therefor. The aircraft landing determination apparatus includes a target processor, a track administrator, and a track reporter. The target processor culls multiple target tracks from a target data transmission provided by a target sensor. The track administrator receives the target tracks, selects at least one arrival track therefrom, and allocates a unique runway assignment to each arrival track. The track reporter receives the arrival track and runway assignment, prepares a track report thereabout, and transmits the track report to a position monitoring apparatus.

Abstract: This optoelectronic device is designed to provide assistance in the piloting of an aircraft under conditions of poor visibility, at the final approach stage, during taxiing and at take-off from a runway equipped with an ILS. The device consists of a collimator displaying a reticule symbolizing a runway centerline that is shown to the pilot during flight as well as on the ground, without any break between the in-flight final approach stage and the ground stage of taxiing on the runway. The device takes account of the pilot's off-centered position in the aircraft cockpit.

Abstract: A method and apparatus for reducing the unwanted sideways motion of an airplane by reducing the lateral side loads and upsets caused by atmospheric turbulence and gusts is disclosed. A rudder modification command that causes the rudder command of the airplane to be changed in a manner that relieves the net force across the vertical stabilizer of the airplane caused by atmospheric turbulence and gusts is produced. More specifically, the pressure differential across opposite sides of the vertical stabilizer is measured and used to produce a rudder deflection value that is roll rate and yaw rate compensated. The compensated deflection value is high-pass filtered with a corner frequency that is twenty-five percent (25%) of the Dutch roll frequency of the airplane. The result is a first rudder deflection value that is subtractively combined with a second rudder deflection value.

Abstract: A method and apparatus for producing a dynamic thrust asymmetry airplane rudder compensation command with no direct thrust measurement is disclosed. An excess thrust estimate based on measured acceleration along the flight path of the airplane and measured vertical speed of the airplane is low pass filtered (11) to attenuate noise from the inertially derived data. The low pass filtered data is further filtered by a washout filter (15) to produce data that is sensitive only to changes in the excess thrust estimate. The input to the washout filter is frozen (13) and the output of the washout filter reduced to zero if the go-around or flare modes of the control system of the airplane are active or if the thrust asymmetry compensation feature of the primary flight computer is armed. The output of the washout filter is also zeroed when both engines are running, or multiplied by either -1 or +1 when one engine is out. Whether the multiplication is by -1 or +1 depends upon which engine is out.

Abstract: An aircraft control system for computing a yaw compensated angle of attack is disclosed. The system includes an angle of attack sensor for generating a signal which represents an aircraft's angle of attack as determined by an angle of attack vane and a transverse accelerometer for generating a signal in response to the lateral acceleration of the aircraft. A summing device sums the two signals to thereby provide a yaw compensated angle of attack signal. The system may also include a rudder deflection sensor for generating a signal which represents the deflection of an aircraft's rudder. The rudder deflection signal is then subtracted from the yaw compensated angle of attack in the summing device. A stall warning system incorporating the yaw compensation system includes a stall warning computer for comparing the yaw compensated angle of attack signal to a predetermined value.

Abstract: Spoilers, mounted on vertical stabilizer tail surfaces of large aircraft release controllable Bernoulli forces to augment rudder and aileron aircraft controls. Data from surface pressure sensors, also mounted on vertical stabilizer tail surface, is computer-interpreted to minimize tail drag and structural flight-stress. The spoilers suitable for this role include parallel lines of controllable height barriers, located on the fore part of each side of symmetrical airfoils of airplane tails.

Abstract: The present invention relates to a process for the automatic control of the control surfaces of an aircraft for the low speed compensation of a lateral path deviation, in which beyond a given deflection threshold (20) of the deflected rudder (17), there is a control of a deflection, proportional to the exceeding of the threshold, of the wing surfaces (13, 14) of one of the two wings (18), said wing being that on the side of the deflected rudder (17), so as to provide a drag supplement to said wing and therefore yaw on the aircraft.Application to the take-off of large civil transport aircraft.

Abstract: A windshear guidance system for guiding an aircraft during a microburst windshear is provided. This system includes a time comparison subassembly, an airspeed comparison subassembly and a pitch command subassembly. The pitch command subassembly includes a calibrated air speed source, latch means connected to the calibrated air speed source, switch means connected to the latch means, and a pitch guidance command unit connected to the switch means.

