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: An aircraft autopilot system provides a wind correction angle which is added to the desired course to provide commanded heading, the wind correction angle being generated as a proportional and integral function of ground track error. When the aircraft is off its desired course, it can be returned to the desired course by means of a course intercept angle, added to the desired course, to provide a corrected commanded course, the course intercept angle being a proportional function of the perpendicular error (the lateral distance of the aircraft from the original course) suitably limited to some angle within 90.degree. of the desired course. The course correction angle may be generated as a proportion of the perpendicular course error determined by the degree of importance of accuracy, and may be switched over to a course directly toward a waypoint in dependence upon the importance of efficiency.

Abstract: In a rotary wing aircraft, longitudinal groundspeed error and lateral groundspeed error are converted to earth coordinates, and the result scaled and integrated, and retransformed into aircraft coordinates for application to a pitch attitude control system and a roll attitude control system. The invention thus instantly transfers wind trim between the longitudinal and lateral channels, whenever a heading change is undertaken.

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: 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: An apparatus and method for detecting vertical gusts of wind on board an aircraft in cruising flight is disclosed. The method includes the steps of: calculating an absolute value (.vertline..alpha.-.theta..vertline.) of a difference between a pair of first differentials with respect to time (.alpha. and .theta.) of a current incidence .alpha. and a current pitch attitude .theta. of the aircraft, comparing the absolute value to an upper threshold (Ss), comparing a current Mach number (M) of the aircraft to a Mach number threshold (Mo), and generating an electrical signal that represents presence of a vertical gust of wind when the absolute value is above the upper threshold, when the current Mach number is above the Mach number threshold, and when aerodynamic flaps and slats of the aircraft are in a clean configuration.

Abstract: A flight control system comprises trailing edge airfoil members pivotally mounted at the trailing edge of each wing of an aircraft and selectively movable on a laterally extending axis between raised and lowered positions for imparting rolling motion to the aircraft. Leading edge airfoil members mounted adjacent the leading edge on each of the wings are movable transversely of the leading edge between a retracted position generally coextensive with the leading edge and an extended position protruding from the leading edge for imparting countervailing aerodynamic forces on the wings to counteract the effects of aeroelastic wing deformation caused in response to operation of the trailing edge airfoil members by increasing the angle of attack of the wing and lift producing surface area of the wing. A pair of actuator mechanisms are mounted on a wing box member at laterally spaced locations for moving the leading edge airfoil members between the retracted and extended positions.

Abstract: An avionic system (102) in which engine thrust rating data is transmitted from electronic engine controllers (108a and 108b) to flight management computer/thrust management computers (120a and 120b) via digital data buses (110a, 110c, and 114) is disclosed. The flight management computer/thrust management computers (120a and 120b) select the proper data set for the specified engine thrust rating upon power-up. If the thrust rating changes, the flight management computer/thrust management computers (120a and 120b) select a new data set corresponding to the new thrust rating as received over the digital data bus (110a, 110c, and 114). This is accomplished without the need to change aircraft wiring. The flight management computer/thrust management computers (120a and 120b) store the thrust rating in nonvolatile memory (126), allowing the flight management computer/thrust management computers (120a and 120b) to use the stored value to initialize relevant settings according to current engine thrust rating.

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: An apparatus (700) for inhibiting operation of an electronic device (702) during take-off and landing of an aircraft (802) has a sensor (704) that measures a lateral acceleration. A control circuit (706) is coupled to the sensor (704) and has an output (708) coupled to the electronic device (702).

Abstract: The present invention relates to a method for three-dimensional flight control based generally upon measuring and comparing actual air pressures at or near various surfaces of an aircraft during flight. Sensors are provided for measuring air pressure acting on the aircraft surface. The method includes measuring air pressure differentials between two or more sensors to evaluate certain critical flight parameters, such as the actual lift being produced, the air direction and speed relative to the aircraft, the air density, and the aircraft position and trajectory. The actual and comparative data provide information about the present flight conditions and performance if the aircraft, such as whether there is ice formed or forming on the wings, the direction and approach of wind shear, whether a stall is approaching, etc. The information can be evaluated by a computer, the aircraft's automatic flight control system ("AFCS"), or flight crew so that appropriate flight control measures can be taken.

