Abstract: A servo actuator apparatus has: an actuator which relatively displaces an operation unit on the basis of an input signal E1; a position sensor which detects a relative position of the operation unit and outputs a detection signal D1; an actuator which relatively displaces an operation unit on the basis of an input signal E2; a position sensor which detects a relative position of the operation unit and outputs a detection signal D2; a difference calculation unit which subtracts the detection signal D2 from a command signal C1 supplied from a flight control computer, to output the input signal E1; and a difference calculation unit which subtracts the detection signal D1 from a command signal C2 supplied from a flight control computer, to output the input signal E2, where a positive displacement direction of the operation unit is reverse and series to that of the operation unit, and the body units of the actuators are integrally movable, and the actuators can always operate while mutually monitoring the actual o

Abstract: A system useable in a jet aircraft having installed therein a pressurized oxygen supply which feeds oxygen into the interior of the plane when it flies at high cabin altitudes, the system indicating the changing status of the supply as oxygen is drained therefrom. The system includes a pressure transducer coupled to the supply and means associated with the transducer to determine the lapse rate at which the pressure of the supply is reduced as oxygen is drained therefrom to yield a first signal representing this pressure lapse rate, and to concurrently determine the lapse rate at which the number of liters of oxygen in the supply is reduced as oxygen is drained therefrom to yield a second signal representing the liter lapse rate. These signals are applied to a microprocessor in whose data base is entered the total oxygen inventory of the supply and the oxygen demand of the plane in which the supply is installed.

Abstract: The disclosed method is implemented by a system aboard an aerodyne comprising a processor, a terminal and memories in which there are stored all the information necessary for the carrying out of a flight, the system being connected to the other on-board electronic equipment. In order to define a new flight plan to meet a new situation caused by an event, the method comprises: the interpretation of the event to determine the corrective actions to be undertaken given the context in which the aerodyne is located, the analysis of the possibilities of reconfiguring the flight plan that correspond to the corrective actions, according to predetermined and modifiable criteria, the selection of the flight plan solutions that meet the predetermined criteria, and the presenting of these flight plan solutions in association with the significant parameters that have served as the grounds for their selection.

Abstract: This invention is an aircraft orientation system that locates the earth's horizon, in order to determine the aircraft's attitude in the pitch and roll axes. The invention uses infrared sensors on the aircraft aimed to the sides, front and rear of the aircraft. These sensors are grouped in pairs, and detect the differences in heat levels between the left and right of the aircraft, and forward and behind the aircraft. The invention utilizes the physical condition that the earth's surface is typically measurably warmer than the sky. This difference is used to provide a reference for an aircraft regarding the attitude of the aircraft. In an alternative embodiment, a third pair of sensors detects the aircraft's vertical orientation (inverted/non-inverted) to the ground. The aircraft's attitude is determined by comparing the electrical outputs of infrared sensors, which can be used to provide a feedback to the aircraft control system to keep the aircraft in a neutral attitude.

Abstract: A process and neural network architecture for on-line adjustment of the weights of the neural network in a manner that corrects errors made by a nonlinear controller designed based on a model for the dynamics of a process under control. A computer system is provided for controlling the dynamic output response signal of a nonlinear physical process, where the physical process is represented by a fixed model of the process. The computer system includes a controlled device for responding to the output response signal of the system. The computer system also includes a linear controller for providing a pseudo control signal that is based on the fixed model for the process and provides a second controller, connected to the linear controller, for receiving the pseudo control signal and for providing a modified pseudo control signal to correct for the errors made in modeling the nonlinearities in the process. A response network is also included as part of the computer system.

Abstract: Method and apparatus for a fault tolerant automatic control system for a dynamic device having a sensor and a predetermined control algorithm include structure and steps for receiving a status signal from the sensor. Structure and steps are provided for transforming the sensor status signal and a predetermined reference signal into a linear time invariant coordinate system, generating a sensor estimate in the linear time invariant coordinate system based on the transformed sensor status signal and the transformed reference signal, transforming the sensor estimate into a physical coordinate system, detecting an error in the sensor status signal based on a comparison of the transformed sensor estimate and the sensor status signal, and reconfiguring the predetermined control algorithm based on the detected error.

