Abstract: A controller for reducing unwanted sideways motion of an aircraft by reducing lateral side loads, resulting from air mass, turbulence and gusts. The controller functions in a manner that in the presence of higher frequency side loads, the rudder is caused to move in a relieving direction so that the net force across the vertical stabilizer is reduced. A pressure differential across opposite sides of the vertical stabilizer is measured and used to generate a first rudder deflection signal. To maintain stability of the aircraft, a beta-dot signal from the yaw damper module is gain adjusted and filtered to generate a second rudder deflection signal which is added to the first rudder deflection signal. The resulting combined signal reduces lateral side loads at higher frequencies without comprising aircraft directional stability.

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: 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: The disclosure relates to systems for piloting aircraft. Equations of the movements of the aircraft are used for computing, in a computer, of the first and second derivatives parameters V.sub.p, .PHI..sub.v, .sigma..sub.a, .beta..sub.v, .alpha., .PHI..sub.v, .sigma..sub.a, .beta..sub.v, of the parameters V.sub.p, .PHI..sub.v, .sigma..sub.a, .beta..sub.v, .alpha. of the movement of the aircraft, and the values thus computed are compared with those computed by means of the values of said parameters V.sub.p, .PHI..sub.v, .sigma..sub.a, .beta..sub.v, .alpha. at the instants t, (t+dt) and (t+2dt). An alarm signal (circuit 23) is generated if equality is not verified to a given precision.

Abstract: A helicopter engine speed reference (66) is increased (117,118) in response to heavy rotor loading (102) in excess of a threshold at the current descent rate (FIG. 4 ). The reference speed is faded up (117,118) at a rapid rate to 107% of rated speed (119). After a fixed time interval (129), reduce rotor loading (122), reduced pitch angle below a threshold magnitude (121) and reduced roll angle below a threshold magnitude (120) will cause the reference speed to be faded down slowly (146,147) to rated speed (148). Automatic increase in engine reference speed is overridden to 107% of rated speed (203) when a battle switch is activated (201) and weapons are ready (202). Increases in engine reference speed are prohibited (207) when the helicopter is operating in quiet mode (206) or the helicopter is resting on its wheels (208).

Abstract: A controller for automatically controlling aircraft thrust during a noise abatement climb operates so as to smoothly reduce aircraft thrust during a reduction in aircraft climb angle at the noise abatement altitude so as to maintain a selected noise abatement climb gradient dependent upon the pilot maintaining a recommended climb airspeed. The selected noise abatement climb gradient is compared with a measured climb gradient to generate an error term. The measured climb gradient is adjusted by a acceleration correction term in order to make the resulting measured climb gradient term independent of airspeed changes. The climb gradient error term is added to the selected climb gradient to generate a commanded climb gradient which is converted to an equivalent thrust (noise abatement thrust). The resulting noise abatement thrust is compared with measured engine thrust to generate an error term for controlling engine thrust so that the aircraft follows the noise abatement climb gradient.

Abstract: A mode control (25) controls a helicopter flight control system (21) between a heading hold mode and a turn coordination mode according to switching patterns resembling the natural response of an expert pilot to flight conditions. Helicopter operating parameters (31,30) are detected (31,29) and each parameter's detected value is converted from a crisp value to a fuzzy input (100). A new mode fuzzy output is provided by applying a compositional rule of inference (105) across each fuzzy input and a composite mode selection rule base (110). The new mode fuzzy output is converted into a crisp value (112) which is used to determine whether the flight control system operates in the heading hold mode or the turn coordination mode.

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: 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: 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: An aircraft flight control system including model following control laws includes improved logic and algorithms to limit the error between a desired parameter value from the output of a model and an actual parameter value. Such logic is operable to sense the amount of said parameter error and to limit the amount of the error if it exceeds a predetermined value. The difference between the predetermined limit and the actual error is fed back to the model such that the output of the model is adjusted so that the error between the desired and actual parameter values does not exceed the predetermined value.