Abstract: A device for controlling a rudder of an aircraft has at least two servo-systems each including at least one electric control input. At least one of the servo-systems is a mixed servo-system that also has a mechanical control input. When an engine fault occurs, an electric control system causes at least two of the servo-systems to simultaneously operate the rudder. When the electric control system is not operational due to an electric fault, the mixed servo-system operates the rudder based on control signals from the mechanical control input.

Abstract: An aircraft control system for controlling an aircraft, particularly a free wing aircraft in low speed or hover regimes. An air speed sensor measures air speed of the aircraft and outputs an air speed signal to a control processor which processes the air speed signal with a speed control input signal. A control actuator actuates an aircraft control surface in response to the control surface control signal. The air speed sensor may include a shaft mounted impeller located in an airstream of the aircraft. A rotational speed sensor, coupled to the impeller, measures a rotational speed of the impeller and outputs a rotational speed signal as the air speed signal. In an alternative embodiment, the air speed sensor may include a vane located in an airstream of the aircraft and deflected in response to air flow in the airstream.

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: 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: A model-following aircraft control system in which a roll rate command in Euler coordinates is integrated to provide a bank angle command, which has actual bank angle subtracted therefrom to provide a command error that is converted back to aircraft body coordinates for use, so long as the pitch attitude of the aircraft does not approach zenith or nadir. But, while the absolute value of the pitch attitude exceeds 85.degree. the last error generated before exceeding 85.degree. is converted to body coordinates for use by the aircraft, and the initial integrated value of attitude command, for use when the pitch angle reverts below 85.degree., is formed as the sum of said last error and the actual attitude angle of the aircraft.

Abstract: A system for stabilizing an aircraft includes sensors positioned and disposed at predetermined locations on the aircraft and structured for sensing external forces acting on the aircraft at the sensor locations as a result of rapid changes in atmospheric conditions such as those associated with air turbulence and wind shear. A computer processor receives data from the sensors and activates one or more thrust generators positioned at predetermined locations of the aircraft to counteract the external forces and maintain the desired attitude and stability of the aircraft.

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: 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: A velocity command system is provided with a velocity stabilization mode wherein aircraft flight path referenced velocities are determined with respect to an inertial frame of reference, the flight path referenced velocities are held constant during pilot commanded yaw maneuvers so that the aircraft maintains a fixed inertial referenced flight path regardless of the pointing direction of the aircraft. Velocity control with respect to an inertial frame of reference is accomplished by controlling the aircraft flight path based on aircraft body referenced commanded lateral and longitudinal acceleration and based on aircraft body referenced lateral and longitudinal centrifugal acceleration. Operation in the velocity stabilization mode is provided in response to the manual activation of the velocity stabilization mode by the pilot, provided that the aircraft is already operating in the ground speed mode and the aircraft is not in a coordinated turn.

Abstract: A flight critical computer system for an aircraft includes dual independent lanes having two processors in each lane. The first lane has a primary processor and a redundant processor and provides a first command signal. The second lane includes a primary processor and a redundant processor and provides a second command signal. A first monitor compares the primary processor of the first lane with the primary processor of the second lane and generates first comparison signals as a function of disagreement therebetween. A second monitor compares the output signals of the redundant processor of the second lane and the primary processor of the first lane and generates second comparison signals as a function of disagreement therebetween. A third monitor compares the primary processor of the second lane with the redundant processor of the first lane and generates third comparison signals as a function of disagreement therebetween.

Abstract: An in-field method for correcting the thermal bias error calibration of the gyros of a strapdown inertial navigation system. The method is begun after initial alignment while the aircraft remains parked with the inertial navigation system switched to navigation mode. Measurements are made of navigation system outputs and of gyro temperatures during this data collection period. A Kalman filter processes the navigation system outputs during this time to generate estimates of gyro bias error that are associated with the corresponding gyro temperature measurements. Heading error correcting is performed after the extended alignment data collection period as the aircraft taxis prior to takeoff. The gyro bias error-versus-temperature data acquired, along with the heading error corrections, are employed to recalibrate the existing thermal model of gyro bias error by means of an interpolation process that employs variance estimates as weighting factors.