Abstract: A flight control limiting system for a rotary-wing aircraft which generates servo-commands for at least one of a main rotor and a tail rotor. A set of first commands represents the position of at least one flight control which is movable between extreme positions, and may be controlled by the pilot. A set of second commands are generated by a flying aid system. The flight control limiting system limits the values of the set of second commands when the position of any of the flight controls is close to an extreme position, so that the servo-commands for that particular flight control corresponds to the set of first commands.

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: A flight control device for an aircraft is provided with a control unit OP and a plurality of N instruction generating systems, each of which generates a first command instruction for the control unit OP and a least one of which performs autosurveillance and generates a corresponding surveillance signal. The flight control device is also provided with a plurality of P servocontrol systems, each of which is coupled to each of the instruction generating systems so that each of the servocontrol systems receives one of the first command instructions from each of the N instruction generating systems. Each of the servocontrol systems receives information identifying a particular position of a flight control unit, and each of the servocontrol systems communicates a second command instruction to the control unit OP based on the information.

Abstract: A stabilization system is disclosed in which stabilization pitch and roll signals effective for stabilizing the aircraft are combined with respective pilot provided elevator and aileron position demand signals in accordance with a function which reduces the effects of the stabilizing signals in dependence on increasing values of the respective elevator and aileron position demand signals.

Abstract: An aircraft system intended for gliders and low power-to-weight aircraft, to aid the pilot in locating updrafts associated with the buoyancy of humid air in the convective layer of the atmosphere, namely moist thermals. The system comprises a pair of laterally spaced hygrometers, generally at the wingtips. Four types of hygrometer units are described; resistive, piezoelectric, spectrometric, and capacitive. Humidity data is transmitted to the cockpit by telemetry links, of which four types are also described; by dedicated wiring, by carrier current using wingtip position lights wiring, by modulated infrared beams, and by low-power radio waves. The system monitors the difference of humidity between left and right sensors, and also the rate of change of humidity of one or both sensors as a function of time.

Abstract: A monitoring and control system for a dual-engine helicopter for OEI flight operations includes a parametric indicator operative in response to sensor signals for monitoring an engine gas generator speed parameter N1 during dual-engine and OEI flight operations. The indicator includes a rotatable needle that provides an analog indication of the current value of the N1 parameter. For dual-engine flight operations, the indicator includes a first DE indicia defining a normal operating range for the N1 parameter, a second DE indicia that defines a precautionary operating range for the N1 parameter, and a DE indicium that defines a take off power limit for the N1 parameter.

Abstract: A system and method for conducting one engine inoperative (OEI) flight procedures training in a dual-engine helicopter includes a multi-function OEI training switch that is operative to initiate OEI flight procedures training by selecting one of said engines as the single operative engine for OEI flight procedures training, and a training function module that is: (1) operative to establish suppressed 30-second, 2-minute, and maximum continuous OEI operating limits for selected engine operating parameters to limit the actual power provided by said powerplant system during OEI flight procedures training; and (2) operative to generate biasing factors to control the operation of the parametric indicators for the selected engine operating parameters during OEI flight procedures training.

Abstract: An improved method and apparatus for determining the amount of turn coordination gain in an aircraft yaw damper during a turn maneuver is disclosed. The yaw damper includes inputs from the inertial reference units of the aircraft and also from the flight management computer of the aircraft. The flight management computer provides to the yaw damper a signal indicative of the position of the flaps of the aircraft. The yaw damper includes a turn coordination gain box that receives the flap position signal and outputs a turn coordination gain value, dependent upon the flap position. Generally, the turn coordination gain value increases as the flap position is more extended. The precise turn coordination gain value for each flap position is dependent upon the particular aerodynamic characteristics of the aircraft.