Abstract: A helicopter engine fuel control anticipates sudden changes in engine power demand during yaw inputs to thereby minimize engine and main rotor speed droop and overspeed during yaw maneuvers. The rate (121,123) of yaw control (107) position change generates (110) a yaw component (104) of a helicopter fuel control (52) fuel command signal (70). The magnitude of the yaw component is also dependant upon the rate of yaw control position change (703). The fuel command signal yaw component (104) is overridden (113,115) when rotor decay rate (209,217) has been arrested during a sharp left hover turn (216); when the yaw component is removing fuel (239) during rotor droop (238); and when the yaw component is adding fuel (228) during rotor overspeed (227).

Abstract: A helicopter flight control system (21) includes a model following control system architecture which automatically provides a coordinating yaw command signal to the helicopter tail rotor to coordinate helicopter flight during a low speed banked turn. The control system processes information from a variety of helicopter sensors (31) in order to provide the coordinating yaw command signal on an output line (72) to the tail rotor (20) of the helicopter.

Abstract: A helicopter flight control system includes a model following control system architecture to control helicopter response about the yaw axis. The control system processes information from a variety of helicopter sensors in order to provide an output command signal to the tail rotor. Yaw control is accomplished by applying a force on a sidearm controller which provides a yaw axis command signal that is used to schedule a desired yaw rate of change. The lateral acceleration and airspeed signals are used to compute a synthesized yaw rate signal which is compared against the desired yaw rate of change, thus allowing the pilot to control lateral acceleration when he applies force on the sidearm controller.

Abstract: When an aircraft is operating in an asymmetrical flight condition, such as during a landing decrab maneuver or during inflight engine out, the pilot is normally required to make precise compensatory roll control inputs to keep the aircraft moving along its original track angle. The pilot is relieved of this compensatory roll control task by providing a control system which automatically banks the aircraft to maintain the flight of the aircraft along the selected track angle while in the asymmetrical flight condition. Further adjustments in the bank angle of the aircraft are made automatically by a track angle feedback control system so that the aircraft remains on the selected track angle even when subject to external disturbances such as crosswind gusts and shear.

Abstract: A method of and apparatus for determining the actual lateral deviation from the centerline of a runway of an aircraft located outside the normal region covered by a glideslope beam is disclosed. The method comprises the steps of scaling a geometrically developed estimated lateral deviation value (Y.sub.estimate) with a compensation factor (K) and passing the result through a complementary filter. The complementary filter produces an estimated lateral deviation rate value (Y.sub.estimate) that is divided into the actual lateral deviation rate value (Y.sub.actual) produced by the inertial navigation system of the aircraft. The result of the division is the compensation factor (K) that is used to scale the geometrically developed estimated lateral deviation value. Preferably, the compensation factor is filtered by a first order lag filter prior to being used to scale the geometrically developed estimated lateral deviation value.

Abstract: A signal selection and fault detection system that provides triple functional redundancy using the outputs (Y.sub.A, Y.sub.B) of two unmonitored primary channels and the output of one or the other of two complementary filters (22, 34) that each receive and process a signal (Y dot IRU) that is derived from a third, independent primary signal source (18) and that is representative of the rate or first derivative of the primary signals (Y.sub.A, Y.sub.B). During non-failure conditions, the system output (Y.sub.CF1) is generated by a normal mode complementary filter 22 from the synthesized rate signal (Y dot IRU) and the selected (10) midvalue one of this output signal (Y.sub.CF1) and the primary signals (Y.sub.A, Y.sub.B). In the event of a suspected failure in one of the primary signals, the output (Y.sub.IRU) of a failure mode complementary filter is substituted (16) for the system output signal (Y.sub.CF1) as the third input (C) to the midvalue selector (10).

Abstract: A system for warning the pilot when the aircraft has reached a high yaw condition. A differential in pressure or angle of attack between the port and starboard sides of the aircraft is measured and converted into a measured sideslip angle. If the measured sideslip angle exceeds a predetermined maximum acceptable sideslip angle, a warning device, such as a rudder pedal shaker, is activated. The maximum acceptable sideslip angle is determined as a function of the ability of the aircraft flight controls to counteract a rolling moment generated by the yaw condition.