Abstract: A method and apparatus are provided for addressing the effect of centripetal acceleration upon estimates of cross-track velocity, for determination of east gyro bias error, generated with a taxiing aircraft. After initial estimates of crab angle, ratio of crab angle to centripetal acceleration and lever arm are provided, velocity, heading angle and heading angle rate are observed as the aircraft taxis. An estimated value of centripetal acceleration is taken as the product of heading angle rate and heading velocity. Cross-track velocity is computed from cross-heading velocity and this is integrated to generate cross-track position. A Kalman filter generates various gains, including one associated with the ratio of crab angle to centripetal acceleration, for error allocation.

Abstract: A device for flight control of a model plane has a plurality of photo cell sensors, preferably, located on the wing tips, nose and tail and connected to a logic circuit inside the body of the plane. The logic circuit, computes the location of the model plane based on the sum and differential output of the sensors, and outputs electronic signals that drive the servo motors controlling the plane. The logic circuit can be calibrated to individually calibrate the rudder, elevator and the throttle electronic signal output, based on the light received by the photo cell sensors. Inside an enclosed space, a focussed beam of light is directed at the model plane. The logic circuit, receives the sensors' output, and using them computes the location of the model plane within the beam of light. The focused beam of light is swivel attached to a motorized control for swiveling the reflector housing in a predetermined pattern of a closed path racetrack holding pattern and a closed path figure eight holding pattern.

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 sensor has a sensing region that responds to a surface property by producing an output signal. The sensor has a spatially distributed shape or sensitivity so that the output decreases away from a central part of the sensor, and thus the outputs of plural sensors combined have finite spatial transform as well as high roll off in spatial frequency. Preferably the output decreases to zero at edges of the sensor, and conditions of continuity or vanishing may be imposed on first or higher order derivatives. An edge sensor suitable for mounting at the edge of the structure has its weight function obtained by processes of reflecting and inverting the weight function at an edge. A sensor system employs plural such sensors and edge sensors to produce bounded spatial transfer functions for characterizing the structure. Embodiments of piezeoelectric, resistive, capacitive and thermal sensors are described.

Abstract: Process for flying an aircraft in the "elevator speed maintenance mode" during altitude changes, in which for acquiring and/or maintaining a nominal speed it is supplied to the flight computer at the same time as the instantaneous speed of the aircraft, said computer producing the elevator control instruction, characterized in that a nominal speed is also simultaneously transmitted in continuous manner to the automatic thrust control member of the engines and in that the process is performed during two successive sequences, namely:a first sequence during which the elevator only receives a nose up instruction (on climbing) or a dive instruction (on descending) and the engines move to full thrust in the first case or to idling speed in the second anda second sequence, following the first, during which the real nominal speed is supplied to the elevator instruction computer and said same nominal speed increased (on climbing) or decreased (on descending) by a margin is supplied to the automatic thrust control memb

Abstract: An adaptive sheet structure with distributed strain actuators is controlled by a dynamic compensator that implements multiple input, multiple output control laws derived by model-based, e.g., Linear Quadratic Gaussian (LQG) control methodologies. An adaptive lifting surface is controlled for maneuver enhancement, flutter and vibration suppression and gust and load alleviation with piezoceramic elements located within, or enclosed by sheets of composite material at a particular height above the structure's neutral axis. Sensors detect the amplitudes of lower order structural modes, and distributed actuators drive or damp these and other modes. The controller is constructed from an experimental and theoretical model using conventional control software, with a number of event recognition patterns and control algorithms programmed for regulating the surface to avoid instabilities.

Abstract: A control system includes a dual actuator (2) and primary and secondary controllers (20, 22), each of which has two control channels. In a normal mode of operation, the primary controller (20) controls both valves (32, 38) of the actuator (2). Each primary channel generates a model signal and a monitoring signal corresponding to expected and actual operation of the actuator, respectively. The two signals from each channel are communicated to the other channel. Each channel monitors itself as well as the other channel by comparing model and monitoring signals; and is also similarly redundantly monitored by the other channel of the controller (20). Each channel independently compares its own signals with the signals from the other channel. Each of the two channels has a deactivating switch responsive to a fault status signal from either of the two channels to thereby allow deactivation of a failed channel even when the failure in such channel prevents it from deactivating itself.

Abstract: A helicopter flight control system (21) includes a model following control system architecture which operates in a velocity command mode at low ground speeds. The control system processes information from a variety of helicopter sensors (31) in order to provide a command signal to the main rotor (11) of the helicopter which results in a ground speed which is proportional to the input provided via a sidearm controller (29).