Abstract: Method and apparatus for a fault tolerant automatic control system for a dynamic device having a sensor and a predetermined control algorithm include structure and steps for receiving a status signal from the sensor. Structure and steps are provided for transforming the sensor status signal and a predetermined reference signal into a linear time invariant coordinate system, generating a sensor estimate in the linear time invariant coordinate system based on the transformed sensor status signal and the transformed reference signal, transforming the sensor estimate into a physical coordinate system, detecting an error in the sensor status signal based on a comparison of the transformed sensor estimate and the sensor status signal, and reconfiguring the predetermined control algorithm based on the detected error.

Abstract: First and second pylon flaps are located on aft portions of the first and second jet engine pylons, respectively. The first and second pylon flaps are moveable between retracted pylon flap angular orientations and extended pylon flap angular orientations. The first and second pylon flaps can assume only fully retracted angular orientations and fully extended angular orientations. When the T-tailed aircraft enters into a deep stall condition, the pilot will more than likely push the control column forward to its maximum position. This full forward movement of the control column commands the first and second pylon flaps to assume fully extended angular orientations. A first pylon flap controller controls the first pylon flap to move between the fully extended and fully retracted angular orientations, and a second pylon flap controller controls the second pylon flap to move between the fully extended and fully retracted angular orientations.

Abstract: A transportable three-dimensional calibration wind tunnel system consisting of a small wind tunnel portion for creating a three-dimensional airflow having a suitable wind velocity, and a two-axis rotational deformation device portion for causing the wind tunnel portion to effect a conical motion with a nozzle blow port being in close proximity to an apex to suitably change a flow angle. The two-axis rotational deformation device includes a .beta.-angle rotational deformation device having a .beta.-angle deformation base supported to be rotated horizontally, and an .alpha.-angle rotational deformation device having an .alpha.-angle deformation base supported to be rotated vertically. A rotational axis of the .alpha.-angle deformation base, a rotational axis of the .beta.-angle deformation base and a center axis of the small wind tunnel portion are arranged so that they intersect at a point.

Abstract: The actuation system of the present invention is based upon sensing relatively small delta errors over relative short strokes as opposed to absolute travel error.

Abstract: This invention is a method and apparatus for determining whether or not auxiliary airfoils on an aircraft wing are skewed or lost. It employs either of two types of systems and their associated computer monitor and control requirements. One system utilizes a cable and a spring-loaded mechanism with a cable displacement position sensor. The second system utilizes a drive system position sensor, proximity sensors and segmented proximity targets. These two systems are capable of skew and loss detection for adjacent or individual auxiliary airfoil arrangements. A computer electronic unit is used to perform logic functions to verify the authenticity of sensor signals, and, if appropriate, to shut down the drive system and to compute new flight control parameters including those relating to stall speed and the stick shaker, while alerting the flight crew.

Abstract: This invention is a method and apparatus for determining whether or not auxiliary airfoils on an aircraft wing are skewed or lost. It employs either of two types of systems and their associated computer monitor and control requirements. One system utilizes a cable and a spring-loaded mechanism with a cable displacement position sensor. The second system utilizes a drive system position sensor, proximity sensors and segmented proximity targets. These two systems are capable of skew and loss detection for adjacent or individual auxiliary airfoil arrangements. A computer electronic unit is used to perform logic functions to verify the authenticity of sensor signals, and, if appropriate, to shut down the drive system and to compute new flight control parameters including those relating to stall speed and the stick shaker, while alerting the flight crew.

Abstract: A fault tolerant controller arrangement for electric motor driven apparatus is implemented by a plurality of control channels, each of which includes a motor and associated drive circuitry. The implementation is such that the rating of each of the plurality of channels is reduced and all channels are operated such that the load is shared between all operating channels. The arrangement is such that all channels share the load dynamically and statically and that each channel in a system of n channels supplies .sup.1 n of the required load driving force in the same direction.

Abstract: Method and apparatus for a fault tolerant automatic control system for a dynamic device having a sensor and a predetermined control algorithm include structure and steps for receiving a status signal from the sensor. Structure and steps are provided for transforming the sensor status signal and a predetermined reference signal into a linear time invariant coordinate system, generating a sensor estimate in the linear time invariant coordinate system based on the transformed sensor status signal and the transformed reference signal, transforming the sensor estimate into a physical coordinate system, detecting an error in the sensor status signal based on a comparison of the transformed sensor estimate and the sensor status signal, and reconfiguring the predetermined control algorithm based on the detected error.