Abstract: A flutter control system of an aircraft wing comprising an actuator mounted on the aircraft wing to control a control surface, flight control command unit for outputting an actuation command signal for actuating the actuator, wing-displacement measurement unit for measuring an actual displacement of a predetermined point on the wing and outputting an actual displacement signal corresponding to the measured actual displacement, normal-wing-displacement prediction unit for predicting a normal displacement of the wing during normal flight in which no disturbance occurs, on the basis of the actuation command signal and of an information signal representing information about atmospheric density, flight speed and actuation position of the actuator and for outputting a predicted normal displacement signal representing the predicted displacement during the normal flight, disturbance-component estimation unit for receiving the actual displacement signal and the predicted normal displacement signal and for outputting a

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: An adaptive gain and phase controller for an autopilot for a flight vehicle, such as a hypersonic glide vehicle, that includes applying a reference excitation signal to the control system of a flight vehicle, measuring the response of the vehicle to that excitation signal, namely the gain and phase losses through airframe, and making adjustments to the gain and phase inputs to the autopilot based on those measurements. A high gain narrow bandpass filter is incorporated so that the test signal can be extracted from the airframe of the flight vehicle.

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: A flight control system of a turboprop airplane includes electronic controlled engines, which are governed by a manual operating device for setting the engine power in order to obtain a certain airspeed, a device for selecting a desired airspeed, and an engine control system for computing and controlling the required engine torque and speed as a function of ambient and engine conditions, the selected engine speed and the setting of said operating device. For automatically controlling the engine speed during the final approach to an airfield, the system includes an electronic approach speed control unit of which the adjustment signal influences the engine control device keeping the speed of the airplane during approach at a selected value whereby said manual operating device has a fixed setpoint. This electronic speed control unit may be carried out as an add-on device for retrofitting on a flight control system.

Abstract: A control system for an aircraft or other man-machine system wherein the usual visual feedback system is characterized and is optimally supplemented by a secondary feel oriented feedback arrangement in which input signals are derived from either of two supplementary feedback signal sources and the resulting algorithms characterized mathematically. The disclosure includes several exemplary arrangements of the feedback systems in which some of the input parameters are of selected value. Mathematical characterization of the feedback paths is used.

Abstract: A control system for the reentry phase of a reentry type vehicle with the vehicle having aerodynamic control surfaces for controlling the attitude of the vehicle during reentry includes an estimating device, such as an extended Kalman filter, for estimating dynamic aerodynamic operational parameters of the vehicle in response to status parameters of the vehicle, a sensing device for supplying the status parameters, an autopilot device for generating a drive command signal in response to the dynamic aerodynamic operaational parameters and an actuator device, which may include an electric motor or smaller hydraulic one than prior systems, for controlling the aerodynamics control surfaces in response to the drive command signal. Computation necessary for estinmating is used to determine a more accurate drive command signal than is typically available from prior systems, thereby permitting a smaller, lighter weight actuator to be used.

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: In order to realize an optimum operating condition of a leading-edge flap, a wind-direction sensor and a load sensor are provided at a leading-edge flap or at each of a plurality of parts of a leading-edge flap divided along the direction of a wing span. An operational control mechanism is also provided for calculating an optimum angle of the leading-edge flap on the basis of information supplied from the wind-direction sensor and the load sensor, and for operating an actuator for varying the angle of the leading-edge flap so as to attain the calculated optimum angle. The wind-direction sensor is preferably either a hot-wire anemometer or a pressure sensor.

Abstract: A synthetic speed stability subsystem for maintaining the speed and phugoid stability of an inherently unstable aircraft is connected to a main flight control system. The speed stability subsystem uses the pitch and roll angles, and the equivalent air speed of the aircraft for calculating a synthetic speed stability signal, which is fed to the main flight control system, for controlling the stability of the aircraft.

Abstract: A system and method for adjusting the camber of an airfoil extending from an aircraft for optimum flight performance. In the method, a velocity signal is produced that represents the longitudinal velocity of the aircraft. Utilizing the velocity signal, it is determined if a change in camber of the airfoil can increase the aircraft's flight performance. Thereafter, the camber of the airfoil is changed until an optimum flight performance is achieved.

Abstract: To reduce buffeting caused by separation on the wings of an aircraft, a parameter representing the buffeting in amplitude, frequency and phase is measured. The measuring signal is subjected to filtering intended to establish the characteristics of at least one mode of vibration of the wings and, by actuating a control surface according to a non-stationary law, in a localized region the wings, of alternate stresses are generated whose amplitude and phase are automatically determined to dampen one or several modes of vibration of the wings.