Abstract: This invention makes use of a pressure sensor, which is installed on the wing's surface of the aircraft, to sense the varied air pressure which comes from ground-effect on the wing's surface. A control apparatus is comprised of pressure sensor, a device of signal treatment and the device of elevator control. This apparatus solves the problems of excessively low ground-effected flying altitude and low efficiency of the whole aircraft.

Abstract: A flexible airfoil section for a wing or a blade comprising a streamlined shape and an elastic structure whose stiffness distribution along its chord and span is tailored to provide a desirable cambered shape with proportional increases in camber with increases in lift, mounted to a supporting structure such that the airfoil sections are free to pivot about axes near their leading and trailing edges. In operation, the foil derives much of its lift from elastic bending deformation of its flexible shape, thereby achieving a higher lift than a symmetric foil at the same angle of attack while postponing the onset of flow separation and stall and, for operation in water, of ventilation and cavitation.

Abstract: A helicopter flight control system includes a model following control system architecture having provisions to compensate for Euler singularities which occur when the pitch attitude of the helicopter starts to approach ninety degrees. The control system processes information from a variety of helicopter sensors in order to provide the control commands to the helicopter main and tail rotors. The present invention synchronizes a sensed attitude signal, and a desired attitude signal as the pitch attitude of the helicopter approaches ninety degrees to compensate for the Euler singularities.

Abstract: An aircraft and a method and system for operating thereof as force/moment sensors integrated in elastic connecting joints of parts and units of the aircraft. The various force/moment components determined by the sensors are processed in order to generate control signals for optimizing the operation of the aircraft.

Abstract: According to the invention each of the pitch and thrust commands is a linear combination, inter alia, of the trim setting and of the speed setting, and the control provided by the aircraft joystick is pitch rate control. Protection may be provided with respect to trim, angle of incidence, and vertical load factor.

Abstract: Hydraulic power supplies for aircraft which have such hydraulically operated systems as flaps, ailerons, spoilers, rudders, elevators, brakes, and retractable and steerable landing gear. Typically, the aircraft will be of the multiengine type; and there will also be a multiplicity of hydraulic power supplies and engines. This scheme can be utilized to provide operating redundancy for the hydraulically operated devices of the aircraft. Each hydraulic power supply has a solid state controller which has a number of inputs as well as control and information providing functions. Control over the functioning of each hydraulic power supply can also be exercised by flight deck crew-operated switches which are connected to inputs of the solid state controllers.

Abstract: Airplanes which feature composite construction and have: a further aft than conventional center of gravity; a clean wing with a high aspect ratio; and all primary flight controls. These include trailing edge control segments (or elements) which extend the full span of the airplane wing. A central processor so schedules the upward and downward deployment of the control elements that: (1) downward and upward displacements of the control elements are respectively accompanied by pitch-up and pitch-down moments; (2) the inboard control elements are used primarily to change the coefficient of lift whereas roll control is effected principally by displacement of the outboard control elements; and (3) each of the control elements is individually displaced by a system which eliminates hydraulic systems and control cables. Hallmarks of these novel airplanes include simplicity, relatively low cost, lower weight, and safety.

Abstract: A system for controlling an aircraft by aeroelastic deflections of the wings which is effective beyond control surface reversal is disclosed. The system includes flexible wings, leading and trailing edge control surfaces attached to the wings, sensors to measure selected aircraft flight parameters, an information processing system to receive and process pilot command signals and signals from the sensors, and control mechanisms in the wing that respond to processed signals from the information processing system. The control mechanisms selectively position the control surfaces to produce loads such that the wings are deflected in a desired manner for aircraft control. The system can be used for aircraft control (including maintaining stability), optimum cruise, and maneuver performance. Augmentation can be added for maneuver load control, gust load alleviation, and flutter suppression.

Abstract: Aircraft engine deactivation apparatus for stopping an aircraft engine while the aircraft is on the ground. The apparatus is for safety purposes and is used to prevent a detected object from coming into contact with an engine driven propeller or a jet propulsion intake. A detector, preferably an infra-red radiation sensor, detects an object or person within a selected distance and within a seleced area about the engine. Upon detection, a mechanical engine deactivator, such as brake calipers engageable with the engine flywheel, or an electronic deactivator, such as an electronic switch operable to ground magnetos, shuts down the engine. A by-pass switch renders the system inoperable, when desired.