Abstract: A method and apparatus for detecting oscillatory phenomena indicative of airflow separation or sensor common mode oscillatory failure is disclosed. The oscillations of a differential pressure transducer that senses the pressure difference on the opposite sides of an airfoil are bandpass filtered to remove oscillations lying outside of a band of interest. Oscillation peaks lying within the passband that exceed a positive or negative threshold produce pulses that are counted. The counter value is decremented each time an alternating peak is not detected within one-half cycle of a minimum frequency. If the allowable count threshold is exceeded, a latch is set. Setting the latch produces a command that can be used to inhibit the operation of systems that rely on data produced by the differential pressure transducer. An alternative path senses the position of a control element (e.g.

Abstract: Arcs of segmented spoilers generate a ring on intake airfoil surface of cantilever-suspended jet engines. Strain and temperature sensors on sides of a cantilever connecting structure feed data to a stress-limiting computer whose output communicates with actuator motors of spoiler segments. This stress-feedback network selects and actuates segmented spoiler surfaces to release radial forces that normally are balanced within the engine intake zone. The stress-limiting computer also takes autopilot data of anticipated attitude change patterns and forecasts anticipated strain patterns of an aircraft. An output communication is integrated with an electric power conditioning that actuates motors to position the segmented spoiler barriers.

Abstract: The device embodying the invention uses at least one digital computer receiving information from a set of sensors and controlling actuators acting on the flight control surfaces, the computer comprising an autonomous means for monitoring the service quality thereof, for disconnecting the device and for recentering the actuators subsequent to detection of a failure. It applies notably to the automatic piloting of a helicopter.

Abstract: An unmanned, remotely controlled microwave-powered aircraft for use as a stationary communications platform. The aircraft is generally a flying wing with a large, flat inner wing having a rectenna on the underside. Rectennas may also be provided on the underside of the wings, the combined output from the rectenna being used to provide power to two electric motors housed within torpedo-shaped nacelles which drive two rear propellers. The rectenna converts a microwave signal at 35 GHz generated by a ground power station utilizing dual gyrotrons and a 34-meter diameter antenna dish. The aircraft has a preferred airfoil cross section throughout and is constructed of lightweight but strong materials in order to provide an enhanced flying time of several months. A power management and distribution system manages the DC power produced by the rectenna to supply power to the flight controls, propulsion system and payload.

Abstract: A redundant, multi-channel fly-by-wire control system for use in an aircraft is disclosed. The system includes a left flight control channel (60), a center flight control channel (80) and a right flight control channel (90). Each control channel is capable of flying the aircraft in the event the other two channels fail. Included within each control channel is an actuator controller electronics unit (ACE (62, 82, 92)), which transmits a series of pilot control transducer signals to a set of primary flight computers (64, 84, and 94). The primary flight computers combine the pilot control transducer signals with data obtained from an air data and inertial reference unit (145) to generate a set of flight surface commands. Each ACE selects a set of flight surface commands to control the movement of a set of flight control surfaces on the aircraft.

Abstract: A system for calibrating the sensors of a redundant sensor system by calculating a first set of error signals based on differences between a combination of the outputs and known values and calculating a second set of error signals based on differences between the parity among preselected sensors and a predetermined value and calculating, from the first and second set of error signals, a plurality of calibration signals to be added to the output of the individual sensors which are in error so as to correct their outputs.

Abstract: A method and apparatus for measuring the lift generated by airfoils of an aircraft. This real-time analysis is accomplished by measuring a differential pressure between the upper and lower lift surfaces of the airfoils. The system comprises the steps of: a) measuring an actual differential pressure between the upper and lower lift surfaces for a given aircraft speed, b) transmitting this actual differential pressure measurement to a computer, c) comparing the actual differential pressure measurement with an optimal pressure differential for the same aircraft speed. The apparatus comprises a fixed array of differential pressure sensor mechanisms for measuring actual pressure differentials and a computer for comparing optimal differential pressure measurements to the actual differential pressure measurements. Each sensor mechanism preferably contains a piezoelectric sensor that communicates with the upper and lower lift surfaces.