Abstract: In a system for controlling a servo actuated controlled device, an input signifying the position of displacement of an aircraft stick or similar manipulatable unit is processed to provide a variable gain to the controlled device. In a predetermined range of positions of the stick in which it is spaced to one side of neutral (or in such a range at each side of neutral) the system operates with low gain during displacement of the stick through a predetermined distance in either direction from a turning point to which the stick had been brought by displacement in the opposite direction. For all other displacements in that range the system operates with high gain. The method thus provides for fast, positive response to coarse stick movements but fine response to small trimming movements.

Abstract: The torque gain of an electric motor assisted power steering system is modified as a function of the estimated level of road load friction such that the relationship between the operator exerted steering torque and the resulting rate of steer is substantially insensitive to a reduction in the level of road load. The road load is estimated as a function of the operator exerted steering torque, the torque assist provided by the electric motor, and the resulting steer rate.

Abstract: An automatic trim circuit for hydrofoil craft which compensates for offsets in the system due to manufacturing and assembly tolerances and wear of parts. This is achieved by integrating an error signal derived by comparison of a reference signal with an actual position signal and applying the integrated output, after amplification, to one or more servo systems which actuate the control surfaces of the hydrofoil craft.

Abstract: The disclosure relates to a self-organizing control of advanced turbine engines wherein thrust is controlled and thrust specific fuel consumption (TSFC) is minimized by control of the fuel flow rate as well as optimization of system geometric parameters. The controller computes the net thrust error by comparing commanded thrust with an inferred or calculated thrust measurement.

Abstract: The operation of an actuator (16) is monitored by comparing its position (21) with the position (31, 136) indicated by a model which integrates (45, 135) a limited amount of the difference between the position command (24) applied to the actuator and the achieved model position (31, 136), the limited amount being variable (63, 67, 124) from a nominal limit (61, 65, 124) in dependence upon limited functions (74, 90, 114, 116) of the difference (33, 109) between the actuator position and the model position, and additionally reduced (80, 94, 122) when pilot input overrides (50, 108) the position of the actuator.

Abstract: A pilot controlled stability control system that employs direct lift control (spoiler control) with elevator control to control the flight path angle of an aircraft. A computer on the aircraft generates an elevator control signal and a spoiler control signal, using a pilot-controlled pitch control signal and pitch rate, vertical velocity, roll angle, groundspeed, engine pressure ratio and vertical acceleration signals which are generated on the aircraft. The direct lift control by the aircraft spoilers improves the response of the aircraft flight path angle and provides short term flight path stabilization against environmental disturbances.

Abstract: The present steering mechanism, especially for an aircraft includes a manly operable steering member and an automatic steering system. A difference signal formed from a pilot input and trim signal and from a gyro feedback signal is supplied to a force control and damping circuit twice, once directly and once through an authority limit circuit. The output of the authority limit circuit is also supplied to the motors forming part of the automatic steering system. The force control and damping circuit also receives further input signal or signals representing steering information. The output of the force control and damping circuit is connected to the feel unit of the steering system, whereby an immediate manual control by the pilot is effective, however limited against an excessive control or so-called manual over control.

Abstract: An aircraft carries a pressure indicator well in advance, which may record barometric pressure of vertical air currents. The indicator reading is converted to electric current and averaged over a short period. The unaveraged current at a given position of the pressure indicator is recorded, delayed, read out, compared with the average current and the error current used to actuate wing flaps at the instant the wings pass over the given position, so as to adjust the wing lift for any difference in conditions as the wings move forward. Similarly the current recorded at the instant the pressure indicator was at a given point is delayed, compared with the average current at the instant the tail structure passes over the given point, and the error current is used to adjust elevators for any difference in pressure as the tail structure moves from its first position to the position of the pressure indicator at the given instant.

Abstract: A performance advisory system which includes a display, a control unit and an onboard computer, the computer receiving inputs from aircraft instruments and gauges and from the control unit and providing advisory output information on the display.

Abstract: A control system of the non-linear type which behaves in a substantially `proportional` manner. The system comprises an error detector which compares the input and output of a plant to obtain an error value, and a control unit which derives from the error value a continuous control signal for the plant. The control signal comprises a linear term which is a linear function of the error signal and a nonlinear damping term which is a continuously variable function comprising the ratio of the rate of change of the error signal and the error itself.