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: A method and device for monitoring the consciousness of an aircraft pilot or operator of any other vehicle are described. The monitoring comprises the control of the steering deflections effected by the operator, the size and direction of which are indicated by a steering signal (DP) emitted from the steering control. For every new steering control deflection effected in the opposite direction to the one immediately preceding the time value (CPT) for the steering control deflection is calculated and this time value is compared with at least one predetermined time limit value (CPTW, CPTA). At an excessive signal amplitude considered to indicate an abnormal steering performance which may be caused by a lowered degree of consciousness a warning signal and/or an auto-steering mode is activated. The monitoring according to the invention may also cover abnormally large and rapid steering control deflections as conditions for the activation of the warning signal and the auto-steering mode.

Abstract: A precision approach control system designed to control the approach of an aircraft during landing to provide a more stable and easier mode of landing during critical landing situations, such as during the landing of an aircraft on an aircraft carrier. During operation, when the aircraft is subjected to vertical or horizontal winds or wind shear, the system controls the aircraft to maintain the inertial flight path angle constant which essentially defines operation in the precision approach control mode. In one disclosed embodiment, the precision approach control system changes the controller in the cockpit that is normally the pitch rate command stick controller during a Power Assist landing into a flight path angle rate controller. The autothrottle system for the aircraft is utilized to maintain the aircraft at a predetermined angle of attack during landing in the precision approach mode.

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: The method of the invention provides for actuating at least one actuator, such as a piezoelectric transducer in response to movement of a waveguide, such as an optical fiber. In a preferred embodiment of the invention, both the piezoelectric member and the optical fiber are connected to a structure. The connection is such that movement of the structure is detected by corresponding variations in and/or back scattering from light transmitted into the optical fiber. The piezoelectric member is actuated in response to the movement of the structural member as detected by the variations in the light in passing through the optical fiber. The actuation of the piezoelectric member may effect a dampening of the initial structural movement. Alternatively, monitoring the light passing through the fiber may be used as feed back on the position of the structural member, which in turn may be used to reposition the structural member by further actuation of a piezoelectric member.

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.

Abstract: Prevention of actuator oscillation due to failure mode in an aircraft pitch axis control system is provided by an oscillatory failure monitor. The oscillatory failure monitor provides power shutdown of the actuator through inner loop modeling and error signal processing of inner loop disturbances.

Abstract: A method for controlling an aircraft to prevent high G-caused pilot unconsciousness. Data defining a state space of acceleration, rate of change of acceleration and duration of acceleration at maximum acceleration within which an aircraft may be operated without causing pilot unconsciousness is provided to an aircraft intelligent flight control system. The flight control system continuously monitors the past and present state of the aircraft and compares to the surface boundaries of the defined safe state space, Whenever the aircraft exceeds those boundaries, the flight control system intervenes to unload the aircraft to within those boundaries. Additional data and measurements may be added to define an n-dimensional state space. Another embodiment unloads the aircraft to a baseline acceleration. A simplified embodiment is described which compares current acceleration to a preselected value of acceleration.

Abstract: Monitoring a pilot's oxygenation during flight using near infrared technology to detect pilot blackout, as for example during high G aircraft maneuvers. Initiating automatic programmed flight control or remote controller programmed flight control to avert loss of life, property or aircraft as a consequence of such detection.

Abstract: A vertical rudder control and drive system for an aircraft, such as an aiane, is equipped with electrically controllable vertical rudder drive systems on each side of the aircraft. Additionally, the control system is equipped with a mechanically controllable auxiliary drive system for the operation of the vertical rudder in response to foot pedals operated by the pilot when there should be a failure in the electrically controlled drive systems. A mechanical control signal transmitting link is provided between the foot pedals and the hydro-mechanical drive for the vertical rudder. Monitoring features enable the pilot to test the mechanical drive system without actually using that system in flight. Preflight tests may be performed.

Abstract: Strain gages embedded in an upper skin structure portion of the wing, near the centerline axis, measures centerline moment and produces a feedback signal used for adjusting the control surfaces of the wing, to produce stabilizing forces. This enables the wings to be designed primarily for strength, not stiffness.