Abstract: A system (1) responds to inputs (2) in parallel with a state estimator (4). The controller (6) provides the inputs (2) in response to ordered state inputs (14) and state estimator inputs. An external disturbance vector (f) affects the system (1) so that the observed values of the state variables are not the same as the estimated values. The estimated and measured values are compared (8) to give an error signal (9). The error values are fed to an Ordered State Corrector (11) which computes the error that the order state variable would have in the steady and applies it in the opposite sense to increment the input ordered state (14). If more than one state variable is to be corrected, the correction is applied directly to the control inputs using Multi-State Correction, in which the steady state errors are used to find the control inputs which would cause the errors and then a correction is applied to the control inputs.

Abstract: A hydraulic control system for aircraft is disclosed, which has a flight maneuver system mounted on the aircraft and provided with a maneuver signal input and output system for controlling the aircraft depending on flight conditions. The system includes a differentiating circuit responsive to a first output signal for calculating an actuating speed of a servo actuator and for generating a speed signal Vc, a load calculating circuit responsive to the load signal for computing the load on the actuator and for producing a load signal, a pressure calculating circuit responsive to the speed and load signals for computing a pressure Ps to the actuator and for producing a pressure signal, and a setting circuit responsive to the pressure signal for deciding an optimum pressure of the hydraulic pump so as to precisely control the flight surface without energy loss. The pressure Ps is calculated by the following formula:Ps=E.multidot.(Vc/Vm).sup.2 .multidot.

Abstract: In an aircraft, there is included a Flight Management System (FMS), Autopilot and Autothrottle for controlling the aircraft. The apparatus has a plurality of outputs for definition of the real-time targets, controlled by the Autopilot and Autothrottle, to guide the vertical position of the aircraft to a desired vertical position along the desired vertical flightplan (or profile) according to a set of operational procedures. The FMS includes an apparatus that comprises an element which provides information denoting actual vertical position of the aircraft, and an element which generates information specifying the desired vertical position of the aircraft along the predetermined desired flightplan.

Abstract: A system for improving the flutter behavior of an aircraft provided with a rudder, said system comprising an accelerometer disposed in the rear portion of said aircraft, filter means, and means for adding a correction signal to the piloting signal. According to the invention said filter means comprise a bandpass filter, a phase matching filter, a variable gain amplifier and a switch transmitting the amplified signal to said adder means whenever the air speed of the aircraft is greater than a first speed threshold, said phase matching filter being such that after passing through the control means for the rudder said signal as amplified by said amplifier is retarded by 90.degree. relative to said signal generated by said accelerometer.

Abstract: A control system, having an input command signal and driving a prime mover, is provided with variable magnitude and/or rate input signal limiting that is adjusted by maps 68,70,72,74 which set the limits for a variable magnitude limiter 14 and a variable rate limiter 22 as a function of the prime mover rate. The limiting is adjusted dynamically as the prime mover rate changes. When the prime mover enters an undesirable operating condition, the maps only limit the command signal in the direction of such condition. The adjustable limiting is enabled or disabled based on the position, rate, and/or acceleration of the prime mover.

Abstract: An aircraft rudder command system for allowing a pilot to directly input a sideslip command for yaw-axis control through use of the rudder pedals is disclosed. The aircraft rudder command system includes a pedal input device for receiving a pedal input signal indicative of pilot rudder pedal input, a signal-receiving device for receiving feedback signals indicative of the current state of aircraft operation, a command generator system responsive to the pedal input signal and at least one of the feedback signals for generating a sideslip angle command, command limiting means for generating a limited sideslip angle command, and a feedback control system for transmitting a sideslip control rudder command to a rudder actuation system. The rudder actuation system causes the rudder to move in such a manner so as to produce an actual aircraft sideslip angle which follows the limited sideslip angle command. The aircraft rudder command system may also include a sideslip estimator.