Abstract: This invention is an improvement in aircraft control systems that utilize feedback motion sensors to generate a control signal to control the aircraft. The improvement consists essentially of a complementary filter comprising a simplified model of the aircraft, a high-pass filter, a low-pass filter and a summing amplifier. The control signal is applied to the simplified model of the aircraft which ateempts to compute the vehicle response to the signal. This computed response is then fed into the high-pass filter to eliminate long-term errors in the calculated response, with the result that a good estimate of the high frequency content of the aircraft motion is obtained. In order to obtain a good estimate of the low frequency content of the motion a rate gyro signal is fed through the low-pass filter that eliminates all of the offending noise. The outputs from the high-pass and low-pass filters are summed by the summing amplifier to produce an estimated rate which is then used as a motion feedback signal.

Abstract: Plural servo actuators are coupled to provide a redundant system for operating a common output device. Each servo actuator has plural servo valves each having one active output that causes controlled opposite operation of the output device and one passive output. A fault detector coupled to each servo actuator compares the sum of the passive output pressures with the pressures of two other fluid signals, in one case the supply and return fluid signals and in another case the two active output signals of the servo valves, to detect the condition thereof. A convenient checkout system and a fluid by-pass mechanism also are provided in the fault detector.

Abstract: A method and device are provided for displaying the effect of control action in real time by simultaneously applying the control action both to the system to be controlled and a computer model of the system. The effect of the control action on the model is displayed to the operator in a manner which dynamically decouples the display to the operator in a manner which dynamically decouples the display from the actual system's response over some suitable frequency range, thereby enhancing the stability and performance of the control system's operation. The method and device are described in conjunction with a preferred embodiment of an aircraft gun fire control system and variation of that system. A computer model or pseudo aircraft, termed a command aircraft, is programmed into a computer carried by an actual aircraft. In addition to providing conventional control information to the actual aircraft, the control actions of the pilot "fly" the command aircraft.

Abstract: An autopilot automatically adjusts the sensitivity of a ship's steering system to accommodate changes in speed as well as sea and wind conditions. The autopilot utilizes heading error, speed and speed squared signals to produce a rudder order signal for controlling rudder position. The rudder order signal is developed in a heading keeping circuit unless a heading change greater than a predetermined threshold is commanded, in which event a programmer substitutes a heading change circuit for the heading keeping circuit. The sensitivity of the heading change circuit is automatically adjusted as an inverse function of vessel speed squared, and automatic rudder order limits are established in the same circuit as an inverse function of speed. The sensitivity of the heading keeping circuit is adjusted in accordance with a signal from an automatic gain control circuit which derives a performance index J from ship's speed, heading error and rudder order signals occurring during a given measurement interval.

Abstract: The instant invention relates to a system for estimating aircraft angle of attack .alpha. and sideslip angle .beta. from measured quantities such as angular body rate and linear acceleration. The estimated angle of attack and sideslip signals are generated by means of a Kalman filter configuration which simulates a model of the aircraft angle of attack and sideslip angle dynamics, and may be used either in an aircraft flight control network or for display purposes.

Abstract: A system for rapidly disengaging, with minimum nuisance disengagements, a nventional autopilot system upon detecting hardover failure, which is defined as any autopilot failure which results in rapid and sustained displacement of an aircraft aerodynamic control surface. The aircraft pitch rate is measured and compared with established limits. If these limits are exceeded, the system causes automatic disengagement. Disengagement is inhibited for gust disturbances by deriving a nose up or down signal from the pitch rate signal which is then compared with a signal indicating the direction of elevator hinge moment; if the two directions are opposing, disengagement is inhibited. Disengagement is also inhibited in the event of pilot action. Also, the autopilot pitch axis command line is monitored and a disengagement overriding any inhibits is generated when the signal on that line exceeds certain limits for the altitude hold mode and for the attitude hold mode of the autopilot.

Abstract: The invention relates to a flight control system for an aircraft which maintains reliable control of the angle of attack (.alpha.) in those regions where rate of change of the lift coefficient (C.sub.L) as a function of .alpha. goes to zero or becomes negative. This may cause a condition in which the feedback loop for the pitch moment generating control surface can inadvertently drive the aircraft into a stall or other unstable mode. Reliable control of angle of attack is achieved through a network which produces an estimated angle of attack signal (.alpha.) from the measured pitch rate of the aircraft, the lift acceleration error signal, a gravity control signal and sideslip and roll information. The estimated angle of attack signal is processed in a network which produces a pseudo coefficient of lift signal C.sub.L. The pseudo lift coefficient signal is processed to produce a pseudo lift (or normal) acceleration signal, A.sub.Z, which is supplied to the aircraft control loop.