Abstract: A control system and method for controlling a flight control member employs a hydraulic control valve which is responsive to input from the pilot in a manner to start and stop a hydraulic motor which in turn starts and stops movement of the control member so that the control member assumes the desired flight control position.

Abstract: Accelerations due to excitation of the natural modes of an aircraft's body are suppressed by an active suppression system. Dedicated accelerometers are positioned in the aircraft at optimal locations for sensing modal induced lateral accelerations. The accelerometer produced signals are processed through control logic which, in response thereto, and in response to aircraft velocity and altitude related signals produces output control signals. The control signals effect rudder deployment creating forces to suppress the natural mode induced accelerations.

Abstract: Aircraft/engine performance is determined by the method of the present invention, wherein the performance characteristics are defined by learned performance parameters. The method includes the steps of learning the performance parameters of the aircraft/engine combination, for a given flight, using previously learned performance parameters and current observations. Then, based on the performance parameters learned from the learning step, a prediction is made of the performance of the aircraft/engine combination for subsequent flights.

Abstract: A throttle controller includes a manual mode where engine output is a function of throttle lever angle, and an alternatively selectable Speed Hold/Thrust Hold mode which is initiated when the throttle is placed in a center Hold position. Throttle operation during the Speed Hold mode is governed by (i) a selected speed entered at a mode control panel, or (ii) a speed existing when the lever was placed in the Hold detent, given that no speed is entered and selected at that mode control panel or (iii) a selected speed/thrust designated in a flight plan entered into the aircraft's flight management computer. During the Thrust Hold mode, the throttle controller maintains the level of thrust which existed when the throttle lever was placed in the Hold detent. When out of the Hold detent, the controller operates to control output thrust as a function of Throttle Lever Angle (TLA).

Abstract: An attitude control system for controlling a plurality of attitude control devices, which cooperate to establish the attitude of a vehicle includes redundant control units for outputting attitude control signals to each of the attitude control devices. Additionally, a plurality of discreet power servo units are provided which are located in proximity to and operatively connected to the plurality of attitude control devices. The redundant control units and the plurality of discreet power servo units are connected electrically and/or optically. Each of the discreet power servo units includes an adder for adding an attitude control signal and a positional feed back signal, a rotational drive source for providing a rotational motion in response to the output from the adder, a fluid pressure circuit for hydraulically driving an attitude control device coupled thereto, and a sensor for sensing the position of the attitude control device coupled thereto and for providing to the adder the positive feed back signal.

Abstract: A servo control apparatus comprising: a unit for generating a required position value; an actuator adapted to drive a load and having its maximum displacement in one direction and its minimum displacement in the other direction; a drive unit for driving the actuator; a position detecting unit for detecting a displacement of the actuator; and a control unit for correcting the required position value on basis of the maximum and minimum displacements of the actuator and for computing a deviation between an output of the position detecting unit and the corrected required position value; the load being controlled by the deviation.

Abstract: An aircraft flight simulator system adaptable to an actual inoperative aircraft not in flight provides instrument displays driven by a computer flight simulator program interacting with inputs from sensors attached to the actual aircraft control system and overlay switches. The computer may also drive visual displays placed in the cockpit windshield and side windows, as well as provide audio output and output for the operation of aircraft jacks to simulate inflight motion. All equipment may be relatively quickly and easily installed or removed so as to make maximum use of an otherwise inoperative aircraft on the ground, and by the use of computer program modifications, may be used to simulate a wide variety of aircraft. Realistic control handling is provided by feedback mechanisms attached to various aircraft control surfaces and the control column and which are coupled to the computer.

Abstract: An integrated air data system for use on an aircraft. Airframe air data units (18) are connected to receive a plurality of redundant total pressure sensor inputs (20,22,24), a plurality of redundant static pressure sensor inputs (26,28,30), and a total temperature sensor input (32). Similar redundant sensor inputs are provided a first and second channel in each of two electronic engine controls (34,36). Bidirectional data buses (16) connect the airframe air data system to the electronic engine control system, so that the same sensor signals are available for each of the systems. The airframe air data unit and the primary electronic engine control channel having priority to control each engine are independently operative to select a preferred total pressure, static pressure, and total temperature value according to a common programmed logic scheme. The selection logic gives priority to the use of the same value for common parameters in each of the systems, where possible.

Abstract: System for controlling roll and yaw of an aircraft. The system includes a device capable of elaborating, from electric signals respectively representative of the position of a first voluntary actuation member, the rolling speed, the attitude, the yaw speed, the sideslip and of the position of a second voluntary actuation member, a single electrical order for roll control formed by a linear combination of the electric signals. There is a device capable of elaborating, from the same electric signals, an electrical order for yaw control formed by a linear combination of the electric signals. There is also a device which makes it possible to combine the electrical order for yaw control and a mechanical order coming directly from the second voluntary actuation member.

Abstract: A plurality of two-axis rate gyros are arranged in relation to an aircraft-fixed coordinate system such that they redundantly supply angular rate information. A plurality of accelerometers supply correspondingly redundant acceleration information. First signal processing means comprise means for failure detection and elimination, such that angular rate and acceleration information subjected to error are eliminated. The angular rate and acceleration information thus cleared from errors supply stabilization signals for the autopilot. Second signal processing means obtain the angular rate and acceleration information cleared from errors and integrated if required, and therefrom supply heading and attitude reference signals.

Abstract: A monitoring system for detecting a fault condition in any one of multiple power control actuator units used for positioning an aircraft's control surface includes sensors for producing sense signals corresponding to the load on each actuator unit under a given control surface loading condition. The sense signals are compared with a reference level signal corresponding to a maximum allowable load. An indication is produced in response to the load on any actuator unit exceeding its allowed value.

Abstract: A backup software program is installed in an isolated portion in the memory of at least one of redundant computers. The backup program performs basically the same functions as the prime program but is dissimilarly programmed to prevent a common software error. Switchover to the backup program occurs either automatically in response to monitors, or manually by the operator (i.e. pilot) when he detects an anomaly.

Abstract: A flight control system receiving input flight control signals from a flight control computer to manipulate a flight control surface. The dual actuator has an actuating rod connected to the flight control surface, and a pair of pistons are mounted on the actuating rod in respective first and second cylinders. The flow of hydraulic fluid into and out of the first and second cylinders is controlled by the first and second electrohydraulic servovalves, respectively. The electrohydraulic servovalves are, in turn, operated by respective signals applied to a pair of actuating coils. The system includes two controllers that receive the flight control signals and apply respective output signals to the four coils of the two electrohydraulic servovalves. In one embodiment, the controllers each have a primary channel and a secondary channel which are connected to respective actuating coils of the first electrohydraulic servovalve.

Abstract: An arrangement for controlling the position of a movable member, especially an aircraft control surface member, includes an operational amplifier and a capacitor connected in parallel to the amplifier. The control arrangement further includes a detector which detects the restoration of the supply of electric power following its interruption and generates a limited-duration pulse. This pulse is then used to close a switch, which is interposed between an input terminal of the capacitor and the ground, for the duration of this pulse, so that the capacitor is charged much more rapidly than otherwise. The output signal of the combination of the operational amplifier and the capacitor is supplied to an actuator for the movable member so that the extent of excursion of the movable member out of its desired position after the electric power supply has been restored is minimized.

Abstract: An apparatus and method for compensating for asymmetrically produced total engine thrust caused by an engine failure in a multiengine aircraft. The presence of asymmetric total thrust is detected by monitoring aircraft performance parameters, including engine manifold pressures, airspeed, roll angle and yaw. When an engine failure resulting in the production of asymmetric thrust is detected at air speeds below a minimum controllable air speed for the aircraft, and during large aircraft bank angles, the power output from an operating engine is reduced to regain and maintain controllable flight conditions. Limiting the adverse yaw produced by an engine failure by reducing power output from operative engines reduces the tendency of the aircraft to roll into the inoperative engine, hence the aircraft is